Aerospace Materials and Applications [1 ed.] 9781624104893, 9781624104886

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Aerospace Materials and Applications Edited by Biliyar N. Bhat

P R O G R E S S

I N

A S T R O N A U T I C S

A N D

A E R O N A U T I C S

Timothy C. Lieuwen, Editor-in-Chief Volume 255

Aerospace Materials and Applications

Aerospace Materials and Applications EDITED BY

Biliyar N. Bhat

Volume 255 Progress in Astronautics and Aeronautics Timothy C. Lieuwen, Editor-in-Chief Georgia Institute of Technology Atlanta, Georgia

Published by American Institute of Aeronautics and Astronautics, Inc. 12700 Sunrise Valley Drive, Suite 200, Reston, VA 20191-5807

Cover image credits, left to right, top to bottom: Friendship 7 launch; Credit: NASA X-37B Orbital Test Vehicle; Credit: U.S. Air Force photo/Michael Stonecypher Space Shuttle Discovery, underside view; Credit: NASA Arc jet testing; Credit: NASA DARPA/Air Force Falcon HTV-2 vehicle X-43A research vehicle, from animation video; Credit: NASA American Institute of Aeronautics and Astronautics, Inc., Reston, VA.

Copyright # 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Printed in the United States of America. No part of this publication may be reproduced, distributed, or transmitted, in any form or by any means, or stored in a database or retrieval system, without the prior written permission of the copyright holder. Data and information appearing in this book are for informational purposes only. AIAA is not responsible for any injury or damage resulting from use or reliance, nor does AIAA warrant that use or reliance will be free from privately owned rights. ISBN 978-1-62410-488-6

PROGRESS IN ASTRONAUTICS AND AERONAUTICS

EDITOR-IN-CHIEF Timothy C. Lieuwen Georgia Institute of Technology

EDITORIAL BOARD Paul M. Bevilaqua

Eswar Josyula

Lockheed Martin (Ret.)

U.S. Air Force Research Laboratory

Steven A. Brandt

Mark J. Lewis

U.S. Air Force Academy

Institute for Defense Analyses

Jose´ A. Camberos

Dimitri N. Mavris

U.S. Air Force Research Laboratory

Georgia Institute of Technology

Richard Christiansen

Alexander J. Smits

Sierra Lobo, Inc.

Princeton University

Richard Curran

Ashok Srivastava

Delft University of Technology

Verizon Corporation

Simon Hook

Karen Thole

Jet Propulsion Laboratory

The Pennsylvania State University

Christopher H. M. Jenkins

Oleg A. Yakimenko

Montana State University

U.S. Naval Postgraduate School

TABLE OF CONTENTS

Preface . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . xiii Chapter 1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1 Biliyar N. Bhat, NASA Marshall Space Flight Center

Chapter 2 Aerospace Materials Characteristics . . . . . . . . . . . . . . . . 11 Biliyar N. Bhat, NASA Marshall Space Flight Center

2.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2 Aluminum Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3 Titanium Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.4 Steels . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.5 Superalloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.6 Copper Alloys . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.7 Damage Tolerance Considerations for Metallic Materials 2.8 Hydrogen Embrittlement in Metallic Materials . . . . . . . . 2.9 Material Behavior in Oxygen-Rich Environments . . . . . . . 2.10 Polymers and Composites . . . . . . . . . . . . . . . . . . . . . . . . 2.11 Polymer Matrix Composites in Aerospace Structures . . . 2.12 Aerospace Ceramic Materials . . . . . . . . . . . . . . . . . . . . . . Acknowledgements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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11 13 41 55 62 69 75 87 113 132 158 184 203 203

Chapter 3 Materials Selection for Aerospace Systems . . . . . . . . . . 209 Steven M. Arnold, NASA Glenn Research Center, Cleveland, Ohio; David Cebon and Mike Ashby, University of Cambridge, Cambridge, United Kingdom

3.1 Introduction . . . . . . . . . . . . . . . . . 3.2 Systematic Approach to Materials 3.3 Advanced Selection . . . . . . . . . . . 3.4 Data Issues . . . . . . . . . . . . . . . . . 3.5 Summary . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . .

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Chapter 4 Advanced Nanoengineered Materials . . . . . . . . . . . . . . 275 Brian L. Wardle, Massachusetts Institute of Technology, Cambridge, Massachusetts; Joseph H. Koo, University of Texas at Austin, Austin, Texas; Gregory M. Odegard, Michigan Technological University, Houghton, Michigan; Gary D. Seidel, Virginia Polytechnic Institute and State University, Blacksburg, Virginia

4.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.2 Nanoengineered Structural Materials . . . . . . . . . . . . . . . . . 4.3 Nanoengineered Multifunctional Materials . . . . . . . . . . . . . 4.4 Processing and Manufacturing . . . . . . . . . . . . . . . . . . . . . . 4.5 Modeling of Nanomaterials: Multiscale and Multitechnique 4.6 Future Trends . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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275 277 287 289 293 295 296 296

Chapter 5 Subsonic Aircraft Materials Development . . . . . . . . . . . 305 Michael Mohaghegh, Seattle, Washington; David L. Stone, Redstone Arsenal, Alabama; Antonio F. Avila, Universidade Federal de Minas Gerais, Belo Horizonte, Minas Gerais, Brazil

5.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.2 Subsonic and Supersonic Fixed Wing Aircraft . . . . . . . . . . . . 5.3 Rotorcraft . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.4 Enabling Materials, Structures, and Manufacturing Processes 5.5 Closing Comments . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Acknowledgements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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305 308 368 386 397 397 397

Chapter 6 Materials for Spacecraft . . . . . . . . . . . . . . . . . . . . . . . . 403 Miria M. Finckenor, NASA Marshall Space Flight Center, Huntsville, Alabama

6.1 6.2 6.3 6.4 6.5 6.6 6.7 6.8 6.9 6.10 6.11 6.12

Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Space Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . Spacecraft Design and Materials Requirements . . . . . . Flammability, Toxicity, and Offgassing Considerations Structural Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . Thermal Control Materials . . . . . . . . . . . . . . . . . . . . . . Thermal Protection Materials . . . . . . . . . . . . . . . . . . . . Radiation Shielding . . . . . . . . . . . . . . . . . . . . . . . . . . . Meteoroid/Orbital Debris Shielding . . . . . . . . . . . . . . Optical Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Solar Array Materials . . . . . . . . . . . . . . . . . . . . . . . . . . Lubricants . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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403 404 405 408 412 414 418 420 421 423 423 424

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6.13 Seal Materials . . . . . . 6.14 Adhesives . . . . . . . . 6.15 Lessons Learned . . . 6.16 Concluding Remarks Acknowledgments . . . . . . . References . . . . . . . . . . . . .

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Chapter 7 Materials for Launch Vehicle Structures . . . . . . . . . . . . 435 Grant Henson, Invariant Laboratories LLC, Westlake, Ohio

7.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . 7.2 Launch Vehicle Structures . . . . . . . . . . . . . . 7.3 Basic Material Characteristics . . . . . . . . . . . . 7.4 Structural Design and Requirements . . . . . . 7.5 Pressurized Structure . . . . . . . . . . . . . . . . . . 7.6 Feedlines, Small Lines, and Pressure Vessels 7.7 Unpressurized Structure . . . . . . . . . . . . . . . 7.8 Thermal Protection and Insulation . . . . . . . 7.9 Manufacturing Considerations . . . . . . . . . . . 7.10 Summary, Trends, and Outlook . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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435 438 440 447 458 478 479 485 490 496 498

Chapter 8 Materials for Exploration Systems . . . . . . . . . . . . . . . . 505 Peter A. Curreri, NASA Marshall Space Flight Center, Huntsville, Alabama

8.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.2 Development of a Technical Basis for Materials Processing in Space . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.3 Development of an Economic Basis for Materials Processing in Space: In Situ Propellant Production . . . . . . . . . . . . . . . . . 8.4 Summary and Conclusions . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 9 Thermal Protection Systems and Hot Structures for Hypersonic Vehicles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 531 David E. Glass, NASA Langley Research Center, Hampton, Virginia

9.1 Introduction . . . . . . . . . . . . . . . . . . . 9.2 Background . . . . . . . . . . . . . . . . . . . 9.3 TPS and Hot Structure Components 9.4 Key Technical Challenges . . . . . . . . 9.5 Concluding Remarks . . . . . . . . . . . . Acknowledgments . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 10 Aero Engine Materials . . . . . . . . . . . . . . . . . . . . . . . . 579 James C. Williams, University of North Texas, Denton, Texas

10.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.2 Brief Historical Perspective on Aircraft Engine Evolution . 10.3 Evolution of Jet Engine Materials . . . . . . . . . . . . . . . . . . 10.4 Materials Processing . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.5 Trends and Outlook . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10.6 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . Bibliography . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 11 Materials for Solid Rocket Engines . . . . . . . . . . . . . . . 609 Diep V. Trinh, NASA Marshall Space Flight Center, Huntsville, Alabama

11.1 Introduction . . . . . . . . . . . . . . . . . . 11.2 Solid Rocket Components . . . . . . . 11.3 Motor Cases . . . . . . . . . . . . . . . . . . 11.4 Insulation Materials . . . . . . . . . . . . 11.5 Liner and Inhibitor . . . . . . . . . . . . . 11.6 Nozzle . . . . . . . . . . . . . . . . . . . . . . 11.7 Igniters . . . . . . . . . . . . . . . . . . . . . . 11.8 Hybrid Rocket Propulsion Systems 11.9 Summary, Trends, and Outlook . . . Bibliography . . . . . . . . . . . . . . . . . . . . . . .

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609 610 618 619 626 627 633 634 638 638

Chapter 12 Materials for Liquid Propulsion Systems . . . . . . . . . . . 641 John A. Halchak, Los Angeles, California; James L. Cannon, NASA Marshall Space Flight Center, Huntsville, Alabama; Corey Brown, Aerojet-Rocketdyne, West Palm Beach, Florida

12.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.2 Liquid Rocket Engines . . . . . . . . . . . . . . . . . . . . . . . . . 12.3 General Design Considerations for Materials Selection 12.4 Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.5 Thrust Chamber Materials . . . . . . . . . . . . . . . . . . . . . . 12.6 Turbopump Materials . . . . . . . . . . . . . . . . . . . . . . . . . 12.7 Valves . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.8 Lines and Ducts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12.9 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 13 Advanced Materials for In-Space Propulsion . . . . . . . . 699 Les Johnson and Tiffany Lockett, NASA Marshall Space Flight Center, Huntsville, Alabama

13.1 Introduction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.2 Chemical Propulsion (Liquid Storable and Liquid Cryogenic) 13.3 Gridded Ion Thrusters . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.4 Hall Propulsion Systems . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.5 Solar Thermal Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.6 Nuclear Thermal Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . 13.7 Solar Sail Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.8 Tether Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13.9 Advanced Propulsion Technologies . . . . . . . . . . . . . . . . . . . . 13.10 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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Chapter 14 Materials for Power Systems in Space Exploration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 749 Ajay K. Misra, NASA Glenn Research Center Cleveland, Ohio

14.1 Introduction . . . . . . . . . . . . . . 14.2 Solar . . . . . . . . . . . . . . . . . . . 14.3 Fuel Cells . . . . . . . . . . . . . . . . 14.4 Batteries . . . . . . . . . . . . . . . . . 14.5 Radioisotope Power Systems . 14.6 Fission Power for Space . . . . . 14.7 Concluding Remarks . . . . . . . References . . . . . . . . . . . . . . . . . . . . Index

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Supporting Materials

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PREFACE The present Aerospace Materials and Applications text is the result of an extensive collaboration of the AIAA Materials Technical Committee (MTC) members and invited authors under the leadership of Dr. Biliyar N. Bhat (Bil) of NASA Marshall Space Flight Center (AIAA Progress in Astronautics and Aeronautics Series, vol. 255, B.N. Bhat, Editor, 2018). I personally know Bil has spent countless hours working with all authors and co-authors (I counted there are at least 24 authors and co-authors from the government, industry, and academia) dedicated to complete this time-consuming undertaking that represents the expertise of the AIAA MTC and the aerospace materials community. The present volume is focused on documenting the novel processing, fabrication, characterization, and testing approaches that are unique to aerospace materials/structures/systems. These aerospace materials/structures/ systems consist of metals, alloys, polymers, polymer matrix composites, carbon/ carbon composites, ceramic matrix composites, as well as advanced nanoengineering materials. It attempts to cover the entire field of aerospace materials in a condensed fashion. The scope of this book is to provide both the theory and especially the practice of applying aerospace materials for project managers and technical specialists. The intended outcome of the book is a single repository for all the unique approaches that are taken in validating aerospace materials on the ground and in space. The anticipated audience includes engineers, technicians, and technology managers involved in aerospace materials selection, characterization, modeling, and testing. This volume consists of 14 chapters that are summarized as follows: † † † † †

† †

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Chapters 1 and 2 provide an introduction that summarizes all chapters (Chapter 1) and aerospace materials characteristics (Chapter 2); Chapter 3 covers materials selection for aerospace systems; Chapter 4 describes advanced nanoengineering materials; Chapter 5 discusses subsonic aircraft materials development; Chapters 6– 8 provide detailed discussions on materials for spacecraft (Chapter 6), launch vehicle structures (Chapter 7), and exploration systems (Chapter 8); Chapter 9 covers the thermal protection systems (TPS) and hot structures for hypersonic vehicles; Chapters 10 –12 deal with materials relating to aero engines (Chapter 10), solid rocket engines (Chapter 11), and liquid propulsion systems (Chapter 12); Chapter 13 introduces advanced materials for in-space propulsion (Chapter 13); Chapter 14 completes the volume with materials for power systems in space exploration.

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On behalf of all the authors and co-authors, we wish to express our appreciation to the publications staff at AIAA for their help and guidance, which has been invaluable. We also owe a debt of gratitude to the all the members of the AIAA Materials Technical Committee for their support of this project. Finally, the AIAA MTC wishes to thank the many expert authors who have contributed their vast knowledge and experience – and valuable time – to the production of the present volume. It has been our pleasure to work with them and learn from them. Without their expertise and dedication, this volume would not have been possible.

Joseph H. Koo on behalf of the AIAA Materials Technical Committee The University of Texas at Austin, Austin, Texas July 2018

CHAPTER 1

Introduction Biliyar N. Bhat NASA Marshall Space Flight Center

Advances in aerospace systems such as aircraft, spacecraft, and launch vehicles are strongly dependent on advances in materials and processing technologies. In the past hundred years of powered flight, aircraft structures have evolved around advances in materials that are lighter and stronger. Aircraft propulsion systems are constantly striving to become more fuel efficient, which requires reductions in mass and improved capability for materials to operate at higher temperatures for longer times. Gas turbine engines that power modern aircraft are being designed to run at higher pressures and temperatures to generate more thrust per pound of engine mass. Similar considerations apply to rocket engines where the power densities are much higher. Spacecraft are designed to operate in the harsh radiation environments of outer space. In-space propulsion and power systems are key components of spacecraft, and advanced materials enable these systems. Hence it is important for aerospace systems designers to have a good understanding of how the materials will perform in their systems. Although there are many books on materials and applications in the open literature, none is specifically tailored to meet the needs of aerospace applications. Because of the unique requirements of aerospace hardware, the preferred approach to selecting materials in the design and construction of aerospace systems deserves its own guidebook. This book is intended to fill that need. Modern aerospace systems in general are designed and developed by a team of engineers representing different disciplines, such as structures, materials, manufacturing, and so on. The team is often called a product development team, and it takes a systems engineering approach to design and development. Designers often tend to focus on the function of the system (what does it do?) to meet the customer needs. In the conceptual phase the designer does not always have a clear understanding of the materials requirements and assumes that the materials will be available as and when needed. Materials selection is deliberated later in the

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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process, which may in turn affect the design. In this sense the design process is an iterative process in which the design is refined after each cycle, leading to optimization. Sometimes materials selection may not be made early enough in the design cycle, resulting in a less than satisfactory outcome. Other times the design may look elegant on paper but may present problems during manufacturing. This will drive up the cost and delay the product delivery. In the worst case the system may have to be redesigned completely because it simply could not be built as designed the first time. Getting the design right the first time is challenging and requires good communication among the various disciplines involved. This is particularly true with materials selection. Most engineers are not trained to think in an interdisciplinary fashion, especially about materials. By the same token, materials engineers are generally not trained in systems engineering. They need to have a better understanding of the design process to give their input to the design effort in a timely manner. This book is an attempt to bring the aerospace materials and aerospace systems design disciplines closer together. Aerospace structures and propulsion systems applications are emphasized. These are by far the most demanding applications from a materials perspective. Specific strength and stiffness, factors of safety and damage tolerance, are important considerations. Other considerations, such as environmental compatibility and thermal and electrical properties, may be important in certain nonstructural applications. It is important to select the best available material in every case. Cost is always a consideration, and the goal is to develop the best design at an affordable cost. This book presents aerospace materials and their applications in different aerospace systems. It shows how the materials are selected and used in aircraft, spacecraft, and launch vehicles, propulsion systems, and power systems. It attempts to cover the entire field of aerospace materials in a condensed fashion. The subject matter is organized into 13 chapters (Chapters 2 through 14). Each chapter is contributed by one or more authors who are recognized experts in their fields. The chapters are arranged such that the reader is first introduced to basic aerospace materials and their characteristics (Chapter 2) followed by how to select the best material for a given application (Chapter 3). Advanced nanoengineered materials are discussed next (Chapter 4). These materials are on the cutting edge of materials technology. The rest of the book (Chapters 5 through 14) focuses on the applications of materials in aerospace systems. Chapters 5 through 9 cover materials in aerospace structures: Chapter 5, aircraft; Chapter 6, spacecraft; Chapter 7, launch vehicles; Chapter 8, space exploration systems; and Chapter 9, hypersonic vehicles. Chapters 10 through 13 cover materials and applications in propulsion systems: Chapter 10, aero engines; Chapter 11, solid rocket engines; Chapter 12, liquid rocket engines; Chapter 13, in-space propulsion. Chapter 14 discusses materials for power systems in space exploration. In the interest of keeping the book’s size reasonable, the authors have made an effort to present the material in a condensed fashion. For more details, the reader

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should go to the references listed at the end of each chapter or section. The focus is on the state of the art where the technology readiness levels (TRLs) of materials and processes are high (TRL ¼ 8 or 9). Promising materials developments in the TRL ¼ 6 –7 range are also covered in some instances. Areas of lower technology readiness levels (TRL ¼ 5 or lower) are not covered because they are not considered mature enough for application in real systems. It should be noted that TRLs of materials and processes need to be generally high (greater than 6) before they can be seriously considered in a design. Many chapters discuss the trends and outlook for materials and applications in future aerospace systems. A brief description of chapter contents follows.

CHAPTER 2: AEROSPACE MATERIALS CHARACTERISTICS This chapter gives an overview of commonly used aerospace materials—metals, nonmetals (polymers), composites, and ceramics—organized in short sections. The purpose is to help the designer become familiar with aerospace materials in use today. It is not meant to be exhaustive. The materials information available to the designer is extensive and would take too much space. Hence that information is presented in a condensed fashion with applications in mind. Materials properties most relevant to aerospace structures and propulsion systems are presented. For more details the reader should go to the references, which are listed at the end of each section. Metallic materials are covered in Secs. 2.2 through 2.6: 2.2, Aluminum Alloys; 2.3, Titanium Alloys; 2.4, Steels; 2.5, Superalloys; and 2.6, Copper Alloys. Damage tolerance considerations are described in Sec. 2.7. Details on alloy development, properties, processing, and typical applications are presented. Relationships between properties, microstructure, and processing are also described with aerospace applications in mind. Structural properties such as elastic modulus, tensile strength, ductility, and damage tolerance (fatigue and fracture) are emphasized because they are major considerations in design. Manufacturing technologies commonly used to fabricate metallic material components are also described in the context of design for manufacturing. Environmental effects on materials performance are described in Sec. 2.8, Hydrogen Embrittlement in Metallic Materials, and Sec. 2.9, Material Behavior in Oxygen-Rich Environments. The last two topics are of special interest because of their importance in propulsion systems that use hydrogen and oxygen as propellants. Polymers and composites are described in Secs. 2.10 and 2.11. Polymers are different from the other materials in many ways but generally possess lower densities, thermal conductivities, and moduli. The lower densities of polymeric materials offer an advantage for applications where light weight is a requirement. Carbon fiber reinforced composites and their processing are described. Section 2.11 focuses on mechanical behavior of polymer matrix composites and their applications in launch vehicles compared with aircraft.

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Ceramic materials and their characteristics are discussed in Sec. 2.12. The focus is on several important enabling ceramic systems for aerospace applications, particularly turbine engine thermal and environmental barrier coating systems. This section also emphasizes ceramic materials for various applications, material system properties, and durability performance associated with processing.

CHAPTER 3: MATERIALS SELECTION FOR AEROSPACE SYSTEMS Materials selection is an important subject for materials experts and systems designers alike. Materials selection can be a daunting task due to wide range of choices available. Typically, the job of a designer is to balance a variety of functional requirements, that is, types of loading conditions (tension, compression, bending, vibration, cyclic, etc.) with constraints (manufacturability, geometric limits, environmental aspects, etc.) to arrive at the “optimum” choice of structural concept and materials selection for a given weight and/or cost. The chapter presents a systematic methodology to select the best materials for a given application. The design process typically involves multiple iterations due to inherent interactions between function, material, shape, and process. Material screening criteria are described and tools are provided to design specific components. Examples are given from aerospace applications.

CHAPTER 4: ADVANCED NANOENGINEERED MATERIALS Nanoengineered materials are on the cutting edge of aerospace materials used today. This chapter focuses on the advanced nanocomposites that contain carbon nanotubes, which improve polymer matrix composites’ properties in significant ways. Both current applications and future developments are covered. Topics covered include nanoengineered structural materials (nanocomposites, nanostructured fibers and sheets, nanoparticle enhanced adhesives), nanoengineered multifunctional materials, processing and manufacturing, and multiscale modeling of nanomaterials. Examples of nanomaterials applications in aerospace are presented.

CHAPTER 5: SUBSONIC AIRCRAFT MATERIALS DEVELOPMENT This chapter addresses materials development in aircraft structures. Both fixed-wing and rotary-wing aircraft are addressed. Chronological development of aircraft materials is presented from the Wright brothers’ first powered flight to the Boeing 787 aircraft of today. The trend is toward the use of higher specific strength and specific stiffness materials, from wood and fabric to aluminum and titanium alloys to carbon fiber reinforced polymer composites. Evolution of fiber

INTRODUCTION

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metal laminates is presented. Manufacturing and assembly aspects of aircraft production and in-service repair techniques are discussed. In the rotorcraft section, the focus is on the rotor blade structure and materials. Their performance in the field is discussed, including blade erosion in harsh operating environments. Damage tolerance of aircraft components and rotorcraft blades is addressed. The chapter presents an integrated approach to design, manufacture, and materials development. It shows how the relationship between the airframe designer, materials scientists, and manufacturing engineer is becoming closer and more integrated as the need for improved weight/cost performance grows. The chapter concludes with discussion of enabling materials and structures and additive manufacturing technologies for future aircrafts.

CHAPTER 6: MATERIALS FOR SPACECRAFT This chapter discusses materials for spacecraft, both manned and unmanned, from low Earth to geosynchronous orbit, cis-lunar, lunar, planetary, and deep space exploration. It discusses the damaging effects of the space environment on various materials, what has been successfully used in the past, and what may be used for a more robust design. The material categories covered are structural, thermal control for on-orbit and reentry, shielding against radiation and meteoroid/space debris impact, optics, solar arrays, lubricants, seals, and adhesives. Manned spacecraft components must meet toxicity and flammability requirements. Requirements such as fracture control and contamination control are examined, with additional suggestions for manufacturability.

CHAPTER 7: MATERIALS FOR LAUNCH VEHICLE STRUCTURES This chapter concerns materials for expendable and reusable launch vehicle structures. An emphasis is placed on applications and design requirements and how these requirements are met by the optimum choice of materials. Structural analysis and qualification strategies, which cannot be separated from the materials selection process, are described. Verification criteria, qualification strategies, and analysis methods that have matured over the past few decades are described here. These practices have a strong influence on materials selection. The major structural elements covered are pressurized structures including stable metal tanks, balloon tanks, composite tanks, solid rocket motor cases, feedlines, and composite overwrapped pressure vessels. Unpressurized structures include intertanks, skirts, adapters, payload fairings, and nose cones. Other important topics covered include composite sandwich structures and their applications and nonstructural thermal protection and insulation materials and their applications. Manufacturing and assembly of large structures such as cryogenic tanks and stages is discussed, including welding technology.

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CHAPTER 8: MATERIALS FOR EXPLORATION SYSTEMS This chapter delves into the future of space exploration when people routinely travel out of Earth’s orbit to settle in distant places such as the moon and Mars. Taking supplies from Earth is impractical, and in situ resource utilization is required. Materials processing in space is an important part of this scenario, and this chapter develops a technical basis for it. It gets into the details of materials science and processing in space, considerations for using nonterrestrial materials, materials availability, and extraction and processing in lunar and Martian environments with emphasis on solidification processing. In situ propellant production on the moon and Mars is discussed. In addition, the chapter presents an economic basis for materials processing in space, production of energy from space sources, and eventual establishment of largely self-sufficient human bases off Earth.

CHAPTER 9: THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES This chapter discusses thermal protection systems (TPSs) and hot structures for hypersonic vehicles. The focus is on air-breathing hypersonic vehicles in Earth’s atmosphere. This includes single-stage to orbit and two-stage to orbit accelerators, access to space vehicles, and hypersonic cruise vehicles. The chapter opens with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocketbased vehicles to air-breathing vehicles, we need to move away from the “insulated airplane” approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the chapter discusses issues and design options for ceramic matrix composite (CMC) TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state of the art is briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and they are discussed briefly.

CHAPTER 10: AERO ENGINE MATERIALS This chapter traces the history of the aircraft gas turbine engine from its inception in the late 1950s to the present. The evolution of jet engine materials is discussed for different sections of the engine: fan, compressor, combustor, high-pressure turbine, and low-pressure turbine. The emphasis is on the contributions of materials and processing to the substantial progress that has been made, but

INTRODUCTION

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the combined results of design and improvements in materials and materials processing methods are the real story. The increases in thrust/weight and the improvements in specific fuel consumption nicely summarize the performance improvements. For commercial aircraft the improvements in reliability and reduction in cost of ownership are equally important. These improvements are also discussed and illustrated. The quest for better materials and designs will continue. The aircraft gas turbine is a monumental success story of engineering teamwork.

CHAPTER 11: MATERIALS FOR SOLID ROCKET ENGINES This chapter describes rocket engines that use solid fuels and oxidizers as propellants. These are also known as solid rocket motors (SRMs). The chapter focuses on the Space Shuttle SRM, which serves as a good example of generic SRMs. Materials and process details for various SRM components are discussed, including propellant grain formulation, fuels, oxidizers, binders, insulation and liner, igniters, rocket motor cases, and the rocket nozzle. Insulation and liner performance parameters are described. Materials used in hybrid rocket propulsion systems, in which one of the two propellants is solid and the other is liquid, are discussed. Current trends and the future outlook for solid rocket materials are presented.

CHAPTER 12: MATERIALS FOR LIQUID PROPULSION SYSTEMS This chapter describes rocket engines that use liquid fuels and oxidizers, with a focus on Earth-to-orbit propulsion. It describes how liquid rocket engines work and explains different rocket cycles commonly used today. A brief history of liquid rocket engines is given followed by a description of various types of liquid propulsion systems. The chapter describes design considerations for the materials used in the various components of liquid rocket engines and provides examples of usage and experiences in each. Components covered are thrust chamber, including injector, nozzle, and nozzle extension, turbopumps for fuel and oxidizer, turbine drive methods and components, propellant valves, and lines and ducts. The chapter contains many illustrations of components and materials used. Advances in rocket engine materials and designs are presented chronologically from the early days of rocketry to modern high-performance rocket engines. The chapter also discusses additive manufacturing technology development for rocket engine components.

CHAPTER 13: ADVANCED MATERIALS FOR IN-SPACE PROPULSION This chapter describes propulsion after reaching the orbit. These propulsion systems perform the functions of primary propulsion, reaction control, station

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keeping, precision pointing, and orbital maneuvering. In addition, the chapter covers advanced in-space propulsion technologies that enable much more effective exploration of our solar system and permit mission designers to plan missions with greater reliability and safety. Materials selection and application are covered for all these systems. Propulsion topics of interest include chemical propulsion, gridded ion thrusters, Hall propulsion systems, solar thermal propulsion, nuclear thermal propulsion, solar sail propulsion, and tether propulsion. More advanced propulsion technologies are briefly discussed, including plasma rockets, pulsed inductive thrusters, beamed energy propulsion, fusion propulsion, and antimatter propulsion.

CHAPTER 14: MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION Power systems are key components of spacecraft. The amount of power and the source of power are different for different missions. Materials enable space power systems. This chapter provides an overview of materials used in various space power systems. The focus is on materials systems for solar power systems (photovoltaic cells), fuel cells, batteries (lithium-ion batteries), radioisotope power systems (radioisotope thermoelectric generators), and nuclear power systems (space nuclear fission reactors). It also discusses current materials research activities to meet future space power needs and long-term research needed to increase power density and energy density of future space power systems. The chapters in this book are independent of each other, and the reader can start with any chapter. However, someone new to the field of aerospace materials will benefit by starting with Chapter 2. Designers who have some familiarity with materials can go straight to Chapter 3 on materials selection in design. Crossreferences to other chapters are made where appropriate. The authors have attempted to give a historical perspective on materials development and applications in aircraft and aerospace applications, which tells us how much we have progressed in the last hundred years. It is my earnest hope that this book will help to heighten the awareness of materials technology for designing future aerospace systems.

ACKNOWLEDGMENTS The editor expresses his sincere thanks to the many individuals and their organizations who contributed to this book. They include the lead authors: Steven M. Arnold, Brian L. Wardle, Keith B. Bowman, Miria M. Finckenor, Grant Henson, Peter A. Curreri, David E. Glass, James C. Williams, Diep V. Trinh, John A. Halchak, Les Johnson, and Ajay K. Misra. They also include the contributing authors: Awadh B. Pandey, Sesh Tamirisakandala, Michael V. Nathal, David L. Ellis, Preston B. McGill, Jonathan A. Lee, Samuel E. Davis, Joseph H. Koo, Alan Nettles, Dongming Zhu, David Cebon, Mike Ashby, Gregory

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M. Odegard, Gary D. Seidel, Michael Mohaghegh, Antonio F. Avila, David L. Stone, Clyde S. Jones, James L. Cannon, Corey Brown, Tiffany Russell, Harold Gerrish, Dan Goebel, Hani Kamhawi, Mike Houts, Dave Edwards, Roy Young, Rob Hoyt, Sonny White, Mark Carter, James Polk, Richard Hofer, Yiangos Mikellides, Creon Levit, Eric Davis, Jeff George and Bob Frisbee. All have contributed their precious time and effort to make this book a success. Thanks are due to members of the American Institute of Aeronautics and Astronautics (AIAA) Materials Technical Committee (MTC), who served as members of the book subcommittee and played an active role in developing the book’s scope and table of contents. Special thanks are due to Leo Daniel, who served as co-chair of the book subcommittee. Several lead authors and contributors are members of the MTC. Other subcommittee members were Donald Jaworske, Ajit Kelkar, John Matlik, Rajiv Naik, Sai Raj, Samit Roy, Greg Schoeppner, Steven Wanthal, Michael Weimer, and Ann Whitaker, who provided valuable guidance. Gary Seidel helped to create a file- and message-sharing website for the book within the AIAA-MTC website to facilitate review and editing. The editor acknowledges the support of David Arthur and Katrina Buckley of AIAA and their staff in editing and publishing the book.

CHAPTER 2

Aerospace Materials Characteristics Biliyar N. Bhat NASA Marshall Space Flight Center

2.1 INTRODUCTION This chapter gives an overview of aerospace materials and their characteristics. It focuses on the most commonly used materials in aerospace structures, including aircraft, space craft, launch vehicles, and propulsion and power systems. The treatment is necessarily brief and serves as an introduction to different classes of materials and their characteristics in the context of aerospace applications. It is not intended to be exhaustive. Excellent information sources for these materials are available elsewhere and are referenced in each section. The reader should go to these references for more details. Aerospace materials can be broadly classified into four classes: metallic materials (metallics), nonmetallic or polymeric materials, composite materials (composites), and ceramic materials (ceramics). Examples from these classes of materials are given in this chapter. Historically, aircraft used the best materials available at the time they were built. The Wright brothers used aluminum alloys in their aircraft to make them lighter (compared with steel) so that they could become airborne more readily. Lightweight nonmetallic materials such as wood and fabric were also used. There has been a continuous improvement in aerospace materials of all classes over the last hundred years. Carbon fiberreinforced composites were introduced some 60 years ago, and their use has become more common because of their lighter weight and higher strength compared with other materials. The goal is to optimize the design by using materials that provide greater performance at lower cost. Metallic materials are the most commonly used materials in building the aerospace systems of today. They are covered in Secs. 2.2, Aluminum Alloys; 2.3, Titanium Alloys; 2.4, Steels; 2.5, Superalloys; and 2.6, Copper Alloys. Damage tolerance considerations are described in Sec. 2.7. Details on alloy development, properties, processing, and typical applications are presented. Relationships between properties, microstructure, and processing are also described with This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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aerospace applications in mind. Structural properties such as elastic modulus, tensile strength, ductility, and damage tolerance (fatigue and fracture) are emphasized because they are major considerations in design. Manufacturing technologies commonly used to fabricate metallic material components are also described in the context of design for manufacturing. Environmental effects on materials performance are described in Sec. 2.8 (hydrogen embrittlement) and Sec. 2.9 (oxygen compatibility). These two topics are of special interest because of their importance in propulsion systems that use hydrogen and oxygen as propellants. Polymers and composites are described next in Secs. 2.10 and 2.11. Polymers are organic compounds that are chemically based on carbon, hydrogen, and other nonmetallic elements (e.g., O, N, and Si). Some of the common polymers are polyethylene (PE), nylon or polyamide (PA), polyvinyl chloride (PVC), polycarbonate (PC), polystyrene (PS), silicone rubber, epoxy, and phenolic. Polymers are different from the other materials in many ways but generally possess lower densities, thermal conductivities, and moduli. The lower densities of polymeric materials offer an advantage for applications where light weight is a requirement. Carbon fiber-reinforced composites and their processing are also described. Section 2.11 focuses on the mechanical behavior of composites and their applications in launch vehicles. Ceramic materials and their characteristics are discussed in Sec. 2.12. The focus is on several important enabling ceramic systems for aerospace applications, particularly turbine engine thermal barrier coating systems and environmental barrier coating systems, and nonoxide-type SiC/SiC ceramic matrix composites. This section also emphasizes ceramic materials for various applications, material system properties, and durability performance associated with processing. Critical thermal and thermomechanical environment life design considerations, and laboratory and simulated operating environment tests are also discussed. Aerospace materials have advanced steadily in the last hundred years. They have become much stronger and lighter, as the reader will see in the following sections. More material choices are available to the designer today than at any time in the past. The selection of the best material for design is a key step in the design process, which is discussed in Chapter 3.

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2.2 ALUMINUM ALLOYS Awadh B. Pandey Pratt and Whitney Aircraft

2.2.1 INTRODUCTION There has been considerable use of aluminum alloys in aerospace applications at moderate temperatures (up to 3008F) due to their high specific strength (strength/density), durability, and damage tolerance. Aluminum alloys demonstrate very attractive mechanical properties, including strength, fatigue resistance, and fracture toughness. The mechanical properties are affected by alloy composition, processing, and heat treatment. Aluminum alloys have good corrosion resistance due to the formation of aluminum oxide on the surface. The primary use of high-strength aluminum alloys is in aircraft; the airframe of modern aircraft is typically 80% aluminum by weight. Aluminum alloys also make up a major portion of spacecrafts and launch vehicles. There are two excellent textbooks available on aluminum alloys: one by Polmear [1] and the other by Hatch [2]. In addition, Starke and Staley [3] have written a comprehensive overview on aluminum alloys for aerospace applications. This section provides an overview of different classes of commonly used aluminum alloys in aerospace applications along with a look at current trends and future developments, including high-temperature aluminum alloys and discontinuously reinforced aluminum.

2.2.2 ALUMINUM ALLOY CLASSIFICATION SYSTEMS There is a broad range of aluminum alloys, and therefore a classification system has been developed for them. Two major types of aluminum alloys are available depending on the processing used to produce the material: wrought alloys and cast alloys. Wrought alloys are produced by casting plus deformation, which may be in the form of extrusion, forging, or rolling. A number of wrought aluminum alloys are available depending on alloy composition, processing, and heat treatment. Because of this, wrought alloys have different designation systems than cast alloys. Cast alloys are produced by casting, and no deformation is imparted to the material or component. Therefore, cast aluminum alloys have limited strength and ductility. The major advantage of cast aluminum alloys is that they can be used for making intricate shaped parts, which are difficult to produce by wrought processing. Wrought aluminum alloys are normally designated by a four-digit number; there can be an additional letter and number to indicate temper and condition developed by the Aluminum Association. This classification has been adopted by different parts of the world and is known as the International Alloy  United

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Copyright # 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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TABLE 2.1

ALUMINUM ALLOY DESIGNATION SYSTEM FOR WROUGHT ALLOYS

Four-digit Series Alloys

Major Alloying Elements

1xxx series

99% aluminum

2xxx series

Copper (most also contain magnesium)

3xxx series

Manganese

4xxx series

Silicon

5xxx series

Magnesium

6xxx series

Magnesium and silicon

7xxx series

Zinc (most also contain magnesium and copper)

8xxx series

Others including lithium and iron

Designation System [3]. The designation system for wrought aluminum alloys is given in Table 2.1. For wrought alloys, the first digit indicates the major alloying element, the second digit indicates modification of the original alloy or impurity limits, and the last two digits indicate the specific aluminum alloy. For instance, the 2xxx series alloys contain copper as the main alloying element and may contain magnesium and manganese as additional alloying elements. The designation system for cast aluminum alloys is shown in Table 2.2 [3]. In this three-digit system, the first digit refers to the major alloying element and the next two digits denote a specific composition. The zero after the decimal point indicates casting, and other numerals indicate ingots. For instance, the 3xx.x series alloys contain silicon as the major alloying element with magnesium and/or copper as additional alloying elements. TABLE 2.2

ALUMINUM ALLOY DESIGNATION SYSTEM FOR CAST ALLOYS

Three-digit Series

Major Alloying Elements

1xx.0

99.00% minimum aluminum

2xx.0

Copper

3xx.0

Silicon with added copper or magnesium

4xx.0

Silicon

5xx.0

Magnesium

6xx.0

Unused

7xx.0

Zinc

8xx.0

Tin

9xx.0

Others

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A letter prefix is used to identify the impurity level or the presence of an additional alloying element. These letters are given in alphabetical order starting with A and ending in I, O, Q, or X. X is normally used for experimental alloys. The designation A201.0 is used for a higher purity version of 201.0, and A357.0 is used for a higher purity version of 357.0.

2.2.3 ALUMINUM ALLOY TEMPER DESIGNATION SYSTEMS The heat treatment or temper designation system for wrought aluminum alloys is given in Table 2.3 [3]. The nomenclature has been developed by the Aluminum Association. The letters are used as suffixes to the alloy number. As shown in the table, F indicates “as fabricated,” and O indicates the annealed condition of the alloy. The letter H indicates hardened condition by cold deformation, and the first digit following H indicates cold work and partially annealed/or stabilized; the second digit indicates how much deformation is given. The alloys that are solution treated but not aged are designated as W, and the alloys that are solution treated and aged are identified as T. Numbers following T identify the type of aging condition given to the alloy. T3 indicates solution treatment plus cold worked, T4 indicates solution plus natural aging, and T6 indicates solution plus artificial aging to achieve peak strength. T8 is used for solution plus cold working plus artificial aging. H is used for strain hardened alloys such as 3xxx and 5xxx series alloys. T3, T4, and T8 are usually applicable to 2xxx series alloys. The cold working in T8 temper introduces dislocations that act as sites for nucleation of precipitates that provide an increased strength. T6 is applicable to 6xxx, 2xxx, and 7xxx series alloys to indicate the highest strength conditions. T3 is used for 2xxx series alloys to indicate higher damage tolerance resistance. T7 is an over-aged condition with improved stress corrosion resistance and is applicable to only 7xxx series alloys. Deformation is often given to the material after quenching from solution treatment to relieve residual stresses, which could have deleterious effects on machining, fatigue, and stress corrosion cracking. The stress relieving can be performed either by stretching, which is denoted by number 1, or by compressing, which is denoted by number 2. Extrusions may be mechanically straightened after stretching, which is denoted by 1; if not mechanically straightened after stretching, then it is denoted by 0.

2.2.4 ALLOYING ELEMENTS AND HEAT TREATMENT Aluminum alloys are formulated by using combination of alloying elements as shown in Fig. 2.1 [2]. There are three main types of aluminum alloys depending on the presence of alloying elements: 1) age hardening (also known as precipitation hardening) alloys, 2) casting alloys, and 3) strain hardened alloys. Age hardening alloys consist of Al-Cu, Al-Cu-Mg, Al-Mg-Si, Al-Zn-Mg, and Al-ZnMg-Cu. Casting alloys consist of Al-Si, Al-Si-Cu, and Al-Si-Mg. Strain hardened

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TABLE 2.3 Suffix Letter F, O, H, T, or W Indicates Basic Treatment Condition

TEMPER DESIGNATION SYSTEM FOR ALUMINUM ALLOYS First Suffix Digit Indicates Secondary Treatment Condition to Influence Properties

Second Suffix Digit for Condition H Only Indicates Residual Hardening

F ¼ as fabricated O ¼ annealed, wrought products only H ¼ cold worked, strain hardened

1 ¼ cold worked only

2 ¼ 1/4 hard

2 ¼ cold worked and partially annealed

4 ¼ 1/2 hard

3 ¼ cold worked and stabilized

6 ¼ 3/4 hard 8 ¼ hard 9 ¼ extra hard

W ¼ solution heat treated T ¼ heat treated stable

1 ¼ partially solution þ natural aging 2 ¼ annealed cast products only 3 ¼ solution þ cold worked 4 ¼ solution þ natural aging 5 ¼ artificially aged only 6 ¼ solution þ artificial aging 7 ¼ solution þ stabilizing 8 ¼ solution þ cold work þ artificial aging 9 ¼ solution þ artificial aging þ cold work

alloys include Al-Mg and Al-Mn alloys. Age hardening alloys are strengthened by precipitates, which are produced by heat treatment including solution and aging treatments. The 2xxx, 6xxx, and 7xxx series alloys are age hardening alloys. Casting alloys are also strengthened by precipitates produced by heat treatment. Casting alloys include 2xx.x and 3xx.x series alloys. Strain hardened alloys are non-heat treatable and strengthened by dislocations introduced through cold working.

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Precipitation hardening is a process whereby a fine precipitate structure is formed in the alloy matrix following a heat treatment process (Figs. 2.2 and 2.3). For an alloy to be precipitation hardened, it requires: 1) decreased solubility with decreasing temperature, 2) the ability to suppress the formation of precipitates by quenching from a solid solution, and 3) the formation of metastable coherent precipitates. The precipitation hardening process consists of the following three steps: 1. Solution treatment: The alloy is heated above the solvus temperature to dissolve all precipitates and ensure that the alloying elements are completely in solid solution. 2. Quench: The alloy is quenched in water so that alloying elements do not have time to diffuse and form precipitates. Thus, the alloying elements remain in solution forming a supersaturated solid solution. 3. Aging: The alloy is heated to an intermediate temperature below the solvus temperature. The alloying elements are able to diffuse to form extremely fine coherent precipitate clusters known as GP (Guinier-Preston) zones. The coherent precipitates increase the strength of the alloy by distorting the crystal lattice and creating resistance to dislocation motion. The number of precipitates increase with increasing aging time, thus increasing the strength of the alloy. However, with increasing aging time, the precipitates become large and incoherent and their strengthening effect decreases. Thus, the following strengthening mechanisms are operative during precipitation hardening: 1) solid solution strengthening in the supersaturated solid solution, 2) coherency stress hardening from the coherent precipitates, 3) precipitation hardening by resistance to dislocation motion, and 4) hardening through resistance to dislocation between precipitates.

Fig. 2.1

Principal aluminum alloys [2].

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Fig. 2.2

Al-Cu phase diagram with markings for heat treatment temperatures [2].

Fig. 2.3 Al-Cu phase diagram showing microstructures resulting from different steps of the precipitation hardening process [4].

AEROSPACE MATERIALS CHARACTERISTICS

19

Table 2.4 shows the precipitation sequence in different precipitation hardening alloys. It shows that Al2Cu (u 0 , theta prime) is major strengthening phase in Al-Cu alloys. Al2CuMg S0 is the main strengthening precipitate in Al-Cu-Mg (2xxx series) alloys. Mg2Si (b0 , beta prime) is mainly responsible for strengthening in Al-Mg-Si (6xxx series) alloys. MgZn2 (h0 , eta prime) is the major strengthening precipitate in Al-Zn-Mg (7xxx series) alloys. In Al-Li alloys, Al3Li (d0 , delta prime) is the main precipitate. In Al-Li-Cu-Mg alloys, S0 , T1, and d0 contribute to strengthening. It should be noted that d0 cannot usually be suppressed during the quench from solution treatment temperature. Further precipitation of d0 occurs during the GP zone formation and normal precipitation sequence of the Al-Li-Cu and Al-Li-Cu-Mg alloys.

2.2.5 STRENGTHENING MECHANISMS Aluminum alloys can be strengthened by a number of mechanisms: solid solution strengthening, grain size strengthening, work (strain) hardening, precipitation hardening, and dispersion hardening. Each mechanism is described briefly in the following sections. 2.2.5.1

SOLID SOLUTION STRENGTHENING

Solid solution strengthening is derived from the presence of alloying elements in aluminum [5]. Two types of solid solution strengthening can be produced in aluminum depending on the presence of alloying elements: substitutional solid TABLE 2.4 Alloys

SOLID-STATE PRECIPITATION SEQUENCE FOR VARIOUS ALUMINUM ALLOYS Precipitation Reactions

Al-Cu

Supersaturated solid solution ! GP Zones ! u 00 ! u 0 ! u (equilibrium) Al2Cu

Al-Cu-Mg

Supersaturated solid solution ! GP Zones ! S0 Al2CuMg ! S (equilibrium) Al2CuMg

Al-Mg-Si

Supersaturated solid solution ! GP Zones ! b0 Mg2Si

Al-Li

d0 þ Supersaturated solid solution ! d0 Al3Li ! d AlLi

Al-Li-Cu

d0 þ Supersaturated solid solution ! GP Zones ! u 00 ! u 0 ! u ! T1 (Al2CuLi)

Al-Li-Mg

d0 þ Supersaturated solid solution ! d0 Al3Li ! Al2LiMg

Al-Li-Cu-Mg

d0 þ supersaturated solid solution ! GP Zones ! S00 ! S0 þ d0 ! S þ d

Al-Zn-Mg

Supersaturated solid solution ! GP Zones ! h0 (hexagonal MgZn2) ! h MgZn2

20

B. N. BHAT

solution and interstitial solid solution. Solid solution strengthening depends on the misfit strain that is caused by a change in the lattice parameter of the aluminum matrix due to the presence of alloying elements. 2.2.5.2

GRAIN SIZE STRENGTHENING

Grain size strengthening is a phenomenon where strength increases with a decrease in grain size (Hall-Petch model [5]). When a stress is applied to the material, dislocations tend to align at the grain boundaries, leading to pileup of dislocations. The finer grains have more grain boundaries; therefore, dislocation pileup is higher in finer grain materials than in coarse grain materials. The dislocation pileup raises the amount of stress required for deformation, which results in increased strengthening in finer grain materials. 2.2.5.3

WORK (STRAIN) HARDENING

Strain hardening (usually through cold working the material) can provide strengthening by increasing dislocation density. As dislocation density increases, the shear stress required to overcome the dislocation barrier increases. Strain hardening is applicable for the non-heat-treatable aluminum alloys, 3xxx and 5xxx series alloys. Substantial strengthening in Al-Mn and Al-Mg based alloys can be produced by cold working. 2.2.5.4

PRECIPITATION HARDENING

Strengthening during the precipitation hardening process is illustrated in Fig. 2.4, which shows the shearing and bypassing of precipitates by dislocations. In the underaged (UA) condition, the precipitates are small and coherent with the aluminum matrix. Dislocations cut these precipitates during deformation of the material, which provides strengthening in the underaged condition. The size of precipitates increases with an increase in aging time during the precipitation hardening process. This process continues until the precipitates attain a critical size where coherency is lost and they become incoherent with the aluminum matrix. The aluminum alloys achieve the highest strength in the peak aged (PA) condition. In the overaged (OA) condition, the size of the precipitates increases with an increase in the aging time. The size of the precipitates is so large that dislocations have to go around (bypass) these precipitates, contributing to strengthening in the OA condition. This mechanism is also known as Orowan strengthening. 2.2.5.5

DISPERSION HARDENING

Dispersion hardening can provide substantial strengthening in aluminum alloys from individual dislocation-particle interaction. The microstructure of dispersion-strengthened materials usually contains dispersed particles within

AEROSPACE MATERIALS CHARACTERISTICS

Shearing up to peak aged condition

Fig. 2.4

21

Bypassing in overage condition (Orowan)

Shearing and bypassing of precipitates by dislocation [1].

grains and on the grain boundaries. When dislocations interact with particles within grains, they provide strengthening from dislocation looping. When dislocations move through material with particles, they leave loops around these particles. These dislocation loops exert back stress, providing strengthening in aluminum alloys. Orowan strengthening is applicable to dispersion-strengthened materials. For a given alloy composition, the volume fraction of dispersoids or precipitates will be fixed, which means the particle size must be reduced to reduce the interparticle spacing. Most dispersion-hardened materials are nonheat treatable. For example, the Al-Al2O3 system is a dispersion-strengthened alloy system where considerable strengthening is achieved by the Orowan mechanism.

2.2.6 PROPERTIES, MICROSTRUCTURE, AND PROCESSING Figure 2.5 shows the variation of fracture toughness with tensile yield strength for different heat treatment conditions for 6061 and 2014 alloys [2]. Fracture toughness is the highest in the UA condition and the lowest in the PA condition. It starts increasing again in the OA condition but does not recover completely for the same strength level observed in the UA condition. The variation of fracture toughness is similar to that of ductility in different aging conditions. The 2xxx series alloys are usually stretched after solution treatment to improve strength. Figure 2.6 shows the effect of stretching on the fracture toughness–strength relationship. The fracture toughness decreases considerably with stretching, whereas strength increases. This result is due to an increase in the dislocation density from stretching because

22

B. N. BHAT

dislocations act as nucleation sites for precipitation. The 7xxx series alloys are often used in T7 condition (stabilization by OA) to improve stress corrosion resistance where strength is lower than that in the PA condition. Therefore, recent efforts have focused on improving the strength of these alloys by using stretching before aging, which provides higher strength while maintaining stress corrosion resistance. Fatigue crack growth rate, da/dn vs DK, is shown for 7075-T6 and 2024-T3 in Fig. 2.7 [2]. There are three regimes in crack growth rate curves: I, the threshold regime where there is no crack growth until a threshold stress intensity factor, K, is applied; II, the Paris regime where crack growth is controlled by the power law (da/dn ¼ A DKm); and III, where the crack grows very rapidly, leading to failure. The 2024-T3 alloy has higher crack growth resistance compared to the 7075-T6 alloy in all three regimes. Generally, 2xxx series alloys have better damage tolerance resistance than do 7xxx series alloys. The fatigue strength generally decreases with an increase in the cycles to failure. The fatigue strength depends on the tensile strength of the material, showing higher values for higher tensile strength conditions. Fatigue failure

Fig. 2.5 Variation of fracture toughness with tensile yield strength for 6061 and 2014 alloys [2].

AEROSPACE MATERIALS CHARACTERISTICS

23

Fig. 2.6 Effect of stretching on the fracture toughness–yield strength relationship for 2xxx series alloy [2]. occurs by nucleation, growth, and coalescence of voids where nucleation of the cracks controls the majority of the life. Aluminum alloys do not have an endurance limit; therefore the strength at 107 cycles is considered as the fatigue strength of the materials. The variation of the fatigue endurance limit with tensile strength is shown in Fig. 2.8 for several aluminum alloys. The plot shows a good (positive) correlation between fatigue strength and tensile strength for aluminum alloys.

2.2.7 PROPERTIES OF WROUGHT HEAT-TREATABLE ALUMINUM ALLOYS Wrought heat-treatable aluminum alloys include 2xxx, 6xxx, 7xxx, and 8xxx alloys. The most commonly used aluminum alloys in aircraft applications are 2xxx and 7xxx series alloys. The nominal composition of common alloys is given in Table 2.5 [3]. The 2xxx series aluminum alloys are alloyed with Cu, Mg, and Mn. This series of alloys provide precipitation hardening through formation of S0 (Al2CuMg) for

24

B. N. BHAT

Fig. 2.7

Fatigue crack growth rate da/dn vs DK plot for 2024-T3 and 7075-T6 alloys [2].

Fig. 2.8

Variation of fatigue endurance limit with tensile strength for aluminum alloys [2].

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.5

25

NOMINAL COMPOSITION OF SELECTED AEROSPACE ALUMINUM ALLOYS

Alloy

Compositions in wt %

2004

Al-6.0 Cu-0.4 Zr

2014

Al-4.4Cu-0.5Mg-0.8Mn-0.7Fe-0.5Si

2017

Al-4.0Cu-0.6Mg-0.7Mn-0.7Fe-0.5Si

2024

4.4Cu-1.5Mg-0.6Mn-0.5Fe-0.5Si

2219

Al-6.3Cu-0.3Mn-0.2Zr-0.3Fe-0.2Si

2224

Al-4.1Cu-1.5Mg-0.6Mn-0.15Fe-0.12Si

2324

Al-4.1Cu-1.5Mg-0.6Mn-0.12Fe-0.10Si

2519

Al-5.8Cu-0.2Mg-0.2Zr-0.3Fe-0.2Si

2618

Al-2.3Cu-1.5Mg-0.2 Mn-1.2Ni-0.1Zn-0.2(Ti þ Zr)-1.2Fe-0.25Si

6013

Al-1.0Mg-0.8Si-0.35 Mn-0.3Fe-0.8Si

6113

Al-1.0Mg-0.8Si-0.35 Mn-0.3Fe-0.8Si-0.2O

7010

Al-6.2Zn-2.35Mg-1.7Cu-0.3Fe-0.2O

7049

Al-7.7Zn-2.45Mg-1.6Cu-0.15Cr-0.35Fe-0.25Si

7050

Al-6.2Zn-2.25Mg-2.3Cu-0.1Cr-0.15Fe-0.12Si

7055

Al-8.0Zn-2.05Mg-2.3Cu-0.1Zr-0.15Fe-0.1Si

7075

Al-5.6Zn-2.5Mg-1.6Cu-0.25Cr-0.4Fe-0.4Si

7079

Al-4.3Zn-3.2Mg-0.6Cu-0.2Mn-0.15Cr-0.4Fe-0.3Si

7093

Al-9.0Zn-2.5Mg-1.5Cu0.1Zr-0.15Fe-0.12Si-0.2O

7150

Al-6.4Zn-2.35Mg-2.2Cu-0.1Zr-0.15Fe-0.1Si

7178

Al-6.8Zn-2.8Mg-2.0Cu-0.23Cr-0.5Fe-0.4Si

7475

Al-5.7Zn-2.25Mg-1.6Cu-0.21Cr-0.12Fe-0.10Si

7085

Al-7.5Zn-1.5Mg-1.7Cu-0.08Zr

7068

Al-7.8Zn-2.6Mg-2.0Cu-0.1Mn-0.05Cr-0.1Ti-0.1Zr

higher Mg containing alloys and precipitation of u 0 (Al2Cu) for higher Cu to Mg ratio alloys upon heat treatment. They also contain Zr, Cr, Mn, or Ti to control the grain size. These alloys have lower crack growth rates and thus better fatigue resistance than 7xxx series alloys. The cladding on 2017-T4 and 2024-T3 alloys consists of commercially pure aluminum metallurgically bonded to either one or both surfaces of the sheet. The 2324-T39 and 2224-T3 alloys were developed by modifying the composition and processing of standard 2024 alloy. The amount of cold work applied after quenching from solution and before aging

26

B. N. BHAT

Fig. 2.9

Aircraft internal structure with extrusions and plates of 2xxx alloys [6].

was increased from 1%–3% (for 2024-T351 plate) to about 9%. Processing conditions were also modified for extrusions to retain the deformation crystallographic texture for additional texture strengthening. Alclad 2xxx-T3 alloy (also known as C188 by Alcoa) was developed to provide a combination of strength and fracture toughness to meet the requirements of fuselage skin application. The fatigue crack growth resistance of this alloy is almost two times better than that of 2024-T3 sheet at high level of peak stress pffiffiffiffi intensity factor (greater than 22 MPa m). The chemical composition and processing of the alloy were used to control intermetallic particles to provide higher fracture toughness and fatigue crack growth resistance. The 2219 and 2618 alloys have superior high-temperature capability compared to other commercial aluminum alloys. The 2219 alloy has a higher Cu to Mg ratio, which forms u 0 (Al2Cu) precipitate that improves high-temperature capability. The 2618 alloy contains Fe, Ni, and Mn, which forms thermally stable dispersoids such as Al3Fe, Al6Mn, Al3Ni, and Al9FeNi that provide high-temperature capability. Figure 2.9 shows an aircraft integral structure that includes extrusions and plate of 2xxx series alloys such as 2024, 2124, and 2618. The upper and lower wing structures of the Boeing 757 and 767 are manufactured with improved alloys compared with Boeing 747. The 7xxx series alloys containing Zn and Mg additions offer the greatest potential for precipitation hardening through precipitation of h0 (MgZn2) phase. Cu is added in the 7xxx series alloys to improve stress corrosion cracking resistance. Stress corrosion cracking resistance decreases with increasing Zn:Mg ratio. Most aluminum alloys contain a small amount of Zr, Cr, or Mn to control grain growth by forming fine dispersoids on grain boundaries.

AEROSPACE MATERIALS CHARACTERISTICS

27

Figure 2.10 shows the use of advanced aluminum alloys in aircraft over time [7]. The 7075-T6 alloy, which was introduced in 1943 and first used on the U.S. Navy’s P2V patrol bomber, is one of the oldest alloys. Subsequently, a number of 7xxx series aluminum alloys were developed with improved mechanical properties. The X7080-T7 alloy has higher stress corrosion resistance than the 7079-T6 alloy. It was developed for thick parts because it is relatively insensitive to quench rate and can provide higher strength, even for thicker parts. The 7050-T74 alloy was developed for thick section applications and has been used extensively in aircraft applications. It has high strength, good stress corrosion cracking resistance, and good fracture toughness and fatigue resistance. The 7050 alloy contains Zr instead of Cr (which is present in the 7075 alloy) to control grain size through formation of Al3Zr dispersoids, which help in achieving higher toughness. A slightly higher amount of Cu and slight modification in the Zn:Mg ratio improves the strength and stress corrosion resistance of the 7050 alloy compared with the 7075 alloy. Therefore, 7050-T74 plates and forgings are the standard material for thick section parts in aircraft applications. The 7150-T6 alloy plates and extrusions were used as wing skins and stringers in the upper wing structures. Figure 2.11 shows critical aircraft wing structures That are made of 7xxx series alloy sheet or integrally stiffened extrusion construction. The 7150-T61 plate and extrusions are used on the McDonnell Douglas MD-11. The 7150-T77 alloy was developed with a more balanced combination of strength, stress corrosion cracking resistance, and fracture toughness. The higher fracture toughness of 7150-T77 was related to the controlled volume fraction of intermetallic particles and the unrecrystallized grain structure. The higher

Fig. 2.10

Advancement of aluminum alloys in aircraft [7].

28

B. N. BHAT

Fig. 2.11 Critical aircraft wing structures made of 7xxx series alloy sheet or integrally stiffened extrusion construction [6].

strength and corrosion resistance of the 7150-T77 alloy was attributed to the size and distribution of strengthening precipitates. These precipitates reduce stress concentration at the grain boundaries through homogenizing deformation and reduce electrochemical differences between the matrix and grain boundaries. The 7055-T77 alloy has fracture toughness and crack growth resistance similar to that of the 7150-T6 alloy and stress corrosion cracking resistance intermediate to that of 7075-T6 and 7150-T77 alloys. The intermetallics in the 7055-T77 alloy were minimized to provide higher fracture toughness. The ratio of Zn:Mg and Cu:Mg was maintained at high level to provide an excellent combination of strength, corrosion resistance, and fatigue strength and fracture toughness. The 7055 alloy has a strict restriction on solute content and thermomechanical processing to produce a material that has higher strength, fracture toughness, and fatigue resistance than the 7178-T6 alloy along with improved resistance to stress corrosion cracking and exfoliation. Considerable advances have taken place in the development of newer 7xxx series alloys, including 7090, 7091, and CW67, using a powder metallurgy (P/M) approach to improve the strength, toughness, and stress corrosion cracking resistance. This approach uses rapidly solidified powder produced by gas atomization, followed by compaction and extrusion of the powder. These alloys contain Co, Ni, and Zr, which form very fine dispersoids of Al9Co2, Al3Ni, and Al3Zr depending on the composition. The grain size of P/M alloys is finer than that of ingot-based alloy, which provides additional strengthening. However, development of P/M aluminum alloys was not pursued vigorously due to higher cost and powder contamination issues.

AEROSPACE MATERIALS CHARACTERISTICS

29

2.2.8 PROPERTIES OF CAST ALUMINUM ALLOYS A number of cast aluminum alloys are available for aerospace applications (see Table 2.6). The most commonly used cast aluminum alloys (300 series) contain Si as a major alloying element to provide fluidity, which is essential for producing sound castings. Si precipitates as free Si during the aging of Al-Si alloys. Because Si has limited solubility in aluminum, it does not provide substantial strengthening. Therefore, Mg and Cu are added to provide strength through precipitation strengthening. Some of the casting alloys contain Ti and Mn to control grain growth through formation of intermetallic particles at grain boundaries. The 355 and 356 alloys are commonly used in aerospace applications. The 357 alloy has slightly higher strength compared to the 355 and 356 alloys due to the presence of beryllium (Be) and is used in space applications. Typically cast alloys containing 7% Si are preferred because they provide good strength with acceptable ductility. Al-Cu based cast alloys (200 series) have higher strength compared to Al-Si alloys. Usually, cast alloys have limited ductility due to lack of deformation processing. However, casting is generally more economical and sometimes the only way to produce complex shaped products. Cast aluminum alloy components are produced by different casting methods: sand casting, investment casting, die casting, and permanent mold casting. It is an economical process to make good-quality aluminum alloy components even in small numbers. Investment casting is a more expensive process; it has a high tooling cost and also a high labor cost. Investment casting is used for producing very precise parts. Die casting is used for high-precision, complex geometry parts with thin walls. Die casting is suitable for relatively small parts compared with sand casting. Sand casting is more economical for a small number of parts, whereas die casting is more economical for a large number of parts. Die casting is very efficient as it can produce a large number of parts in short time. Permanent

TABLE 2.6 Alloy

NOMINAL COMPOSITION OF SELECTED CAST ALUMINUM ALLOYS Composition in wt %

201

Al-4.6Cu-0.35Mg-0.35Mn-0.2Ti

213

Al-7.0Cu-2Si-2.5zn-0.6Mn-0.1Mg

355

Al-5Si-0.5Mg-0.5Mn-1.2Cu-0.35Zn-0.25Ti

356

Al-7.0Si-0.25Cu-0.3Mg-0.35Mn-0.35Zn-0.25Ti-0.6Fe

357

Al-7.0Si-0.55Mg-0.12Ti-0.055Be

360

Al-9.5Si-0.6Fe-0.35Mn-0.15Mg-0.5Ni-0.5Zn

413

Al-12Si-2Fe-1Cu-0.5Ni-0.5zn-0.35Zn

518

Al-8Mg-1.8Fe-0.35Mn-0.25Cu

30

B. N. BHAT

Fig. 2.12

Thixoformed A356.0-T6 inner turbo frame for Airbus aircraft [6].

mold casting uses metal as split mold instead of expendable materials. The casting process is selected depending on the size, shape, and number of parts to be produced for aluminum alloys. The philosophy of most aircraft manufacturers has been to use cast aluminum alloys only in the structures where failure of parts cannot cause loss of the aircraft. Figure 2.12 shows the thixoformed A356.0-T6 inner turbo frame for the Airbus family of aircraft.

2.2.9 LOW-DENSITY, HIGH-MODULUS Al-Li ALLOYS Aluminum–lithium (Al-Li) alloys offer a significant opportunity for weight savings due to their lower densities and higher moduli. Each 1 wt % of Li added in aluminum reduces alloy density by about 3% and increases modulus by about 6%. Li has high solid solubility (4 wt.% at 6108C, or 11208F) in aluminum and responds to age hardening, due to precipitation of an ordered metastable phase d0 (Al3Li) that is coherent and small misfit with the matrix. Because of these characteristics, Al-Li alloys attracted considerable attention for the development of a new generation of low-density, high-modulus alloys for aerospace applications. Specifically, Al-Li alloy 2195 has been used for cryogenic propellant tanks in the space shuttle. Table 2.7 lists the compositions and properties of selected Al-Li based alloys. Most of the effort has been concentrated on adding Li to 2xxx series alloys containing Cu, Mg, and Zr. Initial alloy development efforts included the addition of high percentages of Li (2 to 3 wt.%) in alloys such as 2090, 2091, 8090, and 8091 to provide maximum benefit from density reduction. The Al-Li-Cu-Mg

Alloy

2091 8090

NOMINAL COMPOSITION AND TENSILE PROPERTIES OF SELECTED Al-Li based alloys [8, 9]

Composition, wt % Al-2Li-2.1Cu-1.5Mg-0.1Zr Al-2.4Li-1.3Cu-0.9Mg-0.16Zr

Temper

0.2% Yield Strength, MPa

Ultimate Tensile Strength, MPa

Elongation, %

Fracture Toughness, MPa m0.5

T8X (UA)

370

460

15

40

T851

475

525

9

25

T81 (UA)

360

445

11

45

T6

400

470

6

35

T851

455

510

7

30

2090

Al-2.3Li-2.7Cu-0.3Mg-0.5Zr

T83

510

565

5

—

8091

Al-2.6Li-2Cu-0.85Mg-0.16Zr

T851

515

555

6

22

1421

Al-5Mg-2Li-0.2Mn-0.2Sc-0.1Zr

T8

330

470

10

65

Weldalite 049

Al-6.3Cu-1.3Li-0.4Mg-0.4Ag-0.18Zr

T8

725

797

9.8

—

2099

Al-2.7Cu-1.8Li-0.7Zn-0.3Mg-0.3Mn0.09Zr-0.07Fe-0.05Si

T83

490

545

6

—

T8E79

345

400

8

42

2199

Al-2.6Cu-1.6Li-0.55Zn-0.23Mg-0.3Mn0.09Zr-0.07Fe-0.05Si

380

428

8

42

Al-4.0Cu-1.0Li-0.6Mg-0.25Mn-0.4Ag0.12Zr-0.15 max Fe-0.12max Si

T84

530

556

9

37

2050

Al-3.55Cu-1.0Li-0.4Mg-0.35Mn-0.25 max Zn-0.45Ag-0.1Zr-0.1Fe max 0-0.08 max Si

T84

476

503

8

36 31

T8E80 2195

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.7

32

Fig. 2.13

B. N. BHAT

Fatigue crack growth rate behavior of 2090 alloy vs 2124 and 7150 alloys [10].

alloys are strengthened by three types of precipitates: d0 (Al3Li), T1 (Al2CuLi), and S0 (Al2CuMg). Most of the Al-Li alloys contain a small amount of Zr, similar to other aluminum alloys. The role of Zr is to twofold: Al3Zr controls recrystallization and grain growth, and Al3Zr particles have a similar L12 structure as d0 , substituting the Al3Li precipitates forming Al3(Li, Zr). The 2091, 8090, and 8091 alloys contain S0 precipitate, which is resistant to shearing by dislocations, promoting more homogeneous deformation. These alloys provide improved precipitation in the T8 condition. Figure 2.13 indicates that Al-Li alloys have superior crack growth resistance compared to 2xxx and 7xxx series alloys [10]. None of these high-Li alloys showed widespread applications due to property anisotropy, low toughness, and poor corrosion resistance. The chemical composition of Weldalite was modified to reduce Cu and Li (2195) to provide a more balanced combination of strength, fracture toughness, corrosion resistance, and fatigue crack growth resistance. The 2195 alloy was developed to replace the 2219 alloy for space shuttle hydrogen tank application. The 2195 alloy was first applied in NASA’s mission STS-91, which required significant development and characterization of the material in different processing conditions. The 2195 alloy today shows excellent strength, fracture toughness, and corrosion resistance. The 2195 alloy was produced by Alcan (Constellium), and the space shuttle external tank was produced by Lockheed Martin (Fig. 2.14).

AEROSPACE MATERIALS CHARACTERISTICS

33

The 2050 alloy has received significant attention due to its attractive properties for medium and thick sections, where it outperforms 2024 or 2027 alloys in strength, fracture toughness, fatigue, and corrosion resistance in addition to density and modulus. For higher thicknesses, the 2050 alloy offers a low-density alternative to the 7050 alloy. Compared to 7050-T74, the 2050-T4 alloy shows a better strength-toughness combination at 5% lower density and improved stress corrosion resistance. Compared to the 2024-T351 alloy, 2050-T4 shows significantly higher strength and corrosion resistance in addition to lower density. Table 2.7 shows the composition and tensile properties of 2099 and 2199 alloys. Figure 2.15 is a schematic representation of precipitates and dispersoids that contribute to strength and toughness in 2099 and 2199 alloys. Although strengthening in 2099 and 2199 alloys is achieved through T1, d0 , and u 0 precipitates, the contribution of the T1 phase is the highest. The T1 phase is strongly affected by stretching imparted to the alloy after quenching from solution and before aging. It has been shown that as stretching increases, the strengthtoughness relationship improves in Al-Li alloys. Stretching increases the volume fraction and reduces the size of strengthening precipitates, and hence improves the strength-toughness relationship. It is preferred to use the T8 condition instead of peak-aged T6 condition for Al-Li alloys because T8 condition provides a higher strength and toughness combination. The processing of 2099 and 2199 alloys is performed in such a way they have an unrecrystallized microstructure to provide higher fracture toughness. The 2199 plate with thickness 0.5–1.5-in. in T8E79 or T8E80 conditions has better properties than the 2024-T351 plate,

Fig. 2.14

Fuel tank of the space shuttle fabricated from Al-Li based 2195 alloy [6].

34

B. N. BHAT

Fig. 2.15 Precipitates and dispersoids that contribute to strength and toughness of 2099 and 2199 alloys [8]. which is used in lower skin wing applications for Bombardier. The 2099-T83 extrusions are commercially available in the thickness range 0.05–0.3-in., which is used on the Airbus A380 in fuselage and floor applications. Russian Al-Li based alloys 1441 and 1421 (Table 2.7) have attractive mechanical properties. The 1441 (Al-Cu-Mg-Li) alloy is used in fuselage applications for the Russian B-103 aircraft. It is cold-rollable and has several attributes that make it attractive for fuselage skin applications, such as lower density and higher specific modulus with similar strength as compared to conventional Al-Cu-Mg alloys. Another Russian alloy, 1421 (Al-Li-Mg-Sc-Zr), does not contain Cu but has lower density, excellent weldability, and superior corrosion resistance. This alloy has been used in a number of applications, including fuel tanks, fuselage stringers, cockpits, and other aircraft parts.

2.2.10

HIGH-TEMPERATURE ALUMINUM ALLOYS

Considerable work has been done recently in developing aluminum alloys for higher temperature applications, up to 6008F [11–15]. Among these alloys, Al-Fe-V-Si, Al-Fe-Ce, Al-Fe-Ce-W, Al-Fe-Mo-V, and Al-Cr-Zr-Mn are the most notable. A few selected alloys are listed in Table 2.8. In addition, Al-Zr-V and Al-Ti were also investigated for such applications. Most of these alloys have incoherent dispersoids with volume fractions of 15%–25%. The tensile strength of these alloys degrades considerably at higher temperatures due to coarsening p of dispersoids. The fracture toughness of most of these alloys is low, 8–15 MPa m, except for Al-Fe-V-Si, which had a toughness of

Alloy Composition, wt % Al-8Fe-4Ce Al-8Fe-2Mo-1V

Test Temperature, C

Al-8Fe-1.4V-1.7Si Al-5Cr-2Zr

Yield Strength, MPa

Ultimate Tensile Strength, MPa

Elongation, %

Fracture Toughness (K1Ca), MPa m0.5

25

418.9

484.9

7

8.5

316

178.1

193.8

7.6

7.9

25

323.5

406.6

6.7

9

170

187.5

7.2

8.1

25

464.1

524.5

4

5.7

316

206.3

240

6.9

6.1

316 Al-10.5Fe-2.5V

TENSILE PROPERTIES OF HIGH-TEMPERATURE ALUMINUM ALLOYS

25

362.5

418.8

6

36.4

316

184.4

193.8

8

14.9

326

351

18

11.2b

25

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.8

a

Computed K1C from J1C data. KQ measured.

b

35

36

B. N. BHAT

p 36.4 MPa m. The higher toughness of Al-Fe-V-Si is due to delamination at stringers of intermetallic particles. Figure 2.16 shows the variation of fracture toughness with yield strength for the Al-Fe-V-Si alloy. The fracture toughness decreases sharply with an increase in the yield strength of the alloy, similar to other aluminum alloys. Two-phase alloys have a higher fracture toughness and strength combination than three-phase alloys.

2.2.11

DISCONTINUOUSLY REINFORCED ALUMINUM (DRA)

Discontinuously reinforced aluminum (DRA) offers superior mechanical properties, including higher modulus, tensile strength, fatigue strength, and creep resistance, compared with unreinforced aluminum alloys [16–20]. However, the ductility and the fracture toughness of DRA are lower than that of unreinforced matrix alloys. The properties of DRA can be tailored to meet the requirements for the target applications. DRA consists of aluminum alloys as a matrix and SiC, B4C, TiB2, TiC, and Al2O3 particulates as reinforcement to provide improved properties. The properties of DRA depend on three constituents: matrix alloys, reinforcements, and matrix-reinforcement interface. The matrix alloys are typically 6061, 2124, 7075, 7093, or A360. The selection of the matrix alloy depends on its compatibility with ceramic reinforcements for the processing techniques used to produce DRA materials. In addition, a good interface between matrix and reinforcement is necessary for improved mechanical properties. DRA can be produced either by a powder metallurgy process or by a casting process. Powder metallurgy process includes the conventional powder process,

Fig. 2.16

Variation of fracture toughness with yield strength for Al-Fe-V-Si alloy [11].

AEROSPACE MATERIALS CHARACTERISTICS

37

reaction dispersion (XD) method, and spray deposition (Osprey). In the conventional powder metallurgy process, a matrix alloy and reinforcement powder are blended together and compacted to produce dense billet. In the XD process, the reinforcement is created in situ by reacting Ti with B to form TiB2 reinforcement. In spray deposition, matrix and reinforcement powders are mixed in the atomization zone and the mixture is deposited on the substrate, which solidifies to form dense billet. The casting process includes a simple mixing method where reinforcement is mixed in molten liquid and poured in the mold to solidify. This process is suitable for a relatively small-volume fraction of reinforcements. It is hard to keep uniform distribution of reinforcements in molten metal due to density differences. Most ceramic reinforcements have higher density compared to the aluminum alloys matrix. In the pressure infiltration casting method, preform is produced with ceramic particles and molten metal is forced into the preform by applying pressure. This method is applicable to composites having a relatively high-volume fraction of reinforcement, which provides a higher modulus and lower thermal expansion coefficient. The powder metallurgy process can provide more uniform distribution of reinforcements and better mechanical properties than casting. Two powder-based materials, 6092 Al/ 17%SiCp and 2009 Al/20%SiC, are used in commercial applications [21]. The 6092/17% SiCp is used for ventral fin and fan exit guide vane applications, and 2009/15% SiC is used for rotating components in European helicopters. The elastic moduli of DRA materials are compared with those of Al and Ti alloys in Fig. 2.17. The modulus of DRA with 15 vol. % B4C is considerably higher than those of aluminum alloys as reinforced particles provide higher modulus through load transfer. The modulus depends on the volume fraction

Fig. 2.17

Elastic moduli of aluminum, DRA, and titanium alloys [21].

38

B. N. BHAT

Fig. 2.18

Yield strength of aluminum composites and aluminum alloys [21].

of reinforced particles. The rule of mixture can be used to predict the modulus of the DRA materials [19, 20]. Figure 2.18 shows the yield strength of 2009/15% SiC, 6092/17.5% SiC, 6092/ 25% SiC, and 7xxx/15 vol. % SiC composites along with unreinforced aluminum alloys. DRA materials have significantly higher yield strength than aluminum alloys and generally follow the matrix yield strength [16–19]. The ceramic reinforcement provides higher yield strength in the aluminum matrix through composite strengthening by load transfer from high modulus reinforcement to the matrix aluminum alloy. The ductility of DRA is generally reduced due to the presence of ceramic reinforcement. The ductility depends on the size, volume fraction, and distribution of the reinforcement. The size and volume fraction of the reinforcement are selected in such a way that they provide a balanced combination of strength and ductility required for the application.

ACKNOWLEDGMENTS The author would like to thank Mike Maloney, Chris Rhemer, Dave Furrer, and Frank Preli from Pratt and Whitney Aircraft for their support and approval to publish this chapter. The author would also like to thank Dan Miracle, Kevin Kendig, and Jonathan Spowart from the U.S. Air Force Research Laboratory; Biliyar Bhat from NASA Marshall Space Flight Center; and Mark van den Bergh from DWA Aluminum Composites for useful discussions.

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39

REFERENCES [1] Polmear, I. J., “Light Alloys,” Metallurgy of the Light Metals, Edward Arnold, London, 1989. [2] Hatch, J. E., Aluminum: Properties and Physical Metallurgy, ASM International, Materials Park, OH, 1984. [3] Starke, E. A., Jr., and Staley, J. T., “Application of Modern Aluminum Alloys to Spacecraft,” Progress in Aerospace Sciences, Vol. 32, No. 2, 1996, pp. 131–172. [4] Aluminum and its alloys used in Aircraft. [5] Hertzberg, R. W., Deformation and Fracture Mechanisms of Engineering Materials, 3rd ed., Wiley, New York, 1989. [6] Kaufman, J. G., Introduction to Aluminum Alloys and Tempers, ASM International, Materials Park, OH, 2000. [7] Sanders, R. E., “TMS Presentation on the Influence of Disruptive Innovation on Today’s Aluminum Products,” Alcoa Technical Center, New Kensington, PA, 2006. [8] Giummarra, C., Thomas, B., and Rioja, R. J., “New Aluminum-Lithium Alloys for Aerospace Applications,” Proceedings of the Light Metals Technology Conference, CANMET, Ottawa, ON, Canada, 2007. [9] Lequeu, P. H., Smith, K. P., and Danielou, A., “Aluminum-Copper-Lithium Alloy 2050 Developed for Medium to Thick Plate,” Journal of Materials Engineering and Performance, Vol. 19, No. 6, 2010, pp. 841–847. [10] Pandey, A. B., “Composites,” ASM Handbook, Vol. 21, ASM International, Materials Park, OH, 2006, pp. 150–159. [11] Skinner, D. J., “Dispersion Strengthened Aluminum Alloys,” Proceedings of the Six Session Symposium on “Dispersion Strengthened Aluminum Alloys,” held at the 1988 TMS Annual Meeting, Phoenix, Arizona, January 25–29, 1988, edited by Y.-W. Kim, and W. M. Griffith, The Minerals, Metals and Materials Society, Warrendale, PA, 1988, p. 181. [12] Kim, Y.-W., “Dispersion Strengthened Aluminum Alloys,” Proceedings of the Six Session Symposium on “Dispersion Strengthened Aluminum Alloys,” held at the 1988 TMS Annual Meeting, Phoenix, Arizona, January 25–29, 1988, edited by Y.-W. Kim, and W.M. Griffith, The Minerals, Metals and Materials Society, Warrendale, PA, 1988, p. 157. [13] Palmer, I. G., Thomas, M. P., and Marshall, G. J., “Dispersion Strengthened Aluminum Alloys,” Proceedings of the Six Session Symposium on “Dispersion Strengthened Aluminum Alloys,” held at the 1988 TMS Annual Meeting, Phoenix, Arizona, January 25–29, 1988, edited by Y.-W. Kim, and W. M. Griffith, The Minerals, Metals and Materials Society, Warrendale, PA, 1988, p. 217. [14] Chan, K. S., “Dispersion Strengthened Aluminum Alloys,” Proceedings of the Six Session Symposium on “Dispersion Strengthened Aluminum Alloys,” held at the 1988 TMS Annual Meeting, Phoenix, Arizona, January 25–29, 1988, edited by Y.-W. Kim, and W. M. Griffith, The Minerals, Metals and Materials Society, Warrendale, PA, 1988, pp. 283–308. [15] Kendig, K. L., and Miracle, D. B., “Strengthening mechanisms of an Al-Mg-Sc-Zr alloy,” Acta Materialia, Vol. 50, No. 16, 2002, pp. 4165–4175. [16] Pandey, A. B., Kendig, K. L., and Miracle, D. B., “Development of a New Discontinuously Reinforced Aluminum for Space Applications,” Affordable Metal

40

[17]

[18]

[19]

[20]

[21]

B. N. BHAT

Matrix-Composites for High Performance Applications I, edited by A. B. Pandey, K. L. Kendig, and T. J. Watson, The Minerals, Metals and Materials Society, Warrendale, PA, 2001, pp. 36–45. Hunt, W. A., Jr., “Processing and Fabrication of Advanced Materials III,” Symposium Proceedings, edited by V. A. Ravi, T. S. Srivatsan, and J. J. Moore, The Minerals, Metals and Materials Society, Warrendale, PA, 1994, pp. 1663–681. Pandey, A. B., Shah, S. S., and Shadoan, M., “High Strength Discontinuously Reinforced Aluminum for Rocket Applications,” Affordable Metal-Matrix Composites for High Performance Applications II, Proceedings of Symposium held on November 9–12, 2003, edited by A. B. Pandey, K. L. Kendig, J. J. Lewandowski, and S.R. Shah, The Mineral, Metals and Materials Society, Warrendale, PA, 2003, pp. 3–12. Pandey, A. B., Majumdar, B. S., and Miracle, D. B., “Effects of Thickness and Precracking on the Fracture Toughness of Particle-Reinforced Al-Alloy Composites,” Metallurgical and Materials Transactions A, Vol. 29A, 1998, pp. 1237–1243. Pandey, A. B., Majumdar, B. S., and Miracle, D. B., “Deformation and Fracture of a Particle-Reinforced Aluminum Alloy Composite: Part I. Experiments,” Metallurgical and Materials Transactions A, Vol. 31A, 2000, pp. 921–936. DWA-Aluminum Composites USA, Inc.: https://dwa-usa.com/aluminum-mmcs. html

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2.3 TITANIUM ALLOYS Sesh Tamirisakandala Ernie M. Crist and Patrick A. Russo RTI International Metals, Inc.

2.3.1 INTRODUCTION Titanium (Ti: atomic number 22) is the ninth-most abundant element in the Earth’s crust at a level of about 0.6% and is the fourth-most abundant structural metal after aluminum, iron, and magnesium. High strength, low density, excellent corrosion resistance, and biocompatibility are the principal characteristics that make titanium attractive for a variety of applications. Examples include aircraft (high strength in combination with low density), aeroengines (high strength, low density, fatigue and good creep resistance up to about 5508C), biomedical devices (corrosion resistance, high strength, low elastic modulus, and biocompatibility), and chemical processing equipment (corrosion resistance). As a structural metal, Ti is still in its infancy compared with steel and aluminum. Its main use (approximately 60% of the Ti manufactured worldwide) is still in airborne applications such as aeroengines, airframes, missiles, and spacecraft. This section provides a brief overview of characteristics of Ti alloys to set the stage for aerospace applications described in later chapters in this book. Some of the basic characteristics of Ti and its alloys are listed in Table 2.9 and compared to those of other structural metallic materials based on Fe, Ni, and Al. Although Ti has the highest strength-to-density ratio, it is the material of choice only for certain niche application areas because of high cost. This high cost is mainly a result of the high reactivity of Ti with oxygen and high raw material cost. The use of vacuum or inert atmosphere is required during the production processes of Ti metal extraction as well as during melting processes to avoid oxygen contamination. The relatively high cost of Ti has hindered wider use, for example, in automotive applications. To minimize the inherent cost problem, successful applications must take advantage of the special features and characteristics of Ti that differentiate it from competing engineering materials, thereby making its use cost-effective. Doing so requires a more thorough understanding of Ti alloys as compared to other, less expensive materials, including the interplay between cost, processing methods, and performance. The classic comparison between Ti and other aerospace metals is on a strength-to-weight ratio basis as shown in Fig. 2.19 [1]. Tensile strengths as a function of temperature for commercial purity (CP) titanium and workhorse Ti alloy Ti-6Al-4V (Ti-64) are compared with those of 7075-T6 aluminum alloy, precipitation hardening (PH) stainless steel (17-4PH), and nickel base superalloy IN718 in Fig. 2.19a. Note that strength comparisons without density considerations are not so favorable for Ti alloys. However, the use of Ti in aircraft applications is greatly dependent on these density-adjusted properties and makes Ti alloys very attractive (Fig. 2.19b). Unalloyed Ti (CP Ti), on the other hand, is Copyright # 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

42

B. N. BHAT

TABLE 2.9

CHARACTERISTICS OF Ti ALLOYS COMPARED WITH OTHER STRUCTURAL METALLIC MATERIALS

Characteristic

Ti

Fe

Ni

Al

Melting point (8C)

1670

1538

1455

660

Crystal structure

bcc (b), hcp (a)

fcc (g), bcc (a)

fcc

fcc

Phase transformation (8C)

882 (b ! a)

912 (g ! a)

—

—

Elastic modulus (GPa)

115

215

200

72

Yield stress (MPa)

1000

1000

1000

500

Density (g/cm )

4.5

7.9

8.9

2.7

Thermal expansion coefficient (1026/8C)

9

11.8

13.4

23.1

Thermal conductivity (W/m8C)

7

80

90

237

Heat capacity (J/kg8C)

530

450

440

900

Electrical resistivity (mVm)

1.67

0.09

0.07

0.03

Corrosion resistance (relative)

Very high

Low

Medium

High

Reactivity with oxygen (relative)

Very high

Low

Low

High

Metal cost (relative)

Very high

Low

High

Medium

3

not attractive for aerospace structural applications due to low strength (absolute as well as specific). Some Ti alloys provide outstanding specific strength advantages at cryogenic temperatures also, as shown in Fig. 2.19b. The much higher melting temperature of Ti as compared to aluminum (Al), the main competitor in lightweight structural applications, gives Ti a definite advantage at application temperatures in excess of about 1508C. The high reactivity of Ti with oxygen limits the maximum use temperature of Ti alloys to about 6008C. Above this temperature, rapid ingression of oxygen through the surface occurs that leads to the formation of oxide scale and a brittle subsurface oxygen-enriched layer (known as alpha case) underneath the scale. The high reactivity with oxygen, on the other hand, leads to the immediate formation of a stable, adherent, and epitaxial oxide layer on the surface when exposed to air, resulting in the superior corrosion

AEROSPACE MATERIALS CHARACTERISTICS

b) Specific ultimate tensile

a) Ultimate tensile strength

Fig. 2.19

43

Properties of Ti and other structural metal alloys.

resistance of Ti in various kinds of aggressive environments, especially in aqueous acid environments. Superior corrosion resistance, rather than strength, is the primary reason for many industrial applications (e.g., chemical, oil and gas, etc.) of Ti. Emerging energy industry applications attempt to take advantage of the higher strength of Ti alloys combined with their superior corrosion resistance in hot aqueous corrosive environments (e.g., offshore risers, deep well tubulars).

2.3.2 TITANIUM ALLOYS CLASSIFICATION Ti has an unfilled d electron shell, thereby classifying it as a transition metal. It is allotropic, existing in two allotropic forms, a hexagonal close-packed (hcp) crystal structure a phase that is stable up to 8808C, which transforms to a body-centered cubic (bcc) b phase, which is stable up to the melting point (16708C). Alloying elements added to Ti are categorized according to how they affect the b transus temperature (a ! b phase transformation temperature), as shown in Fig. 2.20.

Fig. 2.20

Influence of alloying elements on phase diagrams of Ti alloys.

44

B. N. BHAT

The most common a stabilizers are Al, which forms a substitutional solid solution, and O, N, and C, which form interstitial solid solutions. Beta phase can be stabilized via two types of elemental additions, isomorphous (V, Mo, Nb, Ta) and eutectoid (Fe, Cr, Ni, Co, Cu, Si, Mn, H). Of these b stabilizers, only hydrogen (H) is an interstitial addition. The eutectoid additions develop a eutectoid reaction with the presence of compounds in the system at equilibrium. The isomorphous systems show continuous presence of b phase without the formation of any compounds. Some reasons for adding b stabilizers include solid solution strengthening, strengthening via microstructure refinement, and microstructure modification by heat treatment and thermomechanical processing (TMP), which is important for secondary property improvement, improved hardenability, enhancement of workability, and improved tolerance to H. The neutral additions, Sn and Zr, are important to develop an enhanced base alloy especially designed for high-temperature usage. Commercial Ti alloys are classified conventionally into three basic groups: a and near-a, a þ b, and near-b (or metastable b) and b, according to their position in a pseudo-binary section of phase diagram schematically shown in Fig. 2.21a [2]. Each group has a distinct set of properties qualitatively compared in Fig. 2.21b with respect to the workhorse a þ b alloy Ti-6Al-4V. This figure indicates a few important trends: 1) alloys with higher beta content can be heat treated to higher strengths (however, this trend comes with a deficit in weldability with a alloys having better weldability than a þ b or b alloys); 2) alloys leaner in b content generally have better high-temperature creep resistance; and 3) at equivalent strength levels, metastable b alloys have better formability. A clarification of the term metastable beta alloy is needed. A beta Ti alloy technically would be an alloy that is shown to contain only beta phase at room temperature under equilibrium conditions. However, most Ti alloys are not enriched enough in

Fig. 2.21 a) Schematic pseudo-binary section of b isomorphous phase diagram (Ms: Martensite start) and b) qualitative comparison of main characteristics of different Ti alloy family groups.

AEROSPACE MATERIALS CHARACTERISTICS

45

beta-stabilizing elements to create such an alloy except in cases requiring corrosion resistance beyond what conventional alloys can offer. In addition, because beta Ti alloys cannot be heat treated to high strength and are quite dense due to the addition of large quantities of heavy alloying elements, they have a very small niche in the overall Ti alloy systems. Metastable b Ti alloys are defined as alloys that can be quenched from above their b transus and retain all b phase in a small sample. Because the alloy system indicates that under equilibrium the alloy should contain both a and b phases, this alloy group is termed metastable beta. The potential to precipitate the a phase during heat treatment gives this alloy group high-strength capability as compared with stable b alloys. In addition to the substitutional alloying elements, changes in mechanical properties of Ti can be altered quite markedly by interstitial elements such as O, C, N, and H. These elements can enter into Ti as impurities in raw materials or can be introduced as controlled additions within specification limits for beneficial effects. Generally, they increase strength at acceptable levels of reductions in ductility and damage tolerance. For example, about 0.1% N (by weight) more than doubles the strength of Ti but reduces ductility in half. Large additions of the interstitials (still less than 1%) can embrittle Ti to the point of engineering uselessness. Thus, in seeking beneficial changes in properties by alloying, the interstitial additions must be carefully controlled. Several strength levels of unalloyed commercially pure Ti are thus produced by controlling the level of interstitials. Similarly, several grades of selected Ti alloys (e.g., commercially pure Ti, Ti-6Al-4V, Ti-5Al-2.5Sn) are produced by controlling the oxygen content. The one interstitial that offers no mechanical property improvement in Ti and must be maintained at very low levels, usually less than 125 ppm, is H. It can be absorbed by Ti during hot processing, especially when reducing furnace atmospheres are present, and during the acid pickling of Ti for surface cleaning. Unlike the other interstitials, H can be easily removed by vacuum annealing at a lower temperature. However, this process is expensive and the controls to limit absorption of H are favored. For a given Ti alloy, properties can be further improved via processing. Alloying lays the basis for an increase in strength (via solid solution strengthening, age hardening, etc.), allows the generation of ordered structures (e.g., Ti3Al), determines most physical properties (density, modulus, CTE, etc.), and largely controls the chemical resistance (corrosion, oxidation). Processing allows the careful balancing of property combinations of Ti alloys. Depending on the property desired in the final application, different microstructures can be generated in Ti alloys by means of TMP to optimize for strength, ductility, toughness, durability and damage tolerance, creep resistance, formability, and so forth. The microstructure of conventional Ti alloys is primarily described by the size and arrangement of the two phases, a and b. The two extreme cases of phase arrangements, shown in Fig. 2.22, are the lamellar microstructure, which is generated upon cooling from the beta phase field, and the equiaxed microstructure, which is a result of a

46

B. N. BHAT

Fig. 2.22 Three typical microstructural classes of Ti alloys (backscattered electron images, contrast: a, gray; and b, white). recrystallization process. Both types of microstructures can have a fine as well as a coarse arrangement of their two phases. A mix of lamellar plus equiaxed, referred to as bimodal (or duplex), provides the required balance of properties for certain applications. Table 2.10 shows qualitatively how the size and arrangement of the phases influence various properties. A list of the most important commercial Ti alloys belonging to each of the three different groups is shown in Table 2.11. Alpha alloys are primarily chosen for corrosion resistance and fabricability (formability and weldability) applications (chemical processing equipment, heat exchangers, pumps, piping, etc.). Alpha þ beta alloys are by far used in the greatest quantities and are applied when higher strength and better fracture resistance, fatigue strength, or elevated temperature capability with good corrosion resistance are needed (aircraft structures, aircraft engines, steam turbine blades, spacecraft components, etc.). These are most readily grouped by strength, use temperature, and microstructural TABLE 2.10

Fine

INFLUENCE OF MICROSTRUCTURAL PARAMETERS ON MECHANICAL PROPERTIES (QUALITATIVE) Lamellar

Equiaxed

0

Coarse 0

Modulus

Property

0

þ/2 (texture)

þ

2

Strength

2

þ

þ

2

Ductility

2

þ

2

þ

Fracture toughness

þ

2

þ

2

Fatigue crack initiation

2

þ

2

þ

Fatigue crack growth rate

þ

2

2

þ

Creep strength

þ

2

þ

2

Superplasticity

2

þ

þ

2

Oxidation behavior

þ

2

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.11 Common Name

47

IMPORTANT COMMERCIAL Ti alloys Alloy Composition (wt %)

Beta Transus (8C)

a Alloys CP Ti Grade 1

Ti-0.2Fe-0.18O

890

CP Ti Grade 2

Ti-0.3Fe-0.25O

915

CP Ti Grade 3

Ti-0.3Fe-0.35O

920

CP Ti Grade 4

Ti-0.5Fe-0.40O

950

CP Ti Grade 7

Ti-0.2Pd

915

CP Ti Grade 12

Ti-0.3Mo-0.8Ni

880

Ti-5-2.5

Ti-5Al-2.5Sn

1040

Ti-3-2.5 (half Ti-6-4)

Ti-3Al-2.5V

935

a 1 b Alloys Ti-811

Ti-8Al-1V-1Mo

1040

IMI 685

Ti-6Al-5Zr-0.5Mo-0.25Si

1020

IMI 834

Ti-5.8Al-4Sn-3.5Zr-0.5Mo-0.7Nb-0.35Si-0.06C

1045

Ti-6242

Ti-6Al-2Sn-4Zr-2Mo-0.1Si

995

Ti-6-4

Ti-6Al-4V-0.20O

995

Ti-6-4 ELI

Ti-6Al-4V-0.13O

975

Ti-662

Ti-6Al-6V-2Sn

945

Ti-6-22-22

Ti-6Al-2Sn-2Zr-2Mo-2Cr-0.15Si

960

IMI 550

Ti-4Al-2Sn-4Mo-0.5Si

975

Ti-6246

Ti-6Al-2Sn-4Zr-6Mo

940

Ti-17

Ti-5Al-2Sn-2Zr-4Mo-4Cr

890

SP-700

Ti-4.5Al-3V-2Mo-2Fe

900

Beta-CEZ

Ti-5Al-2Sn-2Cr-4Mo-4Zr-1Fe

890

Ti-5-5-5-3

Ti-5Al-3V-3Mo-3Cr

860

Ti-10-2-3

Ti-10V-2Fe-3Al

800

Beta 21S

Ti-15Mo-2.7Nb-3Al-0.2Si

810

Ti-15-3

Ti-15V-3Cr-3Al-3Sn

760

Beta C

Ti-3Al-8V-6Cr-4Mo-4Zr

730

B120VCA

Ti-13V-11Cr-3Fe

700

b Alloys

48

B. N. BHAT

condition. Strength classification includes low-strength alloys: tensile yield strength (TYS)  800 MPa (115 ksi), medium-strength alloys: TYS 825900 MPa (115–130 ksi), and high-strength alloys: TYS 900-1175 MPa (130– 170 ksi). High-temperature alloys are classified by use temperature 3258C but 5508C. Three distinct microstructural conditions achieved by TMP include 1) fully lamellar for high toughness and lower strength, low ductility, high creep strength; 2) fully equiaxed for high ductility, lower toughness than fully lamellar, lower strength than bimodal, lower creep strength than fully lamellar; and 3) bimodal for balance between fully lamellar and fully equiaxed. Metastable beta alloys are applied where very high strength with acceptable fracture resistance is needed (aircraft landing gear, fasteners, springs, heavily loaded parts in jet engines and helicopters, etc.). The most commonly used Ti alloys for aerospace and space applications are: Ti-5-2.5, Ti-3-2.5, Ti-6-4 (regular and extra-low interstitial ELI grades), Ti-6242, Ti-6246, Ti-17, Ti-10-2-3, Beta 21S, Ti-15-3, and Beta C. Conventional alloy development activity in Ti alloys is focused on affordable processing methods via minor chemistry tweaks and incremental property improvements. Recently, a third option of Ti matrix composites (TMCs) has gained importance [3]. TMCs consist of a Ti matrix containing continuous reinforcing fibers (35–40 vol %) such as SiC. The principal attractions of TMCs are strength and stiffness. On a density-corrected basis, SiC-reinforced TMCs have about twice the ultimate strength and stiffness of conventional Ti alloys, measured parallel to the fiber direction. In principle, this makes them the most structurally efficient engineering materials. In practice, it is often difficult to fully capitalize on the unidirectional capability of TMCs in a component, because off-axis loads are usually present. Because of the fiber reinforcement, TMCs are extremely anisotropic, which creates a challenge to maximize the benefits of the longitudinal properties while minimizing the penalties associated with the lower transverse properties. The cost of TMC components is very high. An early, but important, small-volume production application of TMCs was established in 1999 for augmenter actuator links in a large (F-16 class) military engine. The TMC link weight is roughly 50% of that of the superalloy link. Other successful applications include landing gear components and NASA’s Super Lightweight Interchangeable Carrier. An alternate means of using TMCs at lower total cost is the use of selective reinforcement in locations where the greatest benefit can be created. A number of important recent developments were made to commercialize TMCs. Alloy design for high-temperature service requiring creep resistance evolved since the 1950s and is based on four principles: 1) control of the beta-stabilizing elements such as Mo, V, and Nb at levels slightly above their solid solubility in alpha phase, and minimization of Fe and Ni, as they all enhance creep rates due to rapid diffusion rates; 2) addition of trace silicon that forms silicides, which are effective barriers to plastic deformation at elevated temperatures; 3) introduction of Ti3Al precipitates via a final stabilization treatment to create effective barriers to deformation; and 4) microstructure engineering to produce the

AEROSPACE MATERIALS CHARACTERISTICS

49

size and morphology that provide better creep resistance. Intermetallic compounds based on Ti and Al are light (low-density), are relatively stiff (highmodulus), and have attractive high-temperature mechanical properties (tensile and creep strength), as shown in Fig. 2.23 [4, 5]. Historically, the issues associated with reduced ductility and fracture toughness of g-TiAl intermetallic alloys have been viewed as significant hurdles and outweigh the benefits of increased strength. Consequently, the use of intermetallic compounds in structural applications has been very limited. Because of their promise, however, a great deal of work has been performed on these compounds beginning in 1953 and continuing until now. In some regards, the effort to introduce these compounds for structural applications illustrates the difficulty associated with the introduction of any new class of material. That is, more than 60 years have passed, and the titanium aluminides are still considered developmental materials. The primary barrier has been concern about brittleness, but newer alloyed versions of both Ti3Al (a2 aluminide) and TiAl (g-aluminide) have alleviated many of these concerns. The principal concern at present is cost and attainable performance improvement. Perhaps the single most attractive current application for g alloys is for lowpressure turbine (LPT) blades in aeroengines. For LPT blades, g alloys would replace conventionally cast Ni base blades made from superalloys such as Rene 77. The maximum service temperature for these LPT blades is about 7508C, and the g alloys have adequate creep strength up to this temperature. The g alloys also have adequate surface stability at these temperatures and therefore retain their strength for extended service periods without embrittlement. Because these blades are rotating components, the reduced mass translates into lower loads on the LPT disk. Using the g alloys, the original Ni alloy disk can

Fig. 2.23 Specific tensile strength as a function of use temperature of selected structural materials vs Ti alloys and gamma Ti aluminides. (CFRP 5 carbon fiber-reinforced plastics.)

50

B. N. BHAT

be reduced in mass while maintaining constant levels of operating stress. In the weight-critical aeroengine industry, this is considered almost unheard of for a single material change and thus it is extremely attractive.

2.3.3 STRENGTHENING MECHANISMS From the four different hardening mechanisms in metallic materials (solid solution hardening, strengthening by a high dislocation density, boundary hardening, and precipitation hardening), solid solution and precipitation hardening are used in all commercial Ti alloys [6]. Boundary hardening plays a significant role in a þ b alloys cooled at high rates from the b phase field, reducing the a colony size to a few a plates or causing martensitic transformation. The martensite in Ti is much softer than the martensite in Fe-C alloys because the interstitial oxygen atoms only cause small elastic distortion of the hexagonal lattice of the Ti martensite. This is in sharp contrast to carbon and nitrogen that cause severe tetragonal distortion of the bcc lattice in ferrous martensite. The a phase is significantly hardened by the interstitial element oxygen. This is best illustrated by comparing the yield stress values of the CP titanium grades 1–4 with oxygen levels between 0.18% and 0.40%, which have yield stresses in the range 170 MPa (Grade 1) to 480 MPa (Grade 4). In commercial Ti alloys, the oxygen content varies between about 0.08% and 0.20% depending on alloy type. Substitutional solid solution hardening of the a phase is caused mainly by the elements Al, Sn, and Zr, which have fairly large atomic size differences compared to Ti and also have large solid solubilities in the a phase. Precipitation hardening of the a phase occurs by coherent Ti3Al particles above about 5% Al. These Ti3Al or a2 particles have an ordered hexagonal structure. Because they are coherent, they can be sheared by moving dislocations, resulting in planar slip and extensive dislocation pileups against boundaries. With increasing size, these a2 particles become ellipsoidal in shape, the long axis being parallel to the c axis of the hexagonal lattice. This a2 phase is further stabilized by the elements O and Sn; the (a þ a2) phase region is pushed to higher temperatures by these elements. In such cases, Sn substitutes for Al whereas oxygen remains as an interstitial. The Ti3Al phase can be deleterious in two basic situations. The first situation involves Ti being used in seawater under stress, thereby making the given Ti alloy susceptible to seawater stresscorrosion cracking. Alloy compositions and heat treatments must be tailored to reduce the tendency to form the Ti3Al phase. Ti-6Al-4V ELI (extra-low interstitial) alloy historically has been a good choice for such applications. The other situation involves high-temperature creep alloys, which can be embrittled by Ti3Al during long thermal exposures such as those encountered in jet engine applications. These situations are circumvented by the formulation of alloys restricting the Al, Sn, Zr, and O contents to levels based on experience. The Ti-6Al-2Sn4Zr-2Mo-Si alloy has logged hours of in-service usage in the 450–5508C temperature range without difficulty due to Ti3Al embrittlement.

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It is difficult to analyze solid solution strengthening of the b phase in a traditional sense because in fast-cooled metastable b alloys, the precursors of the metastable v and b0 phases cannot be separated from the solute effects. Also, in fully aged microstructures, it is difficult to separate the hardening due to a precipitates from the solute effects. In this case, an important solid solution strengthening effect of the b phase is due to alloy element partitioning that accompanies a precipitation. One way to estimate the solid solution strengthening of the b stabilizing elements is to examine the slope of the lattice parameters vs solute content curves for binary alloys, pointing out in a qualitative way the size misfit parameter. The strengthening effects of beta stabilizers from this analysis are greatest for Fe followed by Cr, V, Nb, and Mo.

2.3.4 PROCESSING OF TITANIUM ALLOYS A typical production flowchart to convert ores of Ti, which exist as titanium black sand, to a range of useful mill products used in aerospace applications is shown in Fig. 2.24 [7]. Metallic Ti, as obtained from the ore, is called sponge because it is porous and has a sponge-like appearance. The starting ore for the production of Ti is either rutile (TiO2) or ilmenite (FeTiO3). The extraction of metallic Ti from these ores occurs in five distinct operations: chlorination of the ore to produce TiCl4, distillation of the TiCl4 to purify it, reduction of the TiCl4 to produce Ti, purification of metallic Ti to remove by-products, and crushing and sizing of the sponge to create suitable product for subsequent melting to make ingots. In the 1930s, Dr. Wilhelm Kroll invented the first viable, large-scale industrial method based on Mg reduction to form Ti sponge, which is still the principal process for Ti production today. Melting of crushed Ti sponge along with alloying elements takes place in either a vacuum arc remelt (VAR) furnace, to produce VAR ingots typically used for aerospace applications, or in a cold hearth (electron beam or plasma arc) furnace, to produce feedstock for a subsequent VAR melt. In the VAR furnace, the electrode is consumable, melted by an arc struck between it and a layer of Ti in a water-cooled copper crucible. The molten Ti on the outer surface solidifies on contact with the cold wall, forming a shell or skull to contain the molten pool. The ingot is not poured but solidifies under vacuum in the melting furnace. A second melt ensures homogeneity of the ingot for industrial purposes, and a triple melting operation is used for Ti destined for critical aerospace applications, such as rotating parts in jet engines. The resultant VAR ingots are cylindrical shapes weighing up to 13,500 kg, which are forged to slabs or billets and then formed to mill products. The alternative sponge-melting process uses a cold hearth furnace. Here, the crushed sponge and alloying elements can also be mixed with inexpensive recycled Ti scrap before melting to reduce costs. The mixture is melted by either electron beams in a vacuum or by plasma arc torches under a positive pressure of helium. The metal flows across the cold hearth, where it forms a pool that allows impurities to sink to

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Fig. 2.24

Typical production flowchart for Ti sponge, mill products, and castings.

the bottom or to be evaporated. Cold hearth melting removes both hard a (enriched interstitial compounds) and high-density (such as tungsten) inclusions and is the preferred method for producing clean Ti for aerospace applications. The molten Ti is direct cast into a near-net shape, which can be slab intended for further processing, or a remelt electrode for subsequent VAR. A combination of cold hearth and VAR melts can eliminate inclusions and defects that even

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triple VAR melting cannot remove, and therefore it has become a common practice for aerospace applications. VAR ingots and cold-hearth melted cast slabs are pressed or rotary forged into slabs (rectangles) or billets (rounds). Then hot forming produces forgings, plates, and extrusions. This can be followed by cold rolling and common processing techniques to create mill products—basic structural shapes with desired properties that maximize metal use. The majority of aircraft components require hot working to overcome spring-back, minimize residual stresses, and reduce forming forces needed. The workhorse of the aerospace industry and a feedstock for additional mill products, forgings are available in a wide range of sizes. New developments in precision forging now provide near-net shapes with improved material efficiencies. Ti billets can have round, square, rectangular, hexagonal, or octagonal cross-sectional shapes. Hot rolling produces plate, which is generally available in thicknesses greater than 4.75 mm and widths greater than 500 mm. Vacuum creep flattening is widely used to achieve critical plate flatness. Sheet is a flat roll product that is typically less than 4.75 mm thick and produced by either hot or cold rolling. Strip is cut from cold-rolled and annealed sheet. Pipe and tube can be manufactured as either welded or seamless product in a variety of standard diameters and wall thicknesses. Ti is cost-effectively extruded into desired near-net shapes by forcing heated metal through a die. Extrusions maximize material usage and reduce the need for downstream milling, welding, and assembly. Other common Ti products include weld wire, fasteners, screws, nuts, bolts, washers, and rod. Casting is the most advanced and diversified of the netshape technologies. It offers greater design freedom and significantly reduces the need for expensive machining and fabrication to attain the desired shape. Commercial casting of Ti began in the late 1960s, and today the technology has matured to routinely supply critical gas turbine engine and airframe products. TMP is used for most fracture-critical components, which require better process control than working and heat treatment to achieve much tighter distributions around higher property values. Heat treatment, for example, annealing, reduces scatter in property values but at lower mean values. A typical TMP sequence is shown in Fig. 2.25. Critical Fig. 2.25 Thermomechanical processing sequences and critical parameters of Ti alloys.

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variables during TMP are: deformation temperature (above or below the b transus), amount of plastic strain introduced, plastic strain rate, deformation method (state of stress), cooling rate after deformation, and final heat treatment parameters.

REFERENCES [1] Wood, R., “Titanium and Its Alloys,” ASM International online course, 1994. [2] Donachie, M. J., Titanium: A Technical Guide, 2nd ed., ASM International, Materials Park, OH, 2000. [3] Lutjering, G., and Williams, J. C., Titanium, 2nd ed., Springer-Verlag, Berlin, 2007. [4] Boyer, R., Collings, E. W., and Welsch, G. (eds), Materials Properties Handbook: Titanium Alloys, ASM International, Materials Park, OH, 1994. [5] Leyens, C., and Peters, M., Titanium and Titanium Alloys, Wiley, New York, 2003. [6] Collings, E. W., Physical Metallurgy of Titanium Alloys, ASM International, Materials Park, OH, 1984. [7] Titanium: The Ultimate Choice, International Titanium Association, Broomfield, CO, 2005.

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2.4 STEELS Biliyar N. Bhat NASA Marshall Space Flight Center

2.4.1 INTRODUCTION Steels are the most common metallic materials used in structural applications globally because of their high strength and ductility and relatively low cost. One drawback is the high density of steels relative to aluminum and titanium, which tends to increase the mass of aerospace systems. On the upside, many alloy steels can be heat treated to very high strength levels (.200 ksi yield strength) such that the specific strengths (strength/density) are high enough to be acceptable to designers. High fracture toughness is another attribute that makes steels attractive in applications that require high damage tolerance. Maraging steels combine strength and fracture toughness. Most steels become brittle at cryogenic temperatures and are not recommended for cryogenic applications. Corrosionresistant stainless steels are the most commonly used steels in aerospace applications, which typically require good corrosion resistance. Nickel-chromium steels are oxidation resistant and are used in oxidizing environments. Table 2.12 lists some of the common steels used in aerospace and their chemistries. These steels serve as examples only, and the list is by no means complete. Many more ferrous alloys and their chemistries and properties are addressed in the references listed at the end of the chapter. Only a brief description of the classes of steels and their aerospace applications is given in this section.

2.4.2 ULTRAHIGH-STRENGTH STEELS These steels typically are capable of minimum yield strength of 1380 MPa (200 ksi). They are low-alloy steels that can be thermomechanically processed (TMP) to very high strength and toughness levels. For instance, AISI/SAE 4340 steel can be heat treated (quenched and tempered) to strength levels of 150 to 280 ksi by controlling the tempering temperature. Higher tempering temperature gives lower strength but higher ductility and fracture toughness. This steel is susceptible to hydrogen embrittlement, and care must be taken to bake out any hydrogen. It has extremely poor resistance to stress corrosion cracking (SCC), especially at high strength levels (1500–1950 MPa, 220–280 ksi). It is usually forged at 1065–12308C. It has good welding characteristics and can be readily gas or arc welded. Because 4340 steel is air hardening, welded parts should be annealed or normalized and then tempered soon after welding. AISI/SAE 4340 steel is produced as forgings, light plate, and castings. Typical applications include landing gear and other critical structural members for aircraft. Alloy 300M is basically silicon-modified (1.6% Si) 4340 steel. It has higher carbon and molybdenum content and also contains vanadium. It has high This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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TABLE 2.12 Designation

Commercial Names

Ultrahigh-strength steels 4340

COMMON AEROSPACE STEELS

Typical Composition

Typical Aerospace Applications

Fe-0.35C-0.7Mn-0.3Si-0.8Cr-1.8Ni-0.25Mo-

Landing gear, fasteners

300M

Fe-0.43C-0.8Mn-1.6Si-0.8Cr-1.8Ni-0.4Mo-0.05V

Fasteners, landing gear, airframe parts

D6-AC

Fe-0.45C-0.75Mn-0.25Si-1.1Cr-0.6Ni-1.0Mo-0.08V

Motor cases for solid fuel rockets

Fe-0.30C-0.25Mn-0.2Si-1.0Cr-7.5Ni-1.0Mo-0.1V-4.5Co

Aircraft structural components

AF1410

Fe-0.15C-0.1Mn-0.1Si-2.0Cr-10Ni-1.0Mo-14Co

Aircraft structural components

18Ni (250)

Fe-18Ni-5Mo-8.5Co-0.4Ti-0.1Al

Aircraft structural parts, shafts, missile cases

Fe-18Cr-8Ni

General-purpose corrosion and heat-resistant steel

304, 304L

Fe-LowC-19Cr-10Ni

General-purpose corrosion resistant; better than 301

316, 317

Fe-18Cr-13Ni-Mo

Tubing in propulsion systems

High fracture toughness steels HP 9-4-30 Maraging steels

Corrosion-resistant steels Austenitic stainless steels 301, 302

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Fe-18Cr-10Ni-Ti

347, 348

Fe-18CR-10Ni-Nb

Aircraft structural tubing

21-6-9

Fe-LowC-20Cr-6.5Ni-9Mn-0.28N

Aircraft hydraulic line tubing

Fe-LowC-12Cr

Engine parts

Fe-HighC-17Cr-0.5Mo

Cryogenic bearings

Martensitic stainless steels 403, 410, 416 440A, B, C and F

Precipitation hardening stainless steels 17-4PH Fe-17Cr-4Ni4Cu

Aircraft structural tubing

Structural parts requiring corrosion resistance

17-7PH

Fe-17Cr-7Ni-1Al

Structural parts requiring corrosion resistance

15-5PH

Fe-15Cr-4.5Ni-0.3Cb-3.5Cu

Structural parts requiring corrosion resistance

Custom 455

Fe-LowC-12Cr-8Ni-2Co-1Ti þ Cb

Structural parts requiring corrosion resistance

Fe-25Ni-15Cr-2Ti-1.5Mn-1.3Mo-0.3V

Propulsion systems

JBK-75

Fe-30Ni-15Cr-2Ti-0.1Mn-1.3Mo-0.3V-0.3Al

Propulsion systems

Incoloy

Fe-34Ni-20Cr

Hot sections of engines

Nickel-chromium steels A-286

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ductility and toughness at tensile strengths of 1860–2070 MPa (270–300 ksi). It is also susceptible to hydrogen embrittlement when heat treated to strength levels greater than 1380 MPa (200 ksi) and requires baking. Processing is similar to 4340 steel. D-6AC steel is specially developed by Ladish Company for aircraft and missile structural applications. It is primarily designed for use at roomtemperature strength levels of 1800–2000 MPa (260–290 ksi). D-6AC is produced by air melting followed by vacuum arc remelting (VAR). It can be forged, rolled, welded, and heat treated similar to 4340 steel. The alloy is susceptible to SCC; its in bothpwater and 3.5% NaCl solution appears to fracture toughness (KIscc) value p be low [less than 16 MPa m (15 ksi in)] at high strength levels.

2.4.3 HIGH FRACTURE TOUGHNESS STEELS These steels are capable of a yield strength p of 1380pMPa (200 ksi) and plain strain fracture toughness (KIc) of 100 MPa m (91 ksi in). They are also resistant to SCC. Both HP-9-4-30 and AF 1410 alloys are Fe-Ni-Co type steels and have similar characteristics. Melt practice requires a minimum of VAR. HP-9-4-30 is usually electric arc melted and vacuum arc remelted. It can develop p a tensile of 100 MPa m (91 ksi strength of 1520–1650 MPa (220–240 ksi) and a K Ic p in). Heat-treated parts can be readily welded. Tungsten arc welding under inert gas shielding is the preferred welding process. Postweld heat treatment is not required, but stress relief at 5408C (10008F) is recommended for relieving residual stresses; it does not have an adverse effect on the strength or ductility of base or weld metal. AF1410 steel is an Air Force-developed Fe-Ni-Co type alloy steel designed for replacing some titanium parts. It has significant SCC resistance. By raising the cobalt and carbon content, the ultimate tensile strength was increased to a p value of 154 m (140 ksi typical 1615 MPa (235 ksi) while maintaining a K Ic p in). This combination of strength and toughness exceeds that of other commercial steels of pcomparable strengths. SCC resistance is very good [KIscc of 66 MPa p m (60 ksi in)]. The preferred melting practice is vacuum induction melting followed by vacuum arc remelting (VIM-VAR). Welding is done under inert gas shielding to avoid oxygen contamination.

2.4.4 MARAGING STEELS The term maraging is derived from martensite age hardening and denotes age hardening of a low-carbon iron-nickel lath martensite. Unlike conventional steels that are hardened by martensitic reaction that involves carbon, maraging steels are strengthened by precipitation of intermetallic compounds at about 4808C (9008F). The yield strengths range from 1030 to 2420 MPa (150– 350 ksi). These steels typically have very high nickel, cobalt, and molybdenum content. Carbon is considered an impurity and is kept very low. 18Ni (250) is a common commercial-grade maraging steel with yield strength of 250 ksi

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(Table 2.12). Compared to conventional steels, maraging steels are more hardenable with only slight dimensional changes. This feature allows intricate parts to be machined in the soft condition and then heat treated with minimum distortion. Fracture toughness of maraging steels is considerably better than that of conventional high-strength steels and hence is attractive for applications where damage tolerance is important. Resistance to corrosion and stress corrosion is also better than that of conventional high-strength steels. Premium grades of maraging steels used in critical aircraft and aerospace applications are generally triple melted using air, VIM, and VAR processes to minimize residual elements (such as carbon, manganese, sulfur, and phosphorous) and gas (O2, N2, H2). The ingots can be hot and cold worked using conventional steel mill techniques. The steels are weldable, and fabrication costs are lower than for conventional high-strength steels.

2.4.5 CORROSION-RESISTANT STEELS Corrosion-resistant steels (aka stainless steels) are iron base alloys that contain at least 10.5% chromium, which imparts corrosion resistance through formation of invisible and adherent chromium-rich oxide surface film. Other elements are added to improve particular characteristics include nickel, molybdenum, copper, titanium, aluminum, silicon, niobium, and nitrogen. Carbon is normally present in amounts ranging from less than 0.03% to more than 1.0% in certain martensitic grades. Corrosion resistance and mechanical properties are the most important factors in selecting these steels for aerospace applications. Other considerations include fabrication characteristics, availability, and cost. It is instructive to look at the phase diagram showing compositions of Fe-Cr-Ni alloys for which austenite persists at room temperature (Fig. 2.26) to gain insight into the properties of stainless steels. In the diagram, point P indicates the position of an alloy containing 18% chromium and 8% nickel. Chromium is body centered cubic (bcc) and tends to stabilize the ferrite phase. Nickel is face center cubic (fcc) and tends to stabilize the austenite phase. A combination of nickel and chromium will give a variety of steels with tailored properties. The stainless steels are classified based on microstructure and properties. Martensitic stainless steels are essentially alloys of chromium and carbon that possess a distorted bcc crystal structure (martensitic) in the hardened condition. They are ferromagnetic and hardenable by heat treatment, and their corrosion resistance is relatively low. Chromium content is in the range of 10.5%–18%, and Fig. 2.26 Iron-chromium-nickel (Fe-Cr-Ni) phase diagram.

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carbon content ranges up to 1.2% or more. The chromium and carbon contents are balanced to ensure martensitic structure after hardening. Excess carbides may be present to increase wear resistance. Austenitic stainless steels have an fcc structure. This structure is attained through use of austenitizing elements such as nickel and manganese. These steels are essentially nonmagnetic in the annealed condition and can be hardened only by cold working. They usually possess excellent cryogenic properties. Molybdenum, copper, silicon, aluminum, titanium, and niobium may be added to impart certain characteristics, such as pitting corrosion resistance or oxidation resistance. Precipitation hardening (PH) stainless steels are Fe-Cr-Ni alloys containing precipitation hardening elements such as copper, aluminum, or titanium. These steels may be either austenitic or martensitic in the annealed condition. Those that are austenitic in the annealed condition are generally transformable to martensite through conditioning heat treatments, sometimes with a subzero treatment. In most case these steels attain their strength by precipitation hardening of the martensitic structure. Corrosion resistance is generally the most important characteristic of a stainless steel, but often it is also the most difficult to assess for a specific application. General corrosion resistance to pure chemical solutions is comparatively easy to determine, but actual environments are usually more complex. General corrosion is much less severe than localized forms of corrosion such as SCC, crevice corrosion (in tight places or under deposits), pitting corrosion, and intergranular attack in sensitized material such as weld heat affected zones (HAZ). Such localized corrosion can cause unexpected and sometimes catastrophic failure while most of the structure remains unaffected. Therefore, care must be exercised in selecting the right grade of stainless steel for a given environment.

2.4.6 NICKEL–CHROMIUM STEELS These are iron-nickel-chromium alloys containing more than 25% Ni and more than 10% Cr, always more Ni than Cr. These austenitic alloys are used for oxidizing as well as reducing atmosphere and are used in high-temperature parts of propulsion systems. For example, A-286 is a high-toughness age-hardenable stainless steel. Applications include aircraft turbine engine components such as discs, vanes and blades, shafts, cases, and combustors. The space shuttle main engine (SSME) nozzle is made from A-286 tubes brazed together. The alloy has good strength and toughness down to cryogenic temperatures. A modification of A-286, called JBK-75, was developed for improved weld-cracking resistance. It has higher hot strength and resistance to hydrogen embrittlement than A-286. Incoloy alloy possesses good oxidation and strength properties at temperatures up to 18008F.

SUGGESTED READING There are numerous good references for steels. A few examples follow this paragraph. Steel properties data can be obtained from steel suppliers, which usually

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provide data on the steels they produce and sell. Aerospace companies generally generate databases for materials used in specific applications, but these databases tend to be proprietary. NASA’s Marshall Space Flight Center maintains a database on materials used by NASA that is stored in the Materials and Processes Technical Information System (MAPTIS), which is accessible to qualified users. Metallic Materials Properties Development and Standardization is also a commonly used source for metallic materials properties.

BIBLIOGRAPHY ASM International, ASM Handbook, Vol. 1: Properties and Selection: Irons, Steels, and High-Performance Alloys, 10th ed., ASM International, Materials Park, OH, 1990. Batelle Memorial Institute, Metallic Materials Properties Development and Standardization (MMPDS), Battelle Memorial Institute, Columbus, OH, 2014. Hucek, H. J., Aerospace Structural Metals Handbook, Battelle Columbus Laboratories for Department of Defense, Columbus, OH, 1990. Lula, R. A., Stainless Steels, American Society for Metals, Materials Park, OH, 1986.

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2.5 SUPERALLOYS Michael V. Nathal NASA John H. Glenn Research Center at Lewis Field

The term superalloys has been applied to the alloys used in some of the harshest environments experienced by materials and components. Superalloys are most often associated with aerospace propulsion systems, such as jet and rocket engines, but are also heavily used in power generation and chemical industries. These applications require a balance of properties, most notably a combination of high temperature strength, resistance to high temperature oxidation and corrosion attack, and significant damage tolerance. Superalloys can be based on Ni, Fe-Ni, or Co as the primary ingredient. Early superalloys were an outgrowth of stainless steels and were thus developed on a Fe-Ni-Cr base. Another early alloy was Ni-20Cr, which relied on Cr to provide oxidation resistance. The strengthening effect of adding Al and Ti to NiCr alloys was discovered in 1929, and new alloys based on this discovery began to appear in the 1940s. Continued alloy development over the past 70 years has resulted in more than 100 different superalloys containing as many as 15 alloying elements. Ni-base superalloys are used in the most demanding components of a jet engine, such as the turbine disks and blades, which are exposed to both high stresses and temperatures in an oxidizing environment. When a component’s stress/temperature requirements are not quite as severe, Fe-Ni base superalloys are preferred over Ni-base alloys. These alloys tend to be less expensive, plus they are more amenable to a wide variety of processing techniques, including welding. The Fe-Ni base alloy IN-718 is the most widely used superalloy in the world, as a result of its excellent mechanical properties up to about 6008C, combined with ease of manufacture. The use of Co-base alloys is restricted by the higher cost of Co compared to Ni, but their unique combination of properties is nevertheless suitable for a substantial portion of aerospace applications.

2.5.1 SUPERALLOY METALLURGY Ni-base superalloys consist of a face-centered cubic (FCC) matrix known as gamma (g) and a closely related FCC precipitate based on Ni3Al, known as gamma prime (g 0 ). Figure 2.27 shows typical superalloy microstructure. In Fig. 2.27b, two different populations of precipitates are present, each population formed at a distinct aging temperature. The long-range ordered crystal structure of g 0 is the source for the excellent strength and creep resistance of superalloys. In addition, small amounts of carbides and borides are usually present. The g phase has key attributes of substantial ductility, plus a capacity to accept numerous ternary elements in solid solution, thus allowing a large alloy design space. The g phase of most alloys contains substantial amounts of Cr in solid solution, This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

AEROSPACE MATERIALS CHARACTERISTICS

a) Low-magnification scanning electron micrograph of grain

Fig. 2.27

63

b) Transmission electron micrograph of gamma prime (γ′) precipitates

Typical microstructures of Ni-base superalloys.

which provides some degree of strengthening but is primarily present for environmental resistance. Cr imparts oxidation resistance by promoting the formation of a slow growing, adherent Cr2O3 oxide scale that protects the bulk alloy from oxidation attack. However, the degree of protection provided by Cr2O3 is limited to temperatures of about 10008C, and the most oxidation-resistant alloys form Al2O3 instead. High Cr levels are still the best remedy for resistance to many hot corrosion attack mechanisms. Over the decades, superalloy compositions have become more complex, as refractory metals were added for improved strength. Mo, W, and Re are elements that partition primarily to g and provide solid solution strengthening. The g 0 phase can be similarly strengthened by ternary additions, and Ti, Nb, and Ta are added for this purpose. Alloys are designed such that the lattice parameters of the FCC crystal structures of g and g 0 tend to be closely matched, and the lattice parameter mismatch rarely exceeds +0.5%. Alloying additions are limited to amounts that can remain in solid solution in the g and g 0 . Exceeding the solubility limits results in the formation of deleterious topologically close packed (TCP) phases such as s, m, a, Laves, and d. Additional constraints on alloying include cost, density, and the need to balance competing properties. For example, W is a very potent strengthener of both g and g 0 , but excessive W levels can result in the formation of a or m phase. Cr levels can be reduced to allow for higher W additions, but Cr can only be reduced so far before environmental resistance is compromised. Alloy design is not stymied, however, as alternative means to achieve the balance of properties can be invoked. For example, in the pursuit of higher strength, oxidation resistance of an alloy can be sacrificed to some extent by relying on protective coatings. The role of minor alloying additions is also critical for optimum properties. Almost all alloys contain small amounts of C (,0.1%) that result in the formation of carbides. The carbides can be of several types, such as MC, M23C6, and M6C (where M denotes a combination of metal species) and can strengthen

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grain boundaries. M ¼ Hf, Ti, Ta, and/or Nb in MC; M ¼ Cr in M23C6; and M ¼ Mo or W in M6C carbides. Zr and B at trace levels (,0.1%) are also considered grain boundary strengtheners, although the mechanisms behind their beneficial effect are still unclear. M3B2 borides, with M ¼ Mo, are usually present in small quantities. Other trace elements, including O, N, S, and Pb, are closely controlled to not exceed parts-per-million levels to minimize property degradation. Both Fe-Ni and Co-base alloys are similar to Ni-base alloys in having FCC matrices. Fe-Ni base alloys can be simple solid solutions or can be precipitation strengthened. Both g 0 and g 00 (Ni3Ti,Nb) precipitates are found in IN-718. Cobase superalloys rely on solid solution strengthening but also contain higher amounts of carbides for both strengthening and grain size control.

2.5.2 SUPERALLOY PROCESSING Alloy design is inextricably linked with processing development as well as the intended application. The strongest, most creep-resistant alloys are too strong to be amenable to wrought processing, such as extrusion, forging, or rolling, even near their melting temperatures. Furthermore, these alloys usually cannot be welded by most commercial welding techniques. Thus, they must be processed by casting to near-net shape. Other alloys exist that are specifically designed for wrought processing, for application as a sheet product, for weldability, and so on. Aerospace-grade superalloys are processed with a strong emphasis on tight compositional control, high purity, and low inclusion contents. Vacuum melting is ubiquitous across the industry. Many components, especially those with large size and modest high-temperature strength requirements, are made by cast and wrought processing. Combinations of vacuum induction melting (VIM) and vacuum arc remelting (VAR) produce refined ingots ranging up to several tons in size. These ingots are then broken down into bar, sheet, and tube product by forging, extrusion, and other hot working methods. Cast and wrought alloys are limited to those alloys with enough ductility to allow substantial deformation without cracking. The degree of alloying is also limited by dendritic segregation, which can be severe enough in large ingots to induce cracking and substantially limit the deformability. Powder metallurgy (PM) processing can extend the limits of wrought processing by producing a homogeneous, finegrained billet with improved workability compared to ingot metallurgy, thus allowing more heavily alloyed material to be processed. Isothermal forging is another technique used to extend the degree of workability of high-strength alloys. In this method, quenching of the billet by the forging dies is minimized by supplemental heat used to heat both the billet and the die. Alloys used for turbine airfoils are usually too strong to be hot worked. In addition, many advanced turbine blades have intricate geometries, including internal cooling passages that are best achieved by near-net-shape casting

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65

methods. Investment casting, sometimes known as the lost-wax method, has been developed to produce extremely complex geometries with excellent dimensional tolerance, smooth surface finish, and low defect rates. In this method, the component to be cast is first produced in wax. The wax parts are then dipped in ceramic slurries to produce a shell mold, which is then heated to melt out the wax and sinter the ceramic shell. The alloy to be cast, in the form of pre-alloyed VIM ingot, is melted in a ceramic crucible and poured into the shell mold under vacuum. If the part has internal cooling passages, ceramic cores are set into the wax preform and remain in the shell mold after the wax is removed. The cores are then removed by chemical leaching from the solid castings. This process is known as conventional casting and produces large, equiaxed grained microstructures with grain sizes on the order of 1–3 mm. Specialized, proprietary methods have been developed to refine the grain size of these castings. The most highly stressed turbine blades are produced by directional solidification. In this method, the molds and melt stock are prepared similarly to conventional casting. However, the mold is heated to maintain the alloy as a liquid, and after pouring, the mold is steadily withdrawn from the hot zone at a rate of approximately 20 cm/h. This produces directionally solidified (DS) turbine blades with elongated grains that extend the entire casting length. The lack of transverse grain boundaries in this structure provides a substantial creep strength benefit at the high temperatures seen in turbine airfoils. A further advance of this technique was introduced in the 1980s, where various geometric constraints and/or single crystal seeds were inserted in the mold to restrict multiple grains from forming, thereby producing single crystal (SX) superalloys. SX alloys offer the best high-temperature creep resistance, although their relatively high cost restricts them to components with the highest stress/ temperature combinations. Superalloys are almost always heat treated before use. Cast alloys are usually solution treated above the g 0 solvus, which both dissolves the coarse g 0 and homogenizes the dendritic segregation present in as-cast microstructures. Rapid cooling from the solution temperature is usually prescribed to maintain a fine optimizing g 0 size, stabilizing carbides and relieving residual stresses. Wrought alloys can be processed similarly, although the high-temperature solution treatment may be designed to be slightly below the g 0 solvus. In this case, the small amount of remaining g 0 serves to limit grain size. Applying protective coatings to superalloys is also a mature technology. Turbine airfoils and combustor parts are usually coated with an oxidation- and corrosion-resistant coating. Coatings can be applied by aluminizing, where a part is subjected to an Al vapor, which enriches the surface with enough Al to form a 25- to 100-mm-thick layer of NiAl. A thin Pt layer can be applied before aluminizing, which adds Pt in solid solution in the NiAl coating. Alternate methods such as plasma spraying can deposit overlay coatings with compositions such as NiCoCrAlY. These coatings offer better resistance to hot corrosion compared to NiAl. Finally, thermal barrier coatings (TBC) are applied on internally

66

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cooled blades and combustors. Yttria-stabilized zirconia is the predominant TBC and is applied via either plasma spray or electron beam physical vapor deposition. See Section 2.12 for details on TBC.

2.5.3 SUPERALLOY PROPERTIES AND APPLICATIONS The two prime factors determining the mechanical properties of a superalloy are its composition and grain size. Composition determines the degree of solid solution and precipitation hardening. Many factors influence the optimization of precipitation hardening, but the most significant include g 0 volume fraction, g 0 particle size, and g–g 0 lattice parameter mismatch. Grain size is also of prime importance. Fine grains are preferred for their high strength and fatigue resistance up to about 7008C. At higher temperatures, creep becomes dominant and coarse grains, elongated grains, and single crystals provide good, better, and best creep resistance, respectively. Table 2.13 lists common superalloys used in aerospace and their compositions. Turbine engines are the propulsion systems most closely associated with superalloys. In the front sections of a turbine engine, the air temperature is initially low, and it gradually rises as it is compressed. The components in these sections tend to be Al or Ti alloys and, most recently, polymer matrix composites, with the selection of Ti alloys becoming more common where higher strength and warm temperatures are needed. When the temperatures exceed approximately 4008C, superalloys are selected. The Fe-Ni base alloy IN-718 is used in numerous compressor disks and blades, usually in cast and wrought form. The strength of IN-718 begins to drop as the temperature exceeds 5508C, which occurs in the last stages of the compressor, requiring a shift to stronger Ni-base alloys, especially for blades. The air is then mixed with fuel and burned in the combustor, which is a very-high-temperature but low-stress environment. In this case, oxidation resistance and processability are preferred attributes, and many relatively weak alloys that can be processed into sheet are employed. The combustor is followed by the high-pressure turbine section, usually requiring the most advanced disk and blade alloys. For the disks, the strongest cast and wrought alloys, such as Udimet 720, are most often selected, although advanced PM alloys are displacing them in modern engines. DS and SX blade alloys are used routinely in modern engines. Beyond the high-pressure turbine, temperatures begin to decrease and a gradual transition to lower temperature components is implemented in the low-pressure turbine and exhaust nozzle. Superalloys are also employed in many rocket engine components. The turbopumps used to raise the pressure of the liquid oxygen and fuel are analogous to the turbines in a jet engine, and similar materials are used for the turbopump blades, disks, and housings. IN-718 and SX superalloys are examples. If hydrogen is the fuel, alloy strength may need to be sacrificed to ensure resistance to hydrogen embrittlement.

COMPOSITIONS (IN WT %) OF TYPICAL SUPERALLOYS

Ni

Fe

Cr

Co

Al

Ti

Nb

Mo

Ta

Re

Hf

IN-718

Bal

18.5

18.5

—

0.5

0.9

5.1

3.0

—

—

—

—

0.040

Ni-base cast & wrought alloy

Waspalloy

Bal

2.0

19.5

13.5

1.4

3.0

—

4.3

—

—

—

—

0.070

0.006

0.070

—

—

Ni-base cast & wrought alloy

Udimet 720

Bal

—

18.0

14.7

2.5

5.0

—

3.0

1.3

—

—

—

0.030

0.033

0.030

—

—

Ni-base powder alloy

Rene 88DT

Bal

—

16.0

13.0

2.1

3.7

0.7

4.0

4.0

—

—

—

0.0

0.0

—

—

Ni-base cast alloy

IN-738

Bal

—

16.0

8.5

3.4

3.4

0.9

1.8

2.6

1.8

—

0.170

0.010

0.100

—

—

Ni-base cast alloy

Mar-M247

Bal

—

8.5

10.0

5.5

1.5

—

0.7

10.0

3.0

—

0.160

0.015

0.040

—

—

Ni-base single crystal alloy

CMSX-4

Bal

—

6.5

10.0

5.6

1.0

—

0.6

6.0

6.0

Co-base sheet alloy

Haynes 188

22

3.0

22.0

Ni-Fe-base sheet alloy

Hastelloy-X

Bal

18.5

22.0

Ni-base sheet alloy

Inconel 625

Bal

2.5

21.5

Bal 1.5 —

—

—

—

14.0

—

—

—

9.0

0.2

0.2

3.6

9.0

W

— 0.6 —

3.0

1.4 0.1

C

—

B —

Zr —

—

Mn

Si

0.2

0.2

—

—

—

—

—

—

—

—

—

0.100

—

—

—

—

—

0.100

—

—

0.5

0.5

—

—

—

0.050

—

—

0.2

0.2

67

Alloy Ni-Fe base cast & wrought alloy

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.13

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SUGGESTED READING The following resources provide good background material. The International Symposia on Superalloys has been held every four years since 1968, most recently in 2016. The proceedings from these valuable conferences, published by The Minerals, Metals & Materials Society, are also good resources.

BIBLIOGRAPHY Reed, R. C., The Superalloys: Fundamentals and Applications, Cambridge Univ. Press, New York, 2008. Sims, C. T., Stolloff, N. S., and Hagel, W. C., Superalloys II: High-Temperature Materials for Aerospace and Industrial Power, Wiley, New York, 1997.

AEROSPACE MATERIALS CHARACTERISTICS

69

2.6 COPPER ALLOYS David L. Ellis NASA John H. Glenn Research Center at Lewis Field

2.6.1 INTRODUCTION Copper-based alloys have unique electrical, thermal, and wear characteristics that make them well suited for certain aerospace applications. Copper-based alloys used for aircraft applications include beryllium copper (C17000, C17100, C17200), nickel aluminum bronzes (C63000, C63020, C95510), silicon (C65100, C65500) and silicon aluminum (C64200) bronzes, manganese bronzes (C67300, C86300), high lead tin bronze (C93700), and aluminum bronze (C95400) [1]. Several standards are applicable to Cu-based alloys in the aerospace industry, including BS B 23, AMS 4640, AMS 4880, AMS 4881, and UA11N [1]. Copper alloys are used in regeneratively cooled rocket engines where pure copper, low-alloy coppers (Cu-Cr, Cu-Zr, GlidCop AL-15, etc.), and specially developed Cu-based alloys (NARloy-Z, GRCop-84) are used. All of the alloys are notable for high thermal conductivity and good elevated temperature mechanical properties. Another important application of Cu and Cu-based alloys is electronics and electrical systems. Electrical wiring, electrical motors, electrical actuators, and so on normally use Cu wire. Cu-based materials are frequently used for cold plates and thermal backplanes of standard electronic modules (SEMs) and thermal control of electronics such as heat sinks for electronics chips.

2.6.2 AIRCRAFT APPLICATIONS Cu-based alloys offer two primary advantages for aircraft structural applications. They are exceptionally resistant to wear and galling, and they are very strong with high specific properties. Some are also nonsparking and can offer an advantage over other alloys such as steel if sparks are a concern, for example, near fuel where the potential for a fire or explosion exists. This leads to their application in wear surfaces and bearing applications and even as structural members. Table 2.14, taken from Glaeser [2], shows four of the more common types of bearing alloys used for aircraft applications and lists their hardness, strength, and wear factor, a measure of their wear resistance. The tin bronzes exemplified by C90500 (Cu-10 Sn-2 Zn) receive their strength from the hard d phase that also aids in wear resistance. The leaded versions of these alloys, here typified by C93200 (Cu-7 Pb-7 Sn-3 Zn), add lead to provide self-lubrication. The lead also improves ductility and lowers strength, which helps to distribute loads better in a bearing application. Aluminum bronzes such as C95400 (Cu-4 Fe-11 Al) are solid-solution strengthened and are much stronger than the tin bronzes. They are also usable at elevated temperatures (up to 2608C). Beryllium coppers This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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TABLE 2.14

TYPICAL PROPERTIES OF FOUR BEARING ALLOYS [2]

Hardness, BHN

Approximate Yield Strength, MPa

Wear Factor,a mm3/ m  10212

Bearing Pressure Range,b MPa

Leaded tin bronze (C93200)

65

120

6.4 33.3

0–14 14–40

Tin bronze (C90500)

75

150

2.7 13.4

0–14 14–40

Aluminum bronze (C95400)

170

370

1.3 6.7

0–100 100–200

Beryllium copper (C82500)

380

790

1.1

0–550

Material

a

Wear factor based upon volume of wear of a cylindrical plain bearing, grease lubricated, operating at slow speed over a given number of cycles. Bearing pressure is defined as the radial load divided by the length  diameter product.

b

are precipitation strengthened by the b Be phase and have strengths that rival those of steels combined with an upper operating temperature around 3158C. Alloy 25 (C17200, Cu-1.9 Be-0.3 Co) and Alloy 20 (C82500, Cu-2.2 Be-0.6 Co) are the two most common beryllium copper alloys used in aerospace applications. In addition to the four bearing alloys highlighted in Table 2.14, many other Cu-based alloys find use in aircraft. Some of these are listed in Table 2.15 along with typical properties and typical applications. Additional information on Cu-based alloys’ properties and applications can be found in Glaeser and online [2, 3]. Landing gear bushings, wear surfaces, bearings, and even the entire landing gear structure can be made from beryllium copper [5], nickel aluminum bronze [1], and manganese bronze [1]. In these applications, the high strength of the alloys provides the load-bearing capability while the low friction and high wear resistance prevent galling and wear damage to the parts through repeated landings. Bearings and bearing cages made from silicon bronze, manganese bronze, and aluminum bronze are frequently used in aircraft due to their wear resistance and good strength [1]. Silicon bronzes provide self-lubrication from the Si, whereas sintered bronzes are particularly good for use with oil and other liquid lubricants [1]. One interesting application for future hypersonic aircraft may be actively cooled structures such as leading edges. The National Aerospace Plane (NASP)

TYPICAL ROOM-TEMPERATURE PROPERTIES AND APPLICATIONS OF SELECTED Cu-BASED ALLOYS USED IN AIRCRAFT [3, 4]

Material

Alloy

Condition/ Temper

Density, g/cm3

Thermal Conductivity, W/m-K

BHN

Elastic Modulus, GPa

Aluminum bronze

C95400

M01

7.45

58.7

170

107.0

—

Berylium copper

Modulus of Rigidity, GPa

Yield Strength, MPa

Ultimate Tensile Strength, MPa

Elongation, %

Aerospace Applications

241

586

18

Bearings, wear plates

C17000

TH04

8.41

107.3

—

128.0

50.3

1000

1310

2

C17200

TH04

8.25

107.3

—

128.0

50.3

1000

1344

4

Bearings, journals, wear surfaces

High lead tin bronzes

C93700

M01

8.86

46.9

60

75.8

—

110

124

20

High-speed and heavy load ampliations

Manganese bronze

C67300

H50

8.30

95.0

—

117.2

—

276

448

11

C86300

M01

7.83

35.5

225

97.9

—

427

821

18

High load gear and bearing application

Nickel aluminum bronze

C63000

HR50

7.58

39.1

—

121.0

44.1

517

814

15

C63200

050

7.64

34.6

—

117.0

44.1

365

724

22

Silicon aluminum bronze

C64200

H04

7.70

45.0

—

110.0

41.4

469

703

22

Valve stems,gears, marine and pole line hardware, wing pylons, bolts, nuts, bushings and valve body components

Silicon bronze

C65100

H04

8.75

57.1

—

117.0

44.1

379

483

15

C65500

H04

8.53

36.3

—

103.4

38.6

379

634

22

Bearings, ball bearing cage assemblies

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.15

Aircraft landing gear bushings, bearings, and other military and aircraft

71

72

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examined such structures. Copper alloys such as those used in regeneratively cooled rocket engines and Cu-based composites were considered likely candidates for the actively cooled structures [6]. NASP also chose NARloy-Z (Cu-3 Ag-0.5 Zr) for the combustor section heat exchangers [7].

2.6.3 SPACECRAFT APPLICATIONS Copper and Cu-based alloys appear in several locations in regeneratively cooled rocket engines. Most notably, they are used for the liners of the main combustion chamber. In the case of the space shuttle main engine (SSME), NARloy-Z was used to provide structural containment of the 20 MPa chamber pressure and approximately 30008C flame [8]. This was achieved through extensive active cooling of the liner, which relies upon the high thermal conductivity of NARloy-Z and the good elevated temperature strength of the alloy [9]. Other high-copper alloys, such as the ones listed in Table 2.16, have potential for rocket engine liner applications. Additional information on several alloys is available in work by Robinson [4]. For hydrogen-fueled spacecraft, hydrogen embrittlement is a concern. Copper is often used as a hydrogen barrier to prevent degradation of Ni-based and other alloys susceptible to embrittlement [8]. The SSME also uses copper for the preburner baffles and partial coating of the main fuel valve housing [8].

2.6.4 ELECTRONICS AND ELECTRICAL APPLICATIONS Both airplanes and spacecraft use electronics extensively. Weight and space constraints demand very compact packaging. This results in very high power densities and large amounts of heat generated in a confined space. One common method of removing the heat found on the military’s standard electronics modules is to mount the electronics on a cold plate that takes the heat to a thermal backplane for removal from the system [12]. Cu-Invar composite plates with the thermal expansion coefficient matched to the electronics packaging can be used for the cold plate. Other potential materials include Cu metal matrix composites using high-conductivity graphite fibers, diamond, or carbon nanotubes. These same materials can be used for heat sinks if a cold plate and thermal backplane system is not used. Another area where beryllium copper has found use is instrumentation cages [5]. Beryllium copper provides a nonmagnetic, high-strength cage to hold electronics and instruments. The machinability and formability of beryllium copper also allows the creation of conformal cages to match the fuselage and other contours as needed. New ground is being broken with high-thermal-conductivity Cu-based composites with additions of diamond, graphene, and other high-conductivity phases. Some of these materials are commercially available, with diamond-reinforced copper heat spreaders achieving 800 W/m-K, twice the conductivity of pure

Material

GRCop-84

Alloy

—

TYPICAL ROOM-TEMPERATURE PROPERTIES OF CURRENT AND POTENTIAL ROCKET ENGINE LINER ALLOYS [10, 11] Condition

Annealed

Composition, Wt %

Coefficeint of Thermal Expansion, 1/K  106

Thermal Conductivity, W/m-K

Compressive Yield Strength, MPa

Tensile Yield Strength, MPa

Ultimate Tensile Strength, MPa

Elongation, %

Cu-6.7 Cr-5.9 Nb

15.311

285.411

3

311.8

196.2

368.0

21.7

3

432.9

464.5

464.5

20.5

GlidCop AL-15

C15715

Hard drawn

Cu-0.3 Al2O3

16.6

365.2

Zirconium copper

C15000

Hard drawn þ aged

Cu-0.15 Zr

16.93

366.93

464.0

501.9

510.9

19.5

Chromium copper

C18200

Hard drawn þ aged

Cu-0.9 Cr

17.63

323.63

450.0

441.8

495.2

18.3

Copper chromium zirconium

C18150

Hard drawn þ aged

Cu-1 Cr-0.1 Zr

16.53

323.93

501.4

549.9

564.4

11.2

Solutioned þ aged

Cu-3 Ag-0.5 Zr

17.2

295.0

192.0

314.0

31.0

NARloy-Z9

—

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.16

73

74

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copper [13]. With the increasing availability of carbon nanotubes and grapheme sheets, higher thermal conductivity copper-based materials are likely to be achieved.

REFERENCES [1]

[2] [3]

[4]

[5]

[6]

[7]

[8]

[9]

[10]

[11] [12]

[13]

“Copper-Based Aerospace Alloys,” Aviation Database.com [online database], http://www.aviation-database.com/Technical_Aviation_Articles/copperbased-aerospace-alloys.html [retrieved 5 April 2012]. Glaeser, W. A., “Wear Properties of Heavy Loaded Copper-Based Bearing Alloys,” Journal of Metals, Vol. 35, No. 10, Oct. 1983, pp. 50–55. “Properties of Wrought and Cast Copper Alloys,” Copper Development Association, 20 Jan. 2012, http://www.copper.org/resources/properties/db/ CDAPropertiesSelectionServlet.jsp?mode¼basic [retrieved 5 April 2012]. Robinson, P., “Properties of Wrought Coppers and Copper Alloys,” Properties and Selection: Nonferrous Alloys and Special-Purpose Materials, Vol. 2, ASM Handbook, ASM International, Materials Park, OH, 1990, pp. 265–345. “Advanced Alloys: Alloy Selection for the Aerospace Industry,” IBC Advanced Alloys, http://www.ibcadvancedalloys.com/clientuploads/Technical%20Resources/ AerospaceWhitepaper.pdf [retrieved 5 April 2012]. McGowan, D. M., Camarda, C. J., and Scotti, S. J., “A Simplified Method for Thermal Analysis of a Cowl Leading Edge Subject to Intense Local Shock-Wave-Interference Heating,” NASA TP-3167, NASA Langley Research Center, Hampton, VA, March 1992. Paul, D. B., “Extreme Environmental Structures,” Structures Technologies for Future Aerospace Systems, edited by A. K. Noor, Vol. 188, Progress in Astronautics and Aeronautics, AIAA, New York, 2000, pp. 145–199. “NASA Relies on Copper for Shuttle Engine,” Discover Copper Online, No. 73, Spring 1992, Copper Development Association, http://www.copper.org/publications/ newsletters/cutopics/Ct73/shuttle_engine.html [retrieved 5 April 2012]. Espisito, J. J., and Zabora, R. F., “Thrust Chamber Life Prediction: Vol. 1, Mechanical and Physical Properties of High Performance Rocket Nozzle Materials,” NASA CR-114806, NASA Lewis Research Center, Cleveland, OH, March 1975, pp. 12–18. DeGroh, H. C., Ellis, D. L., and Loewenthal, W. S. “Comparison of GRCop-84 to Other Cu Alloys with High Thermal Conductivities,” Journal of Material Engineering and Performance, Vol. 17, No. 4, Aug. 2008, pp. 594–606. Ellis, D. L., and Keller, D. K., “Thermophysical Properties of GRCop-84,” NASA CR-210055, NASA Glenn Research Center, Cleveland, OH, June 2000. Ng, L. H., and Field, F. R., “A Quantitative Technique to Analyze Materials Trade-Off for SEM ‘E’,” IEEE Transactions on Components, Hybrids, and Manufacturing Technology, Vol. 15, No. 1, Feb. 1992, pp. 78–86. Sung, J. C., Kan, M.-C., Hu, S.-C., Sung, M., and Monteith, B.G., “Diamond Composite Heat Spreader,” Advanced Diamond Solutions, http://www. advanceddiamond.com/whitepapers/DiamondCompositeHeatSpreader.pdf [retrieved 5 April 2012].

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2.7 DAMAGE TOLERANCE CONSIDERATIONS FOR METALLIC MATERIALS Preston B. McGill NASA Marshall Space Flight Center

2.7.1 INTRODUCTION As the name implies, damage tolerance addresses the sensitivity of a material’s structural or load-carrying capability in the presence of damage or defects. Damage tolerance considerations are applied to components that serve as primary structural members and may include beams, columns, struts, pressurized structure, pressure vessels, pressurized components, fasteners, fittings, clevises, and attach hardware. It includes any structural component whose failure due to a defect would result in a catastrophic event. In metallic materials, damage or defects may be described in terms of cracks, crack-like flaws, or discontinuities in the material. Defects in metallic materials may result from anomalies in the original ingots or billets as well as from discrepant processing of wrought products, castings, or forgings related to thermal processing such as solution heat treatment and quenching, or related to mechanical working such as forging, rolling, or stretching. Though defects do occur in raw materials, more probable are fabrication-related defects from forming and machining operations or joining processes such as welding and brazing metallic materials. Materials initially free of defects may develop defects in service due to cyclic load conditions, corrosion processes, or inadvertent damage. Damage tolerance behavior of a material is typically incorporated into the design, development, and manufacture of a component as part of an overall flight safety approach, sometimes referred to as fracture control, to mitigate risk of failure. Therefore, understanding the damage tolerance capability of materials is one element of an overall fracture control program for safety-critical hardware. In addition to damage tolerance of the material, fracture control generally encompasses system-wide disciplines in design, analysis, nondestructive evaluation, quality control, process control, and materials selection. This section highlights considerations for metallic materials selection with respect to damage tolerant behavior. Detailed information regarding damage tolerance design considerations in terms of fracture control requirements, implementation, guidelines, and fitness for service may be found in NASA-STD-5019A [1], NASA-HDBK-5010 [2], JSSG-2006 [3], and the American Petroleum Institute 579-1/American Society of Mechanical Engineers FFS-1, Fitness for Service [4].

2.7.2 DAMAGE-TOLERANCE BEHAVIOR OF METALLIC MATERIALS Numerous texts and papers have been written to describe methods for analyzing, testing, and characterizing metallic material response in the presence of a crack or This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

76

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crack-like defect, that is, the discipline of fracture mechanics, which provides the framework for understanding the fracture behavior of metallic materials [5–8]. Four areas related to the fracture, or damage tolerance, behavior of metallic materials are summarized in this section: fracture toughness, stable crack extension, crack growth rate, and transferability. 2.7.2.1

FRACTURE TOUGHNESS

The primary parameter used to describe the fracture behavior of a metallic material is its fracture toughness. In its simplest form, the fracture toughness of a material is a measure of the material’s ability to resist failure in the presence of a crack. Physically, it is a measure of the amount of energy required to extend a crack in the material. In linear-elastic fracture mechanics, the fracture toughness is most commonly designated as KIc but may also be denoted as Kc, KIe, KJq, or KJIc depending on the test methodology. In elastic-plastic (nonlinear) fracture mechanics, the fracture toughness is most commonly designated as Jq or JIc. Note all of these designations represent a single, scalar parameter characterization of the fracture toughness of the material; note also the subscripts reflect various assumptions regarding sample geometry, flaw geometry, stress state, and loading configuration. Standard test methods published by the International Standards Organization (ISO) and, more commonly used in the United States, ASTM International (formerly the American Society for Testing and Materials), provide methods for measuring the fracture toughness of metallic materials [9–13]. Zhu and Joyce [14] provide a comprehensive summary and technical review of various fracture toughness test methods. Note that fracture toughness values may vary, even within a given alloy, due to heat treat condition, product form, grain orientation, location within the cross section, service environment, and geometric constraint conditions. Papers, reference handbooks [15–18], textbooks [19], and material specifications may be consulted for fracture toughness values for specific alloys. Although fracture toughness is of primary importance when evaluating a material for damage-tolerant applications, some additional considerations in connection with fracture toughness should be included in the selection process. Conceptually, additional insight into metallic materials selection with respect to fracture toughness may be gleaned from consideration of the ideal case of a through crack in an infinite plate under a remote, uniform, uniaxial stress. A sketch of this configuration is shown in Fig. 2.28. For the crack opening mode in the direction of the applied stress, the linear elastic stress intensity, K, at the tip of the through crack is given by pffiffiffiffiffiffi (2:1) K ¼ s pc where s is the remote uniform stress, and c is one-half of the through-crack length. The limiting value for the stress intensity is the critical stress intensity value, or fracture toughness, of the material KIc. If the material is stressed at or

AEROSPACE MATERIALS CHARACTERISTICS

77

Fig. 2.28 Through-crack in an infinite plate.

near the yield stress, sy, based on linear-elastic fracture mechanics, a limiting estimate for the critical flaw length may be calculated as   2 KIc 2 (2:2) (2c)crit ¼ p sy Note the critical flaw length varies with the square of the ratio of KIc to sy. Low values of this ratio can result in small critical flaw sizes that may be difficult to reliably detect. This simplified assessment leads to a general conclusion that the toughness to yield strength ratio should be considered when evaluating the damage tolerance capability of a material, with preference given to materials with higher ratios. 2.7.2.2

STABLE CRACK EXTENSION

Another important material characteristic is the ability of the material to tear, or undergo crack extension, in a stable manner. Common metallic materials fracture toughness test methods, such as ASTM E399 [9] and ASTM E1820 [12], which measure fracture toughness by loading a test specimen with an induced crack and by measuring the load and displacement during the test, determine the point of initiation of crack extension—the fracture toughness. The behavior of the material following the initiation of crack extension is important to the realized damage tolerance of structures. The crack may continue to extend in a stable fashion, requiring further increases in load to extend the crack, or the crack may fail the material instantaneously with unstable crack extension. In these displacement-controlled tests, the ability of the material to sustain crack extension (tearing) in a stable manner may be inferred from the load versus displacement plot. Two different behaviors for the same alloy processed to two different yield strength levels are illustrated in Figs. 2.29 and 2.30. The behavior of the material is shown notionally in the cartoons depicting the compact tension sample. In Fig. 2.29, the higher yield strength material fails in an unstable manner (fast fracture) at or near the initial crack extension. In Fig. 2.30, the crack grows in a

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stable manner after the initial crack extension. In this case, the load increases as the crack extends before unstable fracture. Material exhibiting unstable fracture behavior may be adequate for the design, that is, the toughness may be such that the critical initial flaw can be reliably detected. However, material exhibiting a stable tearing failure mode will offer greater capability for load redistribution and is the preferred failure mode. It is important to note that factors such as environment and heat treatment can result in a change in failure mode for a given alloy. For example, a transition from ambient temperature to cryogenic temperature in titanium alloys can result in an accompanying transition from a stable mode at failure to an unstable mode. Similarly, changing temper conditions of a precipitation-hardened steel, for example, from H1100 to H900, can also result in a change in failure mode from stable to unstable. 2.7.2.3

CRACK GROWTH RATE

Crack growth rate is a measure of crack extension under cyclic loads. This is expressed in terms of crack extension/cycle (da/dN) as a function of the stress

Fig. 2.29

Unstable fracture.

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Fig. 2.30

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Stable crack growth before fracture.

intensity range (DK). Crack growth rates also vary as a function of cyclic stress intensity ratio, (Kmin/Kmax). Papers, reference handbooks [15–18], textbooks [19], and commercial safe life analysis tools [20, 21] may be consulted for crack growth rate curves for specific alloys and environments. Crack growth rate curves for some aluminum alloys are shown in Sec. 2.2. Standardized test procedures for measuring crack growth rates are provided in ASTM E647 [22]. As with fracture toughness, crack growth rates will vary with alloy, heat treat condition, product form, orientation, service environment, and constraint conditions. Consequently, da/dN-DK curves appropriate to the application must be implemented. 2.7.2.4

TRANSFERABILITY

The fracture toughness in terms of the critical stress intensity value is a function of the constraint conditions at the crack tip. Constraint conditions are a function of the stress state in the vicinity of the crack and vary with test sample geometry. Transferability addresses the extent to which the test results can be used to predict behavior of the structure; in this regard it is important that the constraint conditions in the test sample match or conservatively bound the conditions in the structure. Highly constrained material is not free to plastically deform or flow, and the material may reach very high, localized stresses that

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promote failure. Less constrained material is free to plastically deform or flow (in ductile materials), which may result in crack tip blunting or the development of plastic zones well ahead of the crack tip and may reflect toughness values that are high with respect to the structural application. There are a variety of test sample configurations used to evaluate fracture toughness. Three common test sample geometries used in generating fracture data are compact tension samples, surface crack tension samples, and middle tension samples; they are shown Fig. 2.31. Compact tension samples provide high-constraint conditions. Test samples with larger ligaments, larger thicknesses, side grooves, or a combination of these characteristics tend to provide high constraint and trend toward lower bound toughness properties. Conversely, middle tension samples, that is, samples with through cracks, tend to have low constraint at the crack tip and may result in critical stress intensity measures that are higher than what can be developed in the structure. Surface crack tension samples, samples with partially through flaws that are semicircular or semielliptical in shape, have constraint conditions that vary around the crack length. Critical stress intensity values based on these tests may vary with crack geometry and material thickness. When reviewing fracture toughness and crack growth rate data, the test sample configuration and the crack geometry should be considered when evaluating the data for applicability to the structure. Test data with similar or conservative (higher) constraint conditions

a) Compact tension

Fig. 2.31

b) Surface crack tension

c) Middle tension

Fracture toughness test sample configurations.

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should be used to assess the damage-tolerant behavior of the structure. Wallin [5], Kannenin and Popelar [23], and Barsom and Rolfe [24] provide general considerations for evaluating transferability of test results and structural applications. Xia and Shih [25] provide a detailed discussion on constraint effects in ductile materials.

2.7.3 FACTORS THAT AFFECT DAMAGE-TOLERANCE PROPERTIES OF METALLIC MATERIALS Damage tolerance properties generally vary with service condition, material processing, material thickness, and grain orientation. Considerations related to each of these factors are discussed. 2.7.3.1

SERVICE CONDITIONS

Material compatibility with service conditions is an important consideration when selecting any material. Along with other mechanical properties, the fracture behavior of a material may also be affected by its service environment. Temperature and exposure to gas, liquid, or solid media may adversely affect the damage-tolerant capability of the material. With respect to temperature, for example, some high-strength steel alloys and titanium alloys exhibit large cryogenic yield strength enhancements that make the alloys attractive for cryogenic applications from a strength design perspective. However, many of these same alloys also exhibit diminished fracture toughness characteristics at cryogenic temperatures. Based on the linear-elastic fracture behavior discussed earlier, the combination of strength and fracture toughness changes at cryogenic service temperatures can result in critical flaw sizes that cannot be reliably detected by nondestructive evaluation. Conversely, some aluminum alloys exhibit increases in fracture toughness and strength with cryogenic temperatures. This may improve the feasibility of employing a mechanical proof test at ambient temperatures as an effective screen for critical defects at cryogenic operating conditions. Although there is no direct analytical relationship between fracture toughness and ductile to brittle transition test data, ductile to brittle transition data can be a good starting point for qualitative evaluation of material fracture behavior as a function of temperature, particularly for high-strength steels. Materials in critical applications should be used at temperatures above their ductile to brittle transition temperature. The effects of cumulative exposure at service temperatures should also be considered for material that experiences cyclic use at or near the aging or precipitation heat-treat temperatures of the material. Repeated exposure at these temperatures can lead to changes in material toughness, ductility, and crack growth rates. Aluminum alloys can be particularly susceptible because their aging temperature is relatively low.

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Specific damage-tolerance evaluation should be made for materials in service environments that may degrade material performance, such as hydrogen-rich environments (plating or operational environments; see Sec. 2.8), low-pH environments, chlorine-rich environments, and liquid metal or solid metal embrittling environments. Test methods for determining these effects are detailed in ASTM E1681 [26] and ASTM G129 [27]. These tests may be used to determine environmentally assisted cracking threshold stress intensity factors (KEAC, sometimes referred to as KIscc) for metallic materials. Although there is no direct correlation to KEAC behavior, MSFC-STD-3029 [28] provides qualitative data and guidelines for evaluating material susceptibility to stress corrosion cracking in a sodium chloride environment. These data may be used as a starting point for gauging material response in adverse service environments. Particular caution should be exercised for alloys with moderate or high susceptibility to stress corrosion cracking. 2.7.3.2

PROCESSING

Metallic materials used in the manufacture of parts in critical applications should be purchased per standard aerospace industry material specifications (e.g., Society of Automotive Engineers, ASTM International, ISO, National Aerospace Standards). With respect to mechanical properties, these specifications define minimum room-temperature strength and ductility levels. Some material specifications define minimum fracture toughness levels or provide the option for fracture toughness requirements. To mitigate risk associated with fracture failures, some applications may warrant lot acceptance testing and/or inspection beyond what is required by specification. With respect to testing, it may be prudent to implement fracture toughness lot acceptance testing for alloys exhibiting a strong inverse relationship between strength and fracture toughness, for example, precipitation hardening steels and aluminum-lithium alloys. With respect to inspections, consideration may be given to purchasing material with raw stock nondestructive evaluation (NDE) that interrogates the internal (volumetric) integrity of the material. This typically consists of quantitative ultrasonic and/or radiographic inspections to screen for volumetric defects in the material. Quantitative NDE refers to inspection procedures that have been statistically assessed to be capable of finding a given flaw size with a defined probability of detection and confidence level. Guidelines for nondestructive evaluation may be found in NASA-STD-5009 [29]. Texts by Shull [30], Cartz [31], and ASM International [32] provide technical background and applications for various NDE methods. Consideration should be given to variation in fracture toughness and crack growth rates with product form [2]. Plate, sheet, forgings, extrusions, castings, and spun or mechanically formed products of the same alloy will generally exhibit enough property variation to warrant data for each product form. For example, castings tend to exhibit lower properties and more variability than

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wrought material. Toughness properties may vary with location and product form geometry for a given casting, forging, extrusion, plate or formed product. For example, mechanical properties at the apex of a spun formed dome will likely vary from the properties at the equator. Consideration should be given to metallographic evaluation, tensile testing, and fracture testing of areas with complex geometry, large section thickness, or critical stress locations. Additionally, fracture properties will vary with thermal treatment and mechanical processing condition for a given alloy. The amount of cold work, the precipitation-hardened condition, quench rate, temper, and age condition all affect the strength, ductility, stress corrosion capability, and fracture toughness of an alloy. For example, age practice (overaging or underaging) of aluminum alloys can affect the strength, ductility, and fracture toughness of the material for a given temper (see Sec. 2.2). Annealed titanium has better toughness properties than solution-treated and aged titanium (see Sec. 2.3). Similarly, the interstitial content (chemical composition) for a given titanium alloy can affect the fracture toughness of the material. The ductility and fracture toughness of precipitation-hardened stainless steels varies significantly with the temper condition. With respect to applications, it is important to ensure the fracture data are matched to the material process, chemical composition, and thermomechanical processing of the alloy [2]. 2.7.3.3

THICKNESS

Fracture properties in metallic materials may vary with thickness and location within the thickness. Short transverse (S-T) ductility, fracture toughness, and crack growth rate properties tend to degrade with increasing plate thickness. Also, the in-plane properties may vary with location within the thickness of the plate (t), that is, material at the t/2 location may have properties that are different from material at the t/6 location. Through thickness microstructural variations that affect toughness properties may be influenced by the through hardenability characteristics of the material, such as the quench rate sensitivity in steels and the degree of cold work in aluminums. With respect to material selection, it is important that material representative of the thickness and product form of the raw stock used to manufacture the part is used in the characterization of the material. For example, a 6-mm-thick aluminum membrane machined from 50-mm-thick aluminum plate will likely exhibit different fracture properties than 6-mm-thick aluminum sheet. 2.7.3.4

ORIENTATION

Fracture properties vary with grain orientation depending on the degree of anisotropy in the material [2]. In general, for thin plate products, T-L properties (specimen loaded in the long transverse direction with the crack growing in the longitudinal direction) will be the lowest. However, materials with a high

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degree of anisotropy should be evaluated in the T-L, L-T, and 458 directions to determine the worst-case toughness orientation. In thick-plate products, S-T properties are generally the lowest and through thickness material characterization is required. With respect to materials selection, there is a tendency to use thick-plate product forms for machining fittings, fixtures, lugs, threaded blind holes, and so on. Often these will be loaded in the S-T or ST-45 directions. In these cases, S-T or ST-45 characterization of fracture toughness and ductility should be conducted. Often a forged product will provide better short transverse fracture toughness properties than thick plate and should be considered in design.

2.7.4 AVAILABILITY OF DAMAGE-TOLERANCE DATA Of practical importance is the availability of data to support the damage-tolerance assessment. Generating damage-tolerance data, particularly in the appropriate service environment, can be expensive and time-consuming. When selecting a material for a safety-critical application that may require damage tolerance data, resource estimates should include availability of existing data and the cost of generating test data.

2.7.5 SUMMARY Implementation of damage tolerant design practices for primary or safetycritical structures requires evaluation of damage tolerance performance of the material. Important considerations when evaluating metallic materials for damage-tolerance capability include fracture toughness, fracture toughness to yield strength ratio, crack growth rate behavior, tearing capability, and transferability of test data to the structural configuration. Additionally, parameters that may affect the damage tolerance behavior of metallic materials include service conditions (temperature, environment), thermo-mechanical processing, thickness, product form, and grain orientation. Finally, when selecting a metallic material for hardware requiring a damage tolerance assessment, some consideration should be given to the availability of relevant data and the resources required to develop the necessary data.

REFERENCES [1] NASA-STD-5019A, Fracture Control Requirements for Spaceflight Hardware, Revision A, January 2016. [2] NASA-HDBK-5010, Fracture Control Implementation Handbook for Payloads, Experiments, and Similar Hardware, May 2005. [3] JSSG-2006, Department of Defense Joint Service Specification Guide Aircraft Structures, Oct. 1998.

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[4] [5] [6] [7] [8] [9] [10] [11] [12] [13]

[14]

[15] [16]

[17]

[18] [19] [20] [21] [22] [23] [24]

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API 579-1/ASME FFS-1, Fitness-for-Service, American Petroleum Institute and the American Society of Mechanical Engineers, 2016. Wallin, K., Fracture Toughness of Engineering Materials, Estimation and Application, EMAS Publishing, Warrington, U.K., 2011. Saxena, A., Nonlinear Fracture Mechanics for Engineers, CRC Press, Boca Raton, FL, 1998. Suresh, S., Fatigue of Materials, 2nd ed., Cambridge University Press, New York, 1998. Anderson, T. L., Fracture Mechanics, Fundamentals and Applications, 3rd ed., CRC Press, Taylor and Francis Group, Boca Raton, FL, 2005. ASTM Standard E399, Test Method for Linear-Elastic Plane-Strain Fracture Toughness KIc of Metallic Materials, ASTM International, West Conshohocken, PA. ASTM Standard B645, Standard Practice for Linear-Elastic Plane-Strain Fracture Toughness of Aluminum Alloys, ASTM International, West Conshohocken, PA. ASTM Standard E740, Practice for Testing with Surface-Crack Tension Specimens, ASTM International, West Conshohocken, PA. ASTM Standard E1820, Test Method for Measurement of Fracture Toughness, ASTM International, West Conshohocken, PA. ASTM Standard E2899, Standard Test Method for Measurement of Initiation Toughness in Surface Cracks Under Tension and Bending, ASTM International, West Conshohocken, PA. Zhu, X.-K., and Joyce, J. A., “Review of Fracture Toughness (G, K, J, CTOD, CTOA) Testing and Standardization,” Engineering Fracture Mechanics, Vol. 85, 2012, pp. 1–46. MMPDS-07, Metallic Materials Properties Development and Standardization, Federal Aviation Administration, April 2008. Aerospace Structural Metals Handbook, 39th ed., Center for Information and Numerical Data Analysis and Synthesis, Purdue Research Foundation, West Lafayette, IN, 1999. Damage Tolerant Design Handbook, Center for Information and Numerical Data Analysis and Synthesis, Purdue Research Foundation, West Lafayette, IN, 1994. Wilson, A. D., et al., “Fatigue and Fracture,” ASM Handbook, Vol. 19, ASM International, Materials Park, OH, 1996, pp. 589–882. Sanford, R. J., Principles of Fracture Mechanics, Prentice-Hall, Upper Saddle River, NJ, 2003. NASGROw, Fracture Mechanics and Fatigue Crack Growth Analysis Software, Southwest Research Institute, San Antonio, TX, 2012. AFGRO, Fracture Mechanics and Fatigue Crack Growth Analysis Software, LexTech Inc., Centerville, OH, 2010. ASTM Standard E647, Test Method for Measurement of Fatigue Crack Growth Rates, ASTM International, West Conshohocken, PA. Kannenin, M. K., and Popelar, C. H., Advanced Fracture Mechanics, Oxford Univ. Press, Oxford, 1985. Barsom, J. M., and Rolfe, S. T., Fracture and Fatigue Control in Structures, Applications of Fracture Mechanics, 3rd ed., ASTM International, West Conshohocken, PA, 1999.

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[25] Xia, L., and Shih, C. F., “Ductile Crack Growth—I. A Numerical Study Using Computational Cells with Microstructurally-Based Length Scales,” Journal of Mechanical Physics of Solids, Vol. 43, No. 2, 1995, pp. 233–259. [26] ASTM Standard E1681, Test Method for Determining Threshold Stress Intensity Factor for Environment-Assisted Cracking of Metallic Materials, ASTM International, West Conshohocken, PA. [27] ASTM Standard G129, Standard Practice for Slow Strain Rate Testing to Evaluate the Susceptibility of Metallic Materials to Environmentally Assisted Cracking, ASTM International, West Conshohocken, PA. [28] MSFC-STD-3029, Revision A, Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments, February 2005. [29] NASA-STD-5009, Nondestructive Evaluation Requirements for Fracture-Critical Metallic Components, April 2008. [30] Shull, P. J., Nondestructive Evaluation, Theory, Techniques and Applications, Marcel Dekker, New York, 2002. [31] Cartz, L., Nondestructive Testing, Radiography, Ultrasonics, Liquid Penetrant, Magnetic Particle, Eddy Current, ASM International, Materials Park, OH, 1995. [32] “Nondestructive Evaluation and Quality Control,” ASM Handbook, Vol. 17, ASM International, Materials Park, OH, 1992.

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2.8 HYDROGEN EMBRITTLEMENT IN METALLIC MATERIALS Jonathan A. Lee NASA Marshall Space Flight Center

2.8.1 INTRODUCTION Hydrogen embrittlement is a phenomenon that manifests as a decrease in the fracture toughness or ductility of a metal in the presence of atomic hydrogen. The reduction of fracture loads can occur at levels well below the yield strength of the material. Hydrogen embrittlement usually manifests as a singular sharp crack in contrast to the extensive branching observed for stress corrosion cracking (SCC). The initial crack openings and the local deformation associated with crack propagation may be so small that they are difficult to detect except in special nondestructive examinations. Cracks due to hydrogen embrittlement can grow rapidly with little macroscopic evidence of mechanical deformation in materials that are normally quite ductile. A good understanding of hydrogen embrittlement is necessary for selecting materials for applications in propulsion systems that use hydrogen as propellant. In this section, a review of experimental data for the effects of gaseous hydrogen environment embrittlement (HEE) is presented for several types of metallic materials with a view to guide materials selection. For aerospace applications, the material screening methods are used to rate the hydrogen degradation of mechanical properties that occur while the material is under an applied stress and exposed to gaseous hydrogen as compared to air or helium, under slow strain rates (SSR) testing. Because of the simplicity and accelerated nature of these tests, the results are expressed in terms of HEE index and are not intended to necessarily represent true hydrogen service environment for long-term exposure, but rather to provide a practical approach for material screening, which is a useful tool to qualitatively evaluate the severity of hydrogen embrittlement. The effects of hydrogen gas on mechanical properties such as tensile, ductility, fatigue, and fracture toughness are analyzed with respect to the general trends established from the HEE index values. It is observed that the severity of the HEE effects is also influenced by environmental factors such as pressure, temperature, and hydrogen gas purity.

2.8.2 CLASSIFICATION OF HYDROGEN EMBRITTLEMENT The interaction between hydrogen and metals can result in the formation of solid solutions of hydrogen in metals, solid compounds as hydrides, and gaseous compounds with other elements in the metal. Hydrogen embrittlement, through these hydrogen–metals interactions, can be classified into three broad categories: hydrogen environmental embrittlement (HEE), internal hydrogen embrittlement (IHE), and hydrogen reaction embrittlement (HRE). In general, HEE represents the This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Fig. 2.32 Classification of HEE, IHE, and HRE type based on hydrogen, stress, and material factors.

condition when the materials are being exposed to a high-pressure gaseous hydrogen environment. The definition of IHE often implies that the source of hydrogen is from an electrochemical process such as electroplating, corrosion, or cathodic charging, and even from thermal charging with gaseous hydrogen at relatively low pressures. However, the HEE and IHE are similar in many instances, and both require an external applied stress in order for the hydrogen embrittlement effects to occur. The definition for HRE is usually irreversible hydrogen damage due to a chemical reaction with hydrogen, and that such damage can occur without an external applied stress. Figure 2.32 shows an overlapped region that forms the HEE, IHE, and HRE. The size of this region graphically represents the severity for hydrogen embrittlement. The size of the overlapped region can increase or decrease depending on how significant the intersections are from the three main circles that represent the influence of material type, hydrogen embrittlement, and applied stress. By definition, the intersections of hydrogen, material, and the stress circle produce the HEE and IHE effects within the same overlapped region positioned at the center of this graph. The difference between HEE and IHE is that the source of hydrogen for IHE is not usually from a high-pressure system; rather, the hydrogen is unintentionally produced, and the internally absorbed hydrogen can result in a time-delayed embrittlement effect under an applied stress. Because the HRE type is formed by the intersection between hydrogen and material, the HRE type can exist without the influence of the applied stress circle. 2.8.2.1

HYDROGEN ENVIRONMENTAL EMBRITTLEMENT

HEE is commonly known as the degradation of certain mechanical properties that occur while the material is under the influence of an applied stress and intentionally exposed to a gaseous hydrogen environment. HEE is the most common type of hydrogen embrittlement encountered for aerospace applications. It is noted that the applied stress values required to cause failure are in tension mode and can stay well below the yield strength for highly susceptible materials. In most cases, the constant static loading is considered to be more sensitive to hydrogen

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embrittlement than cyclic or dynamic loading, particularly at high cycle. The residual stress is also important and must be considered in combination with the applied external stress. Material screening tests using slow strain rates (pseudo-static) or static loading on special designed notched coupons are commonly used to determine a threshold stress value. In a well-characterized condition, such threshold stress values may be used to indicate the maximum allowable stress that can be applied to avoid the HEE effects. A comprehensive review of experimental data for the effects of HEE is presented in this section for several types of metallic materials. 2.8.2.2

INTERNAL HYDROGEN EMBRITTLEMENT

IHE is commonly recognized as the result of the unintentional introduction of hydrogen into susceptible metals during forming or finishing operations. For example, IHE is the absorption of atomic hydrogen from common chemical processes such as acid pickling, electroplating, and corrosion. All of these are electrochemical processes involving the discharge of hydrogen ions. However, IHE can also result from exposing the susceptible material to certain aqueous environments, which is also an electrochemical process involving the discharge of hydrogen ions. The IHE cracks usually occur internally and are located near the root of an internal defect, where the localized stress values are high. However, the distinction between IHE and HEE is not always clear for thermal charging of gaseous hydrogen at relatively high temperature. For example, the IHE effects often associate with the absorption of hydrogen in ambient atmosphere by molten metal during the welding or casting process. Upon rapid cooling of the weld or casting, entrapped hydrogen can produce internal fissuring or other damaging effects that often attribute to IHE. Because of the higher solubility of hydrogen at high temperature, rapid cooling of a heavy section of certain materials operating in high pressure and temperature conditions can result in IHE without having to experience a high level of applied external stress. It is recognized that there are some differences between the high-pressure gaseous environment for HEE and the electrochemical environment for IHE; however, once atomic hydrogen has been absorbed by a material, the hydrogen embrittlement effects for HEE and IHE are similar and have been shown to be the case for a number of materials. 2.8.2.3

HYDROGEN REACTION EMBRITTLEMENT

At certain elevated temperatures and pressures, atomic hydrogen can easily diffuse though metal surfaces and chemically react with certain types of elements and compounds in the material. However, hydrogen can also react with the metal matrix itself to form metallic compounds such as metal hydride at relatively low temperatures. This form of hydrogen damage is known as HRE. It can occur in materials such as titanium, zirconium, and even some types of iron- or steel-based

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alloys. For example, a common reaction is between hydrogen and iron carbides to form gaseous methane (CH4), according to the following chemical reaction: 4H þ Fe3 C ¼ CH4 þ 3Fe

(2:3)

Beneath the surface layer and deep within the bulk of the material, the formation and migration of CH4 gas molecules usually concentrate at grain boundaries. Metallurgical features such as inclusions, impurity, and defects can lead to brittle rupture through the formation of voids, blisters, and a network of discontinuous microcracks. Moreover, because the carbide phase is a reactant in the mechanism, its depletion in the vicinity of generated defects serves as direct evidence of the HRE mechanism itself. The HRE sensitivity depends on the amount of carbon or carbide in the alloy, the hydrogen concentration, gas pressure, and temperature usually in the range of 200–6008C (392–11108F). Alloy steels with stable carbides, such as chromium carbides, are less susceptible to this form of hydrogen attack due to the greater stability of Cr3C vs Fe3C found in carbon steels. However, as the pressure and temperature of the hydrogen environment increase, a greater amount of these alloying additions are required to prevent such attack.

2.8.3 HYDROGEN EMBRITTLEMENT MECHANISMS Over the years, there were many proposed theories to explain the hydrogen embrittlement mechanisms. This section provides a brief review of a promising theory, which is based on a concept that hydrogen alters the density of states (HADOS) of the host metal’s electron band, leading to the embrittlement effect [1, 2]. The basis of the theory is summarized as follows: Because the cohesive strength of a transition metal is governed by the bonding energy of the d-electrons occupied near the Fermi energy level, it is postulated that hydrogen embrittlement is a reduction of the bonding energy of the metal atoms by the introduction of the hydrogen’s electrons into the d-band of the host metal. The quantum value that describes the number of electron states that is available to be occupied by hydrogen is called the electron density of states (DOS) at the Fermi energy level. If there are a limited number of states available for the hydrogen electrons to occupy, then hydrogen embrittlement is expected to be minimal. Conversely, hydrogen embrittlement is likely to increase if the value of the DOS from the host metal is increased. Louthan [3] has shown that hydrogen solubility in transition metals will increase with increasing DOS values. Louthan also suggested that hydrogen embrittlement is at least partially dependent on the solubility of hydrogen atoms in the host metal, a notion that supports the basis for the HADOS model. Interestingly, the effect of alloy compositions on hydrogen embrittlement can be examined only in the light of the HADOS theory, where other hydrogen embrittlement theories cannot provide a satisfactory explanation. It was suggested that hydrogen embrittlement is more pronounced for pure metals with higher DOS values than for metals with lower DOS values. This correlation was based

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on the d-band electron filling and charge transfer between the hydrogen electron and the host metals. However, the potential merit for the HADOS model seemed to gain momentum later under several binary alloys studies using Fe-Ni, Fe-Co, Ni-Co, Ni-W, Cu-Ni, and Pd-Ag. Experimentally, it was found that the hydrogen embrittlement effects also correlated well with the DOS values for these binary alloy systems [1, 2]. These are important binary systems for compositional analysis because they can form a continuous range of compositions, having a singlephase microstructure, subjected to similar heat treatment within that particular binary system, so that a reasonable comparative study of hydrogen embrittlement mechanisms can be made based solely on the alloy compositions. On the research aspect, it is evident from the hydrogen embrittlement literature survey that there are many good technical papers, published mostly for industrial complex alloys, with common repeating themes such as dislocation movements, fracture surface analysis, void and microcrack formations, and microstructure. Most of these are typical examples of hydrogen-material damage studies that take place in the confinement of the “microscopic” scale. However, very few in-depth “cross-discipline” studies have been conducted to understand the key connection between the electronic properties (atomistic scale) and the HEE effects on mechanical properties (macroscopic scale). At this time, it appears that there are currently two popular models for hydrogen embrittlement mechanism based on the mechanical properties observed at the macroscopic scale: the hydrogen enhanced decohesion (HEDE) model and the hydrogen enhanced localized plasticity (HELP) model as reviewed by W. W. Gerberich [4]. However, it is important to make a distinction that the proposed electron charge transfer mechanism from the HADOS theory operates at a much smaller scale than what the HEDE and HELP models deal with. Whether the HADOS mechanism at the atomistic scale would lend its support to the HEDE or HELP, or both models operating at the microscope scale, remains to be seen.

2.8.4 MATERIALS SELECTION GUIDE FOR HYDROGEN COMPATIBILITY There are several screening methods available to rate the hydrogen embrittlement effects on material properties in hydrogen environments, as compared to air or inert environments such as helium. This section describes the common methods used to screen materials for hydrogen embrittlement and discusses the property data base, which helps to select materials for applications in hydrogen environments. 2.8.4.1

MATERIALS SCREENING METHODS

Experimentally, it has been found that the general trend for HEE effects on notched tensile strength (NTS), measured from a sharp notched specimen, is well correlated with the reduction of area (RA) and elongation (EL) measured

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Fig. 2.33

Trend line for HEE effects on NTS and RA for electrodeposited nickel.

from a smooth specimen under slow strain rates (SSR) testing. For example, the trend line for HEE effects on NTS and RA as a function of temperature, at hydrogen pressure of 8.3 MPa (1.2 ksi), is shown in Fig. 2.33 for electrodeposited nickel. There are several proposed test standards for the HEE effects; however, the test procedures and specimen preparations as baselined in G129 and G142 from the American Society for Testing and Materials (ASTM) are commonly used as material screening methods for the hydrogen embrittlement susceptibility under SSR testing. These test methods can also be used to evaluate the effects of materials’ composition, processing, and heat treatment when the materials are exposed to specific hydrogen pressure and temperature conditions. According to ASTM G129 [5], at a minimum the HEE effects can be evaluated in terms of the reduction ratio of NTS from a notched specimen and the reduction ratio of RA from a smooth specimen. Because of the simplicity and accelerated nature of these tests, the results are not necessarily intended to represent true hydrogen service environment for long-term exposure, but rather to provide a basis for material screening for HEE. The property ratio for NTS, RA, and EL, tested in a hydrogen environment as compared with helium or air, is commonly used as the HEE index. By using a compact tension, fatigue precracked specimen, an important HEE index based on the threshold stress intensity factor ratio will be discussed in further detail for hydrogen effects on fracture properties. In general, the HEE index is a simple concept to evaluate the severity of hydrogen embrittlement as an initial material screening tool. According to ASTM G129, these commonly used HEE

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indices are defined as follows: NTS ratio ¼ NTS in hydrogen/NTS in air or helium RA ratio ¼ RA in hydrogen/RA in air or helium EL ratio ¼ (EL) in hydrogen/(EL) in air or helium

(2:4) (2:5) (2:6)

In all cases, the material screening for the HEE effects will be based on the reduction in the value of the HEE index from 1 to 0. Therefore, for maximum hydrogen embrittlement resistance, it is important to select materials with a high HEE index, as close to unity as possible. Under the SSR testing procedures, typical strain rates for smooth specimens are 0.0005 in./in./min (8.3  1026 mm/mm/s) and 0.005 in./min (8.3  1025 mm/s) for notched specimens, respectively. For notched specimens, the typical values for stress concentration factor (Kt) are 6–8. The HEE indices based on the NTS ratio is the most preferred for material screening. Historically, hydrogen embrittlement evaluation for a large number of metallic materials has been investigated by Walter and Chandler, who first suggested classifying the susceptibility of HEE index measuring at room temperature, in 10 ksi (69 MPa) hydrogen pressure, into four categories: negligible, slight, severe, and extreme [6]. Their classification was based mostly on the NTS ratio taken from notched tensile specimens and the RA ratio taken from smooth specimens. Because of its usefulness for a qualitative material screening method, a simplified suggested format for hydrogen embrittlement category, based on these early works, is shown in Table 2.17. It must be noted that the proposed HEE index classification is only a qualitative material screening method based TABLE 2.17

MATERIAL SCREENING FOR HYDROGEN EMBRITTLEMENT BASED ON HEE INDEX FROM NTS RATIO

H Embrittlement Category

a

HEE Index (NTS Ratio)

Material Screening Notesa

Negligible

1.0–0.97

Small

0.96–0.90

Materials can be used in the specified hydrogen pressure and temperature range with fracture mechanics and crack growth analysis in hydrogen.

High

0.89–0.70

Cautiously use only for limited applications with detailed fracture mechanics and crack growth analysis in hydrogen.

Severe

0.69–0.50

Extreme

0.49–0.0

Not recommended for use at specific pressure and temperature where the HEE index is measured.

Based on application at specific hydrogen pressure and temperature, where HEE index is measured. In all categories, additional testing and fracture analysis must be peformed beyond the material screening phase.

94

B. N. BHAT

on an accelerated test in a laboratory environment. It should not be used for components design without detailed fracture mechanics and design analysis for safety use in a hydrogen environment, particularly, for materials that are qualitatively rated in the embrittlement categories of high, severe, and extreme. 2.8.4.2

MATERIAL DATABASE FOR HYDROGEN ENVIRONMENT EMBRITTLEMENT

A comprehensive worldwide database compilation, in the past 50 years, has shown that the HEE index for metallic materials is mostly collected at two high hydrogen pressure points of 5 ksi (34.5 MPa) to 10 ksi (69 MPa), near room temperature. This invaluable database is commonly used as a qualitative material screening process for hydrogen embrittlement severity based on the decrease in the value of the HEE index from unity [7, 8]. The available HEE indexes data for several metallic materials, based on the property ratios for NTS, RA, and EL are listed in Tables 2.18 and 2.19. Testing was done at room temperature and in a hydrogen pressure range of between 5 ksi (34.5 MPa) and 10 ksi (69 MPa). Table 2.18 lists the HEE index for selected aluminum, copper, titanium, nickel and iron-based alloys and for several different types of austenitic, ferritic, and martensitic steels. The HEE indices for selected nickel, iron, and cobalt-based superalloys are given in Table 2.19. This database should not be used to estimate the HEE effects at high temperature, nor for components design, without the detailed fracture analysis for safe use in a hydrogen gas environment, particularly for materials that are qualitatively rated in the high, severe, and extreme categories.

2.8.5 GENERAL OBSERVATIONS FOR METALLIC MATERIALS 2.8.5.1

ALUMINUM AND ALUMINUM ALLOYS

A dry hydrogen gas environment has negligible effects on aluminum and its alloys. The major issue with hydrogen arises mostly from the exposure to moisture and the formation of gas-filled voids during the molten, casting, and solidification process from the foundry. These voids are material defects, which affect both cast and wrought products’ mechanical properties such as ductility and fracture toughness. During cooling from the melt, hydrogen diffuses to and precipitates in casting defects, producing cracks from the decreased solubility of hydrogen in solid metal at lower temperature. Dry hydrogen gas near room temperature, at pressure up to 10 ksi (69 MPa), does not cause significant hydrogen embrittlement effect in aluminum alloys. However, when a high-strength aluminum alloy is electrochemically charged by hydrogen in an aqueous solution, its ductility is reduced. The main mechanism of embrittlement for aluminum alloys in an aqueous medium could be the SCC rather than pure HE effect. The combined mechanism of anodic material dissolution SCC or cathodic hydrogen embrittlement remains an open question for aluminum alloys when subjected to an aqueous environment.

HEE INDICES FOR SELECTED METALS TESTED AT 248C UNDER HIGH HYDROGEN PRESSURE H Pressure (MPa)

Alloy System

Material

Qualitative Rating for HEE

(HEE Tested at 2488 C) Austenitic Steels

HEE Index, (Ratio H/He)

Smooth Ductility (%), in Helium or Air EL

RA

Tensile Strengths, in Helium or Air (MPa)

NTS

EL

RA

A302B

69.0

High

0.78

0.85

0.50

NTS

YS

UTS

1564

813

827

A286

69.0

Negligible

0.97

1.10

0.98

26

304L (annealed)

69.0

High

0.87

0.92

0.91

86

47

1606

724

1117

78

703

234

531

304N

69.0

High

0.93

0.84

0.73

43

74

641

848

305

69.0

High

0.89

1.03

0.96

63

78

309S

69.0

Small

0.96

0.97

85

76

351

620

241

558

310

69.0

Small

0.93

1.00

0.96

56

64

799

220

531

316

69.0

Negligible

1.00

0.95

1.04

59

72

1109

441

648

18-2-12 (Nitronic 32)

69.0

Severe

0.64

0.47

75

78

482

861

21-6-9 þ 0.12N (Nitronic 40)

69.0

High

0.89

0.80

65

74

434

744

22-13-5 (Nitronic 50)

69.0

Negligible

1.00

1.00

51

67

586

938

18-18 Plus

69.0

Severe

0.67

63

520

910

18-2-Mn

69.0

Severe

0.65

51

730

1007

18-3-Mn

69.0

Small

0.92

50

530

790

1137

95

(Continued )

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.18

HEE INDICES FOR SELECTED METALS TESTED AT 2488 C UNDER HIGH HYDROGEN PRESSURE (Continued )

TABLE 2.18

Alloy System

Material

Qualitative Rating for HEE

(HEE Tested at 2488 C) Ferritic Steels

Martensitic Steels

HEE Index, (Ratio H/He)

NTS

A212-61T (normalized)

69.0

Severe

0.68

A372 (class 4)

69.0

High

0.74

EL

RA

0.50

0.34

Smooth Ductility (%), in Helium or Air

Tensile Strengths, in Helium or Air (MPa)

EL

RA 57

765

20

53

1378

0.60

NTS

YS

UTS

565

813

A515-Gr. 70

69.0

High

0.73

0.69

0.52

42

67

731

310

448

A517-F

69.0

High

0.78

1.00

0.96

18

65

1532

751

813

A533B

69.0

High

0.78

0.89

0.50

19

66

1564

HY-80

69.0

High

0.81

0.86

0.85

23

70

1311

566

676

HY-100

69.0

High

0.73

0.90

0.83

20

76

1546

669

780

430F

69.0

Severe

0.68

0.63

0.58

22

64

1047

496

551

1020

69.0

High

0.85

0.80

0.66

40

68

724

283

435

1042 (normalized)

69.0

High

0.75

0.75

0.45

29

59

1056

400

621

4140 (high strength)

69.0

Extreme

0.40

0.18

0.18

14

48

2160

1233

1283

4140 (low strength)

69.0

High

0.85

34.5

Extreme

0.35

AerMet 100 (peak aged)

69.0

Extreme

0.15

H-11

69.0

Extreme

410

69.0

Extreme

0.50

820

1660 0.31

0.26

0.25

0.00

0.00

0.20

0.12

0.20

0.01

0.01

440C

69.0

Extreme

17-4 PH

69.0

Extreme

17-7 PH

69.0

Extreme

0.22

0.10

0.06

18Ni-250 Maraging

69.0

Extreme

0.12

0.03

0.05

0.18

12.4

8.8 12 3.2

54.2

2157

8.2

1302

1371

30

1736

1681

2060

60

2730

1324

1524

3.2

1027

6.4 17

930

1627

2075

1076

1145

45

2151

1124

1200

55

2914

1709

1723

B. N. BHAT

4340 (1652 F austen.)

96

H Pressure (MPa)

Titanium Based

Copper Based

Aluminum Based

Nickel (electroformed)

69.0

Extreme

0.31

Nickel 270

69.0

High

0.70

0.92

827

Nickel 301 (annealed)

69.0

Extreme

0.52

0.35

K-Monel (Kt ¼ 4; precipitated)

69.0

Extreme

0.45

0.75

56

89

34

531

34

331

1600

486

791

1729

958

K-Monel (Kt ¼ 4; annealed)

69.0

High

0.73

Titanium (pure)

69.0

Small

0.95

0.96

1.00

32

61

868

365

434

Ti-6Al-4V (annealed)

69.0

High

0.79

1.00

1.00

15

48

1674

909

1075

992

689

69.0

Severe

0.58

0.85

0.95

13

47

1571

1082

1130

Ti-5Al-2.5Sn (ELI)

69.0

High

0.81

0.90

0.86

20

45

1385

730

779

Copper (OFHC)

69.0

Negligible

0.99

600

Al umi num Bronze

68.9

Negligible

Be-Cu alloy 25

69.0

Small

GRCop-84 (Cu-8Cr-4Nb)

34.5

Negligible

NARloy-Z (Cu-3Ag-0.5Zr)

40.0

Negligible

70-30 Brass

69.0

Negligible

1100-T0

69.0

Negligible

2011

69.0

2024

69.0

5086

69.0

Negligible

6061-T6

69.0

Negligible

6063

69.0

7039

69.0

7075-T73

69.0

Negligible

0.93

1.00

1.00

63

94

1.02

1.05

48

67

1.00

0.98

22

72

1.00

1.20

20

42

1.10

290 600

544

648

250

413

24

138

269

1.00

0.98

8

70

572

606

0.93

1.00

42

93

34

110

Negligible

0.95

1.01

18

57

227

296

Negligible

0.95

0.97

19

36

324

441

1.05

1.03

20

55

193

303

1.00

1.08

19

61

227

269

Negligible

1.00

1.01

15

83

158

193

Negligible

1.00

1.01

14

85

152

179

0.80

0.94

15

37

372

455

1.38

1.07

0.98

0.92

1344

269 220

124

496

799

97

Ti-6Al-4V (STA)

AEROSPACE MATERIALS CHARACTERISTICS

Nickel Based

Alloy System

HEE INDICES FOR SELECTED SUPERALLOYS TESTED AT 248C UNDER HIGH HYDROGEN PRESSURE

Superalloys

H Pressure (MPa)

Qualitative Rating for HEE

(HEE Tested at 2488 C) Nickel Based

98

TABLE 2.19

HEE Index, (Ratio H/He) NTS

AF-115 (Powder Metall)

34.5

High

0.80

AF-56 (single crystal)

34.5

High

0.84

EL

RA

34.5

Small

0.94

34.5

Extreme

0.14

0.45 0.14

34.5

Extreme

0.38

34.5

Extreme

0.36

CM SX-4D

34.5

Extreme

0.43

CM-SX5

34.5

Extreme

0.36

Hastelloy X

34.5

High

0.86

Haynes 230

34.5

High

0.76

Haynes 242

34.5

High

0.77

IN 100

34.5

Extreme

Inconel 625

34.5

High

Inconel 700

69.0

Extreme

48.2

High

Inconel713LC

41.3

Extreme

Inconel 718 (ST @17508F)

34.5

Extreme

Inconel 718 (ST @17508F)

69.0

Extreme

NTS

YS

UTS

21

1764

1185

1702

Reference

70

26

37

1805

1096

1523

70

14.3

1495

958

1047

69, 71

999

1089

69, 71

14

1461

10.6

1536

69

1488

69

1523 0.98

0.98

54

63

317

0.41

1068

0.20

1860

0.36

0.36

55

50

0.45

0.32

22

44

0.82

69

1006

0.54

37

717

21

365

827

35

861

1337

35

1736

1151

1612

69

1433

634

992

36

1034

1344

30

696

813

69, 73

1578

69

0.42

0.38

6.9

9.5

0.53

0.24

0.34

20.8

29.5

1723

1102

1364

21

0.46

0.09

0.08

17

26

1888

1254

1426

4, 8

B. N. BHAT

Inconel 706

RA

0.30 0.76

Tensile Strengths, in Helium or Air (MPa)

35

CM SX-2 (single crystal)

CM SX-4C (single crystal)

EL 23

Astroloy (Powder Metall)

CM SX-3 (std)

Smooth Ductility (%), in Helium or Air

34.5

Small

0.92

0.87

0.76

26

Inconel X-750

48.2

Extreme

0.26

Inco 4005 (experiment)

34.5

Severe

0.64

MAR-M200 (Direct Solidify)

34.5

Extreme

MAR-M246 (Hf) (single crys)

48.2

Extreme

0.24

0.33

12

MA 6000 (Transverse)

34.5

Small

0.92

0.50

2

0.86

1.00

1

0.27

36

34.5

High

MA 754 (Transverse)

34.5

Extreme

MA 754 (Longitude)

34.5

Extreme

MERL 76 (Powder Metal)

34.5

Small

0.96

MERL 76

34.5

High

0.85

34.5

Negligible

PWA 1480 (single crystal)

34.5

Extreme

0.49

30 0.25 0.98

PWA 1480E (111 plane)

34.5

High

0.88

Rene 41

34.5

Extreme

0.36

Rene 41

69.0

Extreme

0.27

Rene N-4 (single crystal)

34.5

Extreme

0.46

1075

1295

21

1860

1013

1357

35

9.7

0.19

NASA-HR1

2081

35 0.21 0.23

MA 6000 (Longitude)

50.6

69 1213

63 69 69

1158

47

1178

28

1764

28

1660

12.4

1516

24

69 69 1034

1557

70 69

944

1323

74

1040

1151

71

1564

75 35

0.20

0.38

21

0.48

1929

1123

1350

4, 8

1474

978

1158

71

21

1805

1330

1688

70

34.5

Severe

0.62

RR 2000

34.5

Extreme

0.54

1378

69

Udimet 720

34.5

Extreme

0.53

1771

69

Severe

0.65

34.5

Small

0.95

A286 (ST @16408F)

69

Negligible

0.97

A286 (ST þ Aged)

69 (T.C.)

Severe

76

1.10

0.98 0.51

20

34

1950

1199

1523

70

26

44

1605

847

1089

4, 8 38

(Continued )

99

51.7

Waspaloy (Powder Metall)

18

29 10.6

Rene 95 (Powder Metall)

Udimet 700

AEROSPACE MATERIALS CHARACTERISTICS

Iron Based

Inconel 718 (ST @19008F)

TABLE 2.19

Superalloys

H Pressure (MPa)

Qualitative Rating for HEE

(HEE Tested at 2488 C) 48.2

Negligible

0.99

Incoloy 901

34.5

Severe

0.60

Incoloy 903 (ST only)

34.5

Negligible

0.98

24 (T.C.)

Severe

Incoloy 907

69

Small

Incoloy 909 (ST þ Aged)

34.5

Extreme

JBK-75 (ST only)

34.5

Negligible

JBK-75 (ST þ Aged)

172

EL

RA

Smooth Ductility (%), in Helium or Air EL

RA

Tensile Strengths, in Helium or Air (MPa) NTS

YS

UTS 35

1.00

0.40

21

1688

69

0.96

42

2122

69

0.55

77

0.96

1819 0.36

0.98

High

Reference

0.75

0.39

12

22.4

0.92

28

51

1585

69 1054

1350

78, 35

462

744

79, 35

0.45

79

MA 956 (Longitude)

34.5

Severe

0.58

1364

69

MA 956 (Transverse)

34.5

Extreme

0.34

1151

69

Ni-SPAN-C (alloy 902)

69

Small

Haynes188

48.2

High

0.92

0.93

MP35N

24

High

0.73

MP35N

34.5

High

0.70

MP159

24

High

0.66

MP159

34.5

Severe

0.63

MP98T

34.5

High

0.85

X-45 (superalloy)

34.5

High

0.87

16 0.63

0.85

23

7

0.63

¼ H2 thermally charged at 2008C at indicated pressure for 200 hrs (A286) and 500 hrs (Incoloy 903). HEE tested at 248C.

751 63

58

31

1158

1130

44 63

2425

1385

1433

35

2425

1385

1433

35

2253

1895

1922

35

2253

1585

1895

1922

35

1171

1275

35

462

744

35

B. N. BHAT

 T.C.

NTS

Incoloy 802

Incoloy 903 (ST þ Aged)

Cobalt Based

HEE Index, (Ratio H/He)

100

Alloy System

HEE INDICES FOR SELECTED SUPERALLOYS TESTED AT 2488 C UNDER HIGH HYDROGEN PRESSURE (Continued )

AEROSPACE MATERIALS CHARACTERISTICS

2.8.5.2

101

COPPER AND COPPER ALLOYS

Copper and the copper-rich alloys are usually not susceptible to hydrogen embrittlement unless they contain oxygen or copper oxide. When oxygen-bearing copper and copper alloys are annealed or heated in a hydrogen environment, the atomic hydrogen diffuses into the metals and reacts with the copper oxide or the oxygen to form water, which is converted to high-pressure steam if the temperature is above 3758C (7058F). This is a classic example of HRE, as the steam will induce hydrogen damage in the forms of fissures and blisters appearance, decreasing the fracture toughness and ductility of the metals even without applied external pressure. Tough pitch coppers usually contain small quantities of Cu2O; therefore, they should not be exposed to hydrogen gas at any temperature if they will subsequently be exposed to temperatures above 3708C (7008F). The equation for reaction with cuprous oxide particles is Cu2 O þ 2H ¼ 2Cu þ H2 O (g) 2.8.5.3

(2:7)

NICKEL AND NICKEL-BASED ALLOYS

Nickel and nickel-based alloys have good properties for high-temperature strength, oxidation, and hot corrosion resistance. However, a nickel-based alloy that has good ratings for dry oxidation and chemical corrosive environment is not automatically immune to hydrogen embrittlement. As an element, pure nickel is severely embrittled by hydrogen; therefore, most binary alloys with nickel-rich composition, such as nickel–copper, nickel–iron, nickel–cobalt, and nickel–tungsten, are also found to be highly embrittled by hydrogen in the nickel-rich regions [1, 2]. In some nickel-rich alloy systems, the same observation is held. For example, the nickel-rich alloys known as K-Monel have been known to be embrittled by hydrogen at high pressure. However, the influence of nickel on complex compositions of steels and superalloys is much more difficult to analyze due to factors such as heat treatment and product forms. The HEE indices for several materials that contain nickel and the nickel-based superalloys are given in Table 2.18 and Table 2.19, respectively. 2.8.5.4

TITANIUM AND TITANIUM ALLOYS

In general, titanium alloys have excellent corrosion resistance properties in an aqueous environment. This superior corrosion resistance property is due to a thin, stable, and tenacious titanium-oxide (TiO2) film that naturally forms in air and water under the oxidizing conditions. The naturally formed TiO2 film on titanium appears to inhibit hydrogen uptake effectively under low-to-moderate cathodic charging conditions. However, under high cathodic charging current densities, this protective film can break down and become nonprotective for titanium alloys and will allow atomic hydrogen to penetrate into the bulk of the material. In near-neutral electrolytes such as seawater, galvanic coupling to

102

B. N. BHAT

metals such as zinc, aluminum, and magnesium can induce enhanced hydrogen uptake and hydride formation when coupled with titanium at temperatures above 808C (1758F). On the other hand, in a dry-hydrogen gas environment, titanium alloys will absorb hydrogen readily as the temperatures and pressures increase. Relatively small amounts of titanium-hydride precipitates are not detrimental for most applications, particularly in hydrogen concentrations of 40–80 ppm (part per million). However, excessive titanium hydride can form rapidly when the temperature is above 2508C (480F). This type of hydrogen embrittlement is the HRE type; however, it is also considered as the IHE type by some industries during high-temperature processing such as welding or heat treatment in the presence of hydrogen. Graphically, the distinction between IHE and HRE is not always well recognized for metals that form unstable hydrides, as they are positioned closely in the overlapping region, as shown in Fig. 2.32. 2.8.5.5

STEELS

The HEE susceptibility of steels can generally be viewed in four categories: austenitic, ferritic, martensitic, and precipitation hardening (see Sec. 2.4). In general, most low-strength austenitic steels are less susceptible to hydrogen embrittlement, relative to the ferritic steels. However, the martensitic and precipitation hardening steels are known to be extremely susceptible to the HEE and IHE effects. There are some similarities between the austenitic stainless steels and the Fe-Ni-Cr superalloys in terms of composition vs the HEE effects. Concerning the HRE effects, at certain elevated temperatures and pressures, atomic hydrogen can diffuse through metal and react internally with certain types of elements and compounds in the steel-based alloys. The most common reaction is between hydrogen and iron carbides to form methane gas (CH4) (see Eq. 2.3). Because CH4 cannot diffuse out of steel, an accumulation occurs, which causes fissuring and blistering that can lead to embrittlement and loss of strength and ductility. The addition of chromium and molybdenum is beneficial for many carbon and low-alloy steels to reduce or prevent the HRE effects that result in decarburizing and fissuring. 2.8.5.6

SUPERALLOYS

For aerospace applications, superalloys are common materials used in liquid hydrogen and oxygen propulsion systems. The superalloys are discussed in detail in Sec. 2.5. There are more HEE index data available for nickel-based alloys than for any other types of superalloys. The differences in superalloy heat treatment and product form can have an effect on the degree of hydrogen embrittlement. In general, conventional wrought and PM processed superalloys have been found to be slightly less embrittled in hydrogen than cast polycrystalline superalloys with similar compositions.

AEROSPACE MATERIALS CHARACTERISTICS

2.8.5.7

103

GUIDE FOR SELECTING PROPER MATERIALS

At relatively low hydrogen gas pressures, the degradation of fracture properties for many susceptible materials is not as severe as at high pressure. In addition, at a certain cryogenic temperature range, most materials are negligibly embrittled by hydrogen even when they are exposed to a relatively high hydrogen pressure of greater than 5 ksi (34.5 MPa). Therefore, depending on the hydrogen pressure and temperature range, it is possible to select the proper materials for hydrogen applications based on the qualitative rating method shown in Table 2.17. Historically, notable service failures associated with susceptible materials in a hydrogen gas environment have been linked to a combination of highly stressed components that were operating outside the allowable range of hydrogen pressure and temperature. It must be noted that the material database, using the HEE index classification, is only a qualitative material screening method based on an accelerated test in laboratory and should not be used for components design without detail fracture analysis, particularly for materials that are qualitatively rated in the categories of high, severe, or extreme. The selection of proper materials should be done with design using sufficiently high safety factors, coupled with rigorous material characterization based on fracture mechanics, nondestructive evaluations, and material testing in the operating hydrogen pressure and temperature range. Test data taken from thermal precharging in hydrogen gas or any electrochemical method should not be treated as identical to the data taken from actual testing in high-pressure gaseous hydrogen. Heat treatments can have a profound effect on the degree of hydrogen embrittlement, particularly for materials with complex microstructures. In general, superalloys and steel-based alloys that have low tensile strengths and are heat treated in annealed conditions will tend to have better hydrogen embrittlement resistance than higher strength alloys. Therefore, tests should be conducted to determine hydrogen embrittlement effects to select the proper thermomechanical treatment for susceptible materials. For example, for A-286, JBK-75, and Incoloy 903, the data trend shows that aging treatment can drastically affect their HEE behavior. In general, aging conditions will dramatically increase the yield strength and ultimate tensile strength for many materials. However, these aging treatments will cause most of these high-strength materials to have lower ductility and also make them more susceptible to hydrogen embrittlement.

2.8.6 HYDROGEN EFFECTS ON MECHANICAL PROPERTIES Mechanical properties of aerospace structural materials can be severely degraded by hydrogen. The properties of interest include tensile strength and ductility, fatigue life, crack growth rates and fracture toughness. Details are given below.

104

2.8.6.1

B. N. BHAT

TENSILE PROPERTIES

It is experimentally found that the values for modulus of elasticity (E) and the yield strength (YS) of most materials are not strongly affected by hydrogen. However, the ultimate tensile strength (UTS), taken from smooth tensile specimens, can be moderately reduced if the materials are deemed to be very susceptible to hydrogen embrittlement. Therefore, the ratio of E, UTS, and YS taken from smooth tensile specimens are not to be used as good indicators for HEE index. The basic tensile properties that are strongly affected by hydrogen can be listed as NTS, RA, and (EL); see Tables 2.18 and 2.19. Experimentally, it has been shown that the general trend for hydrogen effects on NTS ratio, measured from a notched specimen, correlates well with the RA ratio and EL ratio, measured from a smooth specimen under SSR testing. 2.8.6.2

FRACTURE PROPERTIES

Hydrogen has a significant influence on crack initiation and growth behavior, particularly when reasonably large surface flaws exist on a susceptible material exposed to a high-pressure hydrogen environment. Therefore, fracture mechanics analysis is usually required to assess the maximum allowable stress and service life of a component, based on the surface flaw sizes and crack growth rates, in a hydrogen environment. Using the ASTM E1681 test standard, the threshold stress intensity factor (KTH) in a hydrogen environment can be determined from a precracked specimen, under static loading. Because of the intrinsic nature of the experimental setup to determine the KTH by using a monitoring system for the “onset” of crack initiation, theoretically, it appears that the KTH measurement may be more sensitive as a hydrogen embrittlement indicator than the simple HEE index based on NTS or RA ratio. In reality, it is difficult to measure KTH accurately in gaseous hydrogen at high-temperature and pressure setup conditions. For material screening purposes, KTH values can also be used as the HEE indices according to the ASTM G129 test standard. There are two types of HEE indices based on the KTH ratios relative to the fracture toughness KIC and KC: Plane strain threshold stress intensity factor ratio, HEE index ¼ KIðTHÞ =KIC Threshold stress intensity factor ratio, HEE index ¼ KTH =KC

ð2:8Þ (2:9)

In these equations the values for KI(TH) and KTH are stress intensity factors measured in hydrogen for plane strain and non-plane-strain specimens, respectively. Accelerated test procedures for KTH measurement have also been proposed in recent years to estimate the KTH values from a precracked specimen, without spending a considerable amount of time for the specimen to be under constant static loading as a baseline in ASTM E1681. These relatively rapid test

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105

Fig. 2.34 Effects of hydrogen pressure on threshold stress intensity (KTH) for Incoloy 718 and 903.

procedures are proposed in terms of using a minute incrementally rising load or displacement based on relatively slow strain rates on the precracked specimens. Figure 2.34 shows the trend behavior of KTH values as a function of hydrogen pressure for Inconel 718 and Incoloy 903 tested at room temperature [7]. 2.8.6.3

LOW-CYCLE AND HIGH-CYCLE FATIGUE

Gaseous hydrogen has a considerable effect on low-cycle fatigue (LCF) properties for susceptible materials tested in strain-controlled mode. The hydrogen degradation in LCF life is also a function of strain range, as shown in Fig. 2.35, for several power metallurgy (PM) superalloys tested at 758F (248C) in 5 ksi (34.5 MPa) hydrogen pressure. In general, the lower the strain range, the lower the effect of hydrogen embrittlement degradation for LCF. By selecting a typical strain range from 1% to 2%, the values for cycles-to-failure (CTF) obtained in

Fig. 2.35

Effects of hydrogen on strain controlled LCF life of superalloys.

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hydrogen and in air can be compared to indicate the severity of HEE. For example, at a strain range of 1.2%, typical CTF ratios for the PM superalloys, as shown in Fig. 2.35, are reduced by nearly a factor of 10 at room temperature. Because the CTF values for fatigue testing are plotted on a log-scale, instead of using a linear scale similar to measuring the notched tensile strength or reduction of area, the ratio of CTF generally is not used as the typical HEE index specified in the ASTM G129 test standard, which is based on a linear scale with a ratio ranking from 1 to 0. However, the trend behavior for hydrogen embrittlement based on LCF correlates qualitatively well with the HEE index values as shown in Table 2.19 for superalloys. In general, the greatest reduction in LCF life ratio is also found to exhibit near room temperature for most superalloys, whereas at cryogenic and at relatively high temperatures the LCF properties are not as severely reduced, similar to the HEE index trend lines for Inconel 718 as shown in Fig. 2.33. Comparing the LCF data against the data in Table 2.19, some superalloys, such as Inconel 100, 718, and 625, or Hastelloy-X, which exhibit severe HEE in high hydrogen pressure at room temperature, are similarly affected in LCF tested in strain-controlled mode at room temperature. It has been found that the strain-controlled LCF test is more sensitive to HEE than the one based on the load-controlled mode. As a general rule, it has been observed that HEE has little effect on high-cycle fatigue (HCF) properties for many metallic materials. Typical HCF regimes are usually defined when the cycles-to-failure are at or above the 106 cycles. Because the maximum stress amplitude stays significantly below yield strength for HCF testing, it is possible that no significant HCF life degradation would be observed because the majority of the test time involved in HCF testing is for the crack initiation mode and not for the crack propagation mode under high cyclic loading at a low stress level. HCF life is generally considered to be around 90% for crack initiation and 10% for crack propagation. In addition, when testing HEE using smooth tensile specimens, the most severe property degradation usually occurs when the specimens enter the high plastic strain regions. In the elastic and low plastic strain region, the HEE effects are usually not detected for smooth specimens. The explanation for this condition is somewhat similar to the axially loaded, smooth HCF test specimens, when the loading is close to fatigue limit (cycles-to-failure near  106 cycles), where the strain range is usually very low. 2.8.6.4

CRACK GROWTH RATE

The HEE effects for cyclic crack growth rates (da/dN) as a function of stress intensity factor range (DK) of several superalloys at 5 ksi (34.5 MPa) to 7 ksi (48.2 MPa) are shown in Fig. 2.36 [7]. With the exception of Mar-M-246 from conventional cast, which was tested at 5388C (10008F), the most rapid crack growth at room temperature was observed for Inconel 718 by the solution treated and aged (STA 2) at a solutionizing temperature of 9408C (17258F). Following Mar-M-246, the third fastest cyclic crack growth rate is Inconel 718 in the

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.36

107

HEE effects for cyclic crack growth rates (da/dN) for several superalloys.

labeled STA 1 heat treatment, which is solutionized at 10388C (19008F). The lowest hydrogen embrittlement effect is with cast Inconel 718 in standard heat treat conditions. Therefore, heat treatment can change the hydrogen embrittlement behaviors significantly for superalloys, such as Inconel 718, which have complex microstructures. These crack growth rates are nonlinear curves and also plotted on a log-scale as a function of either K or DK values; therefore, these data are not commonly used as the HEE index as specified in the ASTM G129 test standard but can be used to conduct damage-tolerance analysis.

2.8.7 CONTROLLING FACTORS IN HYDROGEN EMBRITTLEMENT 2.8.7.1

HYDROGEN PRESSURE

As hydrogen gas pressure increases, the susceptibility for hydrogen embrittlement also increases, resulting in the reduction of the HEE index. However, the HEE

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Fig. 2.37

Effects of hydrogen pressure on HEE index for superalloys at 2288 C.

index seems to decrease exponentially until a saturation pressure is reached. The decaying exponent in this relationship may reflect the kinetic effects or rate limitations in the hydrogen embrittlement process as a function of pressure. High-pressure hydrogen gas has a direct influence on the HEE effects of materials. By reducing the hydrogen gas pressure to below certain levels, susceptible materials will become less embrittled as compared to high pressure. At a constant temperature, the influence of hydrogen gas pressure is qualitatively understood based on the fact that high-pressure hydrogen will increase the number of hydrogen atoms available per unit volume, therefore enhancing the localized HEE effects at the tip of a propagating crack. Figure 2.37 shows the effect of hydrogen pressure on the HEE index based on the NTS ratio for several different superalloy systems, including nickel, Astroloy, Hastelloy X, Haynes 188, Rene 41, and Inco 718. Historically, based on Sievert’s pressure–gas law for hydrogen concentration as a function of pressure, the degree of hydrogen embrittlement was often assumed to be proportional to the square root of the hydrogen pressure; however, recent studies [9] have indicated that hydrogen embrittlement may indeed follow a simple power-law relationship based on the exponential function of hydrogen pressure instead of the square root of hydrogen pressure as shown in the following equation: HEE index ¼ a  (P)n

(2:10)

where

a ¼ proportional constant P ¼ hydrogen pressure at constant temperature, n ¼ material-dependent and decaying exponential value that indicates the embrittlement severity for that particular material

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109

When n ¼ 0.50, this is just a special case for the “square root” of hydrogen pressure P as it was often assumed by early researchers based on Sievert’s pressure–gas law. 2.8.7.2

OPERATING TEMPERATURE

HEE can occur over a wide range of temperatures; however, it is most severe in the vicinity of room temperature for many materials. Based on the hydrogen trapping model, a hydrogen trapping may be considered as the binding of hydrogen atoms to impurities, structure defects, or microstructure constituents in the alloy. The binding of hydrogen in the “trapping-mechanism” model has been proposed to explain why hydrogen embrittlement is most severe near room temperature and becomes less severe or negligible at a higher or lower temperature range. At lower temperatures, the diffusivity of hydrogen is too sluggish to fill sufficient hydrogen traps, but at high temperatures, hydrogen mobility is enhanced and trapping is diminished. At high strain rates, fracture may proceed without the assistance of hydrogen because hydrogen’s mobility is not sufficient to maintain a hydrogen trapped atmosphere around moving dislocations. Figure 2.38 shows the hydrogen embrittlement effect of several steels as a function of temperature [8]. Many austenitic stainless steels (Fe-Ni-Cr) are mostly sensitive to hydrogen embrittlement in the temperature range of 21508C to þ1508C. For some superalloys, their HEE can occur over a wide range of temperatures, from cryogenic to at least 8008C (15008F), but it is most severe in the vicinity of room temperature. This is an important factor to consider for safe design because superalloys are often chosen for high-temperature applications. This temperature

Fig. 2.38

Effects of temperature on HEE index for selected steels.

110

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1.05

HEE Index (NTS rao)

1.00

0.95

0.90

0.85

0.80

Astroloy (PM) MERL 76 (PM) Inco 100 (PM)

0.75

Waspaloy (PM) 0.70 0

100

200

300

400

500

600

700

800

Temperature (°C)

Fig. 2.39

Effects of temperature on HEE index for selected PM superalloys.

effect for superalloys is unlike that for the austenitic stainless steels discussed earlier. For instance, the most severe temperature for the HEE index, based on the NTS ratio, occurs at 5388C (10008F) for Astroloy, Merl 76, Inco 100, and Waspaloy, as shown in Fig. 2.39 [8]. Experimental data for the nickel-based Udimet 700 superalloy has shown that its extent of hydrogen embrittlement as a function of temperature is far greater at high temperature than for any other superalloy. At high temperatures more thermal activation energy is available and there is a potential for the HRE effects from the chemical reaction of hydrogen with certain superalloy constituents or impurities at the grain boundaries. The HRE type of embrittlement can add to the HEE at high temperatures, but not at ambient temperature. By selecting a proper operating temperature range, hydrogen embrittlement can be reduced for many materials. In general, materials are less susceptible to HEE, IHE, and even HRE if the temperatures are in the cryogenic range. At high temperatures, the HEE and IHE effects will also become less severe for many materials. However, with materials that can form significant amounts of stable metal hydrides at elevated temperatures, attention must be given to the HRE effects. Therefore, the HEE index near room temperature should not be used to estimate the hydrogen embrittlement effects at high temperatures. This is an important factor to consider for safe design because superalloys are often used for high-temperature applications. Testing must be conducted in the operating environment to accurately assess the potential for hydrogen damage. 2.8.7.3

REDUCTION IN STRESS LEVELS

Reducing the operating stress level can be a viable method to prevent or control the hydrogen embrittlement effects for most materials. This can be achieved by

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111

increasing the cross section of parts, avoiding the stress raiser in the design, or simply reducing the load on the parts. The constant static loading is usually considered to be more sensitive to hydrogen damage than cyclic or dynamic loading, particularly at high cycle. It should be noted that this method of reducing the external loads is generally applicable to HEE and to most IHE cases, but not necessary for all IHE cases because of the possible combination effects of IHE and HRE. Reducing the externally applied load will not help to alleviate the HRE type of embrittlement because the HRE effects can occur even in the absence of stress. Residual stresses often develop in the metal during manufacturing processes such as heat treatment, fabrication, and welding. These stresses can be high and must be minimized. For wrought products, fabricating methods such as forming, straightening, and stretching that involve localized plastic deformation can exceed the elastic limit of the material due to high residual stress. For casting products, because the hot or molten metal shrinks as it cools, casting components or welding parts that are joined into a more complex structure tend to have high residual stress. Residual stresses can be reduced through thermal treatments such as annealing, preheating the parts before welding, or postweld heat treatment. In addition, surface preparation techniques (such as low-stress grinding) that would impact residual compressive stress to the surface can be used to improve resistance to cracking due to applied tensile stress.

REFERENCES [1] Lee, J. A., “A Theory for Hydrogen Embrittlement of Transition Metals and Their Alloys,” Hydrogen Effects in Materials, edited by A. W. Thompson, and N. R. Moody, Minerals, Metals, and Materials Society, Pittsburgh, PA, 1994. [2] Lee, J. A., “Effects of the Density of States on the Stacking Fault Energy and Hydrogen Embrittlement of Transition Metals and Alloys,” Effects of Hydrogen on Materials, Proceedings of the 2008 International Hydrogen Conference, edited by B. Somerday, P. Sofronis, and R. Jones, 2008, pp. 678–685. [3] Louthan, M. R., Caskey, G. R., Donovan, J. A., and Rawl, D. E., “Hydrogen Embrittlement of Metals,” Material Science Engineering, Vol. 10, 1972, pp. 357–368. [4] Gerberich, W. W., Stauffer, D. D., and Sofronis, P.,“A Coexistent View of Hydrogen Effects on Mechanical Behavior of Crystals: HELP and HEDE,” Effects of Hydrogen on Materials, Proceedings of the 2008 International Hydrogen Conference, edited by B. Somerday, P. Sofronis, and R. Jones, 2008, pp. 38–45. [5] ASTM Standard G12900, Standard Practice for Slow Strain Rate Testing to Evaluate the Susceptibility of Metallic Materials to Environmentally Assisted Cracking, ASTM International, West Conshohocken, PA, reapproved 2006. [6] Jewett, R. P., Walter, R. J., Chandler, W. T., and Frohmberg, R. F., “Hydrogen Environment Embrittlement of Metals,” Rocketdyne, NASA-CR-2163, March 1973. [7] Lee, J. A., “Hydrogen Embrittlement of Superalloys,” Gaseous Hydrogen Embrittlement of Materials in Energy Technologies: The Problem, Its Characterization

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and Effects on Particular Alloy Classes, Vol. 1, edited by R. P. Gangloff, and B. Somerday, Elsevier, New York, December 2011. [8] Lee, J. A., “Hydrogen Embrittlement,” AIAA Guide to Safety of Hydrogen and Hydrogen Systems, AIAA G-095-2004, AIAA, Reston, VA, 2012. [9] Lee, J. A., “Empirical Method to Predict Hydrogen Embrittlement of Metals by High Pressure Hydrogen Gas at Constant Temperature,” presented at the ENERGY 2010 Conference, Sponsored by the American Ceramic Society, Cocoa Beach, FL, 2010.

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2.9 MATERIAL BEHAVIOR IN OXYGEN-RICH ENVIRONMENTS Samuel Edgar Davis NASA Marshall Space Flight Center

2.9.1 INTRODUCTION Oxygen (O2) is a pale-blue odorless and tasteless gas at ambient temperatures. It can also can also exist in a liquid state at temperatures below 21838C (22978F) and as a solid below 22198C (23628F). Oxygen comprises 20.9% of Earth’s atmosphere and exists primarily in the form of O2 molecules. Gas turbine engines draw oxygen from the atmosphere to burn fuel to generate power. Oxygen in its liquid state is important to the aerospace industry because it is used as a rocket propellant usually in combination with hydrogen or hydrocarbon as fuel. Liquid oxygen is easily stored, easily transferred from one tank to another, and readily converted to gaseous oxygen. Oxygen can readily react with metallic materials to produce metal oxide, which is usually a highly exothermic reaction. The highly reactive nature of oxygen requires that special care must be taken when using it to ensure the safety of personnel and equipment. Materials that do not burn in normal atmosphere can burn violently in pure oxygen, sometimes to the point of explosion. Materials selection in oxygen systems must be done with care with system safety in mind. The hazards inherent in the oxygen system come primarily from the risk of fires and/or explosions. The standard fire triangle (Fig. 2.40) demonstrates that a fire requires three separate and independent components, represented as three sides of the fire triangle. The three components are fuel, oxidizer, and ignition source. The oxidizer system obviously has the oxidizer side, but the hardware material itself can actually become the fuel. The pressures and energies generated by fluid flow in a rocket propulsion system are quite high. For example, the space shuttle liquid engine burns more than 1,000 lbs/s (454 kg/s) of liquid oxygen and have high speed turbopumps that can pump this amount. The high velocity of oxygen flow can cause the valve and seal materials, seat materials, and sometimes even the metallic materials to become fuels for combustion in the presence of an ignition source. This section focuses on material behavior in oxygen systems and some of the hazards of the oxidizer in propulsion systems and the means by which these hazards can be mitigated.

2.9.2 EXAMPLES OF OXYGEN HAZARDS Designers, engineers, and technicians sometimes fail to recognize how dangerous oxygen-enriched systems can be, which can lead to disastrous consequences. Following are a few examples of NASA-related oxygen events and their consequences. This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Fig. 2.40

Fire triangle. Fuel

2.9.2.1

Ignition source

APOLLO 1 (AKA APOLLO 204) [1]

Three astronauts were killed by a fire in the Oxidizer command module. The cause of the fire was determined to be related to an electrical wire shortage that occurred after the module was filled with an oxygen-rich mixture. A short in the electrical circuit caused the wire insulation to ignite, which in turn ignited other materials. The burning materials rapidly engulfed the module. It happened so quickly that the astronauts could not escape. The materials that burned rapidly in the oxygen atmosphere of Apollo 1 were actually safe to use in air. 2.9.2.2

SPACE SHUTTLE EXTRAVEHICULAR MOBILITY UNIT [2]

An early design of the Space Shuttle Extravehicular Mobility Unit, better known as the spacesuits worn for spacewalks, ignited during a laboratory test. This suit was the selected design that was to be manufactured and worn by the shuttle astronauts. The circumstances of the event seemed benign. A spacesuit design was being tested on a laboratory table, and oxygen began to flow into the suit in the manner it would when an astronaut is wearing the suit in space. Suddenly, a flash fire occurred as the technicians were standing next to the suit. The fire resulted from oxygen flowing against an aluminum regulator that caught fire. This fire resulted in no serious personnel injuries, but a better spacesuit had to be designed and developed at a high additional cost. 2.9.2.3

APOLLO 13 [3]

This event is the most significant oxygen-related event that occurred in a spacecraft in deep space. Two oxygen storage tanks were severely damaged, and the astronauts’ lives almost lost, when activation of the stirring motor in the oxygen system led to an explosion. The cause of this fire was a damaged electrical contact on the stirring motor. This damage had happened before launch but was undetected until the explosion.

2.9.3 TESTING OF MATERIALS FOR OXYGEN COMPATIBILITY Materials, as discussed earlier in this section, can be hazardous in pure oxygen environments. The earliest oxygen systems workers learned this lesson the hard way. Another lesson that they learned was that testing materials in oxygen environments was the only reliable method to determine if a material would be hazardous for a specific application. Several test methods have been developed to determine the relative safety of materials in various applications.

AEROSPACE MATERIALS CHARACTERISTICS

2.9.3.1

115

MECHANICAL IMPACT TEST

Testing is the only way to select materials that are safe for use in oxygen systems. However, there is no single decisive test or evaluation method that will clearly lead to the selection of the most appropriate materials for use in a given system. Several test techniques have been developed to help determine which materials will be safer in oxygen applications on a relative basis. These tests fall into two categories: 1) ignition tests to determine how easily a material ignites, and 2) burn tests to quantify the burn severity after ignition. Developed in NASA’s early days, the ambient pressure liquid oxygen mechanical impact test is the oldest test and is still in use today. It involves dropping a metal plummet onto a disk of the material that is immersed in liquid oxygen. Materials that are very incompatible with oxygen will burn, or even explode, when the energy from the falling plummet is transferred to the sample. This burn, called a reaction, is witnessed by the test operator. The reaction can be a visible flash, audible report, appearance of charring on the surface of the sample, or a combination of these. Figure 2.41a is a schematic of the test fixture section where the plummet will strike the material sample. Figure 2.41b shows a severe reaction with a material that is not compatible with liquid oxygen. After several years of oxygen systems operations, it was realized that the ambient pressure mechanical impact test, in fact, did not adequately address elevated pressure scenarios because a few materials that passed the test at ambient pressure were burning in high-pressure oxygen systems. In response, the tester was modified to perform the mechanical impacts in high-pressure

a) Test fixture schematic

Fig. 2.41

b) Test showing reaction

Ambient pressure liquid oxygen mechanical impact test.

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oxygen. This new tester, the high pressure liquid and gaseous oxygen mechanical impact tester, performs impact testing at pressures up to 10,000 psi (68,948 kPa). The new tester works basically the same way as the ambient pressure tester, except that this test fixture seals a head onto a base to hold the elevated test pressure (Fig. 2.42). This tester was later modified to test materials in gaseous oxygen and also at elevated temperatures. The mechanical impact test was used by NASA as the standard test for materials in oxygen systems for years. However, inspection of materials that burned in high-pressure oxygen systems revealed that this test alone was also not sufficient. The test provided very useful data for applications where actual impacts occurred within the system, such as a valve shutting against a valve seat. The test did not, however, provide useful information for other ignition scenarios. Furthermore, the mechanical impact test also did not distinguish the metals that were very compatible with oxygen and those that were marginally compatible. To provide more useful data, several new test methods were developed. 2.9.3.2

PROMOTED IGNITION TEST

The promoted ignition test, or combustion of materials in oxygen test (Fig. 2.43), has become the new baseline standard for determining if a material is safe in an oxygen environment. This test uses a 1/8-in. (3.2-mm) diam, 4 to 12-in. (10.2 to 30.5 cm) long rod of the candidate material hanging vertically in a test chamber. The sample material rod is ignited at the bottom by an aluminum or magnesium promoter. The rod is surrounded by a gaseous oxygen environment at the highest

Fig. 2.42

High-pressure mechanical impact tester.

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.43

117

Promoted ignition test fixture schematic.

pressure to which the material could be subjected in its actual use conditions. The promoter is initiated, and the sample rod is observed for its burn characteristics, primarily its burn length and burn rate. If the sample burns more than 1.2 in. (30 mm), it is considered sustained burn. Samples that burn more than this are considered unacceptable for unrestricted use in an oxygen system at the test pressure. However, restricted use of that material can still be allowed under certain circumstances, which will be discussed later in this section. The tester can be modified to heat the rod sample to study burn resistance at elevated temperatures; this modified tester is called the Elevated Temperature Promoted Ignition Tester. A photograph of a sample being loaded into the NASA Marshall Space Flight Center (MSFC) Elevated Temperature Promoted Ignition Tester is shown in Fig. 2.44. 2.9.3.3

IGNITION DATA FOR METALLIC MATERIALS

The mechanical impact and the promoted ignition tests provide the pressures at which a given material will ignite or burn. However, the two tests provide significantly different values for the lowest pressures at which a material will burn in oxygen (called threshold pressure). The determination as to which test is more

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Fig. 2.44 Elevated temperature promoted ignition tester at MSFC with sample rod in heating coil. valuable depends on the type of material being tested and the application. The promoted ignition test is more valuable for metals and materials that will form the basic load-bearing structure of the system. The mechanical impact test is more valuable for nonmetals and for seal or seat materials that will actually be subjected to impact loads during normal system operations. Table 2.20 provides summaries of test data generated by the NASA MSFC for common metallic materials in both of these tests. The data clearly show that the threshold pressure for a given material in a promoted ignition test is much lower than the threshold pressure in a mechanical impact test. It should be noted that most of these materials tend to have occasional test series in which they do ignite at a lower pressure than listed. These are exceptions that indicate that a specific lot or batch of material must be tested to eliminate the hazards of lot and batch composition variability. The promoted ignition test is a harsh test and is not a perfect test for materials in oxygen environments; hence the test data should be applied with caution. A vast majority of materials that are commonly used in oxygen systems, viz., stainless steel (SS) alloys, do not meet this criterion at high pressures. Therefore, a more realistic hazards evaluation process for allowing such materials has been developed. This process, called an oxygen compatibility assessment, will be discussed later in this section.

2.9.4 COMBUSTION PROBABILITY AND SEVERITY FOR MATERIALS IN OXYGEN SYSTEMS The ignition and combustion properties of materials can vary from one oxygen system to another. There are several factors that determine ignition

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119

probability and fire severity in an oxygen system. The most important factors that significantly increase both the probability of ignition and the severity of the fire that ensues are oxygen pressure and temperature. The data in Table 2.21 show the pressure effects on the burning rate of metals. Combustible metals, with few exceptions, burn far more readily and rapidly in higher pressure oxygen environments. TABLE 2.20

MINIMUM PRESSURES REQUIRED TO IGNITE OR BURN COMMON METALSa

Material

Environment !

a

Promoted Ignition Test—Threshold Pressure for Sustained Burning, psi (kPa) GOX

Mechanical Impact Test—Threshold Pressure for Ignition, psi (kPa) GOX

LOX

2024 Al

15 (103)

1,500 (10342)

1,500 (10342)

2090 Al

15 (103)

500 (3447)

500 (3447)

2219 Al

15 (103)

1500 (10342)

50 (345)

5052 Al

15 (103)

1500 (10342)

50 (345)

6061 Al

15 (103)

15 (103)

15 (103)

Brass

10,000 (68948)

10,000 (68948)

10,000 (68948)

Copper 12200

10,000 (68948)

10,000 (68948)

10,000 (68948)

Haynes 214

1000 (6895)

10,000 (68948)

10,000 (68948)

Inconel 718

500 (3447)

10,000 (68948)

10,000 (68948)

Magnesium

,15 (103)

,15 (103)

,15 (103)

Monel K-400/K-500

10,000 (68948)

10,000 (68948)

10,000 (68948)

Nickel

10,000 (68948)

10,000 (68948)

10,000 (68948)

SS 17-4

400 (2758)

5,000 (34474)

5,000 (34474)

SS 304

500 (3447)

10,000 (68948)

10,000 (68948)

SS 304L

250 (1724)

10,000 (68948)

10,000 (68948)

SS 316

400 (2758)

10,000 (68948)

10,000 (68948)

SS 316L

250 (1724)

10,000 (68948)

10,000 (68948)

SS 420

750 (5171)

10,000 (68948)

10,000 (68948)

SS 440C

3000 (20684)

10,000 (68948)

10,000 (68948)

Titanium

,15 (103)

,15 (103)

,15 (103)

Data for comparison purposes only; not to be considered standard values for listed material.

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Temperature effects on combustible materials, with very few exceptions, are analogous to pressure effects in that materials burn far more readily and rapidly in a gaseous oxygen environment as the temperature is increased; see Table 2.22. It should be noted that temperatures can increase substantially within a system during normal operation. For instance, adiabatic heating and frictional heating within the system can raise the oxygen temperature and cause ignition followed by combustion. Although liquid oxygen is much colder than gaseous oxygen, many materials tend to be more reactive and burn more readily in liquid oxygen systems. This result is attributed to the fact that liquid oxygen has more oxygen molecules per unit volume (higher density). The data presented here demonstrate that fire hazards are far greater in systems with higher pressures and temperatures than experienced in normal atmospheric conditions. Oxygen system designers can minimize the dangers that are inherent in oxygen systems by using certain techniques that are discussed in detail next.

2.9.5 FIRE SAFETY IN OXYGEN SYSTEMS DESIGN Designers of oxygen systems look for ignition and burn resistance properties in materials to minimize the risk of fire. In general, the higher the oxygen pressure TABLE 2.21

UPWARD FLAMMABILITY OF METALS VS PRESSURE IN 100% OXYGENa,b,c

Material

a

Test Pressure

12-in. (305-mm) Rod Burn Length

Burn Rate in./s (mm/s)

Bronze/aluminum mixture

50 psi (345 kPa) 100 (689 kPa)

0 in. (0 mm) 12 in. (305 mm)

0.0 (0.0) 0.3 (7.6)

Aluminum 4043

25 psi (172 kPa) 50 psi (345 kPa)

4.4 in. (111.8 mm) 12 in. (305 mm)

0.6 (15.2) 1.1 (27.9)

316 stainless steel

500 psi (3447 kPa) 1,000 psi (6895 kPa)

1.7 in. (43.2 mm) 12 in. (305 mm)

0.3 (7.6) 0.4 (10.2)

347 stainless steel

500 psi (3447 kPa) 650 psi (4482 kPa) 3,000 psi (20684 kPa)

2.7 in. (68.6 mm) 8 in. (203 mm) 12 in. (305 mm)

0.3 (7.6) 0.4 (10.2) 0.5 (12.7)

304 stainless steel

500 psi (3447 kPa) 1,000 psi (6895 kPa)

3.8 in. (96.5 mm) 12 in. (305 mm)

0.3 (7.6) 0.4 (10.2)

303 stainless steel

500 psi (3447 kPa) 1,500 psi (10342 kPa)

1.1 in. (27.9 mm) 12 in. (305 mm)

0.3 (7.6) 0.4 (10.2)

316L stainless steel

250 psi (1724 kPa) 1,000 psi (6895 kPa)

2.6 in. (66.0 mm) 12 in. (305 mm)

0.2 (5.1) 0.4 (10.2)

Data for pressure affects comparison only; not to be considered standard values for listed material. Data based upon rods tested per ASTM G124 and ISO14624-4 with rod dimensions 12-in. length, 0.125-in. diam. Data derived from testing at the NASA MSFC.

b c

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.22

121

UPWARD FLAMMABILITY OF METALS VS TEMPERATURE IN 100% OXYGENa,b,c Burn Rate in./s (mm/s)

Material Identification

Test Pressure psi (kPa)

Test Temperature

12-in. Rod Average Burn Length

304L stainless steel

500 (3447) 500 (3447)

758F (2978K) 10008F (8118K)

4 in. (102 mm) 12 in. (305 mm)

0.35 (8.9) 0.43 (10.9)

15-5 PH stainless steel

500 (3447) 500 (3447)

758F (2978K) 10008F (8118K)

2.9 in. (74 mm) 12 in. (305 mm)

0.33 (8.4) 0.35 (8.9)

Inconel 718

750 (5171) 750 (5171)

758F (2978K) 17008F (12008K)

4.2 in. (107 mm) 12 in. (305 mm)

0.35 (8.9) 0.60 (15.2)

316 stainless steel

500 (3447) 500 (3447)

758F (2978K) 10008F (8118K)

5.6 in. (142 mm) 12 in. (305 mm)

0.34 (8.6) 0.37 (9.4)

a

Data for temperature effects comparison only; not to be considered standard values for listed material. Data based upon rods tested per ASTM G124 and ISO14624-4 with rod dimensions 12 in. length, 0.125 in. diam. Data derived from testing at the NASA MSFC.

b c

and temperature the materials are exposed to, the more readily they ignite and burn. Therefore, the designer should use the lowest possible operating pressure and temperature. Thinner sections ignite and burn more readily than thicker sections, and so thicker sections should be preferred. In addition, there are systemic effects that contribute to fire hazards. For instance, the higher the flow rates of oxygen passing over a material, the higher the likelihood of the material igniting and burning in the direction of the flow. Therefore, flow rates within the system should be kept as low as possible. The recommended practices for the design and development of safer oxygen systems include the following: 1. Minimize oxygen and combustible materials within the system. 2. Maximize the use of the oxygen-compatible materials. 3. Control configuration. 4. Minimize the potential ignition sources and mechanisms that are inherent in all systems. Each practice is described in detail in the following sections. 2.9.5.1

MINIMIZE OXYGEN AND COMBUSTIBLE MATERIALS

The safest systems will minimize the amount of oxygen in the system, which reduces the likelihood of ignition and allows for quenching in case of a fire. This can be accomplished by keeping the system pressure at no higher than absolutely necessary, using the smallest amount of oxygen needed for the application, and designing the system such that the flow path for oxygen is the shortest possible. The safest oxygen systems also minimize the amount of combustible

122

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materials within the system, thus providing less fuel available to burn. It should be noted, however, that thicker sections of a material are far less likely to ignite than thinner sections. Therefore, the ignition hazard of a material must be weighed against the safety of using less material within a system. 2.9.5.2

MAXIMIZE THE USE OF OXYGEN-COMPATIBLE MATERIALS

Poor material choices can greatly increase the likelihood of a fire occurring in an oxygen system. Therefore, careful selection of materials can lessen the chances of ignition and promoted combustion, thus limiting the amount of damage resulting from a potential fire. Despite the fact that heat sources can be inherent to an oxygen system or its surroundings, design elements can limit the amount of, or dissipate altogether, the heat generated within the system. After meeting the functional requirements, the selected materials must be compatible with the system oxygen environment and not undergo any dangerous reaction. The best materials for oxygen environments, from a compatibility standpoint, should 1) be difficult to ignite, 2) self-extinguish quickly if ignited, and 3) burn slowly with low heat-of-combustion. These properties are determined by testing. Extensive test data generated by the government and private industries already exist and are available for use. However, newer materials, and even some older ones, have no test data. Therefore, it is important to generate the required test data for the desired materials. A word of caution is necessary here. When designers do not have the test data on a material of interest, they sometimes tend to predict these properties by “similarity” to other materials whose properties are known. This approach is risky because even small composition variations can drastically change the oxygen compatibility of a material. Because of their mechanical strength, a vast majority of the materials in oxygen systems will be metallic. Typically, the structural materials are metallic, with nonmetals only being designed into the systems as valve seals and seats, lubricants, or other places where rigid materials cannot be used. Fortunately, metals are generally more compatible with oxygen than nonmetals. Unfortunately, metals become more incompatible with oxygen as the pressures within the system increase. Furthermore, metals tend to burn much more violently, burn for longer durations, and burn at much higher temperatures than nonmetals. Therefore, proper selection of metals for specific applications is critical. No single metal or alloy is best suited for all applications. The test data provided in Tables 2.21 and 2.22 clearly show that many materials will work well in oxygen systems, and many others will not and should be avoided. The designer must choose the best materials by considering all of the factors involved, including oxygen compatibility under use conditions. The materials that are most oxygen compatible include nickel alloys, copper, brass, and bronze. The metals that are more easily ignited, and should be avoided, include magnesium, titanium, and aluminum alloys. The maximum use pressure within the system will drive the material choices.

AEROSPACE MATERIALS CHARACTERISTICS

2.9.5.3

123

CONTROL CONFIGURATION

The energy required to ignite a material or sustain its burning is not a fixed value but is influenced by several factors. The amount of energy required to ignite a material, as measured by the minimum temperature and pressure at which it will ignite, is determined by the thickness of the material, the shape of the part produced by the material, and the surface configuration. The flammability of a metal is strongly influenced by its configuration, or size and shape of the part. Solid metals are the most resistant to ignition and burn the least if ignited. Metal parts with large surface areas will burn more easily than those with smaller surface area. If a metal tube and a metal rod of the same material and diameter are tested, the tube will ignite more easily. Metal mesh materials, such as filter materials, are the easiest of all configurations to ignite. For example, a 316 stainless steel rod will not ignite until the pressure is generally more than 400 psi (2758 kPa), but a 316 stainless steel mesh filter material shaped into a rod of the same dimensions will ignite in oxygen below atmospheric pressure. Stainless steels are used extensively in high-pressure oxygen systems, even though they are flammable under these pressures, by making them hard to ignite. Oxygen system components for high-pressure applications typically have very thick walls to withstand the stress. This thickness is also advantageous for ignition resistance. The data in Table 2.21 for promoted ignition test shows the flammability of a 1/8-in. (3.2-mm) diam rod. If the thickness of the metal is higher than this, then the minimum pressure required to ignite the metal will also be higher. 2.9.5.4

MINIMIZE IGNITION SOURCES AND MECHANISMS

An ideal oxygen system will have nothing inside that could create an ignition potential. There would be no contaminant, no floating debris, no metal shavings from pipe threading, and only pure, clean oxygen entering into the system. In reality, however, ignition sources do exist, and therefore steps must be taken to minimize the effects of the ignition sources and mechanisms that will certainly find their way into the system. Ignition sources are materials that can ignite within an oxygen system and, in turn, spread the burning to other parts or materials. The most common ignition sources are contaminants in the system. The contaminant particles are typically incompatible with the oxygen and can ignite even under ideal system operations. Ignition mechanisms are processes by which ignition may occur within the system even in the absence of an outside ignition source. These mechanisms usually result from improper system designs that allow sufficient energy to cause an ignition event to occur in the system. The safest oxygen systems are designed to mitigate the ignition hazards that are evaluated by the following questions: 1. Which ignition sources could be present in the system? 2. Which ignition mechanisms could be present?

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3. What is the severity of the hazards produced? 4. What hazards could surface from the worst-case scenario, such as any danger to humans or systems? Several ignition sources and mechanisms have been determined by oxygen system users after many years of failure investigations. The most common ignition sources and mechanisms are contamination (FOD), particle impact, rapid pressurization, mechanical impact, friction, static discharge, electrical arc, external heat (e.g., lightning, welding on hardware, fires in vicinity, vehicular impact, static discharge, weather damage), and external hazards (less common). NASA/TM-2007-213740 [4] gives the details of these ignition sources and mechanisms. A few selected topics are discussed in the following sections. A contaminant is any foreign particle or debris that enters into a system and that can come into contact with oxygen. Aerospace system workers commonly use the acronym FOD (foreign object debris) when referring to contaminants within a system. FOD is the most common ignition source and is probably the most dreaded term to an oxygen system user or designer. Contamination entering the system typically comes from one of the following: dirt and other debris entering the system during assembly; excess lubricants that leak from threading; metal flakes that emanate from the pipe threading during assembly; particles brought into the system in impure oxygen; particulates that emanate from within the system, such as by a breakdown of seals, flaking of vessel plating, and valve friction shearing metal pieces; and hand tools, writing instruments, articles of clothing, and other materials that have been dropped into the system by workers. Even the most careful assembly and quality controls cannot prevent every contaminant, but only lessen their frequency. The particle impact ignition mechanism involves heat being generated by a particle striking the surface of another material with a velocity that will generate sufficient heat to ignite the particle or the material. This particle can be introduced during the assembly of the system or by one released during system operation. This ignition mechanism generally requires a metal particle and metal surface to strike. Figure 2.45 shows a scenario in which particle impact ignition could result from a particle striking a flammable part of a valve. Note the orifice accelerating the particle, increasing ignition probability. A particle impact ignition hazard can be mitigated by several strategies, including the following: 1. Oxygen system components must be specifically cleaned for oxygen service and must be assembled using clean techniques to lessen the chances of outside particles entering the system. 2. Pipe threading must be done with care so as not to release particles. 3. Internal filter elements can be used for trapping particles, thus preventing them from flowing downstream and impacting highly flammable materials.

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.45

125

Particle impact ignition from a particle striking a flammable part of a valve [4].

4. Oxygen flow rate should be regulated down to the lowest tolerable level, preferably less than the particle impact threshold limit. 5. Systems must be designed to produce flow paths that have minimal points where particles can strike at severe angles. 6. Components that can generate particles, such as rotating valves and sliding parts, must be minimized. The rapid pressurization, also known as heat-of-compression or adiabatic compression, ignition mechanism involves extreme heat generated by the oxygen gas itself undergoing pressurization. The ideal gas law and thermodynamic equations for an adiabatic process, that is, no heat loss, demonstrate that if the oxygen gas pressure increases rapidly then the gas temperature increases (under adiabatic conditions), sometimes beyond the ignition point of the materials. Figure 2.46 illustrates this mechanism in a system. Polymeric materials are more readily ignited in this way than are metallic materials. Rapid pressurization ignition potential must be minimized in oxygen systems by good design practices. The most important design practices include: 1) limiting pressurization rates; 2) minimizing the nonmetals, including contaminants, in

126

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areas affected by pressurization; 3) burying nonmetals behind metal parts in the flow path; and 4) not compressing oxygen in the vicinity of soft goods, such as exposed valve seats or lubricants. A good practice sometimes involves using devices that can slow the oxygen flow within the system, such as filter elements and various types of meters. The mechanical impact ignition mechanism involves heat energy generated when two objects collide. Numerous collisions occur with the normal operation of an oxygen system. Collisions such as valves closing are designed into the system. However, collisions sometimes occur with large pieces of contaminant debris or by parts of the system breaking free. Both of these sources of mechanical impact must be mitigated when designing oxygen systems. Figure 2.47 shows a mechanical impact ignition hazard that has resulted in a number of fires during an oxygen transfer process. The friction ignition mechanism involves the heat energy generated when two objects rub together. Friction ignition can occur if three factors are present: 1. Two metals are rubbing together. 2. The rubbing must be at a high speed to generate enough heat to ignite one of the metals. 3. The metals must be exposed to high loads, that is, press together with enough force to make the rubbing severe. Figure 2.48 illustrates a condition in which friction ignition can occur in an oxygen system. Some components, such as check valves, regulators, and relief valves, may become unstable and chatter during use. Chattering can result in rapid oscillation of the moving parts within these components, creating a friction ignition hazard. For example, damaged or worn soft goods can result in metal-to-metal rubbing

Fig. 2.46 Rapid pressurization ignition from rapidly compressing oxygen against the nonmetal [4].

AEROSPACE MATERIALS CHARACTERISTICS

127

Fig. 2.47 Mechanical impact ignition from a wrench falling onto an asphalt pad covered in liquid oxygen [4]. between the piston and the cylinder of a reciprocating compressor that could lead to frictional ignition. The frictional ignition hazard should be minimized by avoiding the rubbing together of two metals parts, or by slowing the rate at which they rub. 2.9.5.5

IMPORTANCE OF APPROPRIATE CLEANING

The most important factor in the safe use of any oxygen system is appropriate cleaning to remove contaminants that are a potentially hazardous ignition source. Contaminant particles are easily ignited and can also accumulate to dangerous levels. The following are keys to having clean oxygen systems: 1. Use only oxygen-system-approved cleaning agents and techniques. 2. Clean all parts before system assembly. 3. Minimize the number of different materials used for assembly, such as seal materials and lubricants. (This will allow the assembler to more quickly notice an incorrect material.) 4. Visually inspect all components before assembly and reject any component that does not appear to be 100% visually clean. 5. Be extremely cautious of vendor-supplied parts that are listed as “oxygen compatible.” It is highly recommended that the vendor claim is independently verified or the part is rebuilt using oxygencompatible materials. 6.

Protect clean parts from contamination by keeping them covered until assembly and assembling as many parts as possible in a clean area.

Fig. 2.48 Frictional ignition due to damaged or worn soft goods resulting in metal-to-metal rubbing [4].

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B. N. BHAT

7. Assemble carefully so that two surfaces rubbing together will not generate particulates that can enter the system. 8. Use lubricants only when absolutely necessary and ensure that the lubricants are oxygen compatible. Use the smallest amount of lubricant allowable. 9. Ensure that verification testing of the cleanliness level has been performed and that the cleanliness level is within the predetermined tolerance. 10.

Ensure that vent lines are guarded against contaminants entering the system from the outside.

11.

Use an inert gas to blow through the system to remove any stray particles before oxygen is introduced.

NASA requires that all aerospace components that will contact oxygen be cleaned to meet stringent requirements, as called out in MSFC-SPEC-164 and other NASA documents.

2.9.6 OXYGEN COMPATIBILITY ASSESSMENT The safety of propulsion oxygen systems is especially critical because of the complexity and extreme pressures required to create the oxidizer side of a propulsion system. To elevate the safety of the propulsion oxygen systems, a process tool called an oxygen compatibility assessment, or OCA, has been developed, which is a hazards analysis for oxygen systems. More details on an OCA can be found in NASA TM-213740, mentioned earlier. Industries that evaluate oxygen system safety have adopted evaluation methods very similar to those used by NASA. The OCA is a tool that provides a step-by-step procedure for assessing the risks associated with an oxygen system. It is a stepwise approach that uses the system materials list, drawings of components, and system operating conditions to proceed through a structured risk assessment. The OCA helps prevent some areas from being overlooked while providing a structural approach to any necessary remedies. An oxygen system designer will find that the OCA helps outline the known hazards that may be present, suggests areas that need further review, and provides safety improvement strategies. The basic steps involved in this process are as follows: 1. Determine the worst-case operating conditions at each point within the system. 2. Assess the flammability of the oxygen-wetted materials at their worst-case conditions. 3. Evaluate potential ignition mechanisms and determine their probability of occurrence. 4. Evaluate any kindling chain within the system and determine the most severe possible effects.

AEROSPACE MATERIALS CHARACTERISTICS

129

5. Determine the reaction effects, that is, the severity of the worst potential outcome of a hazard, including human casualty and equipment destruction. 6. Compile the list of hazards and determine methods of correcting them. 7. Document the results for each hazard or component. 8. Provide required modifications and recommendations for system improvement. 9. Provide limitations of the assessment. 10.

Determine the safety of the system if it is built exactly as it is designed.

An OCA will not be necessary in cases where all the materials used within the system are oxygen compatible at the maximum use temperature and pressure of the system. This approach has been tried, but it allowed very few materials to be used, and almost all the systems turned out to be extremely heavy. The desire to allow a wider selection of materials, especially lighter weight materials, necessitated the development of the OCA tool. NASA has provided the following important resources to individuals who are interested in doing work related to oxygen system safety: 1. NASA TM-213740, Guide for Oxygen Compatibility Assessments on Oxygen Components and Systems [4]. This is a good starting point for performing an OCA. 2. The NASA Materials and Processes Technical Information System (MAPTIS) database houses the most current, and most extensive, data and information related to aerospace materials. Literally, tens of thousands of pages of materials information are contained within the database. MAPTIS provides materials information for most aerospace vehicle applications, not just oxidizer systems.

REFERENCES [1] NASA-TM-84105, Report of Apollo 204 Review Board, National Aeronautics and Space Administration, NASA Historical Reference Collection, NASA History Division, NASA Headquarters, Washington, DC. [2] NASA, Johnson Space Center Release 80-039 Investigators File Report on Cause of Spacesuit Backpack Fire, National Aeronautics and Space Administration, NASA Lyndon B. Johnson Space Center, Houston, TX. [3] NASA-TM-X-65270, Report of Apollo 13 Review Board, National Aeronautics and Space Administration, NASA Historical Reference Collection, NASA History Division, NASA Headquarters, Washington, DC. [4] NASA TM-2007-213740, Guide for Oxygen Compatibility Assessments on Oxygen Components and Systems, National Aeronautics and Space Administration, NASA Headquarters, Washington, D.C.

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BIBLIOGRAPHY ASTM D 2512, Standard Test Method for Compatibility of Materials with Liquid Oxygen, ASTM International, West Conshohocken, PA. ASTM G 74, Standard Test Method for Ignition Sensitivity of Materials and Components to Gaseous Fluid Impact, ASTM International, West Conshohocken, PA. ASTM G86, Standard Test Method for Determining Ignition Sensitivity of Materials to Mechanical Impact in Ambient Liquid Oxygen and Pressurized Liquid and Gaseous Oxygen Environments, ASTM International, West Conshohocken, PA. ASTM G124, Standard Test Method for Determining the Combustion Behavior of Metallic Materials in Oxygen-Enriched Atmospheres, ASTM International, West Conshohocken, PA. ASTM G128, Standard Guide for Control of Hazards and Risks in Oxygen Enriched Systems, ASTM International, West Conshohocken, PA. ASTM G145, Standard Guide for Studying Fire Incidents in Oxygen Systems, ASTM International, West Conshohocken, PA. ASTM Manual 36 (MNL36), Safe Use of Oxygen and Oxygen Systems: Handbook for Design, Operation, and Maintenance, ASTM International, West Conshohocken, PA. CGA G-4, Oxygen, Compressed Gas Association Inc., Chantilly, VA. CGA G-4.1, Cleaning Equipment for Oxygen Service, Compressed Gas Association Inc., Chantilly, VA. ISO 14624-4, Space Systems—Safety and Compatibility of Materials—Part 4: Determination of upward flammability of materials in pressurized gaseous oxygen or oxygen-enriched environments, International Organization for Standardization, Geneva, Switzerland. ISO 15859-1, Space Systems—Fluid characteristics, sampling and test methods—Part 1: Oxygen, International Organization for Standardization, Geneva, Switzerland. ISO 22538-1, Space Systems—Oxygen Safety, Part 1: Design of oxygen systems and components, International Organization for Standardization, Geneva, Switzerland. ISO 22538-2, Space Systems—Oxygen Safety, Part 2: Selection of metallic materials for oxygen systems and components, International Organization for Standardization, Geneva, Switzerland. ISO 22538-3, Space Systems—Oxygen Safety, Part 3: Selection of non-metallic materials for oxygen systems and components, International Organization for Standardization, Geneva, Switzerland. ISO 22538-4, Space Systems—Oxygen Safety, Part 4: Hazards analyses for oxygen systems and components, International Organization for Standardization, Geneva, Switzerland. ISO 22538-5, Space Systems—Oxygen Safety, Part 5: Operational and emergency procedures, International Organization for Standardization, Geneva, Switzerland. ISO 22538-6, Space Systems—Oxygen Safety, Part 6: Facility planning and implementation, International Organization for Standardization, Geneva, Switzerland. NASA-STD-6001, Flammability, Offgassing, and Compatibility Requirements and Test Procedures, NASA, Washington, D.C. NASA-STD-6016, Standard Materials and Processes Requirements for Spacecraft, NASA, Washington, D.C. NFPA 53, Recommended Practice on Materials, Equipment, and Systems Used in Oxygen-Enriched Atmospheres, NFPA, Quincy, MA.

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SAE AIR825, Oxygen Equipment for Aircraft, SAE International, SAE Aerospace, Warrendale, PA. SAE AIR825/1, Introduction to Oxygen Equipment for Aircraft, SAE International, SAE Aerospace, Warrendale, PA. SAE AIR825/13, Guide for Evaluating Combustion Hazards in Aircraft Oxygen Systems, SAE International, SAE Aerospace, Warrendale, PA. SAE AIR5648, Fuel Versus Oxygen: Evaluations and Considerations, SAE International, SAE Aerospace, Warrendale, PA. SAE ARP1176, Oxygen System and Component Cleaning, SAE International, SAE Aerospace, Warrendale, PA.

SUGGESTED READING The information provided in this chapter is only an introduction to the topic of oxygen system materials behavior and safe system design. More information is required for a proper design of a complex oxygen system. The following is a list of useful references on the subject. They are publicly available and found through most Internet search engines. NASA Materials and Processes Technical Information System (MAPTIS). This is the official NASA materials database that houses volumes of materials data, including oxygen compatibility information and data. This database is useful to the experienced oxygen system designer, but not as much to the new designer. It can be found at http://maptis.nasa.gov. NASA-STD-6016, Standard Materials and Processes Requirements for Spacecraft. This document provides a more in-depth discussion of material selection for NASA missions. NASA-STD-6001, Flammability, Offgassing, and Compatibility Requirements and Test Procedures. As mentioned earlier in this section, materials must be tested to ensure that they can be used safely in oxygen systems. This NASA standard provides the test methods that NASA uses to determine if a material is acceptable for use on a NASA mission, including those used in oxygen systems. ASTM Manual 36, Safe Use of Oxygen and Oxygen Systems: Handbook for Design, Operation, and Maintenance. This manual is the most comprehensive reference source for materials used in oxygen systems.

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2.10 POLYMERS AND COMPOSITES Joseph H. Koo The University of Texas at Austin

2.10.1

INTRODUCTION

Solid materials are often classified as metals, ceramics, and polymers. This scheme is based primarily on chemical makeup and atomic structures, and most materials fall into one distinct grouping or another, although there are some intermediates. In addition, there are the composites, combinations of two or more of the basic material classes. Polymers are organic compounds that are chemically based on carbon, hydrogen, and other nonmetallic elements (e.g., O, N, and Si). They have very large molecular structures, often chain-like in nature, that have a backbone of carbon atoms. Some of the common polymers are polyethylene (PE), nylon or polyamide (PA), polyvinyl chloride (PVC), polycarbonate (PC), polystyrene (PS), silicone rubber, epoxy, and phenolic. Polymers are different from the other materials in numerous ways but generally possess lower densities, thermal conductivities, and moduli. The lower densities of polymeric materials offer an advantage for applications where light weight is a concern. The addition of thermally and/or electrically conductive fillers allows polymers to be used for insulative or conductive applications. As a result, polymers may find applications in lightning strike, electromagnetic interference (EMI) shielding, and antistatic protection applications. Polymers can further be classified into three basic polymeric categories: thermoplastics, thermosets, and elastomers.

2.10.2

TYPES OF POLYMERS

Thermoplastics, consisting of individual long-chain molecules, can be reprocessed; products can be graduated and fed back into the appropriate blending machine. Thermosets contain an infinite 3-D network, which is created only when the product is in its final form, and cannot be broken down by reheating. Elastomers (rubbers) contain looser 3-D networks, where the chains are free to change their shapes. Neither thermosets nor elastomers can be reprocessed. Some polymers, such as polyurethanes, can be produced in both thermoplastic and thermoset variants. 2.10.2.1

THERMOPLASTICS

The relative importance of thermoplastics can be judged from their U.S. annual production, sales, and captive use data (Table 2.23 [1]). The 2016 total thermosets production is 15,969 million lb. (14.42% of the total resins) versus total thermoplastics production of 95,721 million lb (85.29% of the total resins). The production rate increases of thermoplastics and thermoset from 2015 to 2016 are þ2.73% and þ1.4%, respectively. The first five in the table are regarded as Copyright # 2018 by J.H. Koo. Published by the American Institute of Aeronautics and Astronautics, Inc., with permission.

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.23

133

U.S. RESIN PRODUCTION, SALES, AND CAPTIVE USE, 2015 AND 2016 (MILLIONS OF POUNDS, DRY WEIGHT BASIS)(1)

Resin

Production 2016

Epoxy (2)

2015

Total Sales & Captive Use % Chg 16/15

2016

2015

% Chg 16/15

537

544

21.3

517

498

3.8

Other Thermosets (5)

15,969

15,644

2.1

16,012

15,602

2.6

Total Thermosets

16,506

16,188

2.0

16,529

16,100

2.7

LDPE (2)(3)

7,071

7,017

0.8

7,056

7,144

21.2

LLDPE (2)(3)

14,693

14,605

0.6

14,688

14,604

0.6

HDPE (2)(3)

19,250

18,870

2.0

19,264

18,918

1.8

PP (2)(4)

17,152

17,153

0.0

17,292

17,230

0.4

PS (2)(4)

4,381

4,408

20.6

4,396

4,387

0.2

EPS (2)(3)

1,030

967

6.5

1,046

945

10.7

PVC (3)

15,455

14,704

5.1

15,353

14,711

4.4

Other Thermoplastics (6)

16,689

16,566

0.7

18,049

17,873

1.0

Total Thermoplastics

95,721

94,290

1.5

97,144

95,812

1.4

112,227 110,478

1.6

113,673 111,912

1.6

GRAND TOTAL PLASTICS

(1) Except phenolic resins, which are reported on a gross weight basis. (2) Sales and Captive Use data include imports. (3) Canadian production and sales data included. (4) Canadian and Mexican production and sales data included. (5) Includes: polyurethanes (TDI, MDI and polyols), phenolic, urea, melamine, unsaturated polyester and other thermosets. (6) Includes: PET, nylon, ABS, engineering resins, SB Latex, and other thermoplastics. Sources: Plastics Industry Producers’ Statistics Group (PIPS), as compiled by Vault Consulting, LLC; ACC # American Chemistry Council, March 2017, www.americanchemistry.com.

commodity thermoplastics. Many manufacturers complete to supply these polymers. Prices change quite rapidly, in response to the price of crude oil, and so the table indicates relative prices. The remaining thermoplastics in Table 2.23 are called engineering thermoplastics because of their superior mechanical properties. They are produced on a smaller scale and have prices about twice that of the commodity thermoplastics. Finally, there are specialty plastics which only sell a few thousand tons per annum. An example is polytetra-fluoro ethylene (PTFE), which has unique low-friction properties.

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Thermoplastics can be divided into amorphous and semicrystalline solids. The amorphous polymers are glassy at temperatures lower than Tg (glass transition temperature) and rubbery liquids at higher temperatures. Semicrystalline thermoplastics have an amorphous phase, and a crystalline phase with a melting temperature, Tm. The densities, glass transition, and crystal melting temperatures of some commonly used thermoplastics are listed in Table 2.24 [2]. 2.10.2.2

THERMOSETS

The basic thermosetting resins all will, upon exposure at elevated temperature from ambient to upward of 2328C, undergo an irreversible chemical reaction often referred to as polymerization or cure [3]. Each family has its own set of individual chemical characteristics based on its molecular component and its ability to either homopolymerize, copolymerize, or both. This transformation process represents the line of demarcation separating the thermosets from the thermoplastic polymers. Crystalline thermoplastic polymers are capable of a degree of crystalline cross-linking, but there is little, if any, of the chemical cross-linking that occurs during the thermosetting reaction. The important beneficial factor lies in the inherent enhancement of thermoset resins in their physical, electrical, thermal, and chemical properties due to their ability to maintain and retain these enhanced properties when exposed to severe environments. The cross-linking reaction, which occurs during the fabrication of thermosets, also provides good adhesion to other materials. As a result, epoxy and polyester resins are used for fiberreinforced composites, and phenolics are used for bonding fibers to brake pads, and for sand for metal casting. Table 2.25 shows comparisons of thermoplastic and thermoset resin characteristics [4]. A thermosetting matrix is defined as a composite matrix capable of curing at some temperature from ambient to several hundred degrees of elevated temperature and that cannot be reshaped by reheating. In general, thermosets contain two or more ingredients, a resin matrix with a curing agent that causes the matrix to polymerize (cure) at room temperature. Thermosets can also be thermally cured at an elevated temperature with a long curing cycle without a curing agent. Some commercially available resins (matrices) are polyester and vinyl ester, polyurea, epoxy, phenolic, bismaleimide, polyimide, cyanate ester, and phenyl triazine. Polyester and vinyl esters Polyester matrices have had the longest period of use, with wide application in many large structural applications. They will cure at room temperature with a catalyst (peroxide), which produces an exothermic reaction. The resultant polymer is nonpolar and very water resistant, making it an excellent choice for marine constructions. The isopolyester resins are the most water-resistant polymers in this polymer group. They have been chosen as the prime matrix material for a fleet of U.S. Navy mine sweepers.

POLYMER DENSITIES AND MELTING AND TRANSITION TEMPERATURES OF COMMON THERMOPLASTICS

Abbreviations

Polymer

Density (kg m23)

Tg(8C)

Tm(8C)

Event if Bent through 908

Semi-crystalline plastics P4MP

Poly (4-methyl-pentene-1)

PP

Polypropylene

830

25

238

Semi-brittle

900–910

210

170

Whitens

LDPE

Low-density polyethylene

920–925

2120

120

Ductile

MDPE

Medium density polyethylene

935–945

2120

130

Ductile

HDPE

High-density polyethylene

955–965

2120

140

Ductile

PA 6

Polyamide 6

1120–1150

50

228

Ductile

PA 66

Polyamide 6,6

1130–1160

57

265

Ductile

PET

Polyethylene terephthalate

1336–1340

80

260

Ductile

POM

Polyoxymethylene (Acetal)

1410

285

170

Semi-brittle

PVDC

Polyvinylidene chloride

1750

218

205

PTFE

Polytetrafluoro ethylene

2200

273

332

PS

Polystyrene

1050

100

SAN

Styrene acrylonitrile copolymer

1080

100

990–1100

100

Whitens

1200

145

Ductile

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.24

Ductile

Glassy plastics Brittle

Acrylonitrile butadiene styrene copolymer

PC

Polycarbonate

PVCu

Polyvinyl chloride unplasticised

1410

80

Ductile

PMMA

Polymethyl methacrylate

1190

105

Brittle

Tm, crystal melting temperature; Tg, glass transition temperature.

135

ABS

136

B. N. BHAT

TABLE 2.25

COMPARISONS OF THERMOPLASTIC AND THERMOSETTING RESIN CHARACTERISTICS

Thermoplastic Resin

Thermosetting Resin

†

High molecular weight (MW) solid

† Low MW liquid or solid

†

Stable material

† Low-medium viscosity, requires cure

†

Reprocessable, recyclable

† Cross-linked, nonprocessable

†

Amorphous or crystalline

† Liquid or solid

†

Linear or branched polymer

† Low MW oligomers

†

Liquid solvent resistance

† Excellent environmental/solvent resistance

†

Short process cycle

† Long process cycle

†

Neat up to 30% filler

† Long or short fiber reinforced

†

Injection/compression/ extrusion

† Resin transfer molding (RTM)/filament winding (FW)/sheet molding compound (SMC)/prepreg/pultrusion

†

Limited structural components

† Many structural components

†

Neat resin þ nanoparticles

† Neat or fiber reinforced þ nanoparticles

†

Commodity: high-performance areas for automotive, appliance housings, toys

† Commodity: advanced materials for construction, marine, aircraft, aerospace

Epoxy The most widely used matrices for advanced composites are the epoxy resins, even though they are costly and do not have the high-temperature capability of the bismaleimides or polyimide. The selection factors of epoxy are listed in Table 2.26 [3]. When epoxy is reinforced with fibers, it forms polymer matrix composites (PMCs). The properties of continuous and aligned glass-, carbon-, and aramidfiber reinforced epoxy composites are included in Table 2.27 [5]. Thus, a comparison of the mechanical characteristics of these three composites may be made in both longitudinal and transverse directions. Bismaleimides (BMIs) The BMI resins have found their niche in high-temperature aircraft design applications where temperature requirements are approximately 1778C (3508F). BMI is the primary product and is based on the reaction product from methylene dianiline (MDA) and maleic anhydride: bis (4 maleimidophynyl) methane (MDA BMI). Variations of this polymer with compounded additive to improve

AEROSPACE MATERIALS CHARACTERISTICS

137

impregnation are now on the market and can be used to impregnate suitable reinforcement to result in high-temperature mechanical properties. Polyimides (PIs) The PI resins are the highest temperature polymers in the general advanced composite, with a long-term upper temperature limit of 232–3168C (450–6008F). Polyureas Polyureas involve the combination of novel MDA polymers and either amine ore imino-functional polyether polyols. The resin systems can be reinforced with milled glass fibers, flaked glass, wollastanite, or treated mica, depending on the compound requirements as to processability or final product. Cyanate ester (CE) and phenolic triazine (PT) The CE resins have shown superior dielectric properties and much lower moisture absorption than any other structural resin for composites. The physical properties of CE resins are comparable to those of a representative BMI resin. The PT resins

TABLE 2.26

EPOXY RESIN SELECTION FACTORS [3]

Advantages

Disadvantages

1.

Adhesion to fibers and to resins

1.

Resins and curatives somewhat toxic in uncured form

2.

No by-products formed during cure

2.

Absorb moisture; heat distortion point lowered by moisture and change in dimensions and physical properties due to moisture absorption

3.

Low shrinkage during cure

3.

Limited to about 2008C for upper-temperature use (dry)

4.

Solvent and chemical resistance

4.

Difficult to combine toughness and high-temperature resistance

5.

High or low strength and flexibility

5.

High thermal coefficient of expansion

6.

Resistance to creep and fatigue

6.

High degree of smoke liberation in a fire and flammable

7.

Good electrical properties

7.

May be sensitive to ultraviolet light of degradation

8.

Solid or liquid resins in uncured state

8.

Slow curing

9.

Wide range of curative options

138

B. N. BHAT

TABLE 2.27 PROPERTIES OF CONTINUOUS AND ALIGNED GLASS-, CARBON-, AND ARAMID-FIBER REINFORCED EPOXY MATRIX COMPOSITES IN LONGITUDINAL AND TRANSVERSE DIRECTIONSa [5] Property

Glass (E-glass)

Carbon (high-strength)

Aramid (Kevlar 49)

Specific gravity

2.1

1.6

1.4

Tensile modulus Longitudinal [GPa (106 psi)] Transverse [GPa (106 psi)]

45 (6.5) 12 (1.8)

145 (21) 10 (1.5)

76 (11) 5.5 (0.8)

Tensile strength Longitudinal [GPa (106 psi)] Transverse [GPa (106 psi)]

1020 (15) 40 (5.8)

1240 (180) 41 (6)

1380 (200) 30 (4.3)

Ultimate tensile strain Longitudinal Transverse

2.3 0.4

0.9 0.4

1.8 0.5

a

Fiber volume fraction ¼ 0.60 in all cases.

also possess superior elevated temperature properties, along with excellent properties at cryogenic temperatures. They are available in several viscosities, ranging from a viscous liquid to powder, which facilitates their applications that use liquid resins, such as filament winding (FW) and resin transfer molding (RTM). 2.10.2.3

ELASTOMERS

Another important group of polymers is the group that is elastic or rubberlike, known as elastomers. This group of materials includes thermoplastic elastomers (TPEs), melt processable rubbers (MPRs), thermoplastic vulcanizate (TPV), synthetic rubbers, and natural rubbers (NR) [6]. Eastomers are suitable for applications, such as conveyor belts, pressure hoses, and thin layers of either steel wire or polymeric fiber reinforcement that can bear the main mechanical loads. These thin layers, with elastomer interlayers, allow flexibility in bending, whereas the reinforcement limits the in-plane stretching of the product. The applications are dominated by natural rubber (NR) and styrene butadiene copolymer rubber (SBR). Other rubbers have specialized properties: butyl rubbers have low air permeability, nitrile rubbers have good oil resistance, and silicone rubbers have high and low temperature resistance. The rubbery behavior of the amorphous phase in semicrystalline thermoplastics is an important property. 2.10.2.4

FILLERS

The term fillers refer to solid additives that are incorporated into the polymer matrix. Fillers are generally inorganic materials and can be classified according

AEROSPACE MATERIALS CHARACTERISTICS

139

to their effect on the mechanical properties of the resulting polymer blend. Inert or extender fillers are added mainly to reduce the cost of the compound, whereas reinforcing fillers are added to improve certain mechanical properties, such as modulus or tensile strength. Although termed inert, inert fillers can affect other properties of the compound besides cost. They may increase the density of the compound, reduce the shrinkage, increase the hardness, and increase the heat deflection temperature. Reinforcing fillers will increase the tensile, compressive, and shear strengths; increase the heat deflection temperature; reduce shrinkage; increase the modulus; and improve the creep behavior. Reinforcing fillers improve properties via several mechanisms. In some cases, a chemical bond is formed between the filler and the polymer; in other cases, the volume occupied by the filler affects the properties of the thermoplastic. As a result, the surface properties and the interaction between the filler and the thermoplastic are of great importance. The particle shape, the particle size, distribution of sizes, and the surface chemistry of the particle affects the polymer properties. In general, the smaller the particle, the greater the improvement of the mechanical property, such as tensile strength. Larger particle sizes may have adverse effects. Particle shape can also influence the properties. Plate-like particles or fibrous particles may be oriented during processing. This may result in properties that are anisotropic. The surface chemistry of the particle is important to promote interaction with the polymer and to allow for good interfacial adhesion. It is important that the polymer wet the particle surface and have good interfacial bonding so as to obtain the best property enhancement. Examples of inert or extender fillers include china clay (kaolin), talc, and calcium carbonate. Calcium carbonate is important filler with a particle size of about 1 mm. It is a natural product from sedimentary rocks and is separated into chalk, limestone, and marble. Calcium carbonate may improve interaction with the thermoplastic. Glass spheres are also used as thermoplastic fillers. They may be either solid or hollow, depending on the particular application. Talc is a type of filler with a lamellar particle shape. It is a natural, hydrated magnesium silicate with good slip properties. Kaolin and mica are also natural materials with lamellar structures. Other fillers include wollastonite, silica, barium sulfate, and metal powders. Carbon black is used a filler primarily in the rubber industry, but it also finds application in thermoplastics for conductivity, ultraviolet (UV) protection, and as a pigment. Fillers in fiber form are often used in thermoplastics. Types of fibers include cotton, wood flour, fiberglass, and carbon. Table 2.28 shows the fillers and their forms [7].

2.10.3

POLYMER STRUCTURE AND SYNTHESIS

A polymer is prepared by stringing together a series of low-molecular-weight species, such as ethylene, into an extremely long chain, and as polyethylene, similar to how one would string together a series of beads to make a necklace (Fig. 2.49) [7]. The chemical characteristics of these starting species will determine

140

B. N. BHAT

TABLE 2.28

FORMS OF VARIOUS FILLERS

Spherical

Lamellar

Fibrous

Sand/quartz powder

Mica

Glass fibers

Silica

Talc

Asbestos

Glass spheres

Graphite

Wollastonite

Calcium carbonate

Kaolin

Carbon fibers

Carbon black

Whiskers

Metallic oxides

Cellulose Synthetic fibers

the final properties of the polyethylene. When two different low-molecular-weight species are polymerized, the resulting polymer is termed a copolymer, such as ethylene vinylacetate as depicted in Fig. 2.50 [7]. Polymers can be separated into thermoplastics and thermosets as stated in the previous section. A thermoplastic material is a high-molecular-weight polymer that is not cross-linked. It can exist in either a linear or a branched structure (Fig. 2.51) [7]. Upon heating, thermoplastics soften and melt, which allows them to be shaped using polymer processing equipment. A thermoset has all of the chains tied together with covalent bonds in a 3-D network (cross-linked) as depicted in Fig. 2.51. A thermoset material will not flow once it is cross-linked, but a thermoplastic material can be reprocessed by simply heating it to the appropriate temperature. The different types of structures are shown in Fig. 2.51. The properties of different polymers can vary widely. Properties can be varied for each individual polymer material as well, simply by varying the microstructure of the material. There are two primary polymerization approaches: step-reaction polymerization and chain-reaction polymerization [7]. In step-reaction (also referred to as condensation polymerization), reaction occurs between two polyfunctional monomers, often liberating a small molecule, such as water. As the reaction proceeds, high-molecular-weight species are produced as longer and longer groups react together. For example, two monomers can react to form a dimer, then react with another monomer to form a trimer. The reaction can be described as n-mer þ m-mer ! (n þ m) mer, where n and m refer to the number of monomer units for each reactant. The molecular weight of the polymer builds up gradually with time, and high conversions are usually required to produce highmolecular-weight polymers. Polymers synthesized by this method

Fig. 2.49

Polymerization.

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.50

141

Copolymer structure.

typically have atoms other than carbon in the backbone, such as polyesters and polyamides. Chain-reaction polymerization (also referred to as addition polymerizations) require an initiator for polymerization to occur. Initiation can occur by a free radical or an anionic or cationic species, which opens the double bond of a vinyl monomer, and the reaction proceeds as shown in Fig. 2.49. Chain-reaction polymers typically contain only carbon in their backbone and include such polymers as polystyrene and polyvinyl chloride.

2.10.4

PROCESSING OF POLYMERS

Processing is the technology of converting raw polymer, or compounds containing raw polymer, to parts of desired shape. Considering the wide variety of polymer types and even wider variety of components made from them, a complete description and analysis of the processing techniques would be impossible in this section. An outline of the common processing techniques, introducing relevant terminology and considering how the different techniques are used in the fundamentals previously discussed, is provided. Because polymer matrix composite is the dominant composite used for aerospace structural applications, this type of composite will be the focus of the remainder of this section. 2.10.4.1

BASIC PROCESSING OPERATIONS

Polymer processing operations can be classified into five broad categories: extrusion, molding, spinning, calendaring, and coating. Of these techniques, extrusion is perhaps the most widely used for aerospace composites. Applications of extrusion include the continuous production of composite pipe, sheet, and rods.

Fig. 2.51

Polymer structures.

142

B. N. BHAT

Molding is normally a batch process, principally in the form of compressionmolding used to make aerospace components. 2.10.4.2

FIBER-REINFORCED THERMOSET MOLDING

Many polymers do not possess enough mechanical strength for structural applications. When these polymers are reinforced with high-modulus fibers, the polymeric composites have high strength-to-weight ratios and can be fabricated in a wide variety of complex shapes. Glass and other fibers are used extensively to reinforce thermosetting resins, such as epoxies, bismaleimides, polyimides, and cyanate esters. Fiber glass-reinforced plastics are now used for all boats under approximately 40 ft. in length, truck cabs, low-production-volume automobile bodies, structural panels, and aircraft components. Many of these parts are fabricated by a hand lay-up process. The mold surface is often first sprayed with a pigmented but nonreinforced gel coat of the liquid resin to provide a smooth surface finish. The gel coat is followed by successive layers of fiber glass, either in the form of woven cloth or random matting, impregnated with the liquid resin, which is then cured to give the finished product. The molds are relatively inexpensive because no pressure is required. The major drawback of this process is the expense of the hand labor involved. A special gun chops continuous fibers into approximately 1-in. lengths. The chopped fibers are combined with a stream of liquid resin and sprayed directly onto the mold surface. Although the random chopped fibers do not reinforce as well as a woven cloth or random mat, the labor savings are substantial. This process is used to fabricate sinks, bathtubs, recreational vehicle bodies, and so on. In fiberglass or graphite fishing rods, vaulting poles, and golf club shafts, the fibers are arranged along the long axis to resist the bending stress applied. Filament winding extends this principle to more complex structures. Continuous filaments of the reinforcing fiber are impregnated with liquid resin and then wound on a rotating mandrel. The winding pattern is designed to resist most efficiently the anticipated stress distribution. This technique is used to produce rocket motors and missile structures, gun barrels in the defense industry, and tanks and pipes for the chemical process industries. Pultrusion is used to produce continuous lengths of objects with a constant cross section, such as structural beams. Continuous fibers (rovings) and/or mat are impregnated by passing them through a bed of liquid resin and pulling them slowly through a heat die of the desired cross section. The resin cures to a solid in the die.

2.10.5 2.10.5.1

MECHANICAL BEHAVIOR OF POLYMERS STRESS–STRAIN BEHAVIOR

The mechanical properties of polymers are modulus of elasticity, and yield and tensile strengths. For many polymeric materials, the simple stress–strain test is

AEROSPACE MATERIALS CHARACTERISTICS

143

employed for the characterization of some of the mechanical parameters [8]. The mechanical characteristics of polymers are highly sensitive to the rate of deformation (strain rate), the temperature, and the chemical nature of the environment (the presence of water, oxygen, organic solvents, etc.). Typically, three different types of stress–strain behavior are found for polymeric materials, as illustrated in Fig. 2.52 [9]. Curve A illustrates the stress-strain behavior of a brittle polymer, which fractures while deforming elastically. Curve B is similar to that for many metallic materials; the initial deformation is elastic, which is followed by yielding and a region of plastic deformation. The deformation displayed by curve C is totally elastic; this rubberlike elasticity (large recoverable strains produced at low stress levels) is displayed by elastomers. Modulus of elasticity (termed tensile modulus) and ductility in percent elongation are determined for polymers using the stress–strain curve (Fig. 2.52). For plastic polymers (curve B, Fig. 2.52), the yield point is taken as a maximum on the curve, which occurs just beyond the termination of the linearelastic region (Fig. 2.53) [9]. The stress at this maximum is the tensile yield strength (O0y). Tensile strength (TS) corresponds to the stress at which fracture occurs (Fig. 2.53); TS may be greater than or less than O0y. Table 2.29 shows some mechanical properties for several polymeric materials [9]. 2.10.5.2

STRENGTH

There are different kinds of strength. Tensile strength is important for a material to be stretched or under tension. A polymer has compression strength if it is strong when under compression. A polymer has flexural strength if it is strong under compression or bending. A polymer sample’s torsional strength is tested

Fig. 2.52 Stress–strain behavior for brittle (curve A), plastic (curve B), and highly elastic (elastomeric) (curve C ) polymers.

144

B. N. BHAT

Fig. 2.53 Stress–strain curve for a plastic polymer showing how yield and tensile strengths are determined. by twisting it. A polymer sample has impact strength if it is strong when one hits it sharply and suddenly with a hammer or similar. 2.10.5.3

ELONGATION

Elongation is a type of deformation. When a polymer sample deforms by stretching, becoming longer, the deformation is known as elongation. It is usually measured in percent, which is the length of the polymer sample is after it is stretched (L), divided by the original length of the sample (Lo), and multiplied by 100. L=Lo  100 ¼ % elongation Two important elongations that are measured are ultimate elongation and elastic elongation. 2.10.5.4

FRACTURE

If a glassy polymer is stressed very rapidly or is stressed at a temperature that is much lower than its glass transition temperature, it tends to break or fracture in a brittle manner. An amorphous polymer tends to draw down in a homogenous manner if stressed at temperature above its glass transition temperature and shows large strains before fracturing. At intermediate temperatures and low rates of deformation, the polymer can yield before fracture or fracture in a ductile manner by neck formation. At low strain (i.e., ,1%), the deformation of most polymers is elastic, which means that the deformation is homogenous and full recovery can occur over a finite time. At high strains, the deformation of glass polymers occurs either by crazing, characteristic of brittle polymers, or by a process called shear bonding, which is the dominant mechanism for ductile polymers. Such deformations are not reversible unless the polymer is heated above its glass transition temperature.

TABLE 2.29

Tensile Modulus [GPa (ksi)]

Tensile Strength [MPa (ksi]

Yield Strength [MPa (ksi)]

Elongation at Break (%)

Polyethylene (low density)

0.917–0.932

0.17–0.28 (25–41)

8.3–31.4 (1.2–4.55)

9.0–14.5 (1.3–2.1)

100–650

Polyethylene (high density)

0.952–0.965

1.06–1.09 (155–158)

22.1–31.0 (3.2–4.5)

26.2–33.1 (3.8–4.8)

10–1200

Poly(vinyl chloride)

1.30–1.58

2.4–4.1 (350–600)

40.7–51.7 (5.9–7.5)

40.7–44.8 (5.9–6.5)

40–80

Polytetrafluoroethylene

2.14–2.20

0.40–0.55 (58–80)

20.7–34.5 (3.0–5.0)

—

200–400

Polypropylene

0.90–0.91

1.14–1.55 (165–225)

31–41.4 (4.5–6.0)

31.0–37.2 (4.5–5.4)

100–600

Polystyrene

1.04–1.05

2.28–3.28 (330–475)

35.9–51.7 (5.2–7.5)

—

1.2–2.5

Poly(methyl methacrylate)

1.17–1.20

2.24–3.24 (325–470)

48.3–72.4 (7.0–10.5)

53.8–73.1 (7.8–10.6)

2.0–5.5

Phenol-formaldehyde

1.24–1.32

2.76–4.83 (400–700)

34.5–62.1 (5.0–9.0)

—

1.5–2.0

Nylon 6,6

1.13–1.15

1.58–3.80 (230–550)

75.9–94.5 (11.0–13.7)

44.8–82.8 (6.5–12)

15–300

Polyester (PET)

1.29–1.40

2.8–4.1 (400–600)

48.3–72.4 (7.0–10.5)

59.3 (8.6)

30–300

Polycarbonate

1.20

2.38 (345)

62.8–72.4 (9.1–10.5)

62.1 (9.0)

110–150

Source: Modern Plastics Encyclopedia ’96. Copyright 1995, The McGraw-Hill Companies. Reprinted with permission.

145

Specific Gravity

AEROSPACE MATERIALS CHARACTERISTICS

Material

ROOM-TEMPERATURE MECHANICAL PROPERTIES OF SOME COMMON POLYMERS

146

2.10.5.5

B. N. BHAT

MODULUS

Modulus is measured to know how well a material resists deformation. There are different types of modulus. Young’s modulus is the material resistance to tensile deformation in a uniaxial tensile deformation and is calculated by the slope of the stress–strain curve in the initial linear region. Shear modulus is the material resistance to shear deformation. Flexural modulus is the material’s resistance to bending. 2.10.5.6

TOUGHNESS

Toughness is really a measure of the energy a polymer sample can absorb before it breaks. Strength tells how much force is needed to break a sample, and toughness tells how much energy (force x distance) is needed to break a sample. Area under the stress-strain curve is a measure of this energy (see Fig. 2.52). Because a polymer sample is strong does not mean it will be tough as well.

2.10.6

AGING OF POLYMERS

For extended periods of exposure to elevated temperatures, moisture, oxygen, and ultra-violet rays, polymers are known to exhibit changes in their molecular structure that can have profound effects on the long-term durability of the polymer and the associated composite material. This process is known as aging [10]. Aging is classified into three different types: physical, chemical, and hydrothermal. Physical aging is the slow contraction of the molecular network that occurs over time when exposed to sub-Tg temperatures. This densification of the material leads to embrittlement and loss of toughness. Chemical aging is the oxidation and continued curing mechanisms that act on the polymer structure during extreme environmental exposure. This aging mechanism can lead to embrittlement, loss of toughness, and discoloration of the polymer. Hydrothermal aging introduces changes into the polymer structure due to the presence of moisture. As water molecules diffuse into the polymer from the structural surface, the water can change the bulk polymer density, can reduce the mechanical integrity of the material, and can chemically react with the polymer to alter the bulk polymer properties. All three aging mechanisms are generally detrimental to the durability of a polymer. Many polymers are formulated to specifically resist the effects of one or more aging mechanisms.

2.10.7 2.10.7.1

POLYMERS FOR AEROSPACE COMPOSITES HISTORY OF RESINS (MATRICES)

The introduction of fiberglass-reinforced structural application in 1949 resulted in a new polymer application. It began with the consumption of

AEROSPACE MATERIALS CHARACTERISTICS

147

10 lbs. annually and grew to a few billion pounds annually over the next several decades. This use has been and still is taking place in applications that take advantage of the extraordinarily low weight to high strength ratio inherent in these fiber-reinforced composite materials. Polymers used as the matrix component of composites in aerospace structures must be exceptionally durable to resist the cyclic mechanical loads associated with aircraft operation. Furthermore, many aerospace structures must also be tolerant of high temperatures, moisture levels, thermal cycling, and UV radiation. For most aerospace vehicles, cost is not as important a consideration as in other industries, and so higher performing polymers that may be relatively expensive are often chosen over less expensive polymers with lowerperforming qualities. Thermosetting polymers are typically used in aerospace applications where high temperatures are expected. Examples of thermosetting composites include the skin of supersonic aircraft and some reentry vehicles. The most common thermosets used in aerospace vehicles are epoxies, BMIs, CEs, phenolics, and PIs [11, 12]. Epoxy resins are the most commonly used matrix material for composites, but their use in aerospace structures is limited by relatively low service temperatures and susceptibility to brittleness and moisture attack. BMIs are similar to epoxies except for improved temperature performance with a relatively high performance-to-cost ratio. CE resins are more easily processed than BMIs and offer excellent strength, moisture absorption, and electrical properties. Similar to epoxy and BMIs, CE resins can be brittle if not toughened properly. Phenolic resins are low-cost, flame-resistant, and lowsmoke resins that are primarily used in interior aircraft panels, ablative, and rocket nozzle applications. Thermosetting PI resins have good adhesion and heat and chemical resistance as well as superior mechanical properties. The processing of some PI resins can be difficult due to the need for using highly corrosive chemicals. Thermoplastic polymers are a more cost-effective alternative for thermosets when extremely high temperature-resistant requirements are not needed. Thermoplastics also have a much higher impact strength than thermosets. Uses of these materials in aerospace applications include aircraft interiors, wing ribs and panels, and land gear doors. Common thermoplastic polymers used in aerospace structures include poly(ether ketone) (PEEK), polyetherimide (PEI), poly(p-phenylene sulfide) (PPS), and thermoplastic polyimide (PI). PEEK has good resistance to moisture attack, corrosion, abrasion, and chemical and radiation exposure. PEEK has low smoke emission and excellent stiffness. PEI polymer has high heat resistance, strength, and modulus, but can degrade with exposure to aggressive fluids such as hydraulic fluid. PPS has excellent resistance to oils, fuels, solvents, anti-icing agents, and acids/bases. It also has excellent hardness, dimensional stability, and excellent fire resistance. Thermoplastic polyimide, like its thermosetting counterparts, offer excellent performance at elevated temperatures.

148

2.10.7.2

B. N. BHAT

REINFORCEMENTS [12]

Fiberglass Fiberglass possess high tensile strength and strain to failure, but the real benefits of its use relate to its heat and fire resistance, chemical resistance, moisture resistance, and very good thermal and electrical properties. Graphite Graphite fibers have the widest variety of strength and moduli and also the greatest number of suppliers. These fibers start out as organic fiber: rayon, polyacrylonitrile, or pitch called the precursor. The precursor is stretched, oxidized, carbonized, and graphitized. The relative amount of exposure to temperatures from 2,500 to 3,0008C will then determine the graphitization level of the fiber. A higher degree of graphitization will usually result in a stiffer (high-modulus) fiber with greater electrical and thermal conductivities. Aramid The organic fiber Kevlar 49, an aramid, essentially revolutionized pressure vessel technology because of its great tensile strength and consistency coupled with low density, resulting in very weight-effective designs for rocket motors. Boron Boron fibers, the first fibers to be used in the production of aircraft, are produced as individual monofilaments upon a tungsten or carbon substrate by pyrolytic reduction of boron trichloride (BCL) in a sealed glass chamber.

2.10.8

OVERVIEW OF MANUFACTURING OF COMPOSITES

There are different fabrication processes relating to the product forms and their processes to manufacture composite parts, as shown in Table 2.30. The processes include pultrusion, RTM, compression-molding, filament winding, hand lay-up, and auto tape laying [13]. This section includes a brief discussion of these fabrication processes to manufacture composite parts based on F. C. Campbell’s excellent textbook [13]. 2.10.8.1

PREPREG LAY-UP

Prepreg lay-up is a manufacturing process in which individual layers of prepreg are laid up on a tool and then cured, as shown in Fig. 2.54. The layers are laid up in the required direction and to the desired thickness. A thin nylon vacuum bag is the placed over the lay-up and the air is evacuated to draw out the air between the plies. The bagged part is placed in an oven or an autoclave (a heated pressure vessel) and cured under the specified time, temperature, and pressure. If oven curing is used, the maximum pressure that can be obtained is atmospheric (14.7 psia or less). An autoclave (Fig. 2.55) works on the principle

Material Form/Process

TYPICAL MATERIAL PRODUCT FORMS VS PROCESS

Pultrusion

RTM

Compression Molding

Filament Winding

Hand Lay-up

†

†

Auto Tape Laying

Discontinuous Sheet molding compounds

†

Bulk molding compounds

†

Random Continuous Swirl mat/neat resin

†

†

†

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.30

Oriented Continuous Unidirectional tape

†

†

Woven prepreg

†

†

Woven fabric/neat resin

†

†

†

Prepreg roving

Preform/neat resin

†

†

Stitched material/neat resin

Roving/neat resin

†

†

†

†

†

† 149

150

B. N. BHAT

Fig. 2.54

Prepreg hand lay-up process [13].

of differential gas pressure. The vacuum bag is evacuated to remove the air, and the autoclave supplies gas pressure to the part. The pressure vessel contains a heating system with a blower to circulate the hot gas. The autoclave offers the

Fig. 2.55

Principle of autoclave curing [13].

AEROSPACE MATERIALS CHARACTERISTICS

151

advantage that much higher pressures (e.g., 100 psig) can be used, resulting in better compaction, higher fiber volume percentages, and fewer voids and less porosity. Presses can also be used for this fabrication process, but they have several disadvantages, as follows: 1. The size of the part is limited by the heated press platen size. 2. Platens may produce high- and low-pressure spots if the platens are not exactly parallel. 3. Complex shapes are difficult to produce. Automated ply cutting, manual collation or lay-up, and autoclave curing (Fig. 2.56) is the most widely used fabrication process for high-performance composites in the aerospace industry. Although manual ply collation is expensive, this

Fig. 2.56

Traditional lay-up and autoclave cure process [13].

152

B. N. BHAT

Fig. 2.57

Filament winding process [13].

process is capable of making high-quality, complex composite parts. Because cost has become a major driver, a tremendous amount of research is being conducted to identify more cost-effective material product forms and processes (Table 2.30). 2.10.8.2

FILAMENT WINDING

Filament winding (Fig. 2.57) is a process that has been used for decades to build highly efficient structures that are bodies of revolution or near bodies of revolution, such as solid rocket motors. Wet winding, in which dry fiber rovings are pulled through a resin bath before winding on the mandrel, is the most prevalent fabrication process, but prepregged roving or tows can also be filament wound. Curing is usually conducted in an oven with or without a vacuum bag. Hoop windings are often applied over separator sheets to provide compaction pressure during cure. 2.10.8.3

WET LAY-UP PROCESS

The wet lay-up process (Fig. 2.58) is often used to build large structures such as yacht hulls. It can be a cost-effective fabrication process when the quantities required are small. Dry reinforcement, usually woven cloth or mat material, is hand-laid one ply at a time. During or before lay-up, each ply is impregnated with a low-viscosity resin. After the ply is placed on the lay-up, hand rollers are used to remove excess resin and air and compact the plies. After lay-up, cure can be done at room or elevated temperatures. Frequently, cure is conducted

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.58

153

Wet lay-up process [13].

without a vacuum bag, but vacuum pressure helps to improve laminate quality. Because cure is usually conducted at room temperature or low temperature, very inexpensive tooling (e.g., wood) can be used to minimize cost. 2.10.8.4

SPRAY-UP PROCESS

The spray-up process (Fig. 2.59) is a more cost-effective process than wet lay-up, but the mechanical properties are much lower due to the use of randomly oriented chopped fibers. Normally, continuous glass rovings are fed into a special gun that chops the fibers into short lengths and simultaneously mixes them with either a polyester or vinyl ester resin that is then sprayed onto the tool. Manual

Fig. 2.59

Spray-up process [13].

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compaction with rollers is again used to compact the lay-up. Vacuum bag cures can be used to improve part quality but are not normally used. Because the fibers are short and the orientation is random, this process is not used to make structural load-bearing parts. 2.10.8.5

LIQUID MOLDING

Liquid molding covers a fairly extensive set of fabrication processes. In RTM (Fig. 2.60), a dry preform or lay-up is placed in a matched metal die and a lowviscosity resin is injected under pressure to fill the die. Because this is a matcheddie process, it is capable of holding very tight dimensional tolerances. The die can contain internal heaters or can be placed in a heated press for cure. Other variations of this process include vacuum-assisted RTM (VARTM), in which a singlesided tool is used along with a vacuum. Instead of injecting the resin under pressure, a vacuum pulls the resin through a flow medium that helps impregnate the preform. 2.10.8.6

COMPRESSION MOLDING

Compression-molding (Fig. 2.61) is another matched-die fabrication process that uses either discontinuous, randomly oriented sheet molding compound (SMC) or bulk molding compound (BMC). A charge of predetermined weight is placed between the two dies and then heat and pressure are applied. The molding compound flows to fill the die and then rapidly cures in 1–5 minutes, depending on the type of polyester or vinyl ester resins used. Thermoplastic composites, usually consisting of glass fiber and polypropylene, are also compression-molded for the automotive industry.

Fig. 2.60

Resin transfer molding process [13].

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Fig. 2.61 2.10.8.7

155

Compression-molding process [13].

INJECTION-MOLDING

Injection-molding (Fig. 2.62) is a high-volume fabrication process capable of making small- to medium-size parts. The reinforcement is usually chopped glass fibers with a thermoplastics resin, because they process faster and have higher toughness than thermoset resin. In the injection-molding fabrication process, polymer pellets containing embedded fibers or chopped fibers and polymer are fed into a hopper. They are heated to their melting temperature and then injected under high pressure into a matched metal die. After the thermoplastic parts cool, they are ejected and the next cycle is started.

Fig. 2.62

Injection molding process [13].

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Fig. 2.63 2.10.8.8

Pultrusion process [13].

PULTRUSION PROCESS

The pultrusion process (Fig. 2.63) is a rather specialized composite fabrication process that is capable of making long constant-thickness parts. Dry E-glass rovings are normally pulled through a wet resin bath and are then preformed to the desired shape before entering a heated die. Mats and veils are frequently incorporated into the finished part. Cure occurs inside the die, and the cured part is pulled to the desired length and cut off. Quick-curing polyesters and vinyl esters are the predominant resin systems.

2.10.9

SUMMARY

This section provided a brief introduction of the three different categories of polymers: thermoplastics, thermosets, and elastomers. Examples of each were given. The processing of polymers was introduced and an overview of fabrication processes to manufacture composites (prepreg lay-up, autoclave curing, traditional lay-up and autoclave cure, filament winding, wet lay-up, spray-up, resin transfer molding, compression-molding, injection-molding, and pultrusion) was given.

REFERENCES [1] Data available online at www.americanchemistry.com [retrieved March 19, 2018]. [2] Mills, N. J., Plastics—Microstructure and Engineering Application, 3rd ed., Elsevier, Oxford, 2005, p. 15.

AEROSPACE MATERIALS CHARACTERISTICS

[3]

[4] [5] [6] [7] [8] [9] [10]

[11] [12]

[13]

157

Wright, R. E., “Thermosets, Reinforced Plastics, and Composites,” Handbook of Plastics, Elastomers, and Composites, 4th ed., McGraw-Hill, New York, 2002, pp. 109–188. Koo, J. H., Fundamentals, Properties, and Applications of Polymer Nanocomposites, Cambridge University Press, Cambridge, United Kingdom, 2016, p. 112. Floral, R. F., and Peters, S. T., “Composite Structures and Technologies,” tutorial notes, 1989. Margolis, J. M., “Elastomeric Materials and Processes,” Handbook of Plastics, Elastomers, and Composites, 4th ed., McGraw-Hill, New York, 2002, pp. 189–228. Baker, A.-M. M., and Mead, J., “Thermoplastics,” Handbook of Plastics, Elastomers, and Composites, 4th ed., McGraw-Hill, New York, 2002, pp. 1–108. ASTM Standard D638, Standard Test Method for Tensile Properties of Plastics, ASTM International, West Conshohocken, PA. Callister, W. D., Materials Science and Engineering: An Introduction, 7th ed., Wiley, New York, 2007, pp. 525–543. Odegard, G. M., and Bandyopadhyay, A., “Physical Aging of Epoxy Polymers and Their Composites,” Journal of Polymer Science Part B: Polymer Physics, Vol. 49, 2011, pp. 1695–1716. Fried, J. R., Polymers in Aerospace Applications, Smithers Rapra Technology, Shropshire, U.K., 2008. Wright, R. E., “Thermosets, Reinforced Plastics, and Composites,” Handbook of Plastics, Elastomers, and Composites, 4th ed., McGraw-Hill, New York, 2002, pp. 166–172. Campbell, F. C., Manufacturing Process for Advanced Composites. Elsevier, New York, 2004, pp. 17–35.

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2.11 POLYMER MATRIX COMPOSITES IN AEROSPACE STRUCTURES Alan Nettles NASA Marshall Space Flight Center

2.11.1

INTRODUCTION

This section addresses the subject of carbon fiber–reinforced polymer matrix composites (PMC) (aka carbon fiber PMCs or simply carbon fiber composites) in aerospace structures. The subject matter is based on launch vehicle (LV) dry structures such as interstage and payload fairing as examples, although the content can apply to other launch vehicle components (Chapter 7), aircraft (Chapter 5), and spacecraft (Chapter 6). Many of today’s launch vehicles (Fig. 2.64) use carbon fiber composites in their structures. This section will briefly discuss the following topics: differences between composites and metallic materials, composites for LVs and how they differ from use in aircraft, methodology of generating “allowables” for composites, why fatigue is not an issue in most composite structures (and certainly not in LVs), mechanical testing of composites (tension and compression), and predicting composite tensile and compression strength.

2.11.2

CARBON FIBER COMPOSITES VS ALUMINUM

Because the apparent alternative to using composites for most aerospace structures is high-strength aluminum or aluminum-lithium alloys, some of the differences in these materials is highlighted. The approximate cost and classification [1–3] of these materials is presented in Fig. 2.65a. Typical stress–strain curves for Al-Li and carbon/epoxy PMCs are shown in Figs. 2.65a and 2.65b to give the reader an idea of the behavior and relative strengths of these two materials in both their strongest and weakest directions. This chart also notes that Al-Li is like composites in that it is not isotropic and is sometimes considered a “composite” (Fig. 2.65a). Figure 2.65b gives the impression that carbon/epoxy is at least four times stronger than Al-Li, which, given the limited configuration of the test specimens, it is. Figure 2.65c shows the extreme lack of strength in a direction perpendicular to the fibers. In practical applications, loads are rarely totally in one direction. Furthermore, without multi-angle layers, composites tend to split parallel to the fibers. Thus, for composites, fibers are typically placed in a multidirectional pattern as shown in Fig. 2.66, usually in angles of 08, 458, and 908 and 1358 (more commonly denoted by 2458). The multidirectional composite is stronger (120 ksi vs 80 ksi) and lighter (1.3 g/cm3 vs 2.5 g/cm3) than AL-Li, but in practice, no structure made from composites can be designed to these strengths due to allowances in damage, flaws, and/ or other stress concentrations that are inevitable in any practical structure. This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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159

Ariane 5

Atlas 5

GSLV

Falcon 9

Fig. 2.64

Delta 4

H iiB

Examples of launch vehicles that use carbon fiber composites.

For composites, strength vs flaw (or damage severity) looks like Fig. 2.67. In this example, compression strength vs the increasing impact damage severity level is plotted. In general, there is a rapid decrease in strength with increasing

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Fig. 2.65a

Candidate materials for use on launch vehicles.

Fig. 2.65b Stress–strain curves of Al/Li [4] and unidirectional carbon epoxy in strongest direction.

Fig. 2.65c Stress-strain curves of Al/Li [4] and unidirectional carbon epoxy in weakest direction.

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161

Fig. 2.66 Schematic of multidirectional composite and associated stress– strain curve.

120

Stress (ksi)

damage severity. This drop in strength levels out at large damage severity levels in this example (representing penetration of the lamiStrain 1.5% nate due to the impact event). The pristine (undamaged) value of 120 ksi listed in the figure can never be used in design. Knowing this “ideal” value makes one regret what might be realized if not having to deal with damage or stress concentrations, which is an unrealistic scenario. The more realistic and practical value of compression strength at a lower value to account for damage and/or stress concentrations is what will inevitably be designed to on virtually all aerospace hardware. Damage in metallic structure also should be accounted for, but there are major differences. Figure 2.68 is a simple schematic that shows what is of concern in metallic structures. Metallic structures are concerned with the growth of a single crack, whereas composites are immune to single cracks due to the immediate blunting of the crack by the embedded fibers, as shown in Fig. 2.69. References to “crack growth” in current composite documents are a holdover from metallic structure and can be misleading. Rather than a single crack growing, composites are mainly concerned with delaminations, or the separating of the layers making up the laminate as shown in Fig. 2.70. This is an important concept to bear in mind when attempting to

120 ksi

Compression strength

Can never be used!

Highest strength with damage Penetration (hole) Smallest damage being designed to Damage severity

Fig. 2.67 Example of compression strength of a composite laminate vs increasing impact energy.

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Fig. 2.68

Schematic of a crack growing in a metallic structure.

apply metallic-type damage-tolerance requirements to composites. The small microcracks between plies are benign compared with the delaminations. As to determining which material to use (metallic or composite), the first question that needs to be answered is, What do you need the part to do? There is no single material suitable for all structures in a launch vehicle. A plethora of materials are at the designer’s disposal, and carbon fiber laminates are simply one of these engineering materials, suitable for some needs but certainly not for

Gray => Matrix (resin) Top view

Side view

Fig. 2.69

Black => Fiber

Microscopic scale

Fibers arrest cracks

Schematic of crack blunting in a composite.

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Fig. 2.70 Laminate cross-sectional schematic showing the difference between cracks and delaminations. all structural parts. Questions on materials selection such as this one are discussed in detail in Chapter 3. After it is determined that carbon fiber PMC laminates are the material of choice, how to use this material and make aerospace parts from it is then explored. One of the biggest problems the author has noted within the launch vehicle industry is that composites knowledge comes from the aircraft industry, but if one is not building an airplane, much of this knowledge is not applicable.

2.11.3

COMPOSITES FOR LAUNCH VEHICLES

Figure 2.71 presents the basic concept of this section: a rocket is not an airplane.

Fig. 2.71

Basic concept of this section.

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Airplanes are designed for thousands of flights over decades, and most LVs are flown once and then discarded. Thus the methods for designing each are very different. This section will point out many of these differences and how LVs have an advantage over airplanes in that not as much up-front time and money need to be spent in materials characterization. First, several types of composites have been used in manned LVs and spacecraft. For example, Table 2.31 lists some of these composites used in the Space Shuttle Orbiter. Most of the information concerning how to design and build with composites laminates comes from the aircraft industry in general and the Composite Materials Handbook (CMH-17) in particular [5]. However, LV structure is typically stiffness driven and aircraft structure is mostly strength driven. Aircraft can be designed for “get-home” loading cases, whereas there is no control over a LV after launch. LVs are inspected in detail before each flight, whereas aircraft can go many flights before an inspection. Finally, LVs are not (or should not be) subjected to foreign object impacts, whereas airplanes are under a constant barrage of foreign object impacts (hail, runway debris) and must continue to fly after these events. These are some of the differences that simplify what needs to be done as far as material characterization when building an LV vs building a large airplane. The concept of lifetimes and inspection intervals, so often mentioned in CMH-17, is relevant to airplanes and helicopters and emphasizes the economic drivers of airplane design. Rockets are much easier to deal with because they TABLE 2.31

COMPOSITES FLOWN IN SPACE SHUTTLE ORBITERA

Composite Material Type

a

Application

Reinforced carbon–carbon

Fuselage nose cap and wing leading edge Fuselage chin panel

Graphite/epoxy

Payload bay doors and Orbital Maneuvering System (OMS) pod shells Wing intermediate spar webs Orbital Vehicle (OV)-103 and subsequent)

Boron/epoxy

Thrust structure stiffening

Boron/aluminum

Midfuselage frame tubes

Aluminum honeycomb sandwich

Upper wing panel Elevons and body flap Rudder/speed brakes Nose Landing Gear (NLG) and Main Landing Gear (MLG) doors Internal airlock

Orbiter Structural Design, NASA-Johnson Space Center (https://www.nasa.gov/centers/johnson/pdf/584733main_Wingsch4g-pgs270-285.pdf)

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Fig. 2.72

165

Example of “lifetimes” of an airplane wing and a launch vehicle component.

typically have a lifetime of one flight, not thousands, between either the next flight (such as the shuttle) or discarding (such as the Space Launch System). An example of the differences in “lifetimes” of an airplane and a rocket are shown schematically in Fig. 2.72. It has been established that LV hardware sees just a fraction of the number of load cycles that an aircraft part does. Thus, the problem of fatigue is eliminated and need not be a part of any future discussion with regard to LVs. Figure 2.73

Fig. 2.73

Example of fatigue data with reference to launch vehicle design space [7].

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is experimental evidence of this fact. Note that for the major design driver for LV hardware and that mentioned in MSFC-RQMT-3479 [6], namely compression after impact, the impact damage does not reduce the fatigue life of the material until at least one million cycles at or above 60% of ultimate load (UL). This kind of severe high-amplitude loading is not an issue for most aircraft parts and certainly not for any LV hardware. In fact, the “design space for launch vehicles” area of Fig. 2.73 is not even populated with any data as it is known that no detrimental effects to the laminate can occur in this region and thus it is simply not tested for. It should also be noted that the motivation for using composites is different for aircraft and for LVs. In aircraft, there are many reasons that composites are preferred over metallic, weight reduction being just a small driver. Composites allow a commercial airplane to operate with higher pressure and humidity in the cabins, increasing passenger comfort. The excellent corrosion and fatigue resistance of composites allow longer intervals between costly inspections and larger windows to exist in the fuselage. For LVs, weight reduction is the only reason to consider composites over metallic. Until the differences in damage-tolerance requirements between aircraft and LVs are considered, composite LV parts will not be as light as their metallic counterparts. Figure 2.74 shows two photographs of the types of damage that are in the design space of aircraft. Extreme amounts of damage must be tolerated for an aircraft. In contrast, Fig. 2.75 shows the types of superficial damage that will ground a LV until the damage is dispositioned. In both of these instances, the hardware was scrapped. Another benefit of not copying what aircrafts do is that the best fiber/resin system for the given part can be used. Consider, for instance, the CAI strength

Fig. 2.74

Examples of the extreme damage aircraft are designed to withstand.

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Fig. 2.75

167

Example of minor damage that will ground a launch vehicle.

data of various fiber/resin systems shown in Fig. 2.76. Note the superior strength of the 8551-7 resin. For a piece of hardware driven by CAI strength (e.g., LV interstage), the IM7/8551-7 systems appears to be the best. However this system is rarely used on aircraft because it has inferior “hot/wet” strength, and the hot/ wet environment is required for aircraft components. The hot/wet environment is not applicable to virtually all LV hardware because the “wet” portion for airplanes is driven by takeoff/landing cycles, which launch vehicles do not have.

Fig. 2.76

Compression after impact (CAI) strength data of various fiber/resin systems.

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This is often overlooked, and a system such as IM7/977-3 is often chosen because “airplanes use it a lot” or because “the F-22 program has produced a large database of coupon test data.” Again, a launch vehicle is not an airplane. Simply copying what the aircraft industry has done should be questioned, and requirements should be chosen that are launch-vehicle specific.

2.11.4

ALTERNATE METHOD FOR GENERATING ALLOWABLES

This section is an attempt to raise awareness about generating costly and timeconsuming “allowables” for pristine lamina and laminate strength properties. The author is suggesting that this lower level of the “building block approach” can be eliminated because the data will never be used; instead strength data should be developed on the damaged composite, because this is what the vehicle is being designed to. In fact, a common methodology used with carbon fiber laminates is to set the maximum allowable strain at 0.4%; this is well below what any of the measured allowables on coupon specimens will give. One of the concepts of this alternative methodology is that if a loading case is not governing the structure being designed, then developing allowables for this benign, or even nonexistent, type of loading is a waste of time and resources. For example, if a rocket motor case is the hardware in question, then developing compression and shear allowables may not be warranted as the structure is so heavily tension dominated. Figure 2.77 is a visual representation of the basic concept of this alternative methodology. As damage severity increases, the strength that one uses for design decreases, as noted previously in Sec. 2.11.2. If a certain amount of damage must be assumed to exist in the structure, then testing for higher strength

Fig. 2.77

Strength data showing level at which allowables should be developed.

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169

values in pristine configurations is useless and only serves to cause regrets about the drastically higher values that could have been used if not for having to deal with damage. Figure 2.77 uses compression as the measure of strength and impact energy as the measure of damage severity, noting where to obtain the allowables. If damage greater than what was designed to exists, then the part is repaired or rejected before flight. The basic steps typically taken to arrive at “allowables” for a composite component include generating strength allowables for all loading cases (or all failure loads). This means that a statistically significant amount of strength values for tension, compression, and shear are determined for unidirectional lamina no matter the critical loading of the hardware in question. The analysts require these lamina-level properties to put into some commercial code to predict the strength of subsequent laminates, and these laminates are then tested to verify the code. Once the strength predictions have been validated, some optimum lay-up is then defined by the designer. These optimum laminates are then tested for strength, again, typically at all failure modes just like the lamina specimens. Once these data are at hand, then the details of the loading cases, operating environment, and damage tolerance are considered for the already defined and unchangeable lay-up. The lay-up is unchangeable at this point because all laminate data are for a certain lay-up, and any change in lay-up would require the entire allowables program to be repeated (a cost-prohibitive option). So even if a more damage tolerant lay-up is identified, it would not be used at this point due to the time and expense invested in the “pristine allowables” generation. In the alternate method of generating allowables, only load cases (failure modes) that are known to be critical are tested for. And then only laminate data with whatever level of maximum damage is allowed in the structure (usually barely visible impact damage) is tested. It is common practice to set the “damage” as a 14-in. hole and use strength values based on that. Note that both methodologies end with the same basic information (data used for final production design), but the alternative methodology has laminates that are optimized with the damage-tolerance requirements considered rather than whatever pristine lay-up of test coupon gives the highest strength value. Also, with regard to allowables, it is interesting to note the fidelity of the data. Figure 2.78 is a schematic of the increasing fidelity of strength data with increasing number of specimens tested. One of the hallmarks of A and B basis allowables is that more specimens give higher fidelity (and thus a slightly higher allowable). This is typically interpreted as “more specimens means higher reliability for the final structure.” However, note that the A and B basis coupon strength data are eventually reduced to far below their values by arbitrary amounts such as factor of safety, which has no mathematical basis. At this point, assuming that coupon strength data actually has meaning for the final structure, the high fidelity gained by testing many more specimens is almost miniscule by the time the average strength value is reduced due to these arbitrary “knockdowns.” This

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Fig. 2.78

Schematic of fidelity of strength data.

raises the question of using statistically significant A or B basis values over something much simpler and more cost-effective, such as average strength. Testing only damaged laminates also has other advantages besides the obvious one of not spending time and money generating data that will never be used. One of these advantages is obtaining strength data that is directly applicable without having to rely on questionable theoretical concepts such as interactive failure criteria and/or first ply failure. The fallacy of first ply failure will be covered in detail in Sec. 2.11.6. Using Hart-Smith’s 10% rule [8] results in a good enough estimate of pristine strength, particularly given what happens to the pristine strength data as shown in Fig. 2.78. This is covered in more detail in Sec. 2.11.7.

2.11.5

FATIGUE IN COMPOSITE STRUCTURES

This section touches on the heavy emphasis on fatigue testing and terminology that has found its way into composite materials documents, undoubtedly as a result of the importance of fatigue in metallic structure. This terminology still persists today in Federal Aviation Administration (FAA) documents and CMH-17, even though composites are far superior in fatigue resistance than metals and the only practical implications of studying fatigue in composites are in rare applications that experience millions of load cycles near limit load (such as helicopter rotors). Generally speaking, LVs would not experience such fatigue loading as mentioned in Sec. 2.11.3. The gradual reduction in strength of a metal (due to fatigue) vs the sudden decrease in strength of a composite (due to a discrete source event) is shown

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171

schematically in Fig. 2.79 [6]. For the metallic material, it is possible to catch and repair the growth of the crack during an inspection interval and prevent the metal from not being able to carry UL. The composite, on the other hand, has an immediate drop in load-carrying capability after the impact event, and there is no time to catch and repair this damage before the next scheduled inspection interval. All of this is irrelevant to many aerospace structures because the inspection interval is set at every flight, something unheard of in the aircraft industry. Throughout CMH-17 and FAA documentation, there is much made of damage growth and how this growth affects certification. The next figures provide data of damage growth and how it manifests itself in carbon composite laminates. It was noted decades ago that if any damage growth could be measured, this growth was sudden and rapid, not gradual as a crack grows in metals [9]. In analogy with metallic structure, any growth that occurs is supposed to be monitored during inspection intervals, and when the growth (crack length in metals) reaches some “critical” stage, the damage is repaired. Note that if any growth is sudden and rapid, it can easily occur between inspection intervals and cause catastrophic failure. Relying on detecting delamination growth before failure is an obvious safety issue. Thus damage growth is a poor measurement parameter in composites as applied to safety and certification. Figure 2.80 shows an example of damage growth that is simply an artifact of testing small coupons and not “true” damage growth [10]. Despite the obviousness that the initial damage area does not grow, these data are reported as damage growth. Unlike metallic structures, fatigue cycled composites actually get stronger after a fatigue spectrum. This is due to the large number of crack paths that can be formed and then

Fig. 2.79 Chart showing decreased strength as a function of damage for metals (fatigue damage) and composites (impact damage).

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Fig. 2.80 Example of reported damage growth that is not. (From Han et al. [10]. Text in italics is the author’s and not part of the original figure.)

immediately arrested, allowing for stress relaxation around any points within the material that experience stress concentrations that would normally lead to premature static failure.

2.11.6

MECHANICAL TESTING OF COMPOSITES

This section is presented to give the reader insight into some of the peculiarities of mechanical tests commonly performed to generate composite material allowables. By understanding these tests, and their limitations, it will be revealed that even if a pristine structure could be designed to, the allowables are dubious and do not represent what they are supposed to be, namely, material variability. A no costsolution as to what to use for allowables (as first proposed by Hart-Smith [11]) is presented at the end of this section. For tension and compression testing, the advantages of testing a notched specimen (specimen with a hole) are presented in Fig. 2.81. Many carbon fiber composite structures are designed in such a way that a 14-in. hole is assumed to exist anywhere in the structure. This accounts for a vast array of damage-tolerance concerns and represents conservative strength values, although if one is designing to reduce weight, these values may result in an overly heavy structure. Because laminates tend to be weaker in compression than tension, the design value of interest is often open-hole compression (OHC) strength. It is interesting to note that for IM7 carbon fiber laminates, the OHC strength of a quasi-isotropic lay-up is about 48 ksi and because the modulus is about 8.4 megapounds per square inch (MSI), the strain to failure is about 0.57%. Applying a 1.2 safety factor lowers this to 0.47%, and using the often-assumed 85% knockdown for B-basis ANscatter gives a final design strain of 0.40, which is what is commonly designed to for carbon fiber laminated structures (as mentioned previously in Sec. 2.11.4).

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173

As an example of how test technique can affect the measured tensile strength of a laminate and thus nullify the whole notion of allowables, Fig. 2.82 shows tensile strength data generated by the author. All values are the tensile strength averages from the same fiber/resin panel. The only difference is that the 734 MPa value average coupons were cut with the panel rotated -458 which now places the load bearing (08) fibers on the outer surface (rather than at the center). Note that doing this gives a significantly lower measured tensile strength value (734 vs 814 MPa), despite the panel not changing. Also noted in this figure is how testing laminates with a hole, which is what is typically designed to anyway, removes the problem of lay-up orientation giving artificially low values. If the alternative method of generating allowables as outlined in Sec. 2.11.4 is used, many of the problems with generating allowables will be eliminated. The measured scatter in strength would decrease, and the erroneous attribution of high scatter to material variability and not test technique would be avoided.

Fig. 2.81 Schematic explaining why testing notched specimens reduces variability in strength measurements.

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Fig. 2.82

Tensile strength of quasi-isotropic laminates.

A mechanical coupon level test worth noting is the transverse lamina tensile strength. Transverse (908) tensile strength is a test with no fibers carrying any load because the load is applied perpendicular to the fibers. The meaning of this test and some potential pitfalls of performing this test will be discussed. The strength values determined by this test are extremely low, as might be expected, because the fibers are useless (or even detrimental) to the strength of the bulk polymer matrix material in this configuration. Figure 2.83 shows the typical tensile strength values determined by performing this test. The values are about 7–8 ksi (or about 5000 microstrain). This is about one-third the tensile failure strain of most practical laminates in use. The specimens will

Fig. 2.83

Typical transverse (9088 ) tensile strength of composites.

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.84

175

Transverse (9088) tensile strength in realistic laminate.

show a clean straight break parallel to the fibers because any fracture path would not cross through fibers because the fibers are so strong. Figure 2.84 points out that when 908 plies are embedded between load bearing plies, nothing happens in tension until near failure (120 ksi), but according to the 908 tensile test, the 908 plies should have “failed” at about 7–8 ksi. “Failure” (as in first-ply-failure) is assumed to mean matrix cracking, which can develop in cross-ply laminates if enough 908 plies are clumped together. Figure 2.85 is presented to show that not only has it been qualitatively known that first-ply-failure will be lower as more 908 plies are clumped together, but also quantitative data has been in the literature since at least 1982 [12]. As pointed out in this figure, because real engineering laminates do not clump plies together, the “real” value of first-ply-failure is closer to 15 ksi rather than the oft-quoted 7–8 ksi. Figure 2.86 notes the criticality of the gage length when compression testing laminates. Long gage lengths will cause global buckling, and very short, “stubby” gage lengths can cause artificially high values. Thus something in-between is needed, but what defines this gage length value that gives a “true” compressive strength value? Testing notched laminates will give you more useful information and reduce the influences of test parameters. Figure 2.87 notes that if a hole were present in the coupons in the previous figure, then the compression strength would all tend toward the same value regardless of gage length because the failure is localized to the hole. In the figure, as long as the specimen width/hole diameter is greater than 6, similar values are obtained because failure is forced at the holes. By testing specimens with holes, the scatter goes down, which helps prove that much of the measured scatter in strength is not from the material but from the

176

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Fig. 2.85 Relationship between transverse ply “failure stress” and thickness. (From Flaggs and Kural [12]. Text is the author’s, not a part of the original figure.) specimen preparation. The reduced scatter of specimens with holes can deliver B-basis allowables near or above those for specimens without a hole. The sketch in Fig. 2.88 is from a John Hart-Smith publication [11] in which he made this same point. Because tabs must be used for pristine specimens, premature

Fig. 2.86 Geometry dependence of compression test strength values.

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.87

177

Open-hole specimens would eliminate geometry effects.

breaks due to poor tab bonding can occur and give values lower than specimens with holes because these specimens do not need tabs, thus eliminating these premature breaks due to poor tab preparation. In closing, the variability obtained by the type of tests presented in this section is more a measure of specimen preparation and test technique than of material variability [11], despite all scatter being attributed to material variability. Simple proof that the material is not responsible for all scatter is the lower scatter that is always obtained for open-hole testing. The material does not magically become more uniform because a hole is drilled in a specimen, yet that is what is assumed in large databases of allowables in which the high scatter from testing smooth specimens is attributed entirely to the prepreg material.

Fig. 2.88

Hart-Smith’s schematic on coupons with holes having less scatter [11].

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A simple solution to obtaining pristine lamina data in which no tests need be performed was first proposed by Hart-Smith [11]. He suggests using vendorsupplied data. The most reliable (and highest strength) data usually comes from the material manufacturers. All end users now use the same database for a given material and then appropriately attribute their lower measured strength values and higher scatter to errors in processing the test specimens and the tests themselves. In the author’s experience, this is usually the case because the first thing done after performing a series of tests is to compare the results to others’ and reject the results if they fall on the low side.

2.11.7

STRENGTH OF COMPOSITE LAMINATES

As has been the theme throughout this section, pristine strength allowables of multidirectional laminates are of little practical use, but should they be needed, this section shows a no-cost way to obtain these values for tension and compression. The methodology to do this has been independently arrived at by two of the most respected composites engineers in the field [8, 13]. Before entering into discussions of strength of multidirectional laminates, it is important to know the terminology about stacking sequence used in this section. Usually, the stacking sequence is written with the numbers in the brackets indicating angles in degrees from some reference point. For laminates that only contain plies with angles of 08, þ458, 2458, and 908 (p/4 angles) with no other angles present, the stacking sequence can be notated by the percentage of each of these four angles in the laminate. The percentage of 08 plies is noted first, followed by the percentage of þ458 and 2458 plies combined. Finally, the third number indicates the percentage of 908 plies in the laminate. A laminate with more 08 plies will have a higher first number at the expense of lowering one (or both) of the other two. For “quasi-isotropic type” laminates (i.e., laminate only contains some combination of 08, 458, 2458, and 908 fibers), the nomenclature is sometimes written in the form [% 08 plies=% 458 plies þ % 458 plies=%908 plies] Thus, a [þ458/908/2458/08]S laminate can be written as [25%/50%/25%], and a [þ45/90/245/02]S laminate can be written as [40/40/20]. Figure 2.89 presents the tensile strength and failure strain of multidirectional laminates of various IM7 fiber/polymer matrix systems. The lay-ups are noted with the percentage type designations just discussed and are noted in the brackets above each value bar. As expected, a vast range of stress values is realized because the laminates have a differing percentage of 08 plies, which are responsible for most of the strength contribution. When plotted as strain, the values all tend toward 1.5% failure strain despite some of the resins being of different classes [epoxy and bismaleimide (BMI)].

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Fig. 2.89 Tensile data for strength of multidirectional laminates. Figure 2.90 contains an example of a stress–strain curve of a multidirectional laminate generated by the author. It shows the linear behavior of the laminates that is routinely seen despite first-ply-failure theory predicting a “knee” at the lower end of the curve. This linear behavior will help to simplify strength predictions. In this figure the typically used design strain of 0.4% is highlighted to show that for any practical applications, the stress strain behavior is linear even if some nonlinearity appears near the failure stress. Note that if a linear stress–strain relationship holds, then the tensile strength of a laminate made of IM7 Fig. 2.90 Typical stress-strain curve for a multidirectional laminate.

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carbon fiber is simply the modulus multiplied by 0.015 failure strain. For AS4 carbon fiber laminates, this value would be 0.012. Table 2.32 shows how much error is present if all IM7 carbon fiber laminates are assumed to fail at 1.5% strain. With one exception the data are within 10%; arguments have been made about the difficulty in testing pristine laminates, and so this variation between measured values is not surprising. Thus, for any lay-up, the IM7 fiber composite with any polymer matrix can be expected to fail at about 1.5% strain with little error by this assumption. Although the p/4 type of laminates are by far the most commonly used, sometimes the strength of a laminate with plies at some angle other than 08, 458, or 908 may be desired. Classical lamination theory works very well to calculate the modulus of any given lay-up if the unidirectional elastic constants are known. Vendor data are typically not difficult to obtain; thus, the modulus of any lay-up can easily be calculated with a basic computer code that can do classical lamination theory. Because in compression the resin helps support the fibers and prevent them from micro-buckling, the resin does play a larger role in determining strength TABLE 2.32

Fiber/Resin

PERCENT ERROR IF ALL IM7 CARBON FIBER LAMINATES ARE ASSUMED TO FAIL AT 1.5 PERCENT STRAIN Source

Lay-Up

Estimated Ef

Actual Ef

Error

IM7/5250-4

(1)

[25/50/25]

1.5

1.61

28.1%

IM7/977-3

(1)

[25/50/25]

1.5

1.52

22.0%

IM7/8552

(4)

[25/50/25]

1.5

1.50

20.6%

IM7/8552

(2)

[25/50/25]

1.5

1.51

21.3%

IM7/5250-4

(1)

[33/58/8]

1.5

1.52

22.0%

IM7/977-3

(1)

[33/58/8]

1.5

1.46

þ2.0%

IM7/5250-4

(1)

[42/16/42]

1.5

1.58

26.0%

IM7/977-3

(1)

[42/16/42]

1.5

1.54

23.4%

IM7/5250-4

(1)

[50/43/7]

1.5

1.48

21.0%

IM7/977-3

(1)

[50/43/7]

1.5

1.45

þ6.0%

IM7/5250-4

(1)

[58/34/8]

1.5

1.45

þ2.7%

IM7/977-3

(1)

[58/34/8]

1.5

1.31

þ12.1%

IM7/MTM-45

(3)

[50/0/50]

1.5

1.54

23.4%

(1) Anon, “Lockheed Aeronautical Systems Company; F-22 Structural Development Tests Coupon Tests Results and Evaluation Volume 5, Graphite Epoxy”, December 1994. (2) Reference [15] in “References”. (3) Ridgard, C. “Complex Structures for Manned/Unmanned Aerial Vehicles,” AFRL-RX-WP-TM-2008-4054, January 2008. (4) Wisnom, M.R., Khan, B. and Hallett, S.R., “Size Effects in Unnotched Tensile Strength of Unidirectional and Quasi-isotropic Carbon/Epoxy Composites,” Composite Structures, Vol. 84, June 2008, pp. 21–28.

AEROSPACE MATERIALS CHARACTERISTICS

Fig. 2.91

181

Compression data for strength of multidirectional laminates [14].

in compression than strength in tension. Also, compression testing of laminates is more difficult due to specimen buckling and obtaining a uniform application of load, especially as the laminate becomes stronger and stiffer (more 08 plies). Consider Fig. 2.91, which shows compression strength data from IM7 carbon fiber laminates with various percentages of 08 plies [14]. Note that when plotted as failure strain, the laminates all tend to the same value of 1.3%, with those with higher percentage of 08 plies giving slightly lower values due to the test limitations of laminates with more 08 plies, as mentioned previously. The measured values in Fig. 2.91 are close to those predicted if the vendor data [15] for compression of a uniaxial laminate is used along with classical laminated plate theory and a cutoff failure strain of 1.3%. Considering these values will be greatly reduced, as mentioned many times in this section, these values can be used as preliminary design estimates if need be. The B-basis strength values can be estimated by taking 85% of the average strength values, as mentioned in Sec. 2.11.6. This is a common practice for carbon fiber composites. To obtain an A-basis strength value, typically 70% of the average strength value is used.

2.11.8

SUMMARY

This section discusses the behavior of composites as structural materials and compares them with those of metallic materials such as aluminum and Al-Li alloys. It makes the point that composites and metals are different and the damage and failure mechanics are entirely different. Very often, aircraft-type requirements are levied upon launch vehicles when composites are used, despite the two vehicles

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having vastly different purposes and lifetimes. The aircraft industry has far more stringent requirements for the many reasons mentioned, especially the fact that aircraft must fly with damage. The argument is made that a costly building-block approach to generating material allowables using pristine lamina and laminate-strength properties is not necessary. Strength data must be developed on damaged structures, because this is what the vehicle is designed to. Fatigue testing important for metals is a nonissue for most applications of carbon fiber composites. Further, the section sheds light on what “allowables” values are, the tests that generate them, and why they are far from the high fidelity they claim. It presents a simple method to determine the tensile and compressive strength of any carbon fiber laminate with no testing needed.

REFERENCES [1]

Anon, “Future Materials: Aluminum-Lithium, Standard Metals or Composites,” Leeham News and Comment, https://leehamnews.com/2012/07/10/futurematerials-aluminum-lithium-standard-metals-or-composities/ [retrieved online 1 Nov. 2013]. [2] Hicks, J. P., “Business Technology: New Materials Altering the Aircraft Industry,” New York Times, 20 Dec. 1989. [3] Anon, “The Impact of Composites on the Aerospace and Defense Industry,” Aerospace and Defense Advisory, Summer 2008. [4] Hales, S. J., and Hafley, R. A., Structure-Property Correlations in Al-Li Alloy Integrally Stiffened Extrusions. NASA TP-2001-210839, 2001. [5] Composites Material Handbook 17G, Vol. 3, Chap. 12, SAE International, Warrendale, PA, 2013 Pages 12.1–12.182 [6] NASA Marshall Space Flight Center, “Fracture Control Requirements for Composite and Bonded Vehicle and Payload Structures,” MSFC-RQMT-3479, Marshall Space Flight Center (MSFC), Huntsville, AL, 2006. [7] Nettles, A. T., Hodge, A. J., and Jackson, J. R., “An Examination of the Compressive Cyclic Loading Aspects of Damage Tolerance for Polymer Matrix Launch Vehicle Hardware,” Journal of Composite Materials, Vol. 45, No. 4, pp. 437–458. [8] Hart-Smith, L. J., “The Ten-Percent Rule for Preliminary Sizing of Fibrous Composite Structures,” Weight Engineering, Vol. 52, No. 2, pp. 29–45. [9] Horton, R. E., Whitehead, R. S., Kan, H.P., and Graves, M. J., Damage Tolerance of Composites, Vol. I, AFWAL-TR-87-3030. [10] Han, H. T., Mitrovic, M., and Turkgene, O., The Effect of Loading Parameters on Fatigue of Composite Laminates: Part III, DOT/FAA/AR–99/22, 1999. [11] Hart-Smith, L. J., “An Account of One Engineer’s Long Term Involvement with Aerospace Applications of Composite Structures, Part I: Analytical Developments,” Boeing paper PWDM05-0089. [12] Flaggs, D. L., and Kural, M. H., “Experimental Determination of the in situ Transverse Lamina Strength in Graphite/Epoxy Laminates,” Journal of Composite Materials, Vol. 16, pp. 103–116.

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[13] Tsai, S. W., and Melo, J. D., Composite Materials Design and Testing: Unlocking the Mystery with Invariants, Composites Design Group, San Diego, CA, 2015. [14] Nettles, A. T., “Notched Compression Strength of 18-Ply Laminates with Various Percentages of 08 Plies,” Journal of Composite Materials, Vol. 49, No. 4, pp. 495–505. [15] Marlett, K. “Hexcel 8552 IM7 Unidirectional Prepreg 190 gsm and 35% RC Qualification Material Property Data Report,” National Institute for Aviation Research Report, CAM-RP-2009-015, 2011.

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2.12 AEROSPACE CERAMIC MATERIALS Dongming Zhu NASA John H. Glenn Research Center at Lewis Field

2.12.1

INTRODUCTION AND SUMMARY

Ceramics are important materials for aerospace applications because of their high temperature capability (high melting point), high stiffness and strengths, and excellent resistance to oxidation and corrosion. Ceramic materials also generally have lower densities and specific strengths as compared to metallic materials. Currently used engineering or structural ceramics (aka crystalline inorganic nonmetallic materials) in aerospace include ceramic thermal and environmental barrier coatings for protecting hot-section components of aircraft turbine engines from high heat flux in high-temperature combustion environments, rocket exhaust nozzles, and thermal protection systems for space vehicles. Ceramic matrix composites (CMCs), including nonoxide and oxide CMCs, are also being incorporated in turbine engines in high-pressure and high-temperature section components and turbine exhaust nozzles with long-duration design operating lifetimes. Although ceramic materials have many attributes that make them excellent materials for high-temperature and ultra-high-temperature protective coatings and structural materials, the current uses have been limited due to their low toughness, large variability in mechanical properties, and complex environmental effects in harsh operating conditions. The complexity and variability of aerospace ceramic processing methods, compositions, and microstructures also make the material designs and validations a more challenging task. This section presents an overview of aerospace ceramic materials and their characteristics. The focus is on important enabling ceramic systems for aerospace applications, particularly turbine engine thermal and environmental barrier coating systems: nonoxide-type silicon carbide (SiC)/SiC CMCs. This section also covers the ceramic materials for various applications; material system properties and durability performance associated with processing; and thermal, thermomechanical, and environmental life design considerations. A brief discussion of laboratory and simulated operating environment tests is included. This section is divided into the following subsections: thermal barrier coatings, environmental barrier coatings, and particularly SiC/SiC ceramic matrix composites. Monolithic ceramics have limited fracture toughness, they are used as constituents of the ceramic surface coatings, ceramic matrices, fibers or fiber coatings, and therefore some of their pertinent properties are also described or compared with processed coatings or composites. In each of the subsections, a brief history and material system improvements will be presented. In the CMC section, the ceramic fiber attributes will also be briefly discussed. The current status and future trend of CMC development and applications are described. This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

AEROSPACE MATERIALS CHARACTERISTICS

2.12.2

185

THERMAL BARRIER COATINGS

The performance and efficiency of aero propulsion turbine engines are directly related to the operating temperatures. Ceramic thermal barrier coatings (TBCs) are technologically important because of their ability to increase turbine engine operating temperatures and reduce cooling requirements, thus helping to achieve engine performance and emission goals [1–10]. The advances in ceramic material and processing technologies, particularly for zirconia-based ceramics, have resulted in the applications of ceramic TBCs on aircooled, critical turbine engine hot-section components, such as combustors and high-pressure turbine vanes and blades, as shown in Fig. 2.92. Since the initial entry into commercial service in the 1980s [2, 9], TBCs have achieved significant temperature benefits that are surpassing other materials, including nickel-based single-crystal superalloys, and cooling technology advances achieved in the last three decades [6, 9]. TBCs have provided high-pressure turbine (HPT) component metal temperature reduction up to 1008C, and future potential of greater than 2008C reductions is expected [9], particularly when more advanced low-thermal-conductivity coatings are incorporated. TBCs are complex, two-layer or multilayered, multimaterial systems. A typical TBC system consists of two layers: a ceramic zirconia (ZrO2) coating top coat and a metallic bond coat (either NiCrAlY, NoCoCrAlY, CoNiCrAlY, or PtAl) that are deposited on a nickel-based superalloy substrate. In addition, low diffusion and protective Al2O3 scales, thermally grown oxide (TGO) on the TBC bond coat, are also critical for thermal barrier coating technology. The TGO scales, formed on the bond coats between the alloy bond coat and ceramic thermal barrier

Fig. 2.92 High bypass turbofan engine with examples of thermal-barrier-coated turbine components (inset).

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B. N. BHAT

coatings, the bond coat temperatures, and its cyclic endurance are important design parameters in determining the thermal barrier coating life [11]. TBCs and the material systems are processed and integrated using various processing methods, including commonly used plasma spray and electron beamphysical vapor deposition (EB-PVD). Other thermal spray processing methods, such as suspension plasma spray (SPS) and plasma spray-physical vapor deposition (PS-PVD) are also being developed and used [12]. Turbine engine airfoil components (turbine vanes and blades) typically have thermal barrier coating thicknesses ranging from 100 to 250 mm, whereas combustor or other nonrotating components have a coating thickness ranging from 250 to 500 mm. The coating system compositions, microstructures, and properties can greatly affect the coating durability during engine operations. Microstructures of plasma-sprayed and EB-PVD thermal barrier coatings are shown in Fig. 2.93. It can be seen that microcracks and porosity in the plasma-sprayed splat type of coatings, and intracolumnar and intercolumnar porosity in EB-PVD columnar coatings, are present in the microstructure, which help to increase the thermal strain tolerance and reduce the thermal conductivity of the coating systems. Zirconia (ZrO2) is used for thermal barriers because it has a high melting point (approximately 27008C), low intrinsic thermal conductivity (approximately 2.0– 2.5 W/m-K), and relatively high coefficient of thermal expansion, and thus is an ideal ceramic material for protecting nickel-based superalloy components for high-temperature TBC applications. Oxide alloy dopants such as Y2O3 or rare earth (RE) oxides (e.g., Yb2O3, Gd2O3) are added to stabilize the zirconia and retain the high-temperature phases, particularly favorable metastable tetragonal phase structure, or cubic phase structure, which suppresses the detrimental martensitic tetragonal to monoclinic phase transformation during service [13, 14]. The current state-of-the-art ZrO2-(6-8) wt % Y2O3 TBC compositions have metastable tetragonal (t0 ) phase structures, which possess high toughness and excellent cyclic durability [9, 15]. NASA’s early coating TBC development showed the best

Ceramic coating Ceramic coating

Bond coat

100 μm

a) Plasma-sprayed coating

Fig. 2.93

Bond coat

25 μ m

b) EB-PVD coating

Microstructures of TBCs on metallic substrates.

AEROSPACE MATERIALS CHARACTERISTICS

187

Fig. 2.94 Best furnace cyclic life and durability for TBCs with a composition of ZrO2-6-8 wt % Y2O3 [15]. furnace cyclic life and durability for the TBC coating systems with a composition range of ZrO2-(6–8) wt % Y2O3, as illustrated in Fig. 2.94. More recent work by Mercer et al. [16] showed that the metastable tetragonal coating composition 7YSZ (ZrO2-7 wt % Y2O3) has a higher fracture toughness value (3.0 MPa m0.5) compared to another commercially available higher Y2O3 content TBC composition ZrO2-20 wt % Y2O3 (1.0–1.2 MPa m0.5). Although t0 phase coatings have the advantages of higher toughness and generally more durability for rotating components where erosion and impact resistance can be of major concerns, the coating materials intrinsically are metastable, and therefore their use temperature is limited to 1200–12508C for long-term operation in turbine engine environments. The t0 coatings also have higher thermal conductivity and fast sintering (ceramic coating densification) that can reduce the coating initial porosity, resulting in a significant thermal conductivity increase and reduced cyclic durability [17, 18]. To further increase the turbine engine efficiency and operating temperatures, higher temperature and lower thermal conductivity coating systems have been in development in last two decades [19–23]. Among the advanced low-thermal-conductivity thermal barrier coatings are the multicomponent defect-clustering coatings [23]. The advanced oxide coatings were designed by incorporating multicomponent, paired-cluster dopants in conventional zirconia– yttria oxides. The dopant oxides were selected based on the cation–anion interatomic and chemical potentials, lattice elastic strain energy, polarization, and the electroneutrality of the oxides. Because defect clusters can attenuate and scatter lattice phonon waves as well as radiative photon waves at a wide spectrum of frequencies, the coatings have significant reductions in the oxide intrinsic lattice and radiation thermal conductivity. The creation of the thermodynamically stable and highly distorted lattice structures, with essentially immobile defect clusters and/or

188

b)

2.0

Plasma-sprayed coatings

1.8 1.6

Advanced ZrO2based coatings

1.4 1.2 1.0 0.8 0.6 t ′ phase region

0.4 2

4

Cubic phase region

Thermal conductivity, W/m-K

1.8

7YSZ k

1.6

7YSZ k

k20 -YSZ-Nd-Yb

20

k20 -YSZ-Gd-Yb 5

ZrO2-Y2O3 coatings

k20 -YSZ-Sm-Yb k20 -ZrO2-Y2O3

0

k5 -YSZ-Gd-Yb

1.4

k5 -YSZ-Gd-Yb-Sc k5 -YSZ-Nd-Yb-Sc

1.2

k0 -7YSZ k5 -7YSZ

1.0 0.8 0.6 t ′ phase region Cubic phase region

0.4

6 8 10 12 Total dopant concentration, mol%

c)

2.0 7YSZ k

8YSZ k0 (2500F) 8YSZ k20 (2500F) 8YSZ k0 (2600F) 8YSZ k20 (2600F) k0 Refractron k0 (1371C) k20Refractron k20 (1371C) k0 Refractron k0 (1482C) k20Refractron k20 (1482C) (1371C) k0 Praxair k0 k20Praxair k20 (1371C) (1371C) k0 NASA k0 (1371C) k20

Thermal conductivity, W/m-K

Thermal conductivity, W/m-K

a)

B. N. BHAT

0

14

5

10 15 20 25 Total dopant concentration, mol%

30

2.2 EB-PVD

7YSZ

2.0

7YSZ

1.8 1.6

7YSZ +Yb +Er +Gr +Nd

YSZ 20YSZ

YSZ

1.4

Y+Hf

1.2

37Hf Nd

1.0

Gd

0.8 0.6

31Hf +(Y+Yb)

Nd Plasma Assisted (PA) PVD 7YSZ

0

Gd2Zr2O7

10 20 30 40 50 Total dopant concentration, at% cation

60

Fig. 2.95 Thermal conductivity of various thermal barrier coatings. (a) Plasma-sprayed TBCs, thermal conductivity tested at 137188 C and 148288 C [24]. (b) EB-PVD TBCs, thermal conductivity tested at 131688C [9, 23]. (c) Thermal conductivity comparisons for advanced composition TBCs [9]. nanoscale ordered phases, effectively reduces the mobile defect concentration and suppresses the atomic mobility and mass transport, thus significantly improving the oxide sintering-creep resistance and mechanical properties. Thermal conductivity of various multicomponent defect cluster thermal barrier coatings, along with other advanced low-thermal-conductivity thermal barrier coatings, are summarized in Fig. 2.95a, b, and c [23, 24, 9]. These figures show thermal conductivity of various multicomponent defect cluster and other types of advanced low thermal conductivity thermal barrier coatings, as function of dopant concentrations, and compared with the t0 phase ZrO2-(6-8) wt % Y2O3. The sintering-induced thermal conductivity increases are also shown as the coating thermal conductivity values for as-processed (k0), after 5 and 20 hours at temperatures (k5 and k20). A low thermal conductivity defect cluster coating, ZrO2-9.5wt%Y2O3-5.6Yb2O3-5.2 Gd2O3, rare earth zirconate (Gd2Zr2O7 and Sm2Zr2O7), and a few other compositions and their specifications may be found in the literature [12, 25–27]. Future TBC systems will be more aggressively designed for the thermal protection of engine hot-section components, thus allowing significant increases in

AEROSPACE MATERIALS CHARACTERISTICS

189

engine operating temperatures, fuel efficiency, and engine reliability. However, the coating reliability and durability under high-temperature, high-thermal gradient cyclic conditions still remain as major challenges [18, 28, 29]. Particulate erosion, impact, and engine-ingested low-melting calcium magnesium alumino-silicate sand dusts or volcanic ashes during service have further complicated the life designs of turbine airfoil TBCs [30–34]. Figure 2.96a and b shows cyclic lives of turbine airfoil thermal barrier coatings under various temperature conditions and degradation mechanisms. The coating life is exponentially reduced with increasing temperature, because interface damage effects are significantly increased from the accelerated oxide scale growth and increased cyclic stress-temperature amplitudes during the cycling, as shown as Arrhenius behavior of furnace cyclic lives. The erosion – bond coat oxidation based failure map for 7YSZ and defect cluster low conductivity thermal barrier coatings, also showing the Mach 0.3 particulate erosion is a more dominated failure mode at lower temperatures while the surface heat flux can reduce thermal barrier coating cyclic life and durability. The erosion-based coating failure life can be increased with temperature because the t0 phased coating toughness and plasticity increase with temperature [32]. The coating failure modes are usually complex, largely depending on envisioned engine operating conditions, processed coating composition, architecture, and microstructures. The increases in engine temperature, pressure, and heat flux can raise durability issues for current coating systems. The development of next-generation advanced TBCs will greatly rely on better understanding of the coating behavior and failure modes under the high-temperature, high-thermal gradient cyclic conditions.

2.12.3

ENVIRONMENTAL BARRIER COATINGS

The ever-increasing demands for developing more efficient, low-emission and high-performance aircraft and space vehicle propulsion engines have required new hot-section component materials that are significantly lighter and have higher temperature capabilities. Current nickel-based superalloys are reaching the upper limit of their temperature capabilities, and therefore SiC fiberreinforced SiC/SiC ceramic matrix composites (CMCs) have been envisioned as alternative next-generation turbine engine hot-section materials [7, 8, 35, 36]. Silicon-based ceramics and composites, such as SiC/SiC ceramic matrix composites, are desirable because they have low-density, high-temperature creep strength and oxidation resistance in dry oxidizing environments. Environmental barrier coatings (EBCs) are required to prevent the SiC/SiC CMCs from water vapor attack in the engine combustion environments, due to volatilization of the protective silica (SiO2) scales on SiC when reacting with water vapor during engine operation [37]. The loss of SiO2 from the ceramic surfaces leads to the accelerated strength degradation under combined thermal and mechanical loading conditions in the engine during operation. Therefore, environmental

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B. N. BHAT

Fig. 2.96 Thermal barrier coating cyclic life and durability. (a) Arrhenius behavior of furnace cyclic lives of EB-PVD TBCs [29]. (b) The erosion–bond coat oxidation based failure map for 7YSZ and low conductivity TBCs [32].

AEROSPACE MATERIALS CHARACTERISTICS

191

barrier coatings are considered essential in enabling the CMC component technologies for next-generation aerospace propulsion engine systems. The SiO2 surface recession occurs when the SiO2 scale reacts with water vapor in engine combustion environments, forming gaseous Si(OH)4 and thus resulting in the volatilization of the SiO2 scales of SiC/SiC CMC materials [37]. The volatilization and reaction with the water vapor occurs according to the following equation: SiO2 þ 2 H2 O (gas) ¼ Si(OH)4 (gas)

(12:1)

The surface recession mechanism is dependent on the gas velocity and can be dramatically accelerated as the velocity increases [37, 38]. For turbine engine conditions, the velocity factor is related to the gas Mach number and/or heat transfer coefficient, and film-cooling can also be integrated, studied, and modeled [39]. Figure 2.97 depicts a schematic diagram of the SiC/SiC CMC recession due to SiO2 volatility in convective and convective plus film cooling conditions [39]. In a film cooling case, more complex analysis including Computational Fluid Dynamics (CFD) analysis is needed to understand the local gas flow velocity, pressure and water vapor fractions [39]. For the SiC forming SiO2 scale with the unit silica activity, the surface recession rate Krecession, due to the silica volatilization by the reaction with water vapor, has the velocity and pressure dependence according to the following equation [37]: Krecession ¼ C V 1=2 P(H2 O)2 =ðPtotal Þ1=2

(12:2)

where C ¼ constant V ¼ gas velocity P(H2O) ¼ total partial pressure of water vapor Ptotal ¼ total combustion chamber pressure

(H2O)

a)

b)

Fig. 2.97 Surface recession of SiO2 scales on a SiC/SiC ceramic matrix composite specimen [39]. a) Recession in a convective combustion gas flow. b) Recession in a convective combustion gas and film-cooling air flow.

192

B. N. BHAT

The early generation EBCs consist of Si bond coat, mullite-based intermediate coat, and barium-strontium-aluminosilicate (BSAS, Ba1-xSrxAl2Si2O8; 0 , x , 1) top coat, developed in the NASA Enabling Propulsion Materials (EPM) Program [12, 40–42]. The EBC material systems have shown a good compatibility with SiC/SiC CMC systems, and the coating feasibility and durability have also been tested in various land-based turbine validation tests [43–45]. The temperature limits and stability of the first-generation environmental barrier coating systems have also been well studied. The use temperature is generally limited to 13008C in contact with the silicon bond coats [46–48]. As an example, for BSAS, Ba1-xSrx Al2Si2O8; 0 , x , 1 environmental barrier coating case, the surface recession rate in micrometer per hour is determined by the following equation [7, 45]: Krecession (BSAS) ¼ 105654=½T

8 Cþ273

 x P1:5 x V 0:5 x a

SiO2

(12:3)

where aSiO2 ¼ the SiO2 activity in the EBC. The lower activity of BSAS resulted in a reduced recession rate of the EBC-coated SiC/SiC CMC systems. The next-generation engine systems currently envisioned with higher component operating temperatures demand more advanced environmental barrier coating systems. The second-generation EBCs have temperature capability up to 14828C, using rare earth metal or transition metal disilicate and monosilicate compositions with low silica activity (e.g., Yb2Si2O7, Gd2Si2O7, Er2Si2O7, Y2SiO5, Yb2SiO5, Gd2SiO5) for protecting SiC/SiC ceramic matrix composites [49]. Third-generation coatings include advanced thermal and environmental barrier coating systems with surface temperature capability up to 16508C [7, 50]. Some recession rates for selected environmental barrier coatings including BSAS are shown in Fig. 2.98 and compared with the SiC/SiC CMCs and Si3N4

Fig. 2.98 Environmental barrier coating recession determined under high gas velocity and high-pressure conditions [39].

AEROSPACE MATERIALS CHARACTERISTICS

193

Fig. 2.99 SiC/SiC recession rates at 130088C under various gas velocity and pressure conditions [39]. monolithic ceramics. The recession rates are determined using NASA’s HighPressure Burner Rig in conditions of up to 16 atm and 200 m/s combustion gas velocity, at various temperatures. The SiO2 recession rates on SiC/SiC CMC have also been determined at 13008C, as shown in Fig. 2.99, with the approximate gas velocity dependence of 0.46–0.6 and total pressure dependence of 1.97 (at water vapor partial pressure approximately 9%) [39]. Multicomponent rare earth–silicates, hafnium-aluminate silicates, and hafnium rare earth silicate environmental barrier coating systems, along with HfO2-silicon, and rare earth–silicon based bond coat systems, have also recently been used for 15008C capable environmental barrier coating systems to significantly improve temperature capability and calcium magnesium alumino-silicate (CMAS) resistance [42, 51–54]. Table 2.33 illustrates the advanced environmental barrier coating materials and multilayer structures that have evolved for improving temperature capability, environmental stability, and toughness for turbine engine applications. The EBCs have evolved from BSAS to rare earth silicates to rare earth–hafnium–silicate multicomponent EBCs, while the EBC bond coats have also advanced from silicon to hafnia (HfO2)–silicon, rare earth–silicon to rare earth–hafnium–silicon systems with controlled oxygen and silicon activities for ceramic matrix composite turbine airfoil applications. The rare earth silicate–mullite or alumina coating systems have also been recently considered beneficial because of significantly lower oxygen permeability for mullite and Al2O3 as compared with that for the ytterbium disilicate EBC [55, 56].

EVOLUTION OF NASA EBC TECHNOLOGY FOR SiC/SiC ceramic matrix composites: EBC system developments

EBC Configurations

Gen I 1995–2000 R&D Award

Engine Components:

Combustor

Top Coat:

BSAS (APS)

Gen V 2007–2012

194

TABLE 2.33

Gen VI 2009–present Patent S/N: 13/923,450 PCT/US13/ 46946

Gen III 2000–2005 US 7,740,960 B1

Gen IV 2005–2011 R&D Award (2007) Turbine Airfoil EBC Development US 7,740,960 B1

Combustor/ (Vane)

Combustor/Vane (Hybrid Plasma Spray EB-PVD processing)

Turbine Vane/Turbine Blade

– Vane/Blade EBCs – Equivalent APS combustor EBCs

Airfoil components

RE2Si2O7 or RE2SiO5 (APS)

– (Hf,Yb,Gd,Y)2O3 – ZrO2/HfO2-RE silicates – ZrO2/HfO2 þ BSAS (APS and EB-PVD)

RE-HfO2-Alumino Silicate (APS and EB-PVD)

RE-Hf-Silicate; RE-HfO2-graded Silica processing (EB-PVD)

Advanced RE-Hf þ X silicates

Gen II 2000– 2004

–

–

RE-HfO2/ZrO2aluminosilicate layered systems

Nanocomposite graded oxide/silicate

Gen IV interlayer not required (Optional)

EBC:

Mullite þ BSAS

BSAS þ Mullite

RE silicates or RE-Hf mullite

RE doped mullite-HfO2 or RE silicates

Multi-Component RE silicate systems

Multicomponent REsilicate/self grown

Bond Coat:

Si

Si

Oxide þ Si bond coat

HfO2-Si-X, doped mullite/Si SiC nanotube

Optimized Gen IV HfO2-Si-X bond coat 14828C bond coats

RE-Si þ X based systems

Thickness

250–400 mm

250–400 mm

250–500 mm

250 mm

127 mm

25–100 mm

Surface Temperature capability:

Up to 13168C

13168C

16508C with 13168C CMC

14828C with 13168C CMC

Up to 16508C with 13168C,14828C CMC

Up to 16508C with 13168C,14828C CMC

Bond Coat Temperature Capability

Limited to 13508F

Limit to 13508F

Limit to 13508F

14208C þ ; Advancement to 14828C

14828C (2011 Goal)

14828Cþ

Interlayer:

–

B. N. BHAT

AEROSPACE MATERIALS CHARACTERISTICS

195

The design of environmental barrier coatings for aerospace propulsion engines have significantly benefited from early-generation monolithic coating developments, and also from the turbine thermal barrier coating experience. In general, coating material system design considerations should include coating temperature capability, environmental and mechanical stability, chemical compatibility (e.g., no reactions occurring that form low melting phases), and phase stability during the service operation. High toughness and low thermal conductivity are important properties for turbine airfoil environmental barrier coatings, where a thin coating is required for aerodynamic requirements, ensuring durability in high-pressure and high-velocity gas flow, and inparticulates or molten sand impingement conditions during service. As shown in Fig. 2.100, coatings should be designed to be operated in a safe region (shadow area) where lower stiffness coatings should be designed to increase the strain tolerance when higher thermal stresses from a larger thermal expansion mismatch or large thermal gradient is present. With increasing distance from the bond coat toward surface EBC layers, the coating system generally has increased thermal expansion mismatch stresses or thermal gradients (due to lower thermal conductivity of the top coatings). Thus the stiffness should be reduced to ensure higher strain tolerance and to maintain the coating stability and durability. The EBCs have demonstrated the feasibility and significantly improved hightemperature and environmental stability of the EBC-CMC systems in laboratory testing and simulated engine rig conditions [42, 57, 58], and SiC/SiC CMC turbine shrouds have been incorporated in the engine design and applications. Challenges still remain for significantly improved environmental durability and thermomechanical fatigue resistance in turbine engine high-heat flux environments to achieve the prime-reliant environmental barrier coating designs in aero turbine environments.

Fig. 2.100

EBC mechanical stability safe design approach.

196

2.12.4

B. N. BHAT

SiC/SiC CERAMIC MATRIX COMPOSITES

SiC/SiC CMCs reinforced by continuous-length, polycrystalline high-strength SiC fibers are revolutionary hot-section materials for aerospace propulsion engines. Significant progress has been made in the development and recent implementation of SiC/SiC CMC materials for engine hot-section components [7, 36]. The engine performance benefits are attributed to the low-density, higher temperature capability, and resistance to oxidation and corrosion of SiC/SiC composites, thereby allowing designs with higher component temperature and reduced cooling as compared to nickel-based superalloys. The weight reductions realized by applying SiC/SiC CMC to engine rotating components can further reduce the design complexity and weight of engine structures. The temperature capability of the current-state-of-the-art SiC/SiC CMC systems have the temperature capability of 13168C, which is an improvement with a potential temperature benefit over superalloys exceeding 2238C. As shown in Fig. 2.101, the development of enabling CMC and environmental barrier coatings will result in a step increase in the temperature capability of gas turbine components. Generation II CMCs have a temperature capability of 13168C, which is an improvement, whereas future generation CMC materials are envisioned to be capable of 14828C when advanced SiC fibers and matrix materials are used. The SiC/SiC CMCs generally have two processing routes. The first route was developed and demonstrated in the NASA Enabling Propulsion Material (EPM) Program under a CMC combustor program [36, 59]. The method emphasizes chemical vapor infiltration (CVI) by first applying a CVI–boron nitride (BN) fiber interphase coating (0.1–0.5 mm thickness) for the SiC fiber tow preform, followed by a thin CVI–SiC matrix layer over the BN interphase coating. This CVI process is then followed by a SiC fine particulate slurry infiltration at room

Fig. 2.101 Development and implementation of SiC/SiC CMC and advanced ceramic environmental barrier coatings.

AEROSPACE MATERIALS CHARACTERISTICS

197

temperature, and finally ends with a silicon melt infiltration (SMI) at 14008C. This process is often referred to as the CVI-SMI or CVI-MI process and is illustrated in Fig. 2.102a. a) CVI+SMI

b) Prepreg + MI

Fig. 2.102

Two representative SiC/SiC ceramic matrix composite processing routes.

198

B. N. BHAT

The second CMC process route was developed by General Electric for processing SiC/SiC (HiPerCompTM) turbine engine components [60]. This method fabricates prepregged unidirectional 2-D tapes using a polymer-based binder containing SiC, Si, and carbon particulates, along with CVI processing precoated fiber tows. The 2-D tapes can be stacked into 3-D preforms, heated to high temperature (1400–14508C) for casting and melt-infiltration of CMC components by a reactive melt-infiltration process. The HiPerCom SiC/SiC CMCs have residual silicon levels of 5–15 vol % and also generally require thicker fiber coatings (1 mm, typically BN, Si-doped BN, SiC, and Si3N4) to prevent the fibers to be reacted with silicon at the high-temperature reactive process. The detailed HiPerCom CMC thermophysical, elastic, and fracture strength properties are given in note [43]. Figure 2.102b illustrates the prepreg–melt infiltration processes for SiC/SiC ceramic matrix composites. Advances have been made in various materials constituents of CMCs. Ceramic fibers, in particular, play a significant role in high-performance CMCs. Factors affecting SiC fiber creep resistance include grain size; impurities, such as the residual oxygen content (in particular affecting the primary creep stage, grain boundary relaxation); porosity; and surface roughness and defects. The steady-state creep strain rate e˙ (percent strain per hr) of the ceramic fibers can be written as [36, 62] ˙esteady-state ¼ C s n exp(Qs =RT)

(12:4)

where C s n Qs R T

¼ an empirical constant ¼ applied stress ¼ stress exponent ¼ secondary or steady-state creep activation energy ¼ gas constant ¼ temperature in kelvin

Table 2.34 summarizes the key fiber elastic and creep related properties for two important fiber types, Hi-Nicalon-Type S and Sylramic-iBN SiC fibers. Figure 2.103 shows the creep-stress rupture curves for selected materials. Hi-Nicalon-Type S and Sylramic-iBN fiber reinforced composites have a good rupture life at the envisioned 13008C. The 2-D 0/90-balanced fabric, fiber volume fraction is approximately 18% [36, 62]. CMC development has been continuing for achieving 14828Cþ temperature capabilities [63]. The polycrystalline fiber development efforts focused on thermal–chemical treatments to increase fiber grain size and grain size uniformity, reduce porosity and defects by sintering, and reduce and modify the grain boundary low-viscosity phases. Advanced CMC architectures for achieving higher fiber volume fractions, 3-D fiber architectures, and CVI for improved rupture and interlaminar strengths have also been in development [36, 64]. Some examples

HIGH-TEMPERATURE SiC/SiC CMM SiC-BASED FIBER CREEP PROPERTIES

Maximum Use Temperature (88 C)

Elastic Modulus E (GPa)

Tensile Strength (GPa)

Creep Constant C

Stress Exponent n

Grain Size (nm)

Qs (kJ/ Mole)

Hi-Nicalon-Type S (Nippon Carbon)

16508C

355

2.6–2.8

2.2  1016

3

20

814

Sylramic-iBN (NASA)

18008C

380

3.1

7.0  1017

3

250

814

Fiber Types

AEROSPACE MATERIALS CHARACTERISTICS

TABLE 2.34

199

200

B. N. BHAT

Fig. 2.103

Larson-Miller plot for CVI-MI SiC/SiC CMCs with various fiber types.

of 3-D architecture CMCs are shown in Fig. 2.104. Hybrid CMC processing by chemical vapor infiltration–polymer-infiltration and pyrolysis (CVI-PIP), with multiple PIP cycles, has achieved a silicon-free, dense matrix and has shown improved temperature capability and rupture to 14828C. It looks promising with improved through-thickness thermal conductivity [36, 65]. Figure 2.105 shows an example of high heat flux, thermal gradient testing results of advanced EBC coated CVI þ SMI (average CMC and hybrid CVI-PIP CMCs), demonstrating 300–400 h creep and fatigue durability at 27008F (14828C) [66]. The advanced CVI-PIP CMC material showed low creep strains at significantly higher temperatures as compared with the CVI-SMI CMCs. The high heat flux simulated engine environment fatigue life durability of environmental barrier coated prepreg melt infiltration (MI), CVI-SMI, and CVI-PIP SiC/SiC CMC systems has also been highlighted in the literature [42, 58, 67].

2.12.5

SUMMARY

High temperature ceramics materials are crucial for aerospace applications because of their very unique properties. Advanced ceramics possess high temperature or ultra-high temperature capabilities, low density, high temperature creep rupture strengths, as well as excellent oxidation and corrosion resistance. Engineered ceramic structural materials are continued to be a main focus for advanced propulsion engine and air-vehicle structural applications because of the everincreasing needs for efficiency and higher temperature operations.

AEROSPACE MATERIALS CHARACTERISTICS

201

Zirconia (ZrO2) based thermal barrier coatings have been among the most successful applications of modern ceramic materials. Thermal barrier coating systems of ZrO2-(6-8)wt%Y2O3, along with the improvements of nickel-based single crystal superalloy and bond coat technologies, have revolutionized turbine engine industries. Advanced low conductivity multicomponent ZrO2-9.5wt%Y2O3-5.6Yb2O3-5.2Gd2O3, and rare earth zirconate (such as Gd2Zr2O7 and Sm2Zr2O7), have also been incorporated into engine applications. TBCs with approximately 100 to 150 mm thickness have provided a temperature reduction up to 1008C for high pressure turbine airfoils, with further advancement towards 2008C tempertaure reductions when low conductivity TBCs are implemented. The durability and reliability of turbine engine hot-section components have also been significantly improved largely due to the advances in high stability, high toughness and high strength top coats and bond coats. Ceramic thermal and environmental barrier coatings have been successfully developed and implemented for protecting emerging SiC/SiC ceramic matrix composite turbine engine components in high temperatures combustion environments. Fundamental degradation mechanisms have been extensively studied due to the stability concerns of the volatile SiO2 scales for the Si-based ceramics and

a) 2D Five-harness Satin

b) 3D Orthogonal

c) Angle Interlock

d) Braid

CMC fiber architecture design and property modeling

Fig. 2.104 Various SiC/SiC CMC architecture designs for improved rupture and interlaminar strengths [64, 68].

202

Fig. 2.105

B. N. BHAT

Creep and fatigue durability of a turbine airfoil EBC system, on different CMCs.

silicate-based environmental barrier coatings in turbine engine combustion moisture environments. Environmental barrier coating materials, including BASA, mullite, rare earth silicates, rare earth aluminate silicates, and hafnium-rare earth silicates have been developed to provide up to additional 3008C temperature capability over the current state-of-the-art thermal barrier coating systems. Advanced hafnia–silicon and rare earth–silicon, rare earth–hafnium–silicon based bond coats with controlled oxygen activities have also been developed for 14828C temperature capable turbine airfoil environmental barrier coating systems. SiC CMCs reinforced by continuous, polycrystalline high strength SiC fibers are another revolutionary applications of engineered ceramic materials. The turbine engine performance with the CMC components has benefited largely from the materials high temperature capability and low density, thus allowing the design exceeding 2008C temperature increases compared with nickel base superalloys with significantly reduced cooling. The weight reductions realized by applying SiC/SiC CMCs to engine rotating components can further reduce the design complexity and weight of engine structures. The development and

AEROSPACE MATERIALS CHARACTERISTICS

203

implementation of commercial grades prepreg – MI and CVI-SMI SiC/SiC CMCs have achieved the temperature capability of 13168C, while the future SiC/SiC CMC materials are being developed to be capable of 14828C when advanced SiC fibers, 3-D fiber architectures and matrix materials are incorporated. Although ceramic materials have many attributes for high temperature and ultra-high temperature applications, the applications have been limited to less critical components due to their relatively low toughness, low damage tolerance, large variability in their mechanical properties. Complex environmental effects on component design and durability are not well understood in thermal gradient and fatigue operating conditions. Advances in testing, modeling and validation methodologies in combined high-heat-flux, simulated thermal gradient environment and fatigue conditions will significantly facilitate the developments of design databases and simulation tools. The understanding of ceramic material characteristics, their underlying degradation mechanisms and interactions will help revolutionize the component design methodologies and component life prediction.

ACKNOWLEDGEMENTS The author is specifically grateful to Dr. James L. Smialek and Dr. James A. DiCarlo (retired), both Senior Technologists at NASA John H. Glenn Research Center, for helpful discussions. The author also expresses his gratitude to Dr. Biliyar N. Bhat, NASA Marshall Space Flight Center, for helpful discussions and his guidance in this Ceramic Materials section.

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Materials for Advanced Industrial Gas Turbines (AMAIGT) Program Final Report, December 2010. Roode, M. V., Price, J., Kimmel, J., Miriyala, N., Leroux, D., Fahme, A., and Smith, K., “Ceramic Matrix Composite Combustor Liners: A Summary of Field Evaluations,” Journal of Engineering for Gas Turbines and Power, Vol. 129, No. 1, 2007, pp. 21–30. Miriyala, N., and Price, J. R., “The Evaluation of CFCC Liners After Field Engine Testing in Gas Turbine-II,” ASME Turbo Expo 2000: Power for Land, Sea, and Air, ASME 2000-GT-648, Munich, Germany, 2000. Lee, K. N., Fox, D. S., Eldridge, J. I., Zhu, D., Robinson, R. C., Bansal, N. P., and Miller, R. A., “Upper Temperature Limit of Environmental Barrier Coatings Based on Mullite and BSAS,” Journal of the American Ceramic Society, Vol. 86, No. 8, 2003, pp. 1290–1306. Zhu, D., Lee, K. N., and Miller, R. A., “Thermal Conductivity and Thermal Gradient Cyclic Behavior of Refractory Silicate Coatings on SiC/SiC Ceramic Matrix Composites,” Ceramic Engineering and Science Proceedings, Vol. 22, 2001, pp. 443–452. Zhu, D., Choi, S. R., Eldridge, J. I., Lee, K. N., and Miller, R. A., “Surface Cracking and Interface Reaction Associated Delamination Failure of Thermal and Environmental Barrier Coatings,” Ceramic Engineering and Science Proceedings, Vol. 23, 2003, pp. 469–475. Lee, K. N., Fox, D. S., and Bansal, N., “Rare Earth Silicate Environmental Barrier Coatings for SiC/SiC Composites and Si3N4 Ceramics,” Journal of the European Ceramic Society, Vol. 25, 2005, pp. 1705–1715. Zhu, D., Bansal, N. P., and Miller, R. A., “Thermal Conductivity and Stability of HfO2–Y2O3 and La2Zr2O7 Evaluated for 16508C Thermal/Environmental Barrier Coating Applications,” in Advances in Ceramic Matrix Composites IX, Wiley, New York, 2004, pp. 341–343. Zhu, D., “NASA’s Advanced Environmental Barrier Coatings Development for SiC/ SiC Ceramic Matrix Composites: Understanding Calcium Magnesium Alumino-Silicate (CMAS) Degradations and Resistance,” Thermal Barrier Coatings IV Conference: An ECI Conference Series, Irsee, Germany, 22–27 June 2014. Ahlborg, N. L., and Zhu, D., “Calcium Magnesium Alumino-Silicate (CMAS) Reactions and Degradation Mechanisms of Advanced Environmental Barrier Coatings,” Surface and Coatings Technology, Vol. 237, 2013, pp. 79–87. Poerschke, D. L., Sluytman, J. V., Wong, K., and Levi, C., “Thermochemical Compatibility of Ytterbia-(Hafnia/Silica) Multilayers for Environmental Barrier Coatings,” Acta Materialia, Vol. 61, No. 18, 2013, pp. 6743–6755. Poerschke, D. L., Seward, G. E., and Levi, C. G., “Influence of Yb:Hf Ratio on Ytterbium Hafnate/Molten Silicate (CMAS) Reactivity,” Journal of the American Ceramic Society, Vol. 99, 2016, pp. 651–659. Kitaoka, S., Matsudaira, T., Yokoe, D., Kato, T., and Takata, M., “Oxygen Permeation Mechanism in Polycrystalline Mullite at High Temperatures,” Journal of the American Ceramic Society, Vol. 100, No. 7, July 2017, pp. 3217–3226. Wada, M., Matsudaira, T., Kawashima, N., Kitaoka, S., and Takata, M., “Mass Transfer in Polycrystalline Ytterbium Disilicate Under Oxygen Potential Gradients at High Temperatures,” Acta Materialia, Vol. 135, No. 8, 2017, pp. 372–381.

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[57] Halbig, M. C., Jaskowiak, M. H., Kiser, J. D., and Zhu, D., “Evaluation of Ceramic Matrix Composite Technology for Aircraft Turbine Engine Applications, AIAA 2013-0539,” 51st AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, Aerospace Sciences Meetings, Grapevine, TX, 7–10 Jan. 2013. [58] Zhu, D., Hurst, J. B., and Jaskowiak, M. H., “Advanced Environmental Barrier Coating Development and Validation for SiC/SiC Ceramic Matrix Composite Turbine Engine Components,” 36th Annual Conference on Composites, Materials, and Structures, Cocoa Beach, FL, 2012. [59] Brewer, D., “HSR/EPM Combustor Materials Development Program,” Materials Science and Engineering, Vol. A261, 1999, pp. 284–291. [60] Gorman, G. S., and Luthra, K. L., “Silicon Melt-Infiltrated Ceramic Composites (HiPerCom),” Handbook of Ceramic Composites, Boston, Kluwer Academic Publishers, 2005, pp. 99–115. [61] Luthra, K., “Emerging Applications and Challenges in Using Ceramics at General Electric,” Ceramic Leadership Summit 2011, Baltimore, MD, 1–3 Aug. 2011. [62] Dicarlo, J. A., “SiC Fiber Creep and Rupture Models for Understanding CMC Behavior Above 14008C,” Workshop on the Design of Ceramic–Fiber Based Composites for Service Above 14008C, Boulder, CO, June 2012. [63] Dicarlo, J. A., Jacobson, N. S., Lizcano, M., and Bhatt, R., “Ultra High Temperture (UHT) SiC Fiber (Phase II),” NASA TM-2015-218883, Cleveland, OH, 2015. [64] DiCarlo, J., and Bhatt, R., “Modleing SiC/SiC Creep Rupture Behavior from 2400 to 3000F,” 35th Annual Conference on Composites, Materials, and Structures, Coca Beach, FL, 24–27 Jan. 2011. [65] Bhatt, R., DiCarlo, J. A., and Kiser, J. D., “Tensile and Creep properties of Hybrid CVI-PIP SiC/SiC Compsoites at High Temperatures in Air,” 36th Annual Conference on Composites, Materials, and Structures, Cocoa Beach, FL, 23–26 Jan. 2012. [66] Zhu, D., Bhatt, R., and Harder, B., “Combined Thermomechanical and Environmental Durability of Environmental Barrier Coating Systems on SiC/SiC Ceramic Matrix Composites,” 9th International Conference on High Temperature Ceramic Matrix Composites (HTCMC-9), Toronto, 2016. [67] Zhu, D., Harder, B., Hurst, J. B., Costa, B. G. G., Bhatt, R., and Fox, D. S., “Development of Advanced Environmental Barrier Coatings for SiC/SiC Ceramic Matrix Composites: Path toward 2700F Temerature Capability and Beyond,” i41st Annual Conference on Composites, Materials and Structures, Cocoa Beach, FL, 2017. [68] Zhu, D., “Development of Durable Ceramic Matrix Composite Turbine Components for Advanced Propulsion Engine Systems,” 8th Pacific Rim Conference on Ceramic and Glass Technology, Vancouver, British Columbia, Canada, 2009.

CHAPTER 3

Materials Selection for Aerospace Systems Steven M. Arnold NASA Glenn Research Center, Cleveland, Ohio

David Cebon and Mike Ashby University of Cambridge, Cambridge, United Kingdom

3.1 INTRODUCTION The importance of weight reduction in aerospace systems has been a major factor from the very beginning. Consequently, the use of new materials [e.g., dual alloy turbine disks and composite materials (ceramic, metallic and polymer based)] and new types of structural concepts, particularly the thin-walled type (sized primarily based on buckling) have dominated. Clearly, from a designer’s perspective the primary function of a structure is to transmit forces through space with the minimum possible weight and cost to the customer. Typically, the job of a designer is to balance a variety of functional requirements, such as types of loading conditions (tension, compression, bending, vibration, cyclic, etc.), with constraints (manufacturability, geometric limits, environmental aspects, and maintainability, to name a few) so as to arrive at the “optimum” choice of structural concept and materials selection for a given weight and/or cost (Fig. 3.1). This task can often be a daunting one for the inexperienced as well as the experienced practitioner due to the wide range of choices available. Further, depending on the engineer’s background (e.g., a materials- vs mechanics-oriented professional), his or her approach to designing for suitable deflection, for example, may be quite different. A materials-oriented person would typically think he needs a material with high stiffness [i.e., Young’s modulus (E)], whereas a structures-oriented person would naturally think of high rigidity [e.g., EI, the product of Young’s modulus and the moment of inertia (I) given a beam in bending, the first being material property oriented and the other performance oriented. This diversity in perspective is precisely why a robust and systematic methodology for connecting material and application is required, as will be presented and illustrated in this chapter. Further, design is not a linear process; it involves multiple iterations because the consequences of choices (both structural and material) made at the preliminary concept or embodiment stages may not become apparent until

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Fig. 3.1 Intersection of material, manufacturability, and structural concept that determines a system’s engineering. much later downstream when the details are examined more fully. Consequently, iteration (looping back to explore alternatives) is an essential part of the design process, as is materials selection, which enters each stage of the design process. The importance of incorporating as much physics and understanding of the problem as early in the design process as possible is dramatically illustrated in Fig. 3.2, wherein it has been proposed that more than 50% of the lifecycle cost of a given system is locked in by the end of the preliminary design phase, with 85% by the end of the embodiment phase and almost 100% by the time one goes to production, yet only 10% of the life-cycle cost has been expended. This figure strongly suggests, once again, the need for a robust and systematic methodology for connecting materials selection and application requirements as early in the process as possible to reduce the overall life-cycle cost.

Fig. 3.2

Life-cycle cost (Larson and Pranke, 1999) [1].

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Clearly, material information is central to all aspects of engineering products. However, the nature of the data needed in the early stages differs greatly in its level of precision and breadth from that needed later. At the concept stage, the designer requires approximate property values for the widest possible range of materials. All options are open: a polymer may be the best choice for one concept and a metal may be the best choice for another, even though the function is the same. The problem, at this stage, is not precision and detail; it is breadth and speed of access: how can the vast range of data be presented to give the designer the greatest freedom in considering alternatives? At the embodiment stage, the landscape has narrowed. Here data for a subset of materials are now needed, but at a higher level of precision and detail. These are found in the more specialized handbooks and software that deal with a single class or subclass of materials—metals, or just aluminum alloys, for instance. The risk now is that of losing sight of the bigger spread of materials to which one must return if the details do not work out; it is easy to get trapped by a single line of thinking (or set of “connections”) when other combinations of connections offer a better solution to the design problem. The final stage of detailed design requires a still higher level of precision and detail, but for only one or a very few materials. Such information is best found in the data sheets issued by the material producers themselves and in detailed databases for restricted material classes. A given material (e.g., polyethylene) has a range of properties (that are derived from differences in manufacturing) depending on the producers that make it. At the detailed design stage, a supplier must be identified, and the properties of its product must be used in the design calculations, given the slight property differences observed in a material from different material producers. Sometimes even this is not good enough. If the component is a critical one (meaning that its failure could, in some sense or another, be disastrous), then it may be prudent to conduct in-house tests to measure the critical properties, using a sample of the material that will be used to make the product itself. At the core of the materials selection process is the inherent interaction between function, material, shape, and process (see Fig. 3.3). Function dictates the choice of both material and shape; yet the selection of a material and process cannot be separated from the choice of shape. The word shape is used here to include the external macro-shape and, Fig. 3.3 At the core of materials selection is the interaction between function, material, shape, and process.

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when necessary, the internal, or micro-shape, as in a honeycomb or cellular structure. To manufacture a shape, the material is subjected to processes such as primary forming processes (e.g., casting and forging), material removal processes (e.g., machining and drilling), finishing processes (e.g., polishing) and joining processes (e.g., welding). These processes are influenced by the material’s formability, machinability, weldability, heat-treatability, and so on. Also, process obviously interacts with shape; the process determines the shape, the size, the precision, and, of course, the cost. The processing route can also dramatically change the properties of the material. These interactions are two way: specification of shape restricts the choice of material and process, but the specification of process equally limits the materials you can use and the shapes they can take. The more sophisticated the design, the greater the interactions, and, consequently, the tighter the specifications.

3.2 SYSTEMATIC APPROACH TO MATERIALS SELECTION To select means to choose, but from what? Behind the concept of selection lies that of the permissible set of entities from which the choice is to be made. Within an engineering enterprise, the permissible set of materials will probably be the preferred list, specified by the materials authority. However, for truly optimal materials selection, the permissible set of entities must include all materials: all metals, all polymers, all ceramics and glasses, and all composites. This overall set of materials is referred to as the “universe” of materials. If some of these materials are left out, the selection is no longer optimal over all materials but is only optimal over some subset of them. If, for example, the choice is limited to metals from the start, then the selection can only be optimal over the single family of materials, that of metals. There is a second implication to the concept of selection. It is that all members of the permissible set must be regarded as candidates—they are, after all, viable choices—until, by a series of selection stages, they are shown to be otherwise. From this arise several key requirements of a data structure for the selection database. The selection table must be comprehensive (include all members of the permissible set). It should be structured; that is, it must contain attributes that are universal (apply to all members of the permissible set of entities), and the attributes should be discriminating (have recognizably different values for different members of the permissible set). Similar considerations apply to any selection exercise [2]. In the universe of materials, many attributes are universal and discriminating: density, bulk modulus, and thermal conductivity are examples. Universal attributes can be used for screening and ranking, the initial stage of any selection exercise. But if the values of one or more screening attributes are grossly inaccurate or missing, that material will be eliminated by default. It is important, therefore, that the database be complete, having no holes or gaps in the records that would make a material fail a selection by default due to absence of data, and be of high quality,

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Fig. 3.4

Materials selection strategy.

meaning that the data can be trusted. This creates the need for data checking and estimation, which are tackled by methods described by Ashby [3]. There are five main steps to the materials selection strategy: translation, screening/rejection, ranking, supporting information and research, and organizational cultural constraints (see Fig. 3.4). The steps can be likened to those in selecting a candidate for a job. The job is first analyzed (i.e., essential skills and required experience of the candidate are identified; “translation”) and advertised. Some of these are simple go/no-go criteria such as the requirement that the applicant must have a valid driver’s license or a degree in materials science, eliminating anyone who does not (“screening”). Others imply a criterion of excellence, such as “typing speed and accuracy are priorities,” or “preference will be given to candidates with a substantial publication list,” implying that applicants will be ranked by these criteria (“ranking”). Finally, references and interviews are sought for the top-ranked candidates, building a file of supporting information (research and local cultural constraints)—an opportunity to probe deeply into character and potential.

3.2.1 METHODOLOGY 3.2.1.1

STEP 1: TRANSLATION

How are the design requirements for a component (defining what it must do) translated into a prescription for a material? Any engineering component has one or more functions: to support a load, to contain a pressure, to transmit heat, and so forth. Function must be achieved subject to constraints: that certain dimensions are fixed, that the component must carry the design loads or pressures without failure, that it insulates or conducts, that it can function in a certain range of temperatures and in a given environment, and many more. In designing the component, the designer has an objective: to make it as cheap as possible, as light, as safe, or perhaps some combination of these. Certain

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TABLE 3.1 Function Constraints

FUNCTION, CONSTRAINTS, OBJECTIVE, AND FREE VARIABLES What does the component do?

a

What nonnegotiable conditions must be met? What are the negotiable but desirable conditions?

Objective

What is to be maximized or minimized?

Free variables

What parameters of the problem is the designer free to change?

a

It is sometimes useful to distinguish between hard and soft constraints. Stiffness and strength might be absolute requirements (hard constraints); cost might be negotiable (a soft constraint).

parameters can be adjusted to optimize the objective, giving the designer the freedom to vary dimensions that have not been constrained by design requirements and, most important, the freedom to choose the material for the component. These are referred to as free variables. Function and constraints, objective, and free variables (see Table 3.1) define the boundary conditions for selecting a material and—in the case of load-bearing components—a shape for its cross section. The first step in relating design requirements to material properties is a clear statement of function, constraints, objective, and free variables. Function: Structural elements are components that perform a physical function (i.e., they carry loads, transmit heat, store energy, etc.); in short, they satisfy functional requirements. The functional requirements are specified by the design; for example, a tie rod must carry a specified tensile load, a spring must provide a given restoring force or store a given energy, and a heat exchanger must transmit heat at a given heat flux. Furthermore, the loading on a given component/structure can generally be decomposed into some combination of axial tension or compression, bending, and torsion, wherein one mode almost always dominates. So common is this that the functional name given to a component often describes the way it is loaded: ties carry tensile loads, beams carry bending moments, shafts carry torques, and columns carry compressive axial loads. The words tie, beam, shaft, and column each imply a function; in fact, many simple engineering functions can be described by single words or short phrases, saving the need to explain the function in detail. Table 3.2 itemizes some aerospace components and their associated simple idealization, which can be extremely useful for materials selection purposes. Constraints: Constraints are design requirements that must be satisfied: for example, the minimum working temperature of a material must be 2208C, p the strength must be 400 MPa, the fracture toughness must be 15 MPa m, and so on. Constraints screen out unsuitable choices or enable rejection of specific materials for cause. Objectives: Objectives are design criteria that must be maximized or minimized to optimize the performance of a component. Their function is to rank the materials and facilitate selection of the best candidates. For materials selection,

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the objectives can be used to generate material performance indices, which are combinations of material properties that characterize performance in a given application. Typical examples are the specific stiffness of a material E/r, and the specific strength sf/r (E is the Young’s modulus, sf is the failure strength, TABLE 3.2 Application

AEROSPACE APPLICATIONS/IDEALIZATIONS

Idealization

Loading

Constraints

Truss framework

Tie

Tension

Strength

Fuselages

Cylinder

Compression Tension Bending

Buckling Strength Deflection

Shafts

Thin-walled tube/beam

Torque Rotation Bending

Bending Shear Vibration

Combustion chambers

Thin-walled cylinder, internal pressure

Pressure Noise

Thermal Oxidation

Blades (fan, compressor, turbine)

Tie Beam

Bending Tension Fatigue Thermal Creep

Deflection, strength, vibration, erosion, temperature

Disks (compressor, turbine)

Disk

Rotation, thermomechanical fatigue

Strength, ductility, toughness, oxidation

Space truss framework

Tie rod

Axial loading

Strength, buckling

Nozzles

Thin-walled cylinder

Pressure Noise Thermal

Strength, vibration

Wing

Plate

Bending Tension Twist Vibration Fatigue

Buckling Vibration, strength Deflection

Cryogenic tanks

Spherical shell

Internal pressure Thermal

Strength, low temperature Yield before break Leak before break

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and r is the density). These particular indices can be used to select the optimum material for a light, stiff tie rod, or a light, strong tie rod, respectively. Many material performance indices have been derived and tabulated for standard design cases in mechanical, structural, thermomechanical, and electromechanical engineering [4]. Free variables: Free variables are any parameters that are left open to the designer’s free choice. Examples are cross-sectional area, length, thickness, material type, color, and so on. 3.2.1.2

STEP 2: SCREENING/REJECTION

Unbiased selection requires that all permissible materials are considered to be candidates until shown to be otherwise. Rejection eliminates candidates that cannot do the job at all because one or more of their attributes lies outside the limits set by the constraints. As examples, the requirements that “the component must function in boiling water,” or that “the component must be transparent” impose obvious limits on the attributes of maximum service temperature and optical transparency that successful candidates must meet. These constraints are referred to as attribute limits. One should not be too hasty in applying attribute limits, however, as it may be possible to engineer around them; for example, a component that gets too hot may be cooled (or the geometry modified to enable use of the same material at lower stress, EI vs E), or one that corrodes may be coated with a protective film. An aerospace-specific case in point is the desire by many engineers to impose a minimum limit on fracture toughness and strain to failure (e.g., KIC . 15 MPa/m2 and ef . 2%) to ensure adequate tolerance to stress concentrations. This is likely to be a rash step early in the design process, as it would eliminate all polymers, ceramics, and many composite materials from consideration. A more innovative designer may be able to put these materials to good use. Consequently, at this stage in the process, the engineer must try to keep as many options open as possible. 3.2.1.3

STEP 3: RANKING

Attribute limits do not help with ordering the candidates that remain after the screening constraints have been applied. To do this, one needs optimization criteria. They are found in the material indices, a subset of the performance indices described in Sec. 3.2.2, which measure how well a candidate can do the job. Material indices provide criteria of excellence that allow the ranking of materials by their ability to perform well in the given application. Performance is sometimes limited by a single property, sometimes by a combination of them. For example, the best materials for buoyancy are those with the lowest density r; those best for thermal insulation are the ones with the smallest values of the thermal conductivity l. Here maximizing or minimizing a single property maximizes performance. But as will be shown, the best materials for a light stiff tie rod are those

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with the greatest value of specific stiffness E/r, where E is Young’s modulus. The best materials for a spring are those with the greatest value of sf2 =E, where sf is the failure stress. The property or property group that maximizes performance for a given design is called its material index. There are many such indices, each associated with maximizing some aspect of performance. These indices provide significant insight into the types of optimization strategies needed as well as selection guidelines for the given problem at hand. To summarize, material rejection (or screening) isolates candidates that are capable of doing the job; ranking identifies those among them that can do the job best. 3.2.1.4

STEP 4: RESEARCH

The outcome of steps 1 through 3 produces a ranked short list of candidates that meet the constraints and that maximize or minimize the criterion of excellence, whichever is required (this is why a funnel is used in Fig. 3.4). One could just choose the top-ranked candidate, but what secrets might it hide? What are its strengths and weaknesses? Does it have a good reputation? What, in a word, is its pedigree? To proceed further, one must seek a detailed profile of each candidate: its supporting information (see Step 4 in Fig. 3.4). Supporting information differs greatly from the structured property data used for screening. Typically, it is descriptive, graphical, or pictorial: case studies or experience of previous uses of the material, details of its corrosion behavior in particular environments, information of availability and pricing, or experience of its environmental impact. Such information is found in corporate documentation, handbooks, suppliers’ data sheets, CD-based data sources, and the Internet. Supporting information helps narrow the short list to a few final choices, allowing a definitive match to be made between design requirements and material attributes. Why are all these steps necessary? Without screening and ranking (material rejection), the candidate pool can be enormous and the volume of supporting information overwhelming. Dipping into it, hoping to stumble on a good material, gets you nowhere. But once a small number of potential candidates has been identified by the screening and ranking processes, detailed supporting information can be sought for these few candidates, and the task becomes tractable. 3.2.1.5

STEP 5: SPECIFIC CULTURAL CONSTRAINTS

The final choice between competing candidates will often depend on local conditions: in-house expertise or equipment, the availability of local suppliers, and so forth. A systematic procedure cannot help here; the decision must instead be based on local/institutional knowledge. This does not mean that the result of the systematic procedure (describe earlier) is irrelevant. It is always important to know which material is best, even if, for local reasons (e.g., the specific processing method for a possible material is too expensive), you decide not to use it.

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3.2.2 MATERIAL PERFORMANCE INDICES The screening and ranking processes can be made quantitative by linking the technical (e.g., functional, geometric, material) and economic requirements of the design to the attribute profiles stored in a given database. The foundations of a robust and systematic ranking procedure are particular combinations of material properties that embody the performance of the component. These property combinations are called performance indices, as put forth by Ashby [4]. The performance P of a structural element is determined by three typically independent aspects: the functional requirements F; the geometry G; and the properties of the material of which it is made, M. The performance P of the element can often be described by an equation of the form P ¼ f (F, G, M)

(3:1)

where P, the performance metric, describes some aspect of the performance of the component (e.g., its mass, volume, cost, or life) and f ( ) means “a function of.” Optimum design is the selection of the material and geometry that maximizes or minimizes P, according to its functional requirements. When this group of parameters can be assumed to be separable (that is, M, F, and G are independent of each other), then the performance index is merely a product of three functions, f1, f2, and f3: P ¼ f1 (F)  f2 (G)  f3 (M)

(3:2)

where the product f1(F) . f2(G) has been defined as the structural efficiency coefficient (and incorporates both functional requirements and geometry), and f3(M ) is defined as the material efficiency coefficient. With this significant simplification, the overall performance index can be maximized or minimized by selecting a material to minimize f3(M ), independent of the details of the design. This enables the optimal subset of materials to be identified without solving the complete design problem. Although this is clearly a simplification of the full coupled design problem, it can provide a great deal of insight in the preliminary design stage of a project. This powerful and general method is simple, provided one can clearly identify, from the outset, the objective(s) (what you are trying to maximize or minimize), the constraints, which parameters are specified, and which parameters are free. When any component function is combined with specific constraints and objectives, a specific material index [Mp ¼ f3(M )] is produced, as illustrated in Fig. 3.5. Therefore, it is possible to develop numerous material and structural indices that enable one to efficiently and accurately perform materials selection for a variety of design functions (e.g., stiffness, strength, and vibration-limited designs, as well as damage-tolerant, electromechanical, thermal, and thermalmechanical designs) on a number of fundamental structural components described in Table 3.2 (i.e., tie, beam, shaft, column, panel, cylinder/shell with internal pressure, and rotating disks); see Ashby [4] for additional details. The

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Fig. 3.5

219

Specification of function, objective, and constraint leads to a material index [3, 4].

power of this approach is clearly illustrated by examining the differences in the resulting material indices for a tie, beam, and plate (see Fig. 3.6) with a specified r r ; E1=3 , respectively. Clearly, the only difference in the stiffness, that is, Er ; E1=2 indices is the exponent on Young’s modulus; for a tie it is 1, for a beam it is 12, and for a plate it is 13. Consequently, the best materials for a light, stiff tie rod, r r ; E1=3 , respectively. beam, or plate are those with the smallest values of Er ; E1=2 The significance of this difference in exponents on the optimal choice is best seen by creating a plot of Young’s modulus vs density (i.e., a material property chart) as will be described later in this section. a)

b)

c)

Fig. 3.6 (a) A tie with cross-sectional area A, loaded in tension. Its stiffness is S 5 F/d, where F is the load and d is the extension. (b) A panel loaded in bending. Its stiffness is S 5 F/d, where F is the total load and d is the bending deflection. (c) A beam of square section, loaded in bending. Its stiffness is S 5 F/d, where F is the load and d is the bending deflection.

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Fig. 3.7 Influence of the end constraint on the elastic flexural stiffness for beams and plates, see Table 3.3. Before continuing, two types of applications will be introduced that will be used throughout this chapter to demonstrate the various concepts discussed herein; they are those of blades, be they in the fan (low temperature), compressor (mid temperature), or turbine (high temperature) sections of a jet engine (see Fig. 3.8), and that of an internal pressure vessel (e.g., cryogenic tank, fuel lines, or combustion chambers within a jet engine; see Fig. 3.10). The indices that are useful for materials selection will be established. 3.2.2.1

CASE STUDY 1A: BLADES

An important aeronautic application that illustrates the materials selection process is the use of fan blades, compressor blades, or turbine blades in a gas turbine engine, as shown in Fig. 3.8. The main functional distinctions in the blades used in the different locations are size, shape, and temperature capacity. Figure 3.8c illustrates how a typical blade can be idealized as a cantilever beam subjected to a combination of axial tension (due to centrifugal loads), bending (due to pressure variations across the blade), and vibration loadings (due to the motion of the rotor and rapidly changing air flow field in the engine). The length L and force F are specified and are therefore fixed; the cross section A is free. The mass can be reduced by decreasing the cross-sectional area, but there is a constraint: the section area A must be sufficient to carry the tensile load F, given a failure strength sf, thus requiring that A

F sf

(3:3)

Consequently, combining this constraint with the objective function, one obtains ! r (3:4) m  (F) (L) sf Note the form of this result. The first bracket contains the specified load F. The second bracket contains the specified geometry (the length L of the tie/beam).

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Together, they make up the structural index previously discussed. The last bracket contains a specific relationship of material properties that were previously defined as the material index. The lightest beam that will carry F safely is the one made of the material with the smallest value of r/sf. This could be defined as the material index of the problem, seeking a minimum, but when dealing with specific properties, it is more common to express them in a form for which a maximum is sought. To accomplish this, one inverts the material properties in Eq. (3.4) and defines the material index M1 as   sf M1  (3:5) r Therefore, the lightest beam or tie rod that will safely carry the load F without failing is that with the largest value of this index, the “specific strength.” A a) Without high bypass ratios

b) With high bypass ratios

Fig. 3.8

c) Typical blade response due to engine operations

Schematics of turbine jet engines.

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TABLE 3.3

DESIGN REQUIREMENTS FOR LIGHT, STIFF TIE, BEAM, AND PLATE (SEE FIG. 3.6)

Function

Tie

Beam (A ¼ b . b)

Plate (A ¼ h . W )

Constraint Functional constraints

Beam must support a Tie must bending load F without support axial deflecting too much, tensile load F meaning that the without bending stiffness deflecting too S is specified much (S ¼ F/ d  A . E/Lo) (S  C1 . E . I/L 3)

Panel must support a bending load F without deflecting too much, meaning S is specified (S  C1 . E . I/L 3)

Geometric constraints

Length Lo is specified

Length L is specified

Length L and width W are specified

Objective

Minimize mass (m ¼ A . L . r)

Minimize mass (m ¼ A . L . r)

Minimize mass (m ¼ W . h . L . r)

Free variables

Cross-sectional area A Choice of material

Cross-sectional area A Choice of material

Panel thickness h Choice of material

Performance equation

  m  (S)(L2 ) Er

m



1=2

12S C1 L



12S(L3 ) r C1 L E 1=2



m



1=3

12S C1 L



r 12S 2 C1 W (WL ) E 1=3



similar calculation for a light, stiff tie (one for which the stiffness S rather than the strength sf is specified) leads to the preceding index, given in Table 3.3, where C1 is given in Figure 3.7 for various end constraints. 3.2.2.2

PERFORMANCE INDEX FOR BENDING AT ROOM TEMPERATURE

Consider the case of a beam of section area A and length L. Subjected to bending, it must support a specified load F without failing and be as light as possible (see Fig. 3.6). From strength of materials theory, the stress due to bending is

smax ¼

My M ¼ I Z

(3:6)

where I ¼ the moment of inertia y ¼ the distance from the neutral axis to the outer edge of the beam Z ¼ the section modulus Failure occurs if the load exceeds the moment, that is, M ¼ Z smax ¼ Z sf

(3:7)

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where again sf is the stress at which failure of the material will occur. To enable the section shape to change for the same area of material, the section modulus Z is replaced by fBZo, where Zo is the section modulus for a square beam (i.e., Zo ¼ A3/2/6) and fB is the shape efficiency factor. The maximum shape factor can be considered a material property and used profitably in the selection of materials. Note that solid equiaxed sections (circles, squares, hexagons, etc.) all have efficiency factors close to 1, whereas efficient shapes like thin-walled tubes or I-sections can have shape factors of 50 or more. The physical limit to fB is usually set by local buckling of the component. As before in the case of the tie rod, the mass can be decreased by reducing the cross section, but there is a constraint: the section area A must be sufficient to carry the bending stress, given a failure strength sf, thus requiring that !2=3 M (3:8) A 6 fB sf Finally, combining this constraint with the objective of minimum weight, m ¼ rLA, one obtains the following expression: !2=3 r3=2 2=3 (3:9) m  ð6M Þ L fB s f Consequently, the best material and shape combination is that with the greatest value of the material index: M2 

3.2.2.3

ðfB sf Þ2=3 r

(3:10)

PERFORMANCE INDEX FOR VIBRATION AT ROOM TEMPERATURE

There are two ways to mitigate vibration of a mechanical system: by ensuring that the input does not excite a natural frequency, or, if this is not possible, by ensuring that the system is sufficiently well damped to avoid excessive motion at resonance. Let us consider the first of these methods. In many cases, the need is to avoid exciting the lowest natural frequency of the system, and this can be achieved by ensuring that it occurs at high frequency relative to the input. Any undamped system vibrating at one of its natural frequencies can be reduced to the simple problem of a mass m attached to a spring of stiffness k, the restoring force per unit displacement. The natural frequency of such a system is rffiffiffiffi 1 k (3:11) f ¼ 2p m

224

Fig. 3.9

S. M. ARNOLD ET AL.

Natural vibration modes of beams clamped in different ways.

Different geometries require appropriate estimates of the effective stiffness k and mass m. Often these can be estimated with sufficient accuracy by approximate modeling. Depending on the geometry of the beam and the end constraints, the higher natural frequencies of rods and beams can be simple multiples of the lowest. Figure 3.9 shows the lowest natural frequencies of the flexural modes of uniform beams or plates of length L with various end constraints. The spring stiffness k is equivalent to the stiffness S in Table 3.3 for beam/plate bending, and so the natural frequencies can be written as rffiffiffiffiffiffiffiffiffiffiffi C2 EI (3:12) f ¼ 2p mo L4 where C2 (given in Fig. 3.9) depends on the end constraints and mo is the mass of the beam per unit length. The mass per unit length is just the area times the density Ar, and so the natural frequency becomes rffiffiffiffiffiffiffiffiffiffisffiffiffiffiffiffiffiffiffi C2 A E fB (3:13) f ¼ 2p 12L4 r where I ¼ fBIo has been used to enable the most efficient shape to be considered. Thus, when all dimensions are specified or when the stiffness is specified, the frequencies   sffiffiffiffiffiffiffiffiffi C2 S 1=2 EfB (3:14) f ¼ 2p 12L5 r pffiffiffiffiffiffiffiffiffiffiffiffiffiffi scale with M3 ¼ EfB =r. 3.2.2.4

CASE STUDY 2A: PRESSURE VESSELS

Multiple applications/structures can be classified as cylindrical or spherical in nature, particularly in the case of space structures. Often these structures are internally pressurized as well: for example, the feed lines on rocket engines,

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225

composite overwrap pressure vessels, cryogenic fuel tanks, and expendable/reusable launch vehicles, as shown in Fig. 3.10. Consequently, in this case study, performance indices pertinent to cylindrical/spherical pressure vessels will be identified. In the case of cryogenic storage tanks, multiple design configurations can be, and have been, envisioned, from a single tank with insulation to hybrid tanks with either insulating materials or pure vacuum in between walls or various combinations thereof. The overall objective of the designs is to have a safe, lightweight, thermally efficient cryogenic storage system. Some important tank system parameters relative to flight durations are presented in Table 3.4. The materials, tank structural configurations, and insulation system options are numerous and interdependent. The key material indices applicable for both thermal and mechanical issues of interest in the case of cryogenic tanks are shown in Table 3.5. Clearly it is desirable to use materials that possess high strength, high fracture toughness, and high stiffness, as well as low density and low permeation to liquid and gaseous hydrogen; however, no single material provides all these attributes simultaneously. Consequently, material performance indices associated with these properties, such as those given in Table 3.5, must be used to identify the best material candidates for hybrid tank wall construction. Among these parameters, strength and density tend to dominate the design criteria. c) Composite overwrapped b) Ares V heavy lift system

pressure vessels

d) Double-walled cryogenic storage tank

a) Space shuttle main engine (SSME)

Fig. 3.10

Various space structures/components/vessels.

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TABLE 3.4 IMPORTANT TANK SYSTEM PARAMETERS RELATIVE TO FLIGHT DURATION (IN ORDER OF IMPORTANCE) Short flight duration

Mass density Strength and toughness Coefficient of thermal expansion Stiffness Thermal diffusivity Thermal conductivity

Long flight duration

Mass density Thermal conductivity Strength and toughness Coefficient of thermal expansion Stiffness Thermal diffusivity

In the case of thin-walled spherical pressure vessels of radius R, the stress in the wall is given by

s¼

TABLE 3.5

PR 2t

(3:15)

PERFORMANCE INDICES FOR THERMAL AND MECHANICAL COMPONENTS OF CRYOGENIC STORAGE TANKS

Function and Constraints

Performance Index, Maximize

Thermal Minimum heat flux at steady state, fixed thickness Minimum temperature rise in specified time, fixed thickness Maximum energy stored for given temperature rise and time Minimum thermal distortion

1/k 1/a k/a 1/2 k/a

Mechanical Strength-limiting design with minimum mass Damage-tolerant design with minimum mass Deformation-limiting design with minimum mass

sf/r KIc/r E/r

k ¼ thermal conductivity a ¼ thermal diffusivity (k/rCp) r ¼ mass density a ¼ coefficient of thermal expansion sf ¼ strength KIc ¼ mode I fracture toughness E ¼ Young’s modulus

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227

and the mass of the thin-walled spherical tank is given by m ¼ 4pR2 t r

(3:16)

From fracture mechanics, it is known that a crack will propagate by fast fracture when the stress intensity factor, which is a combination of the applied far field stress s and crack length 2c, reaches the fracture toughness of the material KIc. This condition can be expressed by the following relation: pffiffiffiffiffiffi (3:17) KIc ¼ Y s pc where Y is a parameter that takes on various values depending on the changes in crack geometries and loading conditions (see Broek [5] and Anderson [6]). Consequently, the internal pressure at which the working stress of the tank is below the critical stress that would cause a crack to propagate can be obtained by rearranging Eqs. (3.15) and (3.17) and substituting the results:   2t KIc pffiffiffiffiffiffi (3:18) P R Y pc Clearly, the largest pressure (for a given R, t, and c) is obtained for the material with the largest fracture toughness because the pressure in Eq. (3.18) is proportional to KIc. Such a design, however, is not fail safe, and so additional conditions have been traditionally introduced to provide a greater level of safety. These are known as yield-before-break and leak-before-break. Yield-before-break requires that the stress to cause fracture be greater than that to yield the material, thus providing a state of detectable deformation before fracture. This condition is expressed as  2 1 KIc (3:19) Cmax  2 pY s y where the stress in Eq. (3.17) has been replaced by the yield stress sy of the material. Consequently, the tolerable crack size, and thus integrity of the pressure vessel, is maximized by maximizing the following material index: M4 ¼

KIc sy

(3:20)

Leak-before-break, typically used for large-pressure vessels, requires that a stable crack be just large enough to penetrate both the inner and outer surfaces of the vessel (i.e., without fast facture) so that the leak caused by the presence of the crack can be detected. This condition is obtained by requiring that the unstable crack size exceeds the thickness (i.e., 2c . t) and that the wall

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thickness t contain the pressure P without yielding. This condition leads to the following:  2 KIc 4 (3:21) P 2 pRY sy Consequently, the pressure is carried most safely when the material index M5 ¼

KIc2 sy

(3:22)

is the largest possible. Both of these material performance indices, M4 and M5, can be maximized by choosing materials with low yield stresses; however, that could be problematic because the thickness and therefore the mass of the tank is inversely related to the yield stress [see Eqs. (3.15) and (3.16)]. Therefore, for aerospace applications, which are typically weight critical, this demands that sy/r be as large as possible since m

2pR3 P sy =r

(3:23)

These indices will be used later to down-select materials for use with lightweight pressure vessels.

3.2.3 MATERIAL PROPERTY CHARTS Material property charts enable the condensing of a large body of information into a compact but accessible form. They reveal correlations between material properties that aid in checking and estimating data and, in conjunction with performance indices, constitute the backbone for tackling real design problems. They provide a convenient way of mapping out the areas of “property space” occupied by each material class, and the subareas occupied by individual materials, and provide a mechanism for surveying design-limiting properties (be it a single value or a combination of properties). Using the performance indices discussed in the preceding section enables the engineer to choose the chart axes so that function-specific information can be displayed. For example, a chart of Young’s modulus (E) vs density (r) (see Fig. 3.11) not only guides the best choice of material for the stiffness-limited design of a tie rod, beam, or plate, but also shows the longitudinal wave velocity (E/r)1/2 and design for vibration. Similarly, a plot of fracture toughness KIC against modulus E shows the critical strain energy release rate (toughness) GIC; a diagram of thermal conductivity (k) vs diffusivity (a) also gives the volume-specific heat rCv; and tensile strength (sf) vs Young’s modulus (E) shows the elastic energy-storing capacity s2f /E, to name a few.

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

Fig. 3.11

229

Modulus vs density property chart.

The most striking feature of the charts, as pointed out by Ashby [4], is the way in which members of a material class cluster together despite the wide range in various properties. For example, the modulus and density of metals in Fig. 3.11 (denoted by dark gray contour labeled Metals) occupy a field that is distinct from that of polymers (denoted by lighter gray), that of ceramics (lightest gray ellipse), or that of composites (hashed ellipse). The same is true of strength, toughness, thermal conductivity, and the rest. Even though the fields sometimes overlap, they always have a characteristic place within the whole picture, due to the physical makeup of the various materials. All charts have one thing in common: that some areas within the given material space are not populated. Some of these areas are inaccessible for fundamental reasons that relate to fundamental physics, such as the size and packing of atoms and the nature of the forces that bind their atoms together. But other areas are empty even though, in principle, they are accessible and, if accessed, would enable novel design possibilities. It is precisely this clustering, in combination with the material (or more generally the performance) indices, that enables the rejection and ranking of various materials with respect to structural function. Material property charts have numerous applications. One is the checking and validation of data; another concerns the development of and identification of uses for new materials that can fill in empty regions (see Sec. 3.2.2). But most important of all, the charts form the basis for the previously presented procedure for materials selection.

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On a log-log plot, a contour with a constant value of a material index plots as a straight line. For example, the performance index for a light, stiff plate is given by (Table 3.6) M6 ¼ E1=3 =r Taking logs of both sides 1 log(M6 ) ¼ log(E)  log(r) 3 or log(E) ¼ 3 log(r) þ 3 log(M6 ) Consequently, each line in the family of straight parallel lines of slope 3, on a plot of log(E) vs log(r), corresponds to a fixed value of M6. All materials that lie on a line of constant E 1/3/r perform equally well as a light, stiff panel, whereas those above the line perform better (are lighter for a specified stiffness) than those below the line. As can be seen in Fig. 3.11, Mg-alloys, Al-alloys, Ti-alloys, steels, and W-alloys all fall near a line with the same value of E/r (the index for displacement limited tie rod subjected to tensile loading; see the heavy solid line). This raises the question of why aluminum is used in the manufacture of airplanes and not steel. This is immediately obvious when one considers the appropriate material index for plates/shells subjected to bending (i.e., E 1/3/r). This is plotted as a dashed line of slope 3, passing through the aluminum bubble in Fig. 3.11. Aluminum has a much higher value of E 1/3/r than steel and therefore has superior performance for a light, stiff plate. Other attribute limits (e.g., horizontal or vertical lines) can be added to narrow the associated search window (e.g., E . 10 GPa or r . 2 Mg/m3), thereby producing a short list of candidate materials. Material property charts will be used in the next section for two case studies to illustrate the utility of these charts.

3.2.4 MATERIAL CHOICES FOR CASE STUDIES 3.2.4.1

CASE STUDY 1B: BLADES

Returning to our first case study related to jet engine blades (fan, compressor, or turbine), one can quickly identify the various materials that will be best for resisting bending loads. As indicated in Eq. (3.10), the appropriate material index is M2 

ðfB sf Þ2=3 r

(3:24)

for temperatures at which time-dependent behavior can be ignored. This index can also be used to indicate which material would be useful for each region of

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

TABLE 3.6

231

MATERIAL INDICES StiffnessLimited Design

StrengthLimited Design

VibrationLimited Design

E/r

sf =r

E/r

Stiffness, length, shape specified; section area free

G 1/2/r

sf

Stiffness, length, outer radius specified; wall thickness free

G/r

sf =r

Stiffness, length, wall thickness specified; outer radius free

G 1/3/r

sf

E 1/2/r

sf

Function and Constraints Tie (tensile strut) Stiffness, length specified; section area free Shaft (loaded in torsion)

2=3

=r

1=2

=r

2=3

=r

Beam (loaded in bending) Stiffness, length, shape specified; section area free

Stiffness, length, height specified; width free E/r All dimensions specified —

sf =r —

Stiffness, length, width specified; height free E 1/3/r

sf

1=2

E 1/2/r — E/r

=r

Column (compression strut, failure by elastic buckling) Buckling load, length, shape specified; section area free

E 1/2/r

sf =r

E 1/3/r

sf

Panel (flat plate, loaded in bending) Stiffness, length, width specified; thickness free All dimensions specified

1=2

—

=r

—

Plate (flat plate, compressed in-plane, buckling failure) Collapse load, length and width specified; thickness free

1=2

E 1/3/r

sf

E/r

sf =r

E/(1 – v)r

sf =r

=r

Cylinder with internal pressure Elastic distortion, pressure and radius specified; wall thickness free Spherical shell with internal pressure Elastic distortion, pressure and radius specified; wall thickness free

E 1/3/r E/r

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Fig. 3.12

Flexural strength vs maximum service temperature.

the engine if plotted against maximum service temperature (see Fig. 3.12). For example, the fan section of a jet engine experiences temperatures up to 2908C, with the compressor section experiencing temperatures up to 7008C and the turbine experiencing up to 16008C. Because the actual material index M2 is plotted as the ordinate in Fig. 3.12, a horizontal line becomes our selection guideline. Therefore, considering the bending strength alone, it becomes immediately obvious which materials could be the best choice for low-mass blades: for fan blades, Al, Ti, and polymer matrix composites (PMCs); for compressor blades, Ti-6-4, Ti-6-2-4-6, B/Al, and beryllium; and for turbine blades, Inconel 718, Waspaloy, and SiC/SiC. Note that beryllium would be eliminated for toxicity reasons rather than mechanical performance. (This elimination would occur in the research stage of the selection process in Fig. 3.4.) These are but a few of our choices (depending on the required value of the material index) that have been substantiated over the years through direct experience. It is also clear from Fig. 3.12 that if the choice is limited, a priori, to metals, then as temperature increases (as the application region moves toward the rear of engine), the material performance index must be lowered. This indicates that the mass of higher temperature blades must be greater than that of lower temperature blades for a similar duty. However, if the design space is opened up to include nonmetals (particularly ceramics), the index can be maintained (in fact, doubled from 10 to 20) even at elevated temperatures. It is evident why significant investment is being made to develop composite alternatives to enable significantly lighter weight components for elevated temperatures. .

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233

Bending stress is just one type of loading required to be sustained by a given blade design (fan, compressor, or turbine); the others are axial tension (due to centrifugal loads) and vibrational loadings (due to rotational motion and fluctuating gas flows through the engine). As shown previously, to minimize mass (or weight) of a given blade requires that a material with a high specific strength (sf/r) in tension and a high value of E 1/2/r be selected to maximize its fundamental vibrational frequency. Note any weight savings in blades has a multiplying effect on other components, such as significant reduction in compressor and turbine disks as well as shaft loads. To assist in making the trade between several constraints, the CES Selector software package from Granta Design can be used [7]. This software enables engineers to depict graphically the various material indices and overlay them to arrive at a subset of candidate materials that can be investigated further. For a discussion on how to deal with multiple constraints and objectives, the reader is referred to Chapter 9 of Ashby [4]. Figure 3.13 shows the subset of materials after specifying values for the three material indices given tensile loading, bending, and vibration. The values assumed were: M1 ¼ ðsf =rÞ  60, M2 ¼ ðfB sf Þ2=3 =r  10, and M3 ¼ EfB =r  10,000. Note that fB ¼ 1 in both M2 and M3. Figure 3.13a shows the subset of materials when each index is treated independently, whereas Fig. 3.13b shows the corresponding results when interactions between indices are accounted for; that is, if a given material is eliminated due to failure to meet any one criterion, then it is removed in all three property chart spaces whether or not it fails in that particular space. Comparing Fig. 3.13a to Fig. 3.13b, it is immediately obvious that by considering all loading criteria at once (i.e., Fig. 3.13b) leads to a significantly smaller subset of workable material systems for the various temperature ranges or blade types. This assumes the value of the material index (requirement) would be the same in all three temperature regimes. It is interesting to note that Ti-6-4, Ti-6-2-4-6, Inconel 718, and Waspaloy are equal with respect to specific stiffness, but titaniums are significantly better in both bending and specific strength, thus substantiating their use in fan and compressor blade applications. A key factor yet to be considered is inelastic effects that come into play at elevated temperatures; this will be briefly considered in Sec. 3.3. 3.2.4.2

CASE STUDY: PRESSURE VESSELS

The storage of liquid hydrogen in a lightweight tank provides significant challenges. The low density of the hydrogen fuel results in the need for a larger volume storage vessel relative to other fuels. Mechanical tank loads are derived from 1) the difference in pressure within the tank and the ambient conditions, 2) fuel weight, 3) vehicle acceleration loads, 4) fuel slosh due to aircraft maneuvers, and 5) the weight of the tank system and its supports. Fuel slosh is bound to be encountered as the aircraft maneuvers or as it encounters air turbulence during the flight. Furthermore, the internal tank pressure must be

234

S. M. ARNOLD ET AL.

a)

Fig. 3.13 Interaction between three material indices for a blade subjected to tension, bending, and vibration.

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

b)

Fig. 3.13

(Continued).

235

236

S. M. ARNOLD ET AL.

maintained at a constant absolute pressure to maintain the hydrogen in a liquid state. The tank weight increases with increasing operating pressure. Typically, gaseous hydrogen is used as the pressurant for liquid hydrogen. Two important criteria for tank wall design include materials selection and wall architecture, which are not mutually exclusive. Many options are available, and Mital et al. [8] discuss their advantages and disadvantages and provide guidelines for choosing an optimum system. Here the discussion will be limited to selection of the tank wall material, based on mechanical loading alone to illustrate the selection process. Clearly, it is desirable to use materials that possess high strength, high fracture toughness, and high stiffness, as well as low density and low permeation to liquid and gaseous hydrogen; however, no single material provides all these attributes simultaneously. Consequently, material performance indices associated with the functional requirements, such as those given in Table 3.5, must be used to identify the best material candidates for tank wall construction. Among the material parameters, strength and density tend to dominate the design criteria, as the mass of the tank is inversely related to specific strength [see Eq. (3.23)]. Figure 3.14 shows strength vs mass density for various engineering materials. In this case, materials in the upper left corner are preferable. Composite materials exhibit high specific strength relative to metals and are well suited to aerospace

Fig. 3.14

Strength vs mass density for various engineering materials as per Ashby (2005) [4].

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

Fig. 3.15

237

Fracture toughness vs elastic limit.

applications; continuous fiber reinforced polymer (CFRP) composites provide the highest strength yet lightest choice. However, the use of CFRP composite materials most likely will involve higher initial manufacturing costs, and their permeability to hydrogen is a potential complication. As shown in Fig. 3.14, the materials that have the sufficient strength and acceptably low densities are PMCs, ceramic matrix composites (CMCs), and metallic materials. Ceramic materials offer high specific strength, but due to their low fracture toughness (see Fig. 3.15) are not viable for a tank wall material. A potential lower cost alternative to CFRP may be discontinuous reinforced metallic composites (DRX), specifically discontinuous reinforced aluminum (DRA) as described by Miracle [9]. DRAs are essentially isotropic and can be manufactured using less-expensive techniques such as casting. DRX materials have the added benefit of extremely low (if not negligible) hydrogen gas permeability issues typically associated with PMC systems. However, during a detailed study, Arnold et al. [10] showed that minimum gage thickness requirements kept DRA materials from outperforming PMC materials for this application. Two key material performance indices that are generally applicable to the design of high-pressure vessels but may also be applicable to low-pressure cryogenic storage tanks were described previously. These indices are yield-before break, M4 ¼ KIc/sf , and leak-before-break, M5 ¼ K2Ic/sf , and are illustrated in Fig. 3.15. Using the first index ensures that the stress required to propagate a critical

238

a)

S. M. ARNOLD ET AL.

b)

Fig. 3.16 Illustrations of the impact of allowing interaction of material indices for pressurized vessels. flaw is greater than that to yield the material, and the second ensures that the maximum pressure carried will result in a stable crack that will just enable a leak to occur before catastrophic failure. Figure 3.16 again clearly illustrates the difference between treating material indices as independent quantities (a) vs considering the interaction of the material indices during the selection process (b), the latter resulting in a reduction in potential material candidates. In Fig. 3.16, only a subset of the material universe database available within CES Selector (aerospace materials) is examined for illustrative purposes. Here we see that 2if the assumed s K values of the material indices are M1 ¼ ð rf Þ  20 and M5 ¼ sIcy  10, only a finite set of viable materials, for example Ti and PMCs, remains.

3.3 ADVANCED SELECTION 3.3.1 MATERIALS FOR HIGH-TEMPERATURE APPLICATIONS The response of real materials to thermal and mechanical stimuli can vary greatly depending on the magnitude and multiaxiality of loading and the magnitude of the homologous temperature (TH ¼ T/Tm), that is, the ratio of applied temperature [measured on an absolute scale (e.g., K)] to that of the melting point of a given

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239

material. For example, at room temperature, material response is typically time independent and either reversible (i.e., linear elastic) or irreversible (inelastic), depending on whether the yield stress of the material has been exceeded. Alternatively, when TH . 0.25, time-dependent behavior, both reversible and irreversible, can be commonly observed as illustrated schematically in Fig. 3.17, wherein strainrate sensitive, creep, relaxation, thermal recovery, dynamic recovery, and creep/ plasticity interaction response behavior are all shown. Other complex time- and path-dependent behavior, such as cyclic ratcheting, creep/fatigue interaction, and thermal mechanical fatigue, are also often observed depending on the magnitude and type of loading (e.g., thermal or mechanical) being applied [11]. Note that in the case of Ni-based superalloys, this ratio is significantly higher (TH  0.6–0.9) [12]. A prerequisite for meaningful assessment of component durability and life, and consequently design of structural components, is the ability to accurately predict the stress and strains occurring within a loaded structure composed of a given material. As constitutive material models (be they simple or complex) provide the required mathematical link between stress and strain, this necessitates the development and characterization of an appropriate constitutive model for any material before that material can be certified for use by a designer. Thus, constitutive models with varying levels of idealizations have been proposed and used, each with its own shortcomings/limitations. The most well-known and widely used constitutive relation/model is the generalized Hooke’s law

sij ¼ Cijkl ekl

ð3:25Þ

which describes multiaxial time-independent, linear (proportional) reversible material behavior. Up to now, only uniaxial Hooke’s law (s ¼ E e) has been used. Extension into the irreversible regime is accomplished by assuming an additive decomposition of the total strain tensor 1ij into three components, that is, a reversible mechanical eij (i.e., elastic/viscoelastic), an irreversible mechanism 1Iij (i.e., inelastic or viscoplastic), and a reversible thermal strain 1th ij component: 1ij ¼ eij þ 1Iij þ 1th ij

ð3:26Þ

Equation (3.26) can also be rearranged as follows: eij ¼ 1ij  1Iij  1th ij

ð3:27Þ

Substituting Eq. (3.27) into Eq. (3.25) gives a stress–strain relation that incorporates irreversible strains as well as reversible ones, that is

sij ¼ Cijkl ð1ij  1Iij  1th ij Þ

ð3:28Þ

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Fig. 3.17 Schematics showing representative hereditary material behavior at elevated temperature. Numerous models describing the evolution of the inelastic strain have been proposed in the literature. Table 3.7 contains a few representative examples with a brief description of their required material parameters. The well-known Norton–Bailey creep model [13] and its multiaxial generalization proposed by Odqvist [14] were followed by a coupled fully associative

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241

unified viscoelastoplastic model [15, 16] with sufficient generality to permit systematic introduction of multiple mechanisms for both viscoelastic (timedependent reversible) and viscoplastic (time-dependent irreversible) response components. This general, multimechanism, hereditary deformation model has been shown to accurately represent a wide spectrum of material responses under different loading conditions for the case of titanium alloys [17–20]. Examples include 1) rate-dependent (effective) material tangent stiffness during initial loading or any subsequent reversed loading; 2) pure transient response (e.g., in creep or relaxation) within the reversibility region; 3) an elastic behavior upon stress reversal, irrespective of the load level; and 4) other response features common to “unified viscoplastic” formulations (e.g., rate sensitivity, creep–plasticity interaction, thermal recovery, etc.). The following generalized anisotropic material behavior constitutive model results; see Table 3.7 for a definition of the associated flow and evolutionary laws:

s˙ ¼ Eð˙1  1˙ I Þ þ q˙

ð3:29Þ

All of this is said to emphasize that, although materials are often classified into specific idealized groups (e.g., linear elastic, plastic, viscoelastic, and viscoplastic) for mathematical convenience, nothing can compel real materials to behave according to these idealized models. Typically, when dealing with actual complex material behavior as represented in Fig. 3.17, it is not sufficient to describe this behavior with simple single-parameter relationships or point-wise values (e.g., E, n, sy, H, etc.) that have worked so nicely for materials selection at low homologous temperatures (e.g., typically at room temperature) where such complex path-dependent behavior is suppressed. And so what should we do when the application operates at elevated temperatures—throw up our hands and just select materials based on linear elastic behavior? Not necessarily, as there are simplified (albeit less accurate) approaches for high-temperature performance that allow some inelastic (e.g., creep) behavior to be accounted for and therefore to influence the materials selection process. Four such approaches are: 1) maximum service temperature (as discussed previously), 2) allowable stress, 3) design for creep, and 4) deformation mechanism maps. In this section, a brief description of the maximum service temperature and allowable stress approaches will be discussed. This will be followed by a more detailed description of how to account for creep behavior in the materials selection process. It is strongly suggested, however, that a rigorous nonlinear analysis with a proper constitutive model, such as the unified viscoplastic GVIPS model [15, 16, 18, 20], be conducted for the top material candidates to verify the validity/suitability of the chosen material for a given design application. 3.3.1.1 MAXIMUM AND MINIMUM SERVICE TEMPERATURES The simplest measures of tolerance to temperature are the maximum and minimum service temperatures, Tmax and Tmin. The former tells us the highest

242

TABLE 3.7

EXAMPLES OF CONSTITUTIVE MODELS AND THEIR REQUIRED MATERIAL PARAMETERS

Model

Idealized Material Behavior

Mathematical Representation

Ramberg–Osgood [21]

Elastic plastic

1I ¼ H

Norton–Bailey [13]

Uniaxial creep

1˙ I ¼ As n

Odqvist [14]

Multiaxial creep

1˙ Iij



¼

s sy

n

GVIPS Viscoelastoplastic Arnold et al. [15]; Saleeb et al. [17]; Saleeb and Arnold [18]

sy, H, n A, n A, n

BJ0m 2 Sij m ¼ 1=2ðn  1Þ and B ¼ 1=2ð3ðnþ1Þ=2 ÞA

where J20

Required Isotropic Material Parameters

¼ 1=2Sij Sij

with Sij ¼ sij  1=3skk dij h i  ðaÞ ðaÞ  ðaÞ ðaÞ 1 ðaÞ q˙ij ¼ Mijkl 1˙ kl  1˙ Ikl þ Mijkl hklrs qrs pffiffiffiffiffiffiffiffiffi Fn f ðF Þgij if F  0, ; f ðF Þ ¼ ; gðGðbÞ Þ ¼ HðbÞ ð1  GðbÞ ÞbðbÞ 1˙ Iij ¼ 0 otherwise, m 2 ðbÞ

ðbÞ

a˙ ij ¼ Qijkl 41˙ Ikl 

RðbÞ

3

½GðbÞ mðbÞ

HðbÞ ð1 

ðbÞ ðbÞ pffiffiffiffiffiffiffiffiffi pkl 5 if pkl  ðskl  akl Þ  0 GðbÞ ÞbðbÞ

where M P a¼1

ðaÞ

qij ,

aij ¼

N P b¼1

and 3 þ 5N irreversible constants k, n, m H(b), R(b), m(b), k(b), b(b)

ðbÞ

aij

ðbÞ

ðbÞ

gij ¼ Mijkl ðskl  akl Þ, pkl ¼ Mijkl aij

where N defines the number of viscoplastic mechanisms and M the number of viscoelastic

S. M. ARNOLD ET AL.

qij ¼

2 þ 2M reversible constants: Es, M(a), n, r(a)

 1  sij  aij Mijkl ðskl  akl Þ  1, 2k2ðbÞ

G¼

  1  ðbÞ  ðbÞ aij Mijkl akl 2 2kðbÞ ðaÞ

And all three fourth-order tensors, Eijkl, Mijkl , and hijkl , are taken to ðaÞ

ðaÞ

ðaÞ

be coaxial, that is, Eijkl ¼ Es Nijkl , Mijkl ¼ MðaÞ Nijkl , and hijkl ¼ ra Mijkl where Nijkl ¼

n 1 ðdik d jl þ dil d jk Þ dij dkl þ ð1 þ nÞð1  2nÞ ð1 þ nÞ

Finally, ðbÞ Qijkl

¼ HðbÞ



2 pffiffiffiffiffiffiffiffiffibðbÞ ðbÞ 4Zm  1 G

3 b ðbÞ ðbÞ 5 h i pffiffiffiffiffiffiffiffiffi pffiffiffiffiffiffiffiffiffi aij akl 2k2ðbÞ GðbÞ 1  f1  bg GðbÞ

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

F¼

Note Zm is the “generalized” inverse of Mijkl ; see Saleeb et al. [16] for further elaboration.

243

244

S. M. ARNOLD ET AL.

temperature at which the material can reasonably be used without oxidation, chemical change, or excessive deflection or creep becoming a problem (the continuous use temperature, or CUT, is a similar measure). The latter is the temperature below which the material becomes brittle or otherwise unsafe to use. These are empirical, with no universally accepted definitions; therefore, caution should be used to fully understand the underlying assumptions and implications. For example, the minimum service temperature for carbon steels is the ductile-to-brittle-transition temperature, a temperature below which the fracture toughness falls steeply. For elastomers it is about 0.8 Tg, where Tg is the glass temperature. Below Tg, elastomers cease to be rubbery and become hard and brittle. Examining Fig. 3.18 wherein strength vs maximum temperature is plotted, it is apparent that polymers and low melting metals like the alloys of zinc, magnesium, and aluminum offer useful strength at room temperature, but by 3008C they cease to be useful (as inelastic/softening behavior is exhibited). Indeed, few polymers have useful strength above 1358C. Titanium alloys and lowalloy steels can have useful strength up to 6008C; above this temperature, highalloy stainless steels and more complex superalloys based on nickel, iron, and cobalt are needed. The highest temperatures require refractory metals like tungsten or technical ceramics, such as silicon carbide (SiC) or alumina (Al2O3).

Fig. 3.18 Chart showing the strength (elastic limit) vs maximum service temperature of the material.

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

245

Fig. 3.19 Variation of modulus and threshold stress k as a function of temperature for Ti-6-4 material. In Fig. 3.12, the flexure strength vs maximum service temperature was used to identify potential candidate materials for three types of gas turbine engine blades: fan blades, T , 3008C; compressor blades, T , 7008C; and turbine blades that operate above 7008C. In case study 1b, Ti-6-4 was identified as a potential fan or compressor blade material whose maximum service temperature was given a range of approximately 350–4208C (see Fig. 3.13). However, in a recent study [20], the complete time-dependent and rate-dependent temperature regimes for this material were experimentally mapped out. Results demonstrated that this maximum service temperature is too aggressive and should really be limited to 3008C, because only below this temperature can it be assumed that no rate dependence and only a minor amount of time dependence are present (see Fig. 3.19). In Fig. 3.19, the deformation response of Ti-6-4 over a wide range of temperatures is documented. Here, the variation of the moduli and experimentally determined threshold stress k (the stress that truly delineates between the reversible and irreversible strain regimes) is plotted as a function of temperature [19]. The modulus ES represents the “infinitely slow” modulus (i.e., the elastic modulus of the material if it was loaded at an infinitely slow strain rate), whereas the modulus ED represents the “dynamic modulus,” which is the modulus of the material if it was loaded at an “infinitely fast” (e.g., very high, 1  1023 s21) rate (i.e., all time dependence is locked into this modulus). As shown in Fig. 3.19, at elevated temperatures there is a significant difference between the two modulus values, indicating that even in the so-called elastic range there is significant time

246

S. M. ARNOLD ET AL.

dependence (reflected as rate-dependent apparent moduli). Below a temperature of about 3008C, the two moduli are approximately equal, indicating that the response of the Ti-6-4 material is rate independent but not necessarily time independent. Above 3008C, the response is rate dependent and time dependent even below the threshold stress. To appreciate the practical significance of this fact, the operating temperatures typically encountered in aircraft engines are also noted below the horizontal axis in Fig. 3.19, as well as the occasional, higher-temperature regime encountered during overtemperature maneuvers. Clearly, even when one is within the typical engine design range and expecting the material response to be reversible, or elastic, the material behavior (at least in the case of Ti-6-4) would in fact be rate dependent and would generate an additional strain of s/Es over time, where s is the current applied stress that is less than or equal to the threshold stress k denoted by the square symbol line. Consequently, at temperatures greater than 3008C and stresses below k, if classical elasticity methods were used in the design, then this rate dependence (i.e., viscoelastic response) would not be captured. Furthermore, at stresses above k, the material response is viscoplastic due to the rate and time dependence, and using the classical methods of plasticity in analysis and design would also not be accurate. Figure 3.19 further indicates that the value of the threshold stress k decreases significantly as a function of temperature. More important, Fig. 3.19 demonstrates that, at elevated temperatures, the threshold stress (k) is significantly below the traditional “proportional limit” or apparent yield stress of the material (obtained at a total strain rate of 0.001/s) indicated by the dashed line (and  symbols). However, if one obtains the proportional limit from a tensile test conducted at 1.0  1026/s strain rate, the values would correspond very closely to those of the threshold stress. This comparison then leads to the conclusion that irreversible material behavior can take place at stress levels well below those considered using traditional methods for the analysis of metals. If classical design methods were used, the material response below the proportional limit would be assumed to be fully reversible, and only the response after the proportional limit would be assumed to be irreversible. However, as is shown in the figure, that assumption could lead to significant inaccuracies in predictions of the material response. All of this is said to reinforce the fact that, although simplifications can be made to aid the materials selection/rejection process, it remains extremely important when dealing with material behavior at elevated temperatures that significant care be taken to fully understand the various candidate materials that one will be recommending at the end of the process.

3.3.2 DESIGN FOR CREEP How are materials chosen to avoid failure by creep at high temperatures (Fig. 3.20)? Traditionally, materials selection has been based largely on experience. Polymers can be used at room temperature, but—with only a few exceptions—not above 1008C. The most creep-resistant of aluminum alloys are good to about 2008C,

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

a)

b)

c)

d)

247

Fig. 3.20 Creep is important in four classes of design: a) displacement-limited, b) failure limited, c) relaxation-limited, d) buckling limited. titanium alloys to 6008C, stainless steels to 8508C, and so on. But optimal selection requires much more than this. The choice depends not only on material properties, but also on the mode of loading (tension, bending, torsion, internal pressure), on the failure criterion (excessive deflection, fracture, relaxation of stress, buckling, etc.), and on the optimization objective (minimizing weight or cost or maximizing life). A designer who is not a specialist on creep has no easy way to identify the subset of materials best suited to the project’s needs or to predict the ways in which a change in the design might influence the choice. In design against creep, one seeks the material and the shape that will carry the design loads without failure for the design life at the design temperature. The meaning of failure depends on the application. Here four types of failure are distinguished and illustrated in Fig. 3.20: 1) displacement-limited applications, in which precise dimensions or small clearances must be maintained (as in the disks and blades of turbines, see Fig. 3.20a), when design is based on creep rates 1˙ or displacement rates d˙ ; 2) rupture-limited applications, in which dimensional tolerance is relatively unimportant but fracture must be avoided (as in pressure-piping, see Fig. 3.20b), when design is based on time-to-failure tf ; 3) stress-relaxation-limited applications, in which an initial tension relaxes with time (as in the pretensioning of cables or bolts, see Fig. 3.20c), when design is based on a characteristic relaxation time tr; and 4) buckling-limited applications, in which slender columns or panels carry compressive loads (as in the upper wing skin of an aircraft, or an externally pressurized tube, see Fig. 3.20d), when design is based on critical time-to-instability tb. To tackle any of these situations, constitutive equations, as previously discussed, are required that relate the strain rate 1˙ or time-to-failure tf for a material to the stress s and temperature T to which it is exposed.

248

3.3.2.1

S. M. ARNOLD ET AL.

CONSTITUTIVE EQUATION FOR CREEP DEFLECTION

When a material is loaded at constant stress and homologous temperature TH, above approximately 0.25 it will typically experience creep, as schematically depicted in Fig. 3.21. Creep response typically involves a primary extension stage, followed by a steady-state creep stage, and ends with an accelerating tertiary stage. The important parameters are marked: the steady-state creep rate 1˙ ss , the time to fracture tf, and the creep ductility 1f. The characteristics of the curve and the way it changes with temperature and stress are described mathematically by a constitutive equation [22–27]. Many constitutive equations for creep rate have been suggested, some purely empirical and some science based; most are a mix of the two. Those most widely used in engineering design when deflection is important relate the steady-state strain rate 1˙ ss to the tensile stress s and the temperature T, thus   Q ð3:30Þ 1˙ ss ¼ Af ðsÞ exp  RT where A Q R f(s)

¼ a kinetic constant ¼ an activation energy ¼ the gas constant ¼ a function of stress s

The function f(s) can be approximated, over restricted ranges of stress, by the Norton–Bailey power law given in Table 3.7 and slightly rewritten here:  n  n s Q s ¼ 1˙ o exp  ð3:31Þ 1˙ ss ¼ A so RT so where the constant A, the activation energy Q, the exponent n, and the characteristic strength constant so are material properties. Considerable experience has accumulated in the use of Norton’s Law, which has the appeal that it allows analytical solutions to a wide range of Fig. 3.21 Typical creep curve showing primary, secondary, and tertiary stages and the quantities 1ss, 1f, tf.

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

249

engineering problems (e.g., Finnie and Heller [22], Hult [23], Penny and Marriott [24]). For this reason, it will be used here even though, from a scientific point of view, it lacks a completely respectable pedigree, nor is it capable of accounting for creep/plasticity effects that more sophisticated unified viscoplastic models were developed to handle. 3.3.2.2

CONSTITUTIVE EQUATION FOR CREEP FRACTURE

When fracture rather than deflection is design limiting, creep is characterized instead by the time to fracture tf. It too can be described by a constitutive equation with features like those of Eq. (3.31). Here, again, a power law gives an adequate description over a restricted range of s and T: !q   !q Qf s s ¼ tf0 exp ð3:32Þ tf ¼ B RT sf sf with its own values of kinetic constants B, activation energy Qf, exponent q, and characteristic strength sf. 3.3.2.3

CONSTITUTIVE EQUATION FOR RELAXATION

Relaxation requires a constitutive equation that combines creep (inelastic) and elastic responses when total strain rate is held constant at zero. Consequently, from Eq. (3.29) and assuming tension, it takes the form

s˙ ¼ E˙1Iss

ð3:33Þ

where E ¼ Young’s modulus 1˙ ss ¼ given by Eq. (3.31) s˙ ¼ the rate of change of stress with time It is important to realize, however, that although the form of the inelastic strain is the same as Eq. (3.33), the material constants take on completely different values because the relaxation spectrum is typically significantly shorter than that of steady-state creep. For the bending of a beam (as in Fig. 3.22), the equation becomes instead F˙ ˙ ¼ dc ð3:34Þ S where F˙ ¼ the rate of change of force F S ¼ the bending stiffness d˙ c ¼ the creep deflection rate of the beam Similar expressions describe torsion and compression.

250

S. M. ARNOLD ET AL.

a) (a) At room temperature, a plastic hinge forms where the bending moment is highest.

b) (b) At high temperature, creep plasticity is distributed.

c) (c) Creep fracture starts where the local tensile stress is highest.

Fig. 3.22

3.3.2.4

A cantilever beam loaded with an end load F.

MATERIALS SELECTION IN THE CREEP REGIME

The objective here is to derive indices that parallel those used for roomtemperature design [28] but are now tailored for design when creep takes place. The immediate difficulty is that the “strength” is now no longer a fixed material property but depends on temperature and on the strain rate. The following treatment is kept as brief as possible. More details can be found in Abel and Ashby [29] and elsewhere. 3.3.2.5

DEFLECTION-LIMITED DESIGN

Consider first the simple case of a tensile member, a tie rod, of minimum weight, designed to carry a load F for a lifetime t without deflecting more than d at a temperature T. If the tie has length L, the steady strain rate must not exceed 1˙ ¼

d Lt

ð3:35Þ

Inserting this into the constitutive Eq. (3.32) for tensile creep and inverting gives   F d 1=n s ¼ ¼ s0 ð3:36Þ 1˙ 0 Lt A The objective is to minimize the mass of the tie. Solving for A and substituting this into the equation m ¼ ALr for the mass gives m¼

LrF sD

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

251

with 

sD ¼ s0

d 1˙ 0 Lt

1=n ð3:37Þ

Thus, the lightest tie that meets the constraints of F, T, t, and s is that made of the material with the largest value of 

sD M6 ¼ r

 ð3:38Þ

This has the same form as the preceding elastic material index for a tie rod: M1 ¼

sy r

ð3:39Þ

with sy replaced by sD, defined previously. Consequently, it contains both temperature and time. The analysis of beams, shafts, and pressure vessels (and the like) is a little more complex but follows the same pattern. Consider, as an illustration, the cantilever beam of Fig. 3.22b carrying a load F, but now at a temperature such that it creeps. The objective, as before, is to make the beam as light as possible; the constraints (again as before) are that its length L and the proportions of its cross section are fixed and that it must support the load F for a time t at temperature T without deflecting more than d. The design specification constrains the deflection rate 1˙ : it must not exceed d/t. The deflection rate 1˙ for a cantilever beam with end load F, creeping according to the constitutive Eq. (3.32), is 1˙ ¼



n   2 4FL 2n þ 1 1 ¼ 1˙ 0 L2 2nþ1 s0 ðn þ 2Þ 2n bh n

ð3:40Þ

See Appendix B in Arnold et al. [30] for a complete derivation. Eliminating b and h between the equations m ¼ ALr and Eq. (3.40), using b ¼ ah and A ¼ ah 2, gives 3nþ3

nþ1



m ¼ rL 3nþ1 a 3nþ1

 2n  2n þ 1 4FL 3nþ1 : sD 2n

ð3:41Þ

with sD, herein called the design strength, given by 

sD ¼ s0

 n þ 2 d˙ n 2 L˙10 1

ð3:42Þ

252

S. M. ARNOLD ET AL.

Equation (3.41) looks messy, but it is not as bad as it seems. For example, the fully plastic limit can be found by setting n ¼ 1, thereby simplifying Eq. (3.41) to

4FL 2=3 ð3:43Þ m ¼ rLa1=3 sD n o2=3 if sD is which is identical to the mass of the plastic beam m ¼ rLa1=3 4FL sy replaced by sy. The design strength in creep, then, plays the same role as the yield strength in room-temperature plasticity. In fact, Eq. (3.43) is a good approximation to the more complex result of Eq. (3.41) over the entire range of values of the exponent n (3 , n , 20) normally encountered in metals and alloys. Inspection shows that the mass is minimized by maximizing the index M2 given earlier, 2=3

M2 ¼ sy =r, with sy replaced by sD (remember that sD contains temperature and deflection rate). A parallel calculation for a panel (flat plate in bending) gives the equation 2=3 M7 ¼ sy =r, again with sy replaced by sD, appropriately defined. Similar expressions are derived for torsion and for internal pressure; see Appendix B of Arnold et al. [30]. Selection with the objective of minimizing cost rather than weight leads to results identical to those when r is replaced by Cmr, where Cm is the material cost per kg, and the objective of minimizing energy content is achieved by replacing this with qmr, where qm is the energy content per kg. 3.3.2.6

FRACTURE-LIMITED DESIGN

Consider next an application in which fracture, not deflection-rate, is design limiting (Fig. 3.20b). For a bending beam, the largest stresses appear—and creep fracture starts—in the outer fibers at the place where the bending moment M is greatest (Fig. 3.22c). The time t to the onset of failure of the cantilever, using the constitutive relation of Eq. (3.32), is !q 2q þ 1 4FL : ð3:44Þ t ¼ t fo 2q s fo bh2 Writing b ¼ ah, solving for the area A ¼ bh, and substituting in m ¼ ALr gives 1=3

m ¼ r La

   2q þ 1 4FL 2=3 sD 2q

ð3:45Þ

with the design strength 

sD ¼ s fo

t fo t

1 q

ð3:46Þ

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

253

The parallel with deflection-limited design is obvious, and again the result reduces to that for full plasticity in the limit q ¼ 1. The mass is minimized, as 2=3

before, by maximizing the index of equation M7 ¼ sy =r, with sy replaced by this new, fracture-related sD that depends on design life t and on temperature T. Analogous calculations for ties and panels give equations M1 ¼ sy =r and 1=3

M3 ¼ sy =r; only the definition of sD is different. Similar expressions describe torsion and internal pressure (see Appendix B.1 of Arnold et al. [30]) and are modified for cost or energy content by replacing r with Cmr or qmr as before. 3.3.2.7

RELAXATION-LIMITED DESIGN

A tensile cable, or a bolt, that is pretensioned to provide a bearing or clamping force F at an elevated temperature relaxes with time by creep. The calculation is a standard one; elastic strain s/E is replaced over time by creep. The total change in strain in the cable or bolt is zero because its ends are fixed. The governing equation for the stress in the component was given earlier as Eq. (3.33). Integrating this with the boundary condition s ¼ si at t ¼ 0 gives s n1 s n1 E˙10 t o o  ¼ ðn  1 Þ ð3:47Þ s si s0 where

si ¼ the stress to which the cable or bolt was originally tightened s ¼ the stress to which it has relaxed in time t In this case a constraint is specified by defining a characteristic relaxation time tr as the time required for the stress to relax to a specified fraction of its initial value. Inverting Eq. (3.47) gives "

 n1 # s sno tr ¼ 1  ðn  1ÞE˙10 sn1 si

ð3:48Þ

The quantity in the square bracket can be thought of as merely a scaling factor. Consequently, the result is independent of the choice of s/si. For illustrative purposes, the ratio s/si is taken to be 0.5, and for n . 3, the term in square brackets is then close to unity; thus, for simplicity, it shall be neglected. The design specifies the minimum bearing or clamping load F. Writing s ¼ F/A and substituting for A in equation m ¼ ALr gives the mass of the cable or bolt that will safely provide a clamping load greater than F for a life tr: m¼

LrF sD

ð3:49Þ

254

S. M. ARNOLD ET AL.

with 

sD ¼ so

so ðn  1ÞE˙10 tr



1 n1

ð3:50Þ

The mass is minimized by selecting the material with the greatest value of sD/r; that is, the material index is once more that of equation M1 ¼ sy =r replaced by this new sD. Springs also relax their tension with time. Most are loaded in bending, when the constitutive behavior is that of Eq. (3.34). Taking a beam of length L as an example, the stiffness S can be written as S¼

C2 EI L3

ð3:51Þ

where I ¼ the second moment of its area C2 ¼ a constant Integrating with the boundary conditions F ¼ Fi at t ¼ 0 gives a result with the form of Eq. (3.47). Proceeding as before, the minimum weight design of a leaf spring (or any spring loaded in bending) that must not relax its restoring force 2=3

in time tr at temperature T can be found using the index M7, M7 ¼ sy =r, with sD (when n . 3) given by  1 so n1 sD so ð3:52Þ E 1˙ o tr Abel and Ashby [29] list results for other modes of loading. Similar calcu1=3

lations for ties and panels give equations M1 ¼ sy =r and M8 ¼ sy =r again; only the definition of sD is different. The earlier adaptations to cost or energy apply here as well.

3.3.2.8

SELECTION PROCEDURE

Expressions for the indices M and the associated design strengths sD are summarized in Appendix A of Arnold et al. [30]. The close parallel between these results 2=3

and those for room-temperature plasticity (equations M1 ¼ sy =r, M7 ¼ sy =r, 1=3

and M8 ¼ sy =r) suggests a selection procedure. The design temperature T and acceptable deflection rate d˙ , life t, or relaxation time tr are identified. Using this information, values for the appropriate sD are calculated from a database of creep properties for materials (it is necessary to capture the value of sD at the value sy to allow for the change of deformation mechanism to yielding at low

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

255

temperatures). These are used to construct a chart of log(sD) against log(r). This is the creep equivalent of Fig. 3.14 but is specific to the temperature, deflection rate, or life required by the design because these appear in the definition of sD. The indices M1, M7, and M8 can be plotted onto it, allowing optimum selection for each application. This is best understood through examples. Those in the next section are deliberately simplified to avoid unnecessary digression. The method remains the same when the complexity is restored. 3.3.2.9

CASE STUDY

The following example illustrates the selection of materials for structures loaded at elevated temperatures and that are limited by deflection. Components limited by fracture or by stress relaxation can be handled in a similar manner. Remember that many other considerations enter the selection of materials for hightemperature use as well; for example, resistance to oxidation, resistance to thermal shock, and so on. Here the selection given the initial short list of candidates is examined, for which these additional considerations would then be investigated. 3.3.2.10

CASE STUDY 1C: FAN AND TURBINE BLADES FOR GAS TURBINES

A rotating blade of an aircraft turbine is self-loaded (Fig. 3.20a and Fig. 3.23); the centrifugal force caused by its own mass is much larger than that exerted by the gases that propel it. Adiabatic compression of the intake air can heat the compressor fan blade to 4008C or more. The dominant mode of steady loading, therefore, is tensile and proportional (for fixed-blade proportions) to the density of the blade material. It could, then, be anticipated that the appropriate index is that for tensile loading, M1 of equation M1 ¼ sy =r, with the design strength for tensile loading, sD, replacing sy; see Eq. (3.38). More detailed analyses add complexity but confirm this result [29]. The turbine blade is loaded in the same way but is hotter: designers would like to go to 10008C. The task is to select materials to maximize the safe angular velocity of the compressor or turbine blade, designed to operate for a life t of 1000 h without extending by more than d, which is required to be 1.0% of its length, for each of these temperatures, and at the same time to minimize the weight. Fig. 3.23

Schematic of a turbine blade.

256

S. M. ARNOLD ET AL.

The profile and section are determined by the blade design; neither is free. The mass is minimized by minimizing M9 ¼

1 r

ð3:53Þ

Figures 3.24 through 3.26 show sD creep strength, calculated for 400, 1000, and 15008C, respectively, with a value of d˙ =L and T corresponding to the design specification, plotted against density r. Selection lines plotting the appropriate indices are shown, and the corresponding potential materials selections are listed in Tables 3.8 through 3.10. The sweet spot (desired location) for materials selection is the upper left-hand corner of the figures.

3.3.3 ENGINEERED MATERIALS Engineered or “hybrid” materials are combinations of two or more materials, or of materials and space, assembled in such a way as to have attributes not offered by any one material alone (Fig. 3.27). Particulate and fiber-reinforced composites are examples of one type of hybrid, but there are many others: sandwich structures,

Fig. 3.24

Selection chart for deflection-limited design at 40088C for a life of 1000 h.

MATERIALS SELECTION FOR AEROSPACE SYSTEMS

257

Fig. 3.25

Selection chart for deflection-limited design at 100088C for a life of 1000 h.

Fig. 3.26

Selection chart for deflection-limited design at 150088C for a life of 1000 h.

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S. M. ARNOLD ET AL.

TABLE 3.8

CANDIDATE MATERIALS FOR DEFLECTION-LIMITED BLADE DESIGN AT 4008C FOR A LIFE OF 1000 h

Material

Comment

Ti-matrix composites (Ti/Si/Cf)

The ultimate in performance but very expensive.

Ti alloys (e.g., Ti 685, 6242) High creep strength and low density make these the best choice. Ni-based superalloys (e.g., IN738)

Excellent creep strength, but both indices M1 and M4 are inferior to Ti alloys.

Iron-based superalloys, stainless steels

Not as good as Ni-based alloys but cheaper.

Ceramics and CMCs

Excellent values of M1 and M4, but brittleness is a problem.

TABLE 3.9

CANDIDATE MATERIALS FOR DEFLECTION-LIMITED BLADE DESIGN AT 10008C FOR A LIFE OF 1000 h

Material

Comment

Ni-based superalloys

Among metallic alloys, these have the highest values of M1 but are heavy.

Refractory metals

W–Re alloys offer high M1 but are very heavy.

Oxide ceramics and silicon-based ceramics (CMCs)

Large weight saving is possible if design can accommodate brittleness.

TABLE 3.10

CANDIDATE MATERIALS FOR DEFLECTION-LIMITED BLADE DESIGN AT 15008C FOR A LIFE OF 1000 h

Material

Comment

Oxide and silicon-based ceramics (SiC, Al203)

Offer high values of M1 and M4, but design must accommodate brittleness.

Refractory metals

Tungsten–rhenium creep resistant but heavy.

Ceramic composites (SiC-SiC; carbon–carbon)

Potential for gains in performance, but brittleness and chemical stability require attention.

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Fig. 3.27 Hybrid materials combine the properties of two (or more) monolithic materials or one material and space. lattice structures, segmented structures, and more. Hybrids and composites differ from bulk materials in providing an opportunity to tailor the properties of the “material” to the application, in effect designing the material while designing the component. It requires the ability to synthesize properties of candidate hybrid materials during the design process. This in turn requires the engineer to choose the components of the hybrid, their configuration, and their relative fraction. The new variables expand the design space, allowing the creation of new materials with specific property profiles. Consequently, one can view these new materials as ministructures because, internal to the material, the individual constituents will redistribute the local stress and strain fields to carry the globally applied fields most effectively. But how is one to compare a hybrid—a sandwich structure, for example—with monolithic materials such as polycarbonate or titanium? To do this one must think of the sandwich not only as a hybrid with faces of one material bonded to a core of another, but as a material in its own right, with its own set of effective properties; it is these that allow the comparison.

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The approach adopted here is one of breadth rather than precision. The aim is to assemble methods to allow the properties of alternative hybrids to be scanned and compared with those of monolithic materials, seeking those that best meet a given set of design requirements. Once materials and configuration have been chosen, standard methods—optimization routines, finite-element analyses—can be used to refine them. But what the standard methods are not good at is the quick scan of alternative combinations. That is where the approximate methods discussed here pay off. The broad classes of hybrid can be considered in this way: .

Cellular structures are combinations of material and space giving precise control of density, stiffness, strength, and thermal conductivity.

.

Sandwich structures have outer faces of one material supported by a core of another, usually a low-density material—a configuration that can offer a flexural stiffness per unit weight that is greater than that offered by either component alone.

.

Composites combine two solid components: one, the reinforcement, contained in the other, the matrix.

.

Coated materials enable the surface properties of the bulk material to be enhanced by the addition of a thin surface layer.

Continuum and micromechanical models can be used to estimate the equivalent properties of each configuration. These can then be plotted on materials selection charts, which become comparison tools for exploring unique combinations of configuration and material. Examples follow later in this section. 3.3.3.1

HOLES IN MATERIAL-PROPERTY SPACE

As previously described, material properties can be “mapped” as material property charts; Fig. 3.28 is an example. For an up-to-date survey, see Ashby et al. [31]. All the charts have one thing in common: parts of them are populated with materials, but other parts are not. Some parts of the holes are inaccessible for fundamental reasons that relate to the size of atoms and the nature of the forces that bind them together. But others are empty even though, in principle, they could be filled. 3.3.3.2

CRITERIA OF EXCELLENCE

Is anything to be gained by developing materials (or material combinations) that lie in these holes? To answer this, criteria of excellence are needed to assess the merit of any given hybrid. These are provided by the material indices, described previously. If a possible hybrid has a value of any one of these that exceeds those of existing materials, it achieves the goal. The axes of Fig. 3.28 are Young’s modulus E and density r. The property combinations E/r, E 1/2/r, and E 1/3/r are measures of the excellence of material

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Fig. 3.28 Material property chart showing modulus-density space, with contours of specific modulus E/r.

indices for selecting materials for light, stiff structures. A grid of lines of one index, E/r, is plotted on the figure. The arrow lies normal to the index lines. If the filled areas can be expanded in the direction of the arrow (i.e., to greater values of E/r), the materials so created will enable lighter, stiffer structures to be made. The arrow thus defines a vector for material development. One approach to filling holes—the long-established one—is that of developing new metal alloys, new polymer chemistries, and new compositions of glass and ceramic to create monolithic materials that expand the populated areas of the property charts. But developing new materials can be expensive and uncertain, and the gains tend to be incremental rather than step-like. An alternative is to combine two or more existing materials to allow a superposition of their properties—in short, to create hybrids. The great success of carbon- and glass-fiber-reinforced composites at one extreme and of foamed materials at another (hybrids of material and space) in filling previously empty areas of the material property charts is encouragement to explore ways in which such hybrids can be designed. When is a hybrid a material? There is a certain duality about the way in which hybrids are thought about and discussed. Some, like filled polymers, composites, or wood are treated as materials, each characterized by its own set of macroscopic

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material properties. Others, like galvanized steel, are seen as one material (steel) to which a coating of a second (zinc) has been applied, even though it could be regarded as a new material with the strength of steel but the surface properties of zinc. Sandwich panels illustrate the duality, sometimes viewed as two sheets of face material separated by a core material and sometimes (to allow comparison with monolithic materials) as a material with its own density, axial and flexural stiffness and strength, thermal conductivity, expansion coefficient, and so on. To call any one of these a material and characterize it as such is a useful shorthand, allowing designers to use existing methods when designing with them. But if the hybrid itself is to be designed, it must be deconstructed and thought of as a combination of materials (or of material and space) in a chosen configuration. 3.3.3.3

SYNTHESIZING THE PROPERTIES OF CELLULAR STRUCTURES: FOAMS AND LATTICES

This section provides an example of the synthesizing properties of foams and cellular structures. Some simple mathematical models for synthesizing homogenized properties of these materials are discussed first. A numerical example follows. The principal sources of the following models are Gibson and Ashby [32] and Ashby et al. [33]. For a similar discussion of other classes of hybrids, composites, and sandwich panels, see Ashby et al. [31]. Cellular structures—foams and lattices—are hybrids of a solid and a gas. The properties of the gas might at first seem irrelevant, but this is not so. The thermal conductivity of low-density foams of the sort used for insulation is determined by the conductivity of the gas contained in their pores; the dielectric properties, and even the compressibility, can depend on the gas properties. There are two distinct

Fig. 3.29 Cell in a low-density foam. When the foam is loaded, the cell edges bend, giving a low-modulus structure.

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Fig. 3.30 Micro-truss structure and its unit cell. The cell edges stretch when the structure is loaded. species of cellular solid. The first, typified by foams, are bending-dominated structures; the second, typified by triangulated lattice structures, are stretch dominated—a distinction explained more fully as follows. Foams are cellular solids made by expanding polymers, metals, ceramics, or glasses with a foaming agent—a generic term for one of many ways of introducing gas, much as yeast does in bread making. Figure 3.29 shows an idealized cell of a low-density foam. It consists of solid cell walls or edges surrounding a void space containing a gas or fluid. Foams have the characteristic that, when loaded, the cell walls bend. Lattice structures (Fig. 3.30) are configured to suppress bending, and so the cell edges must stretch instead. 3.3.3.4

DENSITY

Cellular solids are characterized by their relative density, which for the structure shown here (with t L) is  t 2 r˜ ¼ C1 ð3:54Þ rs L where

r˜ rs L t C1 3.3.3.5

¼ the density of the foam ¼ the density of the solid of which it is made ¼ the cell size ¼ the thickness of the cell edges ¼ a constant, approximately equal to 1 MECHANICAL PROPERTIES

In Fig. 3.31 the compressive stress–strain curve of a cellular solid is depicted. The material is linear elastic, with modulus E˜ up to its elastic limit, at which point the

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Modulus is determined by cell edge bending or stretching.

cell edges yield, buckle, or fracture. The foam continues to collapse at a nearly constant stress (the plateau stress s ˜ pl ) until opposite sides of the cells impinge (the densification strain 1˜ d ), when the stress rises rapidly. The mechanical properties are calculated as follows. Elastic moduli of bending-dominated foams. A remote compressive stress s exerts a force F / sL2 on the cell edges, causing them to bend and leading to a bending deflection d, as shown in Fig. 3.29. For the open-celled structure shown in the figure, the bending deflection scales as

d/

F L3 Es I

ð3:55Þ

where modulus of the solid of which the foam is made Es ¼ the t 4 ¼ the second moment of area of the cell edge of square cross section, t 2 I ¼ 12 The compressive strain suffered by the cell as a whole is then 1 ¼ 2d/L. Assembling these results gives the modulus E˜ ¼ s=1 of the foam as  2 r˜ E˜ ¼ C2 Es rs

ðbending-dominated behaviorÞ

ð3:56Þ

Because E˜ ¼ Es when r˜ ¼ rs , one could expect the constant of proportionality C2 to be close to unity—a speculation confirmed by experiment. Numerical simulation gives C1 ¼ 0.7, the value used in the following examples. The quadratic dependence means that a small decrease in relative density causes a large drop in modulus. When the cells are equiaxed in shape, the foam properties are isotropic with shear modulus, bulk modulus, and Poisson’s ratio via ˜ ¼ 3 E˜ G 8

K˜ ¼ E˜

n¼

1 3

ð3:57Þ

Elastic moduli of stretch-dominated lattices. The structure shown in Fig. 3.30 is fully triangulated. This means that the cell edges must stretch when the structure is loaded elastically. On average, one-third of its edges carry tension when the

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structure is loaded in simple tension, regardless of the loading direction. Thus,   ˜E ¼ C3 r˜ Es ðstretch-dominated behaviorÞ ð3:58Þ rs with C3 ¼ 1/3. The modulus is linear, not quadratic, in density, giving a structure that is stiffer than a foam of the same density. The structure of Fig. 3.32 is almost isotropic, and so the shear modulus, bulk modulus, and Poisson’s ratio can be approximated by Eq. (3.57). 3.3.3.6

YIELD STRENGTH, FLEXURAL STRENGTH, AND COMPRESSIVE STRENGTH

Strength of bending-dominated foams. When the structure of Fig. 3.29 is loaded beyond the elastic limit, its cell walls may yield, buckle elastically, or fracture as shown in Fig. 3.32. Consider yielding first (Fig. 3.32a). Cell edges yield when the force exerted on them exceeds their fully plastic moment: Mf ¼

sy;s t 3 4

ð3:59Þ

a)

b)

c)

Fig. 3.32 Collapse of foams. a) A metallic foam that deforms plastically, b) An elastomeric foam, c) A brittle foam.

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where sy,s is the yield strength of the solid of which the foam is made. This moment is related to the remote stress by M / FL / sL 3. Assembling these results gives the compressive failure strength when yield dominates, s ˜ c:

s˜ c ¼ C4

 3=2 r˜ sy,s rs

ðyield of foamsÞ

ð3:60Þ

where the constant of proportionality, C4 } 0.3, has been established both by experiment and by numerical computation. Elastomeric foams collapse not by yielding but by elastic bucking; brittle foams collapse by cell-wall fracture (Fig. 3.32b and c). As with plastic collapse, simple scaling laws describe this behavior well. Collapse by buckling occurs when the stress exceeds  2 r˜ s˜ c 0:05 Es rs

ðbuckling of foamsÞ

ð3:61Þ

The compressive strength s ˜ c is identified with the lesser of Eq. (3.60) or ˜ flex are set Eq. (3.61). Further, the yield strength s ˜ y and the flexural strength s equal to s ˜ c. Strength of stretch-dominated lattices. Collapse occurs when the cell edges yield, giving the collapse stress 1 s˜ c 3

  r˜ sy,s rs

ðyield of latticesÞ

ð3:62Þ

This is an upper bound because it assumes that the struts yield in tension or compression when the structure is loaded. If the struts are slender, they may buckle before they yield. They do so at the stress

s˜ c 0:2

 2 r˜ Es rs

ð3:63Þ

The compressive strength s ˜ c is identified with the lesser of Eq. (3.62) and Eq. ˜ flex are identified (3.63). Further, the yield strength s ˜ y and the flexural strength s with s ˜ c.

3.3.3.7

FRACTURE TOUGHNESS

Fracture toughness of bending-dominated foams. Foams that contain long, cracklike flaws compared to the cell size ‘ fail by fast fracture, meaning that the crack propagates unstably, if the stress intensity factor exceeds the critical value K˜ Ic ,

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which is the lesser of  1=2  3=2 ‘ r˜ KIc,s and KIc,s K˜ Ic ¼ 0:5 a rs

ð3:64Þ

where a ¼ the intrinsic flaw size of the cell edge material KIc,s ¼ its fracture toughness sts,s ¼ its tensile strength, wherein typically ‘=a 10 This behavior is not confined to open-cell foams. Most closed-cell foams also follow these scaling laws. At first sight, this is an unexpected result because the cell faces must carry membrane stresses when the foam is loaded, and these should lead to a linear dependence of both stiffness and strength on relative density. The explanation lies in the fact that the cell faces are very thin; they buckle or rupture at stresses so low that their contribution to stiffness and strength is small, leaving the cell edges to carry most of the load. Fracture toughness of stretch-dominated lattices. Lattices that contain long, crack-like flaws compared to the cell size ‘ fail by fast fracture if the stress intensity factor exceeds the critical value K˜ Ic , which is the lesser of  1=2   ‘ r˜ ˜ K Ic ¼ 0:5 ð3:65Þ KIc,s and KIc,s a rs ‘=a is the ratio of the cell size of the foam to the flaw size in the material, again typically having a value of approximately 10.

3.3.3.8

THERMAL PROPERTIES

Specific heat and thermal expansion. The specific heat C˜ p of foams and lattices (units: J/kg . K) and the expansion coefficient a ˜ (units K21) are the same as those of the solid of which they are made. Thermal conductivity. The cells in most cellular structures are sufficiently small that convection of the gas within them is completely suppressed. The thermal conductivity of the foam is then the sum of that conducted through the cell walls and through the still air (or other gas) they contain. An adequate approximation is    3=2 !    1 r ˜ r˜ r˜ þ2 l˜ ¼ ls þ 1  lg ð3:66Þ 3 rs rs rs where

ls ¼ the conductivity of the solid lg ¼ the conductivity of the gas (for dry air it is 0.025 W/m . K)

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The term associated with the gas is important: blowing agents for foams intended for thermal insulation are chosen to have a low value of lg. 3.3.3.9

ELECTRICAL PROPERTIES

Resistivity. The electrical resistivity r˜ elec of foam is given by

r˜ elec ¼

3    3=2 ! relec,s r˜ r˜ þ2 rs rs

ð3:67Þ

Dielectric properties. Insulating foams are attractive for their low dielectric constant 1˜ r falling toward 1 (the value for air or vacuum) as the relative density decreases:   r˜ ð3:68Þ 1˜ r ¼ 1 þ ð1r,s  1Þ rs where 1r,s is the dielectric constant of the solid of which the foam is made. The dielectric loss tangent is independent of foam density and just equal to that of the solid of which the foam is made. 3.3.3.10

EXAMPLE: ALUMINUM METAL MATRIX COMPOSITE (MMC) FOAMS AND LATTICES

The starting point is aluminum 20% SiC(p) [(p) ¼ particulate], a mix that is the basis of one of the Cymat range of metal foams. The properties of two hybrids are compared to a “foam” and an “octet lattice,” which generates two sets of new records. They appear on the modulus–density chart of Fig. 3.33 as lines of white and gray circles, respectively. The starting material, Al 20% SiC(p), is identified at the upper right. The modeled foams, plotted white, are labeled with their relative densities in parentheses. Measured data for real aluminum-SiC(p) foams (shown as black ellipses in Fig. 3.33) are also labeled and underlined, with relative densities in parentheses, allowing a comparison. Note that the lattices, plotted in gray, lie at higher values of modulus for the same density. At a relative density of 3.5% (0.035), the lattice is predicted to be 10 times stiffer than foam of that density. In Fig. 3.34, many other engineering materials have been added. The lattice structures outperform all of them, expanding the populated area of the chart. Therefore, an AlSiC(p) lattice provides one possible way to find a material with superior modulus/density performance to conventional materials. Additional examples can be found in Ashby et al. [31]. 3.3.3.11

CES HYBRID SYNTHESIZER

The approach described has been implemented in the CES Selector [7], a scoping tool that allows fast exploration of hybrid structures [31]. It presents approximate

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Fig. 3.33 CES Hybrid Synthesizer [31] results, when the input window of the synthesizer, set for cellular structures: octet lattice. analyses of the properties that might be achieved by forming a single material into a cellular structure or by combining two materials to form a sandwich or a composite. The purpose of the tool is to encourage innovation by allowing estimated property profiles for virtual novel hybrids to be explored and compared with the property profiles of established engineering materials. If a suitable innovative material can be found in this way, the next steps are more detailed analyses (e.g., using finite element methods followed by manufacturing, testing, and qualifying the new material).

3.4 DATA ISSUES The quantity and quality of the material information is central to all aspects of the materials selection processes previously described and, most important, to the final engineering products. The needs for materials information change throughout the life cycle of an engineering component; see Table 3.11. Likewise, the informatics infrastructure required to handle the potentially massive amounts of materials data should not be overlooked or underestimated. Early in the design process, in the conceptual design phase, information is needed for all possible materials but at a relatively low level of precision, because the objective at this stage is to select a small subset of materials that can optimally perform the function of the component with acceptable cost and

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environmental impact. “Typical” values of the engineering properties and environmental impact data are best suited to this task. Narrowing this subset to one or two preferred candidates requires encyclopedic data sources that can inform the designer of a wide range of (often unquantifiable) characteristics of the material. This supporting information includes details of corrosion, heat treatment, environmental impact, and joining characteristics, as well as previous in-house experience using the material. As the design of the component is finalized, detailed stress analyses and lifing calculations are needed, which in turn require very precise “allowable” design values of the properties of a single material. Next, the component must be manufactured, which has its own data requirements for information about the processing characteristics of the raw material— the viscosity when molten, the strength at high temperature, the hardness and toughness, and so on. When the component is in service, there is a need for materials data for maintenance purposes (e.g., for manufacturing spare parts), for predicting the remaining life, and for investigating service failures when they arise. Finally, at the end of life, the component may be recycled or disposed of in some other way. Therefore, knowledge of the environmental characteristics of the material (e.g., information about toxicity or degradation in landfill and associated legislative requirements) can be essential.

Fig. 3.34 Al-SiC lattices results generated by the CES Hybrid Synthesizer [31] compared with other materials.

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TABLE 3.11 Product Life Phase

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MATERIAL INFORMATION NEEDS THROUGH A PRODUCT LIFE CYCLE Materials Information

Engineering Operation

Design

Selection data (typical properties) Supporting information Environmental impact information Design “allowables” Specific constitutive and failure models for detailed design

Materials selection Environmental analysis Life-cycle assessment Design calculations Stress analysis Lifing, simulation

Manufacture

Processing data

Process modeling Manufacturing simulation

In-service

Durability information

Maintenance Prediction of remaining life Investigation of failures

End of life

Environmental impact information Hazardous materials regulations Recycling information

End-of-life policy implementation

Having a multifaceted, “complete” materials data source in a well-organized, easily retrievable information management system is a critical and often unappreciated aspect of the entire engineering process. The value of a given database depends on its precision and completeness, or in other words its quality. Consequently, numerous procedures of checking and validating data have been devised. See, for example, Ashby [3] and Cebon and Ashby [2, 34] for a discussion of checking methods, including property range and dimensionless correlations concepts. The materials data life cycle is described in Fig. 3.35, wherein data are captured and consolidated from external sources, legacy databases, and internal (possibly proprietary) testing programs. Next, data are analyzed and integrated to create/discover useful information and then deployed (disseminated) to the people who use it to meet their needs. The continual Fig. 3.35 Four aspects of materials data life cycle.

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maintenance of the whole system (the data and information generated as well as the relationships, or links, between them) is the last yet essential stage of the data life cycle. To support the various required activities throughout this data life cycle requires the (preferably seamless) integration of a variety of software tools. These range from data input, reduction/analysis, visualization, and reporting tools to material parameter estimation tools, product life management tools (PLMs), and structural analysis codes that use a central database. These tools should enable material and structural engineers throughout the enterprise to input, manage, and use information in a way that is as efficient, reliable, and user friendly as possible. An information management system that was conceived of by the Material Data Management Consortium (www.mdmc.net) and built by Granta Design provides this type of environment and information management tool set. The tool set is composed of two primary components: 1. GRANTA MI is a comprehensive materials information system accessible from the Web. Its unique database is specifically designed to store materials properties, response curves, and other data relating to materials. It comprises a series of powerful software tools that help control, analyze, and apply data. 2. CES Selector is a PC software application that offers advanced graphical analysis of materials data, plus specialist eco design and modeling tools, in support of materials selection and substitution decisions.

3.5 SUMMARY Tradeoffs are endemic in both everyday life and materials selection. Typically, a designer’s job is to balance a variety of functional requirements (i.e., types of loading conditions: tension, compression, bending, vibration, cyclic, etc.) with constraints (manufacturability, geometric limits, environmental aspects, and maintainability, to name a few) to arrive at the optimum choice of structural concept and materials selection for a given weight and/or cost. This chapter described a systematic, design-oriented, five-step approach to materials selection: 1) establishing design requirements, 2) material screening, 3) ranking, 4) researching specific candidates, and 5) applying specific cultural constraints to the selection process. At the core of this approach is the introduction of performance indices (i.e., combinations of material properties that embody the performance of a given component) in conjunction with material property charts. These materials selection charts, which plot one property against another, have been shown to provide a powerful graphical environment wherein one can apply and analyze quantitative selection criteria, such as those captured in performance indices, and make tradeoffs between conflicting objectives in idealizations representing various engineering problems. Finding a material with a high value of these indices maximizes the performance of the component. Two specific

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examples were examined pertaining to aerospace (engine blades and pressure vessels), both at room temperature and elevated temperature (wherein timedependent effects were accounted for) to demonstrate the methodology. A discussion was also presented with respect to engineered/hybrid materials and how these can be effectively tailored to fill in holes in the material property space to enable innovation and increases in performance over monolithic materials. Finally, data collection, analysis, deployment, and maintenance issues were briefly addressed.

REFERENCES [1] Larson, W. J., and Pranke, L. K., Human Spaceflight: Mission Analysis and Design, McGraw-Hill, New York, 1999, p. 1035. [2] Cebon, D., and Ashby, M. F., “Information Systems for Material and Processes,” Advanced Materials and Processes, Vol. 157, No. 6, 2000, pp. 44–48. [3] Ashby, M. F., “Checks and Estimates for Material Properties,” Proceedings of the Royal Society, Vol. A454, 1998, pp. 1301–1321. [4] Ashby, M. F., Materials Selection in Mechanical Design, 4th ed., ButterworthHeinemann, Oxford, 2011. [5] Broek, D., Elementary Engineering Fracture Mechanics, Martinus Nijhoff Publishers, Leiden, Belgium, 1982. [6] Anderson, T. L., Fracture Mechanics: Fundamentals and Applications, CRC Press, Boca Raton, FL, 2005. [7] CES Selector 2018, Granta Design, Cambridge, England, U.K. [8] Mital, S. K., Gyekenyesi, J. Z., Arnold, S. M., Sullivan, R. M., Manderscheid, J. M., and Murthy, P. L. N., “Review of Current State of the Art and Key Design Issues with Potential Solutions for Liquid Hydrogen Cryogenic Storage Tank Structures for Aircraft Applications,” NASA TM 2006-214346, 2006. [9] Miracle, D. B., “Metal Matrix Composites for Space Systems: Current Uses and Future Opportunities,” Affordable Metal-Matrix Composites for High Performance Applications, edited by A. B. Pandey, K. L. Kendig, and T. J. Watson, Minerals, Metals and Materials Society, Warrendale, PA, 2001, pp. 1–21. [10] Arnold, S. M., Bednarcyk, B. A., Collier, C. S., and Yarrington, P. W. “Spherical Cryogenic Hydrogen Tank Preliminary Design Trade Studies,” NASA TM 2007-214846, 2007. [11] Lemaitre, J., and Chaboche, J. L., Mechanics of Solid Materials, Cambridge Univ. Press, New York, 1990. [12] Chester, T. S., Norman, S. S., and William, C. H., Superalloys II, Wiley, New York, 1987. [13] Skrzypek, J., and Hetnarski, R., Plasticity and Creep, Theory, Examples, and Problems, CRC Press, Boca Raton, FL, 2000. [14] Odqvist, F., “Theory of Creep Under the Action of Combined Stresses with Applications to High Temperature Machinery,” Proceedings of the Royal Swedish Institute for Engineering Research, No. 141, 1936.

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[15] Arnold, S. M., and Saleeb, A. F., “On the Thermodynamic Framework of Generalized Coupled Thermoelastic-Viscoplastic-Damage Modeling,” International Journal of Plasticity, Vol. 10, No. 3, 1994, pp. 263–278. [16] Saleeb, A. F., Arnold, S. M., Castelli, M. G., Wilt, T. E., and Graf, W. E., “A General Hereditary Multimechanism-Based Deformation Model with Application to the Viscoelastoplastic Response of Titanium Alloys,” International Journal of Plasticity, Vol. 17, No. 10, 2001, pp. 1305–1350. [17] Saleeb, A. F., and Arnold, S. M., “A General Reversible Hereditary Constitutive Model: Part I Theoretical Developments,” Journal of Engineering Materials Technology, Vol. 123, 2001, pp. 51–64. [18] Saleeb, A. F., and Arnold, S. M., “Specific Hardening Function Definition and Characterization of a Multimechanism Generalized Potential-Based Viscoelastoplasticity Model,” International Journal of Plasticity, Vol. 20, 2004, pp. 2111–2142. [19] Arnold, S. M., Saleeb, A. F., and Castelli, M. G., “A General Reversible Hereditary Constitutive Model: Part II Application to Titanium Alloys,” Journal of Engineering Materials Technology, Vol. 123, 2001, pp. 65–73. [20] Arnold, S. M., Goldberg, R. K., Lerch, B. A., and Saleeb, A. F., “An Overview of Prognosis Health Management Research at Glenn Research Center for Gas Turbine Engine Structures with Special Emphasis on Deformation and Damage Modeling,” NASA TM 2009-215827/Rev 1, 2009. [21] Ramberg, W., and Osgood, W. R., “Description of Stress Strain Curves by Three Parameters,” NACA TN 902, 1943. [22] Finnie, I., and Heller, W. R., Creep of Engineering Materials, McGraw-Hill, New York, 1959. [23] Hult, J., Creep in Engineering Structures, Baisdell Press, London, 1966. [24] Penny, R. K., and Marriott, D. L., Design for Creep, McGraw-Hill, Boston, 1971. [25] Gittus, J., Creep, Viscoelasticity and Creep Fracture, Elsevier, London, 1975. [26] Frost, H. J., and Ashby, M. F., Deformation Mechanism Maps, Pergamon, Oxford, 1982. [27] Evans, R. W., and Wilshire, B., “Creep of Metals and Alloys,” Institute of Metals, London, 1985. [28] Ashby, M. F., Materials Selection in Mechanical Design, Pergamon, Oxford, 1992. [29] Abel, C. A., and Ashby, M. F., “Materials Selection to Resist Creep and Creep Rupture,” Cambridge Univ. Engineering Department Report CUED/C-EDC/TR18, 1994, pp. 1–32. [30] Arnold, S. M., Cebon, D., and Ashby, M., “Materials Selection for Aerospace Systems,” NASA TM 2012-217411, 2012. [31] Ashby, M. F., Cebon, D., Bream, C., Cesaretto, C., and Ball, N., “The CES Hybrids Synthesizer–A White Paper,” Granta Design, Cambridge, England, U.K., 2010. [32] Gibson, L. J., and Ashby, M. F., Cellular Solids, Structure and Properties, 2nd ed., Cambridge Univ. Press, Cambridge, England, U.K., 1997. [33] Ashby, M. F., Evans, A. G., Fleck, N. A., Gibson, L. J., Hutchinson, J. W., and Wadley, H. N. G., “Metal Foams: A Design Guide,” Butterworth Heinemann, Oxford, 2000. [34] Cebon, D., and Ashby, M. F., “Engineering Materials Informatics,” MRS Bulletin 31, 2006, pp. 1004–1012.

CHAPTER 4

Advanced Nanoengineered Materials Brian L. Wardle Massachusetts Institute of Technology, Cambridge, Massachusetts

Joseph H. Koo University of Texas at Austin, Austin, Texas

Gregory M. Odegard Michigan Technological University, Houghton, Michigan

Gary D. Seidel Virginia Polytechnic Institute and State University, Blacksburg, Virginia

4.1 INTRODUCTION At the time of this writing, nearly two decades of major investment and progress in nanoscience and nanotechnology [1] has followed the initial interest sparked by Iijima’s identification of the useful and unique properties of carbon nanotubes (CNTs) in 1991 [2]. Much like the development of carbon fiber composites, the aerospace sector has provided much of the early impetus for nanomaterials research and has led the way in advanced applications, particularly nanomodified/enhanced/engineered materials for structural and multifunctional uses. The vast majority of extant work in nanomaterials, whether it be for aerospace or other applications, has focused on nanoscale additives into polymer matrices [3, 4], with numerous review articles appearing [5–13]. Therefore, the focus of this chapter is on carbon nanomaterials in polymer matrices in aerospace applications. The carbon nanomaterials that are farthest along in technological development and application are CNTs, and there have been numerous CNTfocused reviews [14–17]. The parallels between carbon fiber and CNTs are many, with lessons from the development and adoption of carbon fiber (in aerospace applications) providing a good perspective on the current state of CNT-based architectures (see Fig. 4.1). The early days of carbon fibers involved development for aerospace applications, with many applications involving undisclosed or recently disclosed military and space projects, in parallel with “lower-grade” materials and applications such as sporting goods. Adoption in commercial aerospace happened over decades, yielding an aerospace-structures S curve [18] for composites as shown in Fig. 4.1. CNT-composed materials follow a similar course, but, with the benefit of the carbon fiber composite experience behind us and the emergence of unmanned aerial vehicles/systems

Copyright # 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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Fig. 4.1 Aerospace commercial transport S-curve showing carbon fiber and carbon nanotube (CNT) development. (Chart courtesy of # MIT. Data from industry sources and [19]). (UAV/S), the expected rate of adoption may be faster. Note that an earlier S curve for U.S. military fighter aircraft starts in 1970 and has 40% composite structural weight for the F-22 in 1990. Nanomaterials are currently defined as having one dimension of nanometer (nm) scale (between 1 and 100 nm), a length regime where defect-free materials can be synthesized and where, due to confinement and other effects, many physical properties behave fundamentally differently (e.g., electron and phonon transport). Global investment in nanoscience and nanotechnology rivals all other major trends in science and engineering. In a survey of the 100 most impactful materials scientists from the first decade of this century, 78% defined their primary field of research as nanotechnology [20]. Further, all major forward-looking research strategies have significant components of nanomaterials research [21, 22], including the U.S. Air Force Office of Scientific Research (AFOSR) and NASA plans for the next decades of research [23, 24] and a recent National Research Council report on vehicle lightweighting [25]. NASA’s current set of roadmaps includes a specific Nanotechnology Roadmap TA-10 (Technology Area, 10 of 14), but it is also clear that nanotechnology and nanomaterials cut across many of the other roadmap areas, including Materials (TA-12), Power (TA-3), and Instruments (TA-8), among others, demonstrating the more general point that there is a wide range of insertion

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points for nanomaterials-based technologies. Although the focus is not entirely on structural materials, this was both the early desire and the area where nanomodified materials can continue to provide new solutions and new capabilities. Significant publicly known efforts in nanotechnology at major aerospace original equipment manufacturers (OEMs) include the Airbus/European Aeronautic Defence and Space (EADS) nano-augmented, -engineered, and -enabled initiative and UTOPIUM project; Boeing’s Atoms to Airplanes initiative; and the Lockheed Martin’s Nanosystems (LMNS) group, among others. The aerospace community is clearly organized in different ways to address the opportunities and challenges of nanomaterials. Combined with proven and potential multifunctional (as used here, anything nonmechanical) benefits, nanomaterials are beginning to see implementation in aerosystems, as will be highlighted in several high technology readiness level (TRL) [26] case examples of applications of nanomaterials in aerospace. We begin by discussing nanoengineered structural materials, proceed to multifunctional nanomaterials, and finish with processing, modeling, and future trends.

4.2 NANOENGINEERED STRUCTURAL MATERIALS Aerospace structures are highly engineered systems involving multi-objective design requirements and generally comprising heterogeneous, layered, and surface-finished (often multiple layers) attributes. It is therefore at one’s peril to oversimplify such materials, but to zeroth order, aerospace designers are generally concerned with strength and stiffness as normalized by weight or mass-specific stiffness and strength. Thus, a useful representation, and one that illuminates the reason for the displacement of metals by fiber-based composites in many applications, is an Ashby chart of specific stiffness vs strength. Discussion of the detailed use of such charts in materials selection is given in Chapter 3. More details are beyond the scope of this chapter but may be found in Ashby [27]. In the Ashby-inspired log-log chart in Fig. 4.2, “better” materials are found at the upper right. Indeed, advanced micron-scale-diameter continuous fibers such as glass and carbon are combined with polymers to create advanced composites that, by weight, are superior to metal alloys. Advanced composites composed of glass (G) and carbon (C) fiber-reinforced plastics (FRP) are shown. Even higher than the advanced fiber properties are the nanoscale properties of nanometer-scale-diameter CNTs. Considering that the combination of polymers with advanced fibers provides us with advanced composites, the straightforward inference is to combine CNTs with polymers to create even better composites. Such an approach has proven quite difficult, due to physics resisting CNT organization, lack of processing and synthesis techniques, and other fundamental limits such as the inability to realize nanoscale strength and other properties of CNTs at macroscopic scales [5, 28, 29]. Composites of advanced fibers and polymers are highly engineered materials, with properties that go beyond

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Fig. 4.2 Specific stiffness and strength for comparing metal alloys, composites, and CNTs. (Figure courtesy of Dr. Roberto Guzma´n de Villoria, # MIT.) simple linear combinations (“rule of mixtures”) of the ingredients but rather act synergistically to yield properties beyond either of the constituents (e.g., in-plane fracture toughness that can be an order of magnitude or greater for the composite than for either of the constituents) [27]. In addition to improving the in-plane properties of composites via nanomaterial addition, some have pursued out-of-plane, or through-thickness, property enhancement. Matrix-dominated properties, such as interlaminar strength and toughness and through-thickness stiffness, fall several orders of magnitude below in-plane properties for even the most advanced composites (see the “Composites through thickness” vs “Composites” labels in Fig. 4.2) and can significantly benefit from nanoscale reinforcement. Employing nanomaterials for structural benefit has received considerable attention from researchers and industry since the 1990s with many challenges identified and a general lack of success in early attempts. Although some challenges remain, as will be discussed in the following sections, many have been solved, and the remaining challenges are more clearly articulated with innovative solutions arriving rapidly. Early challenges focused on dispersion of nanomaterials in polymers, fighting the fundamental very high surface-to-volume ratios of nanomaterials, which increase viscosity and limit processing. Scale was also an early challenge, although currently this seems to be less of a challenge than organization (i.e., controlling the orientation of the nanomaterials relative to one another and/or to micro- or macroscopic elements of larger material assemblies). A comparison of CNTs to carbon fibers and their composites is shown in Fig. 4.3. Focusing on the scale challenge, however, it can also be inferred that organization or control of morphology is perhaps more important. Advanced composites are composed of aligned, and in some cases collimated, arrays of fibers that maximize the contributions of the fibers to strength and stiffness (as in Fig. 4.2) and allow property tailoring in different directions. Mixing approaches in nanomaterials yielding isotropic materials have given way to other assembly strategies that allow the direction of nanofibers to be controlled, in some cases in concert with existing advanced fiber morphologies such as aligned carbon fiber unidirectional and woven systems used pervasively in aerospace structures. The

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majority of the work in the field of nanoengineered materials has focused on modifying polymers (and polymer-based composites) by including nanomaterials in the polymer matrix or as fiber replacements. Existing polymer-matrix composite (PMC) systems have been hybridized with nanomaterials, with the primary interest of increasing the mechanical properties of carbon fiberreinforced plastics (CFRPs)—materials that comprise the vast majority of fiberbased structural composites used in aerospace (see Chapter 3). Although nanomaterial modification of metals and ceramics has also been pursued, most nanostructured materials involve polymer binders or matrices, both thermosets and thermoplastics. In this section we focus on nanomaterial additions to polymer matrices (nanocomposites); assemblies of CNT nanofibers to create fibers and sheets; and hybrid material systems where nanomaterials are combined with existing materials, primarily fiber-reinforced plastic (FRP) composites, to create layered and sometimes hierarchical materials (see Fig. 4.3). Laminated composite plies are mm scale (ant body), carbon fiber diameter is micron scale (ant legs), and CNT diameter is nm scale (hair on ant’s legs). Although few transitions to application have occurred in the aerospace sector to date, several prominent examples have occurred on a relatively aggressive timeline as compared with general materials introduction timelines (see discussion in Chapter 3), particularly in structural load-bearing aerospace applications. One such example is a polymer nanocomposite (PNC) nonstructural application in Lockheed Martin’s F-35 [30]. Factors such as manufacturability and uncontrolled morphologies that result from popular

Fig. 4.3 Scale and organization in CNT and carbon fiber composites. (Original figure courtesy of Dr. Enrique J. Garcia, # MIT.)

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Fig. 4.4 Exemplary architecture types employing carbon nanostructures: polymer nanocomposites, CNT fibers and sheets, and hybrid or nanoengineered composites. (Figure courtesy of # MIT.)

mixing processing methods often decrease, rather than increase, mechanical properties (and increases are typically quite small) [10]. The broad field of nanomaterials can be subdivided in several ways; examples can be found in books [31] and review articles [5]. Here we consider the subdivision in Fig. 4.4 from a composite structure’s perspective, focusing on fibers and matrices, such as the nearubiquitous CFRPs used for structural applications in the aerospace discipline. Thus, nanomaterials can be subdivided as nanocomposites that modify the composite (polymer) matrix, nanostructured fibers and sheets that are seen as replacements for carbon fibers, and hybrid or nanoengineered materials that combine both of the aforementioned materials or embrace new concepts and hierarchical assemblies. For example, one can easily envision hybrid carbon fibers with high loadings of CNTs in a nanocomposite matrix as a hybrid or nanoengineered bulk material.

4.2.1 NANOCOMPOSITES Many mechanical properties of composites have been pursued for enhancement via the addition of nanomaterials into polymers. Properties of interest are diverse, including modulus (especially compressive), fatigue, fracture toughness, surface abrasion and scratch resistance, strength, and damage resistance and tolerance. It is useful to distinguish PNCs and the goal of hybridizing a matrix vs using

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the nanomaterials as structural reinforcement to compete with micron-scale reinforcements such as carbon fibers. The latter is the subject of the next section, where CNT-based fibers and sheets are considered as replacements for the primary (fiber) reinforcement in structural composites. Because of the low volume and weight fractions of nanostructures incorporated into PNCs, here we consider PNCs as either improved or hybridized polymers (as may be used in surfacing films, or as hybrid matrices to complement micron-scale fibers). This distinction was not generally made until recently and was underpinned by advances in scale and hierarchy of reinforcement, as well as advanced processing techniques (i.e., not mixing), such as those employed in the architectures discussed in the next two sections. Therefore, we focus on nanostructures in polymer matrices here (e.g., see Spitalskya et al. [6]), primarily CNTs in thermoset (especially epoxy [16]) polymer matrices. It should be noted that there is a significant body of work on mixing inorganic (e.g., silica, alumina, clays) nanomaterials into polymers. Although some successes have been reported, those works are generally not as far developed for structural aerospace applications as CNT-based nanomaterials concepts, with Case Study 1 being a notable exception.

CASE STUDY 1 Nanoparticle-Enhanced Adhesives and Composite Matrices A fundamental limitation in mixing approaches for incorporating nanomaterials into polymers is the high surface-to-volume ratio leading to large increases in viscosity and the inability to process many product forms. A clever and facile route to overcome this limitation is to synthesize the nanoparticles in situ from precursors (liquids primarily) that combine readily with the precursors of the host matrix. Examples of nanoparticle synthesis via liquid-based routes are abundant [32], but early and elegant work took the significant additional step of developing the chemistry and processing to allow the nanoparticles (e.g., nanoparticle silica, SiO2-reinforced bisphenol A epoxy resin) to form in structural adhesives and polymer matrices [33, 34] (see Fig. 4.5). In addition to the advantages of processing via in situ nanoparticle synthesis, these approaches also allow significant tailoring of nano- and microparticle features that provide control of properties such as tuning of toughening mechanisms, including synergies between rubber toughening and nanoparticles [35, 36]. The processing has also shown compatibility with processing FRP composite laminates [37, 38]. Specific variations of the nanoparticle-enhanced adhesives (DGEBA/LY566 and others) have been commercialized as Araldite EP1000AB by Huntsman and qualified for use by multiple large commercial airframers, with applications that include bonding of metal and composite structures [39].

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Fig. 4.5 Nanoparticle-reinforced epoxy. (Top SEM image of nanoparticles in polymer courtesy of Huntsman, multiscale damage mechanisms ahead of a crack courtesy of Prof. A. Kinloch, and impact damage resistance comparison for FRP laminates courtesy of Dr. S. Sprenger.)

The scale of reinforcement, and its role in process-structure-property relations for polymer matrix-based materials, is a paramount consideration. Nanostructured particles and fibers have 1000 the surface-to-volume ratio of typical micron-scale fibers (see Fig. 4.3) due to 1/r scaling, where r is the radius or characteristic length of the nanomaterial. This scaling reality has many implications for process and structure, thereby dictating properties. Two implications of primary interest are: 1) high surface-to-volume ratios that (problematically) resist mixing processes and 2) high packing fractions of reinforcement that require nanomaterial spacings that are much smaller than can be currently controlled and, further, are on the order of the polymer chain characteristic lengths, such as radius of gyration [40–42]. Both of these scaling realities dictate that it is extremely difficult to create desired structural morphologies of nanomaterials that are dispersed in bulk polymers, which is even more difficult if micron-scale fibers and fiber beds are to be incorporated. If high packings are achieved, questions abound as to the role of an interphase at the hard nanomaterial interface with the polymer

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system. A similar interphase question was pursued vigorously several decades ago for microfiber-reinforced composites such as CFRPs and glass fiber-reinforced plastics (GFRPs), but due to the high surface-to-volume ratio of nanomaterials, it is much more important. For example, some early work with PNCs, especially thermoplastics, was interpreted as nanomaterial reinforcement, whereas in reality the nanomaterials served as nucleation sites to crystallize a significant fraction of the polymer, thereby increasing the PNC modulus. The extent and degree of alteration of the polymer are still active subjects of research and depend on specifics of the nanostructured reinforcement (e.g., presence of molecules adhered to the nanostructure surface), the polymer (thermoplastic or thermoset, etc.), and the processing parameters of the PNC. Currently, thermoplastics as a class seem more sensitive than thermosets. In extreme cases, nanostructures can initiate significant crystallinity in some thermoplastics [43, 44]. A region of altered polymer morphology (and properties) exists local to a nanostructure surface with characteristic lengths estimated to be 10–100 nm [40–42] for both thermosets and thermoplastics. Consider (unrealistic though it may be) collimated CNTs of 7 nm diameter vs carbon fibers of 7 mm diameter. Presuming simple square packing and a conservative estimate of 10 nm altered polymer morphology in the vicinity of the fiber reinforcement, at a CNT volume fraction of only 10% we reach 100% altered polymer matrix. Properties of the neat polymer become irrelevant in assessing the reinforcement of the CNTs (i.e., the polymer is 100% interphase, whereas the interphase is a vanishingly small part of the 7-mm-diam fiber composite). Although this is a simple discussion on reinforcement, the impact of these nanostructure-property interactions is likely more profound for failure and fracture. A related discussion is given by Ashby et al. with reference to “magic numbers” that assess the number of surface molecules vs interior molecules in nanoscale solids [31]. State-of-the-art mixing approaches, including surface functionalizations, to achieve well-dispersed nanomaterials in PNCs appear in numerous references and are reviewed in Sec. 4.4. One of the consistent results from these works is that it is very difficult to achieve desired homogeneous dispersions (nanomaterials with polymer in between each individual nanoparticle or nanofiber) via mixing, and certainly at high nanoparticle loadings, limiting reinforcement to typically less than 5% nanomaterial by weight or volume. Much more success from mixing approaches is noted in multifunctional properties, such as electrical conductivity and thermal degradation resistance, because such properties can be dramatically altered at nanomaterial loadings of much less than 1% (see discussion in Sec. 4.3). Dispersion issues resulting in lack of homogeneity are typically cited as reasons for degradation (or lack of improvement) of basic mechanical properties such as modulus at such low reinforcement loading fractions (e.g., see Guzman de Villoria and Miravete [45] for effects on modulus), but damage to CNT-based PNCs introduced by mixing (often through “high-energy” processes such as shear mixing and sonication) is also a likely contributor. Random orientation, agglomeration/aggregation, and inhomogeneous distributions abound, and

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chemical surface coatings and functionalizations used to improve dispersion can strongly influence many physical properties due to the high surface-to-volume ratios of the nanomaterial reinforcement. PNCs at high volume fractions require extremely straight and tightly spaced (nanometer-scale control) nanofiber reinforcement. Clearly, as the state of the art in processing of PNCs relies on mixing, nanometer control of collimation and packing is a significant challenge in achieving a high packing of CNTs to create a nanoscale analog to aerospace unidirectional carbon fibers. As with the fibers and sheets discussed in the next two sections, PNCs that have morphological order, such as an aligned-fiber nanocomposite, have seen much more impressive improvements in mechanical properties than with mixing, particularly because alternative synthesis approaches allow higher packing fractions as well as control of nanostructured order. Examples include layer-by-layer deposition of nanomaterials [46], filtering and layering approaches, spinning of fibers and yarns (and fiber precursors) with nanomaterial additives, aligned-nanofiber PNCs that are analogous to unidirectional collimated carbon fiber layers [47, 48], and in situ synthesis techniques involving precursor mixing with polymers as in Case Study 1.

4.2.2 NANOSTRUCTURED FIBERS AND SHEETS Continuing to focus on structural materials for aerospace, here we consider assemblies of nanomaterials to replace carbon fibers (see Fig. 4.3) as the reinforcement in PMCs. Referring to Fig. 4.2, the idea is to take advantage of superior fiber properties of CNTs (vs carbon fiber) to create PMCs with CNT reinforcement. Many advances have been made in this direction, with continuous improvement in properties resulting in basic specific stiffness and strength properties that are approaching carbon fibers, with perhaps even more success in high-energy absorbing fibers. It is, for a variety of reasons that depend on the property of interest (strength, phonon conduction, etc.), a continuing challenge to obtain property performance from macroscopic assemblies of nanocarbons (particularly CNTs) comparable to that achieved at small scales [28]. Assemblages of discontinuous (up to millimeter-scale) CNTs in yarns and sheets have been demonstrated via several techniques, including floating-catalyst CNT growth and collection [49, 50], solution deposition to create “Buckypaper” [51], and spinning of yarns from aligned-CNT mats [52]. Case Study 2 contains a discussion of CNT-sheet appliques on structural elements of a spacecraft for multifunctional purposes, with the future vision that the same will serve structural purposes. Properties of the fiber and sheet materials are continually improving through CNT functionalizations to manage CNT–CNT and CNT–polymer interactions, and via drawing and stretching processes that increase volume fraction and alignment. Using such yarn and sheet architectures effectively solves many of the issues of dispersion when forming a PMC but faces similar challenges (packing, polymer morphology alteration, effective wetting and consolidation, alignment, etc.) as those for PNCs

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discussed previously. A recent and important work explores the concept of damaging the CNTs to create more bonding sites for the polymer along the CNT length [53], thereby introducing a useful tailoring scheme to manage the interaction between the polymer and these materials. The approach involves a tradeoff between damaging the reinforcement (CNTs) and the benefit of improved polymer–CNT interaction, as realized by significant relative increases in modulus and ultimate strength. Another approach is to hybridize carbon fibers by incorporating CNTs [54] or to create nanoscale versions of continuous carbon fibers via electrospinning [55]. CNT-based hybrid fibers are derived from the general field of nanostructure incorporation into polymer fibers [56, 57], where the polymer fiber is then pyrolyzed containing the CNTs (or other nanostructures), creating a hybrid fiber [58]. Exemplary work in this area involves the spinning of polyacrylonitrile precursor fiber containing CNTs and then pyrolyzing [59]. Composites based on such hybrid fibers are under development and constitute exciting new research directions as components of the hierarchical materials discussed in the next section.

CASE STUDY 2 Carbon Nanotube Sheet Appliques on Mission to Jupiter One notable example of CNT-based materials in space is the use of CNT sheets as a surface treatment on the struts of NASA’s Juno spacecraft (see Fig. 4.6), which was

Fig. 4.6 CNT sheet application areas on Juno spacecraft, courtesy of NanoComp Technologies. (Illustration of Juno spacecraft at Jupiter, courtesy of NASA. Scale of CNT sheet material, courtesy of NanoComp Technologies.)

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launched in 2011 [60]. At the time of this writing, the large probe, designed and built by Lockheed Martin, is on its way to Jupiter. Large sheets of discontinuous (mm-scale length) multiwalled CNTs (MWCNTs) are produced via a floating-catalyst chemical vapor deposition method by NanoComp Technologies; the sheets have been infused with resin in past work to create CNT-polymer composite layers [53]. The sheets provide electrostatic discharge protection for the spacecraft, a function usually provided by metal foils and similar to approaches taken on many composite commercial aircraft components. Although these CNT sheets are an impressive demonstration as an applique on structural members of the spacecraft, they are not the first in space. CNT-based materials have been studied in space previously [61], and other nanocarbons (C60 and C70 fullerenes and graphene) have been identified as naturally occurring in space [62, 63].

4.2.3 HYBRID OR NANOENGINEERED MATERIALS Nanoengineered structural materials can be created via hierarchical architecting of nanoscale and microscale constituents into hybrid assemblies (see examples in Fig. 4.7). Computational tools for designing such bulk materials are not available but are being developed, and the inputs to power the design tools are still being developed experimentally via detailed PNC morphology characterizations and physical property measurements. Several hybrid nano- and microfiber PMC architectures have been realized [5], the simplest being CNT-modified matrices combined with aligned fibers using various techniques such as electrophoretic deposition of CNTs to fibers [64], mixing CNTs into polymers (e.g., see Kim et al. [65] and some nonaerospace

Fig. 4.7 From nanostructure to aerospace structures: scaling and scaffolding approach to create improved engineered systems. (Figure courtesy of # MIT; image credits to NASA, Terrafugia, MIT, MTU, and NanoComp Technologies.)

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commercial products), spray deposition onto CFRP prepreg interfaces [66], CNT-fiber plies analogous to unidirectional CFRP plies [67], polymerimpregnated CNT sheets [53], and others. Combining these fibers and sheets with existing carbon fibers is one approach that could be taken to achieve hybrid PMCs at scale, a nanoscale analog to fiber-metal hybrid laminates. One architecture that has received considerable attention is the “fuzzyfiber composite,” where CNTs (or other nanowires) are grown on the surface of microfibers to create a two-scale fiber reinforcement (see review in Qian et al. [68] and inset in lower middle of Fig. 4.7). Such two-scale fiber systems are a clear initial step toward hierarchical and three-dimensionally tailored bulk composite materials.

4.3 NANOENGINEERED MULTIFUNCTIONAL MATERIALS Multifunctional materials, here meaning nonstructural functions of structural materials, have been the focus of many hybrid architecture schemes, and some of the earliest work using nanomaterials has been in this area. The approaches range from nanomodified films and surfaces for sensors to hybridized FRP systems with new functionality, such as self-healing [69, 70]. The majority of work in the area of multifunctionality has been in the nanocomposites arena, with filler-like concepts modifying polymers (bulk and films) and being used for multifunctional attributes such as strain and damage sensing [71–76]. Fewer works have incorporated nanomaterials into fiber-based composites, with the focus being to create conductive nanoparticle (especially CNT) networks for strain and damage sensing [77–81]. A recent review highlights the broad field of multifunctional materials, where many of the advances involve materials composed of nanoscale constituents [82]. Indeed, recent nanomaterial reviews highlight multifunctional aspects of bulk materials composed of CNTs [5, 83]; some reviews focus solely on nanomaterials (particularly CNTs) for multifunctionality [83–85], and multifunctional attributes are leading in applications as reviewed in current books on nanomaterials [3, 86, 87] (e.g., see Case Studies 2 and 3). A major area of research focus has been to enhance electrical properties of polymers via conductive nanoparticle inclusion. It is well known that conductive particles at low loadings can modify electrical transport in otherwise insulating systems, displaying percolation behavior. The percolation phenomenon, where nanoparticulate and nanofiber reinforcement seems advantageous [88], is an area useful to aerospace in terms of systems for electromagnetic management [86]. Some efforts have focused on trying to introduce and control anisotropy in the electrical properties of CNT PNCs via preferential alignment achieved during either synthesis (see Secs. 4.2.1 and 4.2.2) or postsynthesis approaches using mechanical [89–92], magnetic [93–95], and electrical fields [96–100].

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CASE STUDY 3 Ice Protection System for Unmanned Aircraft Systems Ice protection systems (IPSs) are used on most aerovehicles to remove and/or prevent the formation of ice on aerodynamic surfaces. Ice buildup introduces increased weight and drag, thereby reducing the performance and efficiency of the vehicle and, in extreme cases, yielding stall conditions or otherwise causing lifting failures. IPSs take different forms, including de-icing fluids used at airports during freezing weather. In noncommercial aerovehicles, bleed-air or thermal-blanket systems are often integrated into the design, requiring power during flight to heat the aerosurfaces (wings and control surfaces). MIT and Metis Design Corp. have collaborated on several U.S. Air Force and U.S. Navy programs to leverage multifunctional nanoengineered composite properties for IPSs and structural health monitoring. A successful round of ice-tunnel testing on a production unmanned aircraft system (UAS) aerosurface (see Fig. 4.8) to validate a CNT-based IPS in actual heavy icing conditions was recently completed, demonstrating a TRL of 7. Aligned CNTs are applied to the exterior aerosurface where ice forms and used as an efficient resistive heater. Prior work has shown that CNT networks in composites allow for very efficient heating [102]. The surface heat flux of the “nanostitched” [101] IPS can be tailored by controlling the CNT growth parameters to produce a desired resistivity. This integrated approach is particularly attractive to mass-critical UASs that depend on low fuel consumption for extended missions and are sensitive to disturbances (e.g., ice) to laminar flow of the highaspect-ratio airfoils. Such nanostitched laminates are being investigated for both structural and multifunctional enhancement, focusing on reinforcing the problematic interlaminar region between plies of composites (see Chapter 2) with many nanofibers, greater than 10 billion CNTs per cm2.

Fig. 4.8 IPS demonstration on UAS aerosurface. (Figure courtesy of # MIT and Metis Design Corporation.)

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In addition to electrical properties of CNT PNCs, others have focused on tailoring and enabling thermal [102–104] and actuation characteristics. Although still at the PNC level, the vision is to create smart nanoenabled materials with new capabilities and performance levels. Most of these efforts have been directed toward the development of enhanced nanocomposite thermal conductivity for thermal management applications (e.g., heat dissipation, or curing efficiency), whereas others have focused on coupled material response for actuation. Examples include energy-harvesting applications based on a thermoelectric response [105], shape-memory PNCs [106, 107], ionic actuators [108], and piezoelectric and electrostrictive PNC actuators [109– 111] where the CNTs primarily contribute by increasing dielectric constants. As a specific example in the shape-memory actuation area, radio frequency response of CNTs has been used in conjunction with enhanced thermal conductivity of shape memory PNCs to improve shape memory response efficiency and enable remote actuation [112]. Although initial proof-of-concept demonstrations of these multifunctional nanocomposite materials have been made for a variety of useful physical properties, integration and demonstration in aerospace systems is generally quite nascent, as is modeling of the physics governing the behaviors.

4.4 PROCESSING AND MANUFACTURING Although many of the hierarchical assemblies (see Sec. 4.2.3) with components discussed previously (see Secs. 4.2.1 and 4.2.2) require specialized processing techniques (reviewed in Chou et al. [5]), by far the largest area of contribution in the processing of nanomaterials for structural and multifunctional materials is mixing. Mixing must overcome the high surface-to-volume ratios of the nanomaterial constituents as discussed previously, which can be avoided by aligning nanoscale fibers to create high surface-to-volume capillaries to assist wetting and avoid mixing. However, the most used and researched processing technique is mixing to incorporate the nanomaterials into the polymer to create a PNC. There are multiple ways to process different PNCs, but the goal of all processing techniques is to disperse the nanomaterials evenly throughout the polymer matrix and break up the nanomaterial (most are agglomerated) so that it is actually a nanoscale component (i.e., the nanomaterial presents its high surface-to-volume ratio to the polymer in the same way microfibers are bonded to their polymer matrix in PMCs). This section will discuss the major processing techniques commonly used today. According to Koo [113], “solid thermosetting reactive prepolymers and thermoplastic polymers with solid nanoparticles” are best suited for the following processing techniques: solution intercalation, melt intercalation, roll milling, and centrifugal mixing. On the other hand, “liquid thermosetting reactive prepolymers and thermoplastic polymers with solid nanoparticles” lend themselves to the following processing techniques: sonication, in situ polymerization, emulsion

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Fig. 4.9 Solution intercalation of organic montmorillonite (OMMT) and PA12 thermoplastic [115]. polymerization, and high-shear mixing. The techniques can be grouped into either primarily mechanical or chemical in nature. This section will briefly describe these processing methods and provide examples of how these techniques have been used to realize different PNC systems, often in combination (e.g., sonication and shear mixing).

4.4.1 CHEMICAL PROCESSING Solution intercalation processing depends on the nanoparticle being a layered silicate. The layered silicate is swollen using a solvent, into which the polymer is also soluble. Swelling the layers of the nanoparticle allows the polymer access to the space in between the silicate layers and thus intercalates them. The solvent is then removed, and the intercalated silicate and polymer structure remain [114]. Figure 4.9 shows how solution intercalation generally works for PNCs. Solution intercalation is a good processing technique because of its relative simplicity. It does not require large amounts of expensive equipment to perform, and it is also a good technique for producing thin films of PNCs with oriented nanoparticulate layers. However, because it does require the use of solvents, many of which are hard to work with, it is not considered an environmentally friendly method. For this same reason, it is not scalable because dealing with large amounts of organic solvents is cost prohibitive on a commercial scale. In situ polymerization initially mixes a monomer with the nanoparticle being used. The addition of the monomer swells the layered silicate, creating small openings between the layers of the nanomaterial. The monomer is then polymerized through a variety of methods. During the polymerization, the growing polymer uses the small openings to grow into the nanomaterial, thus intercalating it. As this progresses and more polymer grows in between layers of the nanomaterial, the nanomaterial becomes more and more dispersed. The polymerization reaction can be driven by a variety of factors, including heat, radiation, or catalysts. As an example, in situ polymerization was used to produce polyethylene-clay

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nanocomposites with Cloisite 93A clay (Southern Clay) [116]. In situ polymerization is an effective method of creating certain PNCs. It has the advantages of being able to be scaled up for large volumes of production. However, it does not always provide the necessary amount of dispersion of nanoparticles because there is no shearing force that separates the tightly bound layers. The emulsion polymerization process is somewhat of a mix of in situ polymerization and solution intercalation wherein a monomer is added to an emulsion that contains the nanoparticles. The nanoparticles are swollen by the emulsion, which then allows the monomer, when an appropriate initiator is added, to grow into a polymer. The growing polymer is then able to intercalate itself throughout the nanoparticle layers that are present (e.g., a polystyrene-SiO2 microsphere PNC was created) [117].

4.4.2 MECHANICAL PROCESSING Roll milling uses multiple cylinders rotating in different directions to mix nanomaterials with a given polymer. Figure 4.10 shows a three-roll mill and how the various cylinders rotate in relation to each other. The roll mill, like melt intercalation with mixing, uses shearing to disperse nanomaterials throughout a polymer matrix and has been used to disperse multiwall CNTs in epoxy, where it was found that higher aspect ratio CNTs generally resulted in greater shear alignment [118]. As another example, roll milling was used to disperse CNTs [multiwall nanotubes (MWNTs) and double-wall nanotubes] within a styrene-free resin (POLIYA 420) and a vinyl-ester-epoxy resin (POLIPOL 701) [119], and the authors noted that roll milling can disperse the CNTs without damage to the CNTs as in in other processing studies. They also note that the three-roll mill used, much like the one pictured in Fig. 4.10, produces almost entirely shear loads with very few compressive loads. Roll milling is an effective processing technique for some polymer nanocomposites and, like melt intercalation, is environmentally friendly in that it uses no solvents. Melt intercalation, unlike solution intercalation described previously, does not rely on a solvent to disperse the nanomaterials but rather uses heat and kinetic energy to break apart the nanomaterial into individual nanoparticles while allowing the polymer to inject itself between the nanoparticles. In addition to this

Fig. 4.10

Roll-milling equipment and schematic.

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Fig. 4.11

Mixing mechanisms of co-rotating screws.

method being compatible with several polymers, including polyethylene oxide, polystyrene, and most styrene derivatives, melt intercalation has the added benefit of being environmentally friendly because it does not require a solvent. As mentioned previously, melt intercalation in addition to heat also uses kinetic energy to disperse the nanomaterials within the polymer matrix. One or two rotating screws with varying configurations usually provide this kinetic energy. As an example, melt intercalation was used to fabricate a MWCNT-thermoplastic elastomer PNC [120]. Poly[styrene-b-(ethylene-cobutylene)-b-styrene] triblock copolymer (SEBS) from Asahi Kasei Corporation (Japan) was mixed with high-purity MWCNTs from Nikkiso Co., Ltd. (Japan). The researchers’ aim was to create an elastic, conductive PNC. They found that they were able to disperse the MWCNTs uniformly throughout the SEBS matrix using high-shear melt intercalation. A related processing route is extrusion, and within the two-screw extruder class there are two primary types: co-rotating screws (Fig. 4.11) or counter-rotating screws. The sizes of twin-screw extruders can vary considerably depending on polymer volume and required shear rates. Overall, melt intercalation is a very powerful form of processing for PNCs. It is environmentally friendly, scalable, and effective at dispersing a variety of nanomaterials in polymer matrices. In high-shear mixing processes, the liquid polymer is mixed with the nanomaterial and then run through a high-shear mixer to disperse and exfoliate the nanomaterial. The mixer does this by shearing the nanomaterial so that the polymer can then incorporate it. An example of a production-scale high-shear mixer is shown in Fig. 4.12, along with two rotor configurations that physically shear the polymer-nanoparticle mixture. The mixing process can be performed in a loop so that the polymer-nanoparticle liquid can be processed as many times as needed. As an example, high-shear mixing and ultrasonication were both used to create a composite out of modified montmorillonite clay and epoxy resin [121]. The processing included high-shear mixing at 15,000 rpm for 30 min in an ice bath to maintain the mixture at 658C, followed by degassing of the PNC in a vacuum chamber for 12 h. The researchers found that the high-shear mixing samples showed a higher degree of dispersion than those created using ultrasonication. High-shear mixing is generally considered an effective processing technique for PNCs as it is very effective in exfoliating nanoparticles within polymers due to the extremely high shear rates involved. It is relatively scalable, but the number of times a polymer-nanomaterial mixture can be cycled through such systems is limited due to viscosity increase.

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Centrifugal processing is used primarily in powder-to-powder processing, whereas sonication is typically used with liquid media. As an example, powder-to-powder blending using a rotation and revolution Thinky ARE-310 mixer (centrifugal processing) was used to combine PA11 thermoplastic and MWNTs [122]. The PA11 powder was mixed gradually in three steps of 30 g, 30 g, and 39 g with 1 g of MWNT (based on the desired final loading of MWNTs, where the total mixture is 100 g in this example), each time rotating at 2000 rpm for 10 s, resulting in PA11-1wt % MWNT mixtures. Substantial improvements in electrical conductivity were observed with increasing loading of MWNT. In general, centrifugal processing is a very effective powder-to-powder and powder-to-liquid mixer and also an environmentally friendly method. Sonication or ultrasonication are mostly laboratory-type processing methods. They may involve solvent as a medium, the sonication time may be lengthy, and they are generally not considered scalable to commercial usage.

4.5 MODELING OF NANOMATERIALS: MULTISCALE AND MULTITECHNIQUE In the last decade, a considerable amount of research has been focused on the modeling and simulation of nanostructured materials [123, 124]. The materials that have generated the most attention in the modeling and simulation community are nanoparticle-reinforced polymers and nanostructured metal alloys. Because of the large differences in length scales between the nanoscopic material

Fig. 4.12 PreMax High-Shear Mixer by Ross and two examples of mixing heads: quad-slot (top) and megashear (bottom), also by Ross.

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Fig. 4.13 Multiscale modeling of structural materials emphasizing time and length scale differences. (Figure from Gates and Hinkley [130].) structure and the macroscopic material behavior, a wide range of modeling tools must be used for developing structure-property relations; particularly difficult is the area of failure. This is illustrated in Fig. 4.13 where, at very small length (10212 m) and time (10215 s) scales, matter is best described in terms of subatomic particles and quantum phenomena. Of greatest concern in this regime is the molecular assembly of materials, and this is often simulated using density functional theory [125]. At length scales of 1029 m and time scales 1029 s, matter is best described in terms of atoms, molecules, and chemical bonds. The prediction of molecular structure and behavior is often achieved in this regime using molecular dynamics (MD) simulations [126]. At microscopic length (1026 m) and time (1026) scales and macroscopic length (.1023 m) and time scales (.1023 s), matter is best described as a continuous medium. Structural mechanics tools (e.g., finite element modeling and micromechanics) are usually used in this regime to predict material behavior [127–129]. Because very different simulation techniques are required to describe the behavior of matter at different length and time scales, multiscale modeling techniques are being developed to relate material structure to macroscopic material behavior. Multiscale modeling techniques often incorporate multiple simulation techniques into a single algorithm. Because the quantum-, molecular-, and continuum-level simulations require very different approaches and assumptions, efficient and accurate multiscale modeling is still limited. The first reason for the difficulty relates to the intensive simulation times that are required for quantumand molecular-level simulations. For example, simulated times of 1 ns in molecular dynamics simulations can take days or weeks to complete for simulations of tens of thousands of atoms. Second, information that is passed between simulations is often lost or recast due to the characteristics of different modeling

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techniques. For example, the smooth transition of stress waves between molecular dynamics and finite element simulations has been a long-standing challenge. Finally, the use of multiscale simulation approaches requires expertise in multiple simulation techniques. Most researchers can spend an entire career mastering a single class of simulation techniques. It is challenging to train researchers who have a deep understanding of multiple techniques when the underlying science of each technique is highly unique. Traditionally, quantum-level computing has been dominated by physicists and chemists, MD by chemists, and finite element modeling by engineers. Despite these challenges, great strides have been made in the multiscale modeling community toward bridging scales across those shown in Fig. 4.13. The first great achievements were in the development of modeling techniques for simple crystalline materials [131]. These approaches related the positions of individual atoms in molecular models to points in continuous media (e.g., nodes in finite element analysis). These modeling approaches are continuously being refined. Other efforts are focused on modeling and simulation of polymer nanocomposites and nanostructured metals [123]. Modeling and simulation of these materials is complicated by the presence of complex polycrystalline or amorphous molecular structures. These efforts are usually focused on predicting the thermomechanical response of materials as a function of molecular structure. Because of the great interest in predictive modeling and development of nanostructured materials, such modeling and simulation efforts are expected to increase in the coming years, as modeling has been identified as a key enabler of nanotechnology innovation [132].

4.6 FUTURE TRENDS The ability to control matter at finer and finer scales and with greater precision in morphology, extending to three dimensions and nm-scale heterogeneity, continues to progress, potentially leading to atomistic-level assembly methods in the near future. Focusing as we have here on CNT-fiber based enhancement of structural and multifunctional properties for aerospace, we should return to Fig. 4.1 and the larger trend in aerospace structures: the continued growth of FRP composites. The S curve for composites in Fig. 4.1 and the complementary one for CNTs would suggest two different technologies. However, it is much more likely that CNTs (and nanomaterials in general) will enhance existing FRP composites and increase their capabilities and therefore use, effectively pushing farther up the technology S curve in all areas of aerospace. Thus, one can envision CNTs and other nanomaterials as continuing composites along the existing technology S curve rather than displacing them. The trend is likely toward greater and greater heterogeneity (graphene is a potentially useful new PNC component [133, 134]) and hierarchy to meet ever-increasing structural and multifunctional performance requirements. Many aspects of nanoscience

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and nanoengineering need to continue to advance, as discussed herein, including manufacturing, modeling, metrology, environmental health and safety understanding [135], and “big data” in both modeling and experimentation.

ACKNOWLEDGMENTS Brian L. Wardle thanks Seth S. Kessler (Metis Design Corp.), A. Kinloch (Imperial College London), David Lashmore (NanoComp Technologies), Suraj Ruwal (Lockheed Martin Corp.), Stephan Sprenger (Evonik), Roberto Guzma´n de Villoria (MIT), and Noa Lachman [Massachusetts Institute of Technology (MIT)] for assistance with some portions of the chapter and the Fall 2013 MIT 16.223 class for critical review of the chapter. Some sections are based on original work by Wardle [136, 137].

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CHAPTER 5

Subsonic Aircraft Materials Development Michael Mohaghegh Seattle, Washington

David L. Stone† Redstone Arsenal, Alabama

Antonio F. Avila Universidade Federal de Minas Gerais, Belo Horizonte, Minas Gerais, Brazil

5.1 INTRODUCTION Since the first days of powered flight, aircraft designers have focused on achieving minimum weight, both in airframes and in propulsion systems. During the first three decades of 20th century, the absolute minimum weight was required for practical flight, due to the limited capability of the available propulsion systems. Thus, the strength/weight ratio was the leading driver for materials selection for both engines and aircraft. Even now, this consideration continues to be a key issue—lightweight is now necessary but not sufficient. The development of new and improving structural materials has played a significant role in the evolution of flight. When the Wright Brothers made the first powered flight over 100 years ago, their airplane structure was made entirely of wood, a natural composite material composed of cellulose and lignin. The only aluminum on that airplane was an aluminum-copper alloy engine crankcase casting. After that, aluminum usage gradually increased and, in 1919, an all-aluminum airplane, the Junkers F-13, was introduced. The selection of structural materials for aircraft then remained contested throughout the 1920s with designers such as Anthony Fokker proposing wood construction, and others like Hugo Junkers advocating for aluminum. Although the structure was sometimes heavier and more expensive to fabricate, aluminum initially benefited from the perception of being modern and the “wave of the future.” Wood and aluminum, however, co-existed as essentially equivalent structural choices until the pendulum eventually swung in the favor of aluminum as larger production airframes began to offset the increased tooling costs required for metal construction.  The

Boeing Company. US Army Aviation & Missile Research Engineering Center.

†

Copyright # 2017 by the American Institute of Aeronautics and Astronautics, Inc. The U.S. Government and The Boeing Company have a royalty-free license to exercise all rights under the copyright claimed herein. All other rights are reserved by the copyright owner.

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Aluminum materials historically have been the primary material for aircrafts construction, but the more demanding requirements of modern aircraft have called for materials that have higher specific strengths and stiffnesses [1]. In some instances, composite materials have exceeded the capability of conventional aluminum in aircraft construction and it is for this reason composite materials have come of age and have continually replaced more aluminum structures. In 1965, a U.S. Air Force study predicted that these high strength/ weight composites would transform airplane structures and bring about a 35% weight savings relative to aluminum structure. The use of composites has gradually increased, especially since the 1970s, when carbon fiber reinforced plastics (CFRP) began to make inroads. In the mid-eighties, with two decades of development beyond what was possible in 1965, the US Air Force believed that a 50% reduction in airframe structural weight could be achieved with novel design practices coupled with advanced materials. That audacious goal drove many of the science and technology efforts in composite materials for the next twenty years. The weight advantage for carbon fiber composites has been typically between 20% and 30% compared to aluminum for idealized stiffened panel designs. However, when the design is fully defined for resistance to environmental exposure, repair, electro-magnetic effects, and aircraft systems interface requirements the expected weight advantage may not be so dramatic [2]. Extensive development programs were introduced in the aircraft industry aimed at a step-wise introduction of CFRP into larger and larger structures. Figure 5.1 captures the material usage by aircraft designed and manufactured by Boeing. Evolution of airframe materials continues as technology changes. Design solutions will change as the competition between the material systems continues. The current focus for aluminum alloy technology development is increased design value properties while the focus for composite materials is lowering costs associated with materials, manufacturing and repair. A more challenging alternative material solution involves the hybridization of metals and composites in sandwich-like constructions. Airbus is using Glass-Fiber Reinforced Aluminum (GLARE) for the A380 fuselage’s upper and lateral shells. There have also been significant developments with a relatively new concept of titanium foil and carbon fiber reinforced plastics (TiGr). The use of the titanium foil in the laminate facilitates the ability of the composite structure to withstand greater fastener bearing loads in joint locations. This chapter will discuss the evolution and future potential of structural materials for subsonic, supersonic, and rotorcraft air vehicles. It will show how the relationship between the airframe designer, material scientists and manufacturing engineer is becoming closer and more integrated as the need for improved weight/cost performance grows. This approach to aircraft material development is shown in Fig. 5.2.

SUBSONIC AIRCRAFT MATERIALS DEVELOPMENT

Fig. 5.1

Fig. 5.2

Structural materials on selected Boeing commercial aircraft.

Integrated design/manufacture/material development.

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5.2 SUBSONIC AND SUPERSONIC FIXED WING AIRCRAFT Subsonic airplanes are used for both commercial and military transport and come in many sizes and configurations. Material performance attributes include strength, stiffness, density, fatigue and fracture toughness, and corrosion resistance. These attributes support the design requirements for static strength, fail safety, large damage capability, durability and damage tolerance. In addition, materials also must be producible with consistent high quality.

5.2.1 PRELIMINARY CONSIDERATIONS Current design criteria have become more complicated in recent years, requiring more innovative solutions and improved materials and processing methods to accommodate the challenges [3]. The material property requirements for use in airframes vary depending on the particular component under consideration. The fuselage can be seen as a semi-monocoque structure that is composed of skins to carry the cabin pressure (tension) and shear loads, longitudinal stringers to carry the longitudinal tension and bending compression loads, circumferential frames to preserve the fuselage shape and reallocate loads into the skins, and bulkheads to clutch concentrated loads including those associated with pressurization of the fuselage. The wing is in effect a beam that is loaded in bending during flight [4]. The wing supports both the static weight of the aircraft and any additional loads caused by maneuvering or turbulence. Extra wing loads originate from the takeoff and landing operations and from the landing gear during taxi. Considering the wing aerodynamics, it is possible to infer that the upper surface of the wing is primarily loaded in compression because of the upward bending moment during flight but can be loaded in tension while taxiing. The stresses on lower part of the wing are just the opposite. The tail of the airplane, also called the empennage, consists of two stabilizers (one horizontal and another vertical), and control surfaces (e.g. elevators and rudders). Structural design of both the horizontal and vertical stabilizers is basically the same as for the wing. When analyzing each major part of an aircraft, performance limiting parameters that drive the design need special consideration and should be addressed by the application of the right material set and the optimum structural design philosophy. For instance, fracture toughness is the primary material driver for the fuselage design, but competing parameters such as strength, stiffness, fatigue and corrosion are also important design considerations. Furthermore, when dealing with fatigue, crack initiation and crack growth rate cannot be neglected. For wings the design is influenced by material properties such as stiffness in compression and compressive yield strength, with fatigue resistance still a critical parameter due to the alternating flight loads. Likewise for the empennage, as a result of the compressive loads in the upper and lower surfaces of the horizontal stabilizer as a result of bending, stiffness in compression is the most important material property. From these examples, one can infer that to be able to develop an integrated

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approach to aircraft material development, the criteria for design and material selection must be taken into consideration.

5.2.2 DESIGN CRITERIA Aircraft design philosophy is focused on their ability to deliver a safe airframe while offering long-term reliable usage, with minimum weight, cost, and maintenance [5]. This consistent top-level philosophy along with well-established practices and procedures have evolved and governed the structures engineering approach to airframe design and validation through generations of jet transports including Boeing commercial aircrafts (707, DC8, 727, DC9, 747, 737, DC10, 767, 757, 717, 777 and 787). The airplane design criteria comprise ten requirement categories as depicted in Fig. 5.3. The categories are: design loads, materials, stiffness, strength, durability, fail safety/damage tolerance, crashworthiness, producibility, maintainability, and discrete effects. These criteria come from highly integrated set of design, manufacturing and certification requirements needed to produce safe and efficient aircraft.

Fig. 5.3

Top level structures design criteria.

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5.2.3 QUALIFICATION OF NEW MATERIALS The Boeing success has been enabled by the company’s leadership role in the partnership with raw material suppliers that directed continuous development activities in materials science [6]. Commercial airplanes are among the most complex pieces of engineered capital equipment. The material selection for each component of an airplane defines the manufacturing and operational performance characteristics. Specifically, the cost and weight of the vehicle will be established by these choices. Materials, fasteners and standard parts selection must be governed by two primary criteria: 1) they must be the best suited for their intended functions and the environment in which they will operate, and 2) a credible and reliable database or alternate means of compliance (e.g., point design) must be in place at the time of design to ensure readiness for implementation and certification. Expanding on the two primary criteria above, there are several major factors to be considered in the selection process including: a reliable supplier base, performance, operating environment, service conditions, producibility, and business and other regulatory issues. Tables 5.1 and 5.2 highlight the material and processes checklist of data that

TABLE 5.1

METALS MATERIAL/PROCESS CHECKLIST

Materials/ Processes

Producibility

Static

Material specification

Forming

Tension

Process specification

Machining

Compression

Corrosion property

Trimming

Shear

Repair specification

Joining

Bearing

Assembly

Buckling

Chemical processing

Crippling

Real time process control

Joint

Inspection

Environmental factors

Disposal Cleaning

Damage Tolerance and Fatigue Fatigue crack Growth rate

Air Environment KIC

Residual strength

Stress corrosion

KA KISCC

Incidental damage Fatigue strength

Open hole

Joints

Fatigue factors

Finish Environment

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TABLE 5.2

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COMPOSITES MATERIAL/PROCESS CHECKLIST

Materials/ Processes

Producibility

Static

Material specification

Layup

Laminate

Process specification

Cure

Part specific layup

Repair specification

Handling

Joint

Finishing

Interlaminar shear

Machining

Crippling

Joining

Environment factors

Assembly

Sandwich

Real time process control

Damage Tolerance and Fatigue Damage tolerance Delamination Damage growth

Impact Notch Delamination

Residual strength

Impact Notch Durability

Chemical safety

Post impact

Inspection

Open hole

Disposal

Bearing Environment factors

are required in order to commit to the material selection of metallic or composite material, respectively.

5.2.4 QUALIFICATION OF NEW DESIGNS Starting with the design loads, the interaction between the basic structure and each system must be established. Airplane structure, including propulsion, landing gear, interior, high-lift and control components, must be designed to sustain applied loads determined in accordance with the requirements of Federal Aviation Regulations (FARs) and Joint Aviation Regulations (JARs) for commercial aircrafts along with the unique original equipment manufacturer Design Requirements and Objectives. These documents describe the desired operational and performance goals for the unique model design. These goals include, but are not limited to, mission capability, payload and range capacity, design life, certification criteria and unique design objectives to establish product differentiation. The loads for ultimate, fail-safe, discrete source damage and operational cases are calculated for structural design. Typically these loads are associated with flight

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(maneuvers, gusts, and engine and control system failures) or ground (landing, taxi, and ground handling) conditions. Analysis models include aerodynamic, structural, and mass and system representations with some data generated in wind tunnel and laboratory testing. Analysis results are generally validated with ground and flight load testing. Airframe structure must be able to support limit loads without detrimental permanent deformation and ultimate loads without failure for at least three seconds. Design is verified by analysis and is further validated by selected detail, panel, component and full-scale airplane tests. Deformation must be fully accounted for when analysis is used. For composite structures, the effect of environment, manufacturing defects, and bolted assembly must be considered. For the analysis of multiple load path structure, “B” basis allowables are used, which are based on 90% reliability and 95% confidence. For single load path structure, the practice is to use “A” basis allowables based on 99% reliability and 95% confidence. Allowables are given in MMPDS (formerly MIL-HDBK-5) and CMH-17 (formerly MIL-HDBK-17) for many existing materials [7, 8]. Even though there is generally a high level of confidence in the static strength analysis, some results are still validated by subcomponent, component and/or full-scale testing. Fatigue is the first consideration in the durability of metallic structure. All primary airframe structure must be designed to resist fatigue damage for the expected full service life, which includes both operational and training flights. Fatigue analysis is validated through the use of small specimens, subcomponents, components and airplane fatigue tests. Lessons learned from testing, service experience and teardown results are continuously incorporated into the durability technology standards documents. This philosophy is based on the development of aluminum primary aircraft structure. Carbon fiber materials are significantly less susceptible to traditional fatigue damage and require different methodology for repeated load evaluation. For carbon fiber, a no-growth philosophy is utilized. With this approach, the composite structure is tested with defects up to the manufacturing process capability and nondestructive inspection (NDI) limits and impact sufficient to produce damage just at the threshold of visibility. This testing must show that the defect does not grow under fatigue cycling and that the structure remains capable of ultimate load after fatigue cycling is complete. For visually detectable impact damage, it must be shown that damage does not grow for two inspection intervals and that limit load capability is maintained with damage present. Corrosion prevention is another key aspect of the durability for aluminum primary structure. The structure must be designed to resist corrosion and wear damage to the levels specified in the design requirements and objectives. Adequate finishes, sealing and drain paths must be specified. The corrosion performance of materials is predicted by small specimen tests in the laboratory and validated by service performance based on operator feedback. Lessons learned are continuously incorporated into designers’ corrosion prevention design guides. The

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structure must have corrosion prevention and control program as part of the specified maintenance program. Carbon fiber structures are not subjected to the same detrimental effects of corrosion as that of aluminum structures, however hot-wet and other environmental conditions, such as ground air ground (GAG), hygrothermal cycling and ultraviolet (UV) light exposure, must be accounted for in the design of the material system. For commercial and some military aircraft, all primary flight-loaded structure must be designed to be damage tolerant. CFRP structure follows the no-growth approach outlined above. For metallic structure, this requires that the structure have sufficient damage growth properties and detection characteristics so that if damage were to develop at single or multiple sites, specified airline inspections would ensure that the damage is detected before the fracture propagation at operating loads reduces the residual strength capability of the structure below regulatory load requirements. Primary flight-loaded structure, including trailing edge flaps and control surfaces, may be required to be designed to be fail-safe, that is, with sufficient residual strength to carry limit load with failure or obvious partial failure of a principal structural element. This requires that the structure have multiple elements or redundant load paths and have adequate damage containment capability for a period of un-repaired use. Limit load conditions must consider tension, compression, shear, internal pressure, combinations of these loads and, when appropriate, the presence of an active crack tip in determining the residual strength. Fastener capability must be evaluated to ensure load redistribution requirements are met. For commercial and some military aircrafts all primary structure must be designed to be damage tolerant and safe life if damage-tolerant design is impractical (i.e. typically landing gear structure). Analysis verified by testing must validate safe-life structure. The typical design life limit will be much smaller than the test demonstrated life as a result of the application of appropriate scatter factors. Design criteria used for validating the ability of an aircraft to withstand an emergency landing are sets of rules accrued from experience; the rules lead to designs that afford occupants with a reasonable chance to escape injury in the event of an emergency landing. Existing designs have evolved over time using regulatory requirements for emergency landings and retentions of items of mass. The modern commercial fleet, in general, has demonstrated a reasonable level of safety in the event of emergency landings. Protection of the occupant is important and was part of the basic design process for the Boeing 787. Material selections must be accounted for in determining the effect on aircraft response during an emergency event. The structure must be designed to ensure a high level of producibility while considering cost, weight, and performance constraints. Design concepts must be coordinated with manufacturing engineering, tooling, and vendors to ensure their concurrence with the product definition before release. As a method of

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controlling cost and aiding operators with multiple airplane model fleets, designs should strive for commonality across models. The airframe must be designed with appropriate access to the structure to allow for inspection. For primary structure, it must be a prime design objective to ensure that the most critical element of any joint, splice, attachment or multilayer structure is accessible and visually inspectable, preferably from outside the airplane. All designs must address the need for repair. Consideration must be given to cost, repair time and spares requirements. Specific requirements address the environment and discrete events an airplane may experience in service. These include provisions for hail, lightning strike, bird strike, rapid decompression, engine blade loss, and tire rupture. The materials and design features of any new or derivative airframe structure will generally be a mixture of existing proven elements and new features being used for the first time. In each case, it is essential to validate that the structure will perform as intended as early as possible in the design process, preferably before the design has been committed to production. The primary effort will be to ensure that the available database justifies the proposed design. For new materials and design features, the available database may well be limited or nonexistent. In this case, the need for validation is more imperative and is more difficult to achieve. Therefore, selection of new materials and concepts should always be made early enough in the airplane program to permit a reasonable level of validation through preliminary analysis and testing prior to commitment to the design. Some risk will always remain until full-scale tests and service experience have fully validated the new concepts. In developing structural substantiation, the building block approach in Fig. 5.4 is utilized. This is a systematic way of obtaining material allowables and larger scale design values, and provides the knowledge necessary to validate the design. The building block approach is particularly important when using new materials, analyzing materials that may have multiple failure modes.

5.2.5 EVOLUTION OF ALUMINUM MATERIALS New alloys and tempers have been continuously evolving since the first use of aluminum in airplane structures. More recently, aluminum alloy research and structural design has focused on matching the alloy’s optimum properties to the various components’ requirements. 5.2.5.1

ALUMINUM DEVELOPMENT

Pure aluminum is too soft for most structural applications and must be alloyed with other elements (e.g. Cu, Mn, Zn, Li) to increase the strength and other engineering properties [9]. The optimum strengthening of aluminum is achieved by alloying and heat treatments that promote the formation of small, hard precipitates which interfere with the motion of dislocations. As those alloys are

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Fig. 5.4

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Building block approach for airplane design.

polycrystalline materials, in most of the cases, the heat treatment leads to hardening by precipitation, which leads to an increase in strength. The down side of such technique is the number of discontinuities created within the material, which act as nucleation sites for pit and crack origins. The application of aluminum alloys for aircraft structures and components has become standard due to their lower cost of manufacturing, life cycle costs, and low substitution risk [3]. Another important issue that promotes the use of aluminum alloys in the aerospace industry is the ability to tailor the microstructure of these alloys during heat treatment. Different microstructures will lead to distinct results with the size of the grain structure influencing the overall fatigue performance. Fine grain size with a uniform dispersion of small, hard particles will inhibit dislocation motion, and consequently increase strength. High durability and toughness in aluminum alloys is possible when fine structures with clean grain boundaries and no large particles are present in these alloys. Microstructure plays a key role in fatigue durability. To avoid crack initiation, a fine grain size and no surface defects is the ideal. However, the inclusion of large particles in the grain structure can provide the benefit of blunting propagation once the crack is present. Details of microstructure-property relationships in aluminum alloys are given in Chapter 2.

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DEVELOPMENT TIMELINE

The Wright brothers used an aluminum-copper casting for the crankcase of their engine. The need for low weight meant that stronger alloys were needed before aluminum could be used as a primary structural material. Such a material, duralumin, became available when Alfred Wilm discovered that aluminum-copper alloys could be precipitation (age) hardened. Duralumin was the alloy used by Junkers for his all-aluminum F.13 airplane. In today’s classification system, duralumin is comparable to 2017-T4. The need for lighter structure has since driven the development of higherstrength alloys. Alloy 2024, which was the structural alloy used for the venerable DC-3 airplane, was introduced in 1931. It is essentially an aluminum-coppermagnesium alloy. A few years later, it was discovered that the aluminum-zincmagnesium system would yield even higher strengths. Alloy 7075, the first Al-Mg-Zn, became available in 1943. Alloys 2024-T3 and 7075-T6 are both still common choices today for damage tolerant and strength designed structure, respectively. Since the 1940s, many aluminum alloys (some more successful than others) have been applied to airplane structure. Figure 5.5 shows the common aircraft alloys along with the year of first flight for the commercial airplanes on which they were introduced. Some of the early lessons learned had to do with stress corrosion and fracture toughness. For instance, alloy 7079-T6 forgings, introduced in 1954 and used in the early 1960s, experienced early stress corrosion cracking (SCC) fractures and had to be redesigned. This was not predicted by the standard stress laboratory tests and the only clue available at the time was some early failures in industrial atmosphere. In addition, alloy 7178-T6 plate was used for the wing upper surfaces of the 707, 727, and the early models of the 737. The use of this alloy was eventually discontinued for new Boeing designs because its damage tolerance was not well balanced relative to its high strength. Alloys 2024-T3 and 7178-T6, were the highest strength structural aluminum alloys in 1965 when the U.S. Air Force made its prediction that CFRP potentially

Fig. 5.5

Introduction of new aluminum alloys.

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• Additive manufacturing • Nanotechnology aluminum alloys

Mostly uncharted Known territory

• New alloys (low Fe+Si standards) • Multi-alloy friction stir welding

• Aluminum mill fabrication processes • High stiffness alloy technology

• Advanced machining processes • Advanced structural adhesive bonding

Weight and cost Improvement

• Advanced welding processes • Target properties for advanced alloys

• Integrally stiffened panels, spars, ribs • Thick aluminum plate (large ingot technology)

• Minimum tooling build processes • High speed machining

• Mechanically fastened spanwise joints • Friction stir welding integrally stiffened panels

• Large, integrally stiffened exhaust panels • High strength and high damage tolerant alloys implemented

• Monolithic subcomponents facilitated by aluminum plate • Determine assembly process

Time

Fig. 5.6

Advancing aluminum technology.

could bring about a 35% weight saving relative to aluminum. Since 1965, several advancements in aluminum technology have taken place, some of which have been implemented and some of which are still to be brought to market. Figure 5.6 shows a graphical representation of the aluminum technology advancement over time. In the late 1960s and early 1970s, Boeing discovered that the damage tolerance of any aluminum alloy could be improved by reducing its levels of the tramp elements Fe and Si [10–11]. This opened the way for further improvements in aircraft aluminum. Boeing, in conjunction with the major aluminum companies in the U.S., developed Alloys 2224, 2324, and 7150 in the late 1970s. Thermo-mechanical processing was introduced to raise the strength of the above alloys beyond what would otherwise be possible. Alloy 2224 extrusions are extruded with speeds and temperatures adjusted such that an unrecrystallized, fibrous, grain structure is maintained through to the final product. The 2324-T39 alloy plate is cold worked after heat treatment to raise strength. The 7150 alloy is an un-recrystallized version of the thick plate alloy 7050, aged to peak strength. The application of these alloys led to a 6% weight reduction on the wing structure on the 757 and the 767 airplanes when compared to previous Boeing aircraft models. Another major step in the evolution of aluminum was the introduction of the -T77 temper [12] (U.S. Pat# 4,477,292, M. H. Brown, 10/15/1981), which is a three-step aging treatment that produces very good corrosion properties

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without the strength drop normally associated with -T7 type tempers. The application of the -T77 temper to Alloy 7150 made possible the use of this highstrength alloy in the fuselage structure as body stringer material on the 777 airplane. The fuselage skin on the 777 is made from another ‘clean’ version of 2024, designated 2524. The 2524 alloy has the same strength as 2024, but the highdamage tolerance of 2524 saves structural weight. Improved knowledge about the Al-Zn-Mg system, combined with the -T77 temper development was taken advantage of for the 777 wing upper surface. The Alcoa developed alloy, 7055-T77, is a high zinc alloy, stronger than 7150, and more corrosion resistant than 7150-T651. 5.2.5.3

CURRENT DEVELOPMENTS

The challenge of aluminum alloys development is to put in practice the basic metallurgical concepts for optimizing properties of those alloys. This task in under way as significant improvement into aluminum alloys has been made since the 2014-T4 aluminum was introduced in the 1920s. Figure 5.7 shows some of the most commonly used aluminum alloys and their corresponding yield strengths [3]. The commercial and military demands for aircraft improvements led to the development of existent aluminum alloys and the development of new alloys with new heat treatments. One example of such development is the

Fig. 5.7

Aluminum alloys and their yield strength.

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Fig. 5.8

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Strength-toughness combinations.

2524-T3 that helped to eliminate the tear straps in a weight-efficient manner on the Boeing 777. Another example is the -T77 temper, developed by Alcoa to improve fracture toughness of 7150-T77. The 7150-T77 improvement can be credited to two factors; the controlled volume fraction of coarse intermetallic particles, and the un-recrystallized grain structure [3]. Figure 5.8 shows the strengthtoughness combination for several aerospace applicable aluminum alloys. Although significant improvement to the 2XXX alloy series has been made, usage of this material for fuselage skins brings the problem of intergranular corrosion. To overcome this for skins, the 2XXX alloys for fuselage skins must be clad. In addition, 2XXX alloys cannot be welded by fusion. An option is to use the 6XXX alloys since they are weldable and cheaper than 2XXX alloys. However, 6XXX alloys with high concentration of copper (e.g. 6013-T6 and 6065-T6) also are subject to intergranular corrosion. The reason for such vulnerability is due to the formation of precipitate-free zones at grain boundaries which are anodic with respect to the grains. An alternative to welding parts is to improve the alloy’s formability. By applying this approach, complex components can be formed from one single sheet. The 7XXX aluminum alloy series are considered superplastic forming alloys, possible of experiencing elongations up to 375% during uniaxial tests at elevated temperatures above 5008C [13]. Although the 7XXX alloy series have the advantage of enabling complex components, there is a weight penalty for their use.

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The search for a low density aluminum alloys with stiffness higher than conventional aerospace aluminum alloys led to development of aluminum-lithium alloys. Lithium can reduce the aluminum density by 3% and increase stiffness by 6% for very low weight percentage additions of lithium [14]. However, high lithium percentages have shown to be disastrous. Al-Li-Cu alloys (e.g. 2090, 2091 and 8090 alloys) with lithium additions of approximately 2% have demonstrated a plethora of problems [3]. Among those problems are crack deviations, low stress-corrosion threshold, and excessive anisotropy. To overcome these problems, the 2094 and 2195 aluminum-lithium alloys were developed. The 2094 alloy has yield strength of around 700 MPa and elongation of 10%, while 2195 offers a 12.5% lighter solution to 2219 with higher yield strength. The 2096, 2097 and 2197 alloys are representative of the third-generation Al-Li alloy. Improved fatigue strength allowed 2097/2197 alloys to replace the 2124 alloys for bulkheads on the F-16 [3]. The 2097 has a 5% density advantage over 2124 and a much larger fatigue stress spectrum. As a consequence of these improvements, the fatigue life of BL19 longeron of the F-16 has increased by approximately 100%. Another successful application of Al-Li alloys is the replacement of F-16 2024 skin by 2098 alloy. The 2098 skin has proved to have a lifespan close to 6 times longer than the 2024 skin used on F-16. The aluminum-lithium alloys C406 T8511 from Alcoa and 2196-T8511 from Alcan, developed for extrusion applications, are comparable to the conventional 7175-T73511 alloy [14]. These alloys were qualified for use in the airbus A380 floor beam applications (e.g. crossbeams, seat rails, false rails, cockpit and emergency bay floor structures), offering considerable weight savings. Another Al-Li alloy with promising properties is the C47A-T851, an AL-Li alloy with properties similar to 2524 aluminum alloy but with significant improvement in fatigue growth. Airbus was considering applying the C47A-T851 to the A380 fuselage. To successfully replace conventional aluminum in aerospace applications, Al-Li alloys must offer the correct balance of properties together with an approximate 10% weight savings [15]. For example, good options for replacing Al 2024-T3 in damage-tolerant structure such as skin and wing skin for transport aircraft and rotorcrafts are 2091-T8X and 8090-T81 (underaged). For control surfaces, stringers and fuselage frames that require medium-strength high stiffness alloys such as Al 2091-T851 and 8090-T6/T85, the Al-Li alloy option would be 2091-T851 and 8009-T6/T851. Likewise, for compression-loaded structures such as upper wing skins and undercarriage parts that call for high-strength materials such as Al 7075-T6/T73, alternate alloys such as 2090-T83 and 8091-T85 offer better resistance to buckling. The major aluminum producers are working to commercialize a completely new set of advanced aluminum alloys, designed to be weight competitive with current CFRP. The target properties for these alloys are published in Boeing documents (referred to as D6 alloys). The damage tolerant lower wing surface plate is targeted to be 40% stronger than 2024-T351, while the upper wing surface plate will be 25% stronger than 7178-T651. The percentage improvement is even

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Fig. 5.9

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Aluminum alloys evolution.

higher for the wing stringers. The fuselage skin and stringers will similarly be weight competitive with CFRP. The candidate materials are all “conventional” direct chill cast alloys taking advantage of the power of computing in determining the phase compositions and properties of multi-component systems. Assuming success in meeting the Boeing document target design values, aluminum will have achieved roughly the 35% weight savings predicted by the U.S. Air Force for CFRP in 1965. Figure 5.9 shows the development of strength and toughness of structural aluminum alloys. 5.2.5.4

FUTURE ALUMINUM POTENTIAL

The potential of aluminum is not expected to be fully realized by the advent of the ‘D6 alloys.’ By reducing the grain size to the nanoscale level, much higher strengths can still be achieved. In addition, if the alloy is produced from the vapor phase, rather than from a melt, the limits of solid solubility no longer apply. For instance, Cr can be introduced into aluminum alloys at much higher concentrations by vapor deposition than through conventional means. It has been known since the 1980s that aluminum can achieve an ultimate strength nearly twice as high as today’s alloys along with a 40 percent higher elastic modulus [16]. Economically scaling up the production of billets from the vapor phase is the key to wider use.

5.2.6 EVOLUTION OF STEEL AND TITANIUM Titanium alloys have seen a growing use over time because of their corrosion resistance and high temperature properties. Steel alloys offer the highest strength for metallic structures, but are used only in limited applications such as landing

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gear components. In contrast to titanium, high strength steels do not have good corrosion resistance and have been replaced with titanium for some applications. Many new alloys with improved performance continue to be developed, and new fabrication methods such as additive manufacturing are evolving. 5.2.6.1

STEEL AND TITANIUM DEVELOPMENT

There are two other alloy systems used extensively on commercial aircraft— titanium alloys and steels. Some Ni-based alloys are used, but these are used for specialized applications in the nacelle area and for high strength fasteners. The steel alloys used today are similar to those used on the first Boeing commercial transport, the 707. They include the 300 series stainless steels (SS), the precipitation hardened SS (15-5PH, 17-7PH, 17-4PH and PH13-8Mo) and the high strength low alloy (HSLA) steels (4330, 4330 M and 4340 M) that have minimum ultimate tensile strengths ranging from 517 to 1,930 MPa. The highest usage by weight would be 15-5PH and 4340 M with 15-5PH usage increasing significantly in the 1980s starting with the 757. These high strength alloys are used mostly in the landing gear and flap track areas (Fig. 5.10), but due to their high strength capability across a large temperature range (i.e. 255 to 2508C), they are found throughout the aircraft [17, 18]. 5.2.6.2

DEVELOPMENT TIMELINE

There has been much activity in the past few years in the development of higher strength stainless steels (SS). With their high corrosion resistance, SS offer some significant advantages over the high-strength low-alloy (HSLA) steels. The machining of an alloy such as 4340 M is highly controlled due to the potential for the formation of un-tempered martensite, which significantly reduces the

Fig. 5.10

737 HSLA flap tracks.

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Fig. 5.11

323

Strength toughness goals for high steel alloys.

ductility and toughness. Due to its propensity for corrosion, 4340 M is plated with a sacrificial coating of Cd, an element considered a suspected human carcinogen. In Europe the use of Cd is not permitted except for “critical aerospace structure.” Cd coatings can be compromised, leading to corrosion pitting of the substrate. The corrosion pits can act as initiation sites for stress-corrosion or fatigue cracks, and consequently will lead to higher maintenance costs. Boeing is pushing for the development of an “ultra-high” strength SS capable of a minimum tensile of 1,930 MPa that could be used as a replacement for 4340 M and eliminate the corrosion problems inherent with the HSLA steels [17]. This ultra-high strength SS could make a significant contribution in reducing life cycle costs of these HSLA steels, particularly in structures such as landing gear. It would also be better for the environment with the elimination of Cd plating for corrosion protection. The gains anticipated for these alloys in terms of strength and fracture toughness are shown in Fig. 5.11. Numerous titanium alloys also are being used on commercial aircraft. These are attractive for aerospace applications due to their high strength, low density, elevated temperature capabilities and corrosion resistance, but their use has always been restricted by their high cost relative to aluminum and steel alloys. They have traditionally been used in corrosion prone areas requiring high strength, such as landing gear support structure, wing actuation devices, and floor support structure in the galley and lavatory areas. Similar to the 2024 and 7075 alloys, one of the first alloys developed was annealed Ti-6Al-4V in the

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Fig. 5.12

Historical titanium alloy improvement.

1950s, and it has been the dominant alloy through the years. Figure 5.12 illustrates the strength-toughness combinations available with titanium alloys. In the 1970s beta-annealed heat treatment was developed. It maximized the damage tolerance properties, fracture toughness, and crack growth and stress corrosion resistance. This damage-tolerant heat treatment has been used primarily in conjunction with composite structure due to its galvanic compatibility with the carbon fiber in the composites. If the carbon fiber were to come into contact with the aluminum in an aqueous environment a galvanic corrosion would be set up which would corrode the aluminum. There are corrosion-protection schemes to isolate the aluminum from the graphite but, in critical structures that are difficult to inspect and replace (such as the empennage attachment fittings on the 777), beta-annealed titanium is used. These compatibilities of titanium with graphite have been exercised on the 787 resulting in very high titanium usage relative to other commercial aircraft. 5.2.6.3

CURRENT DEVELOPMENTS

Over the years several near-beta and metastable beta alloys have been developed because of their higher strength capabilities and, in some cases, processing advantages. Ti-10V-2Fe-3Al is a high strength-forging alloy that was used extensively on the 777 landing gear, replacing almost all of the 4340 M except for the inner and outer cylinders and axles on the main landing gear (Fig. 5.13). Boeing implemented the new alloy on the 787, Ti-5Al-5V-5Mo-3Cr, which is capable of even higher strength (1,100 MPa). Furthermore, this titanium alloy can be heat treated in thicker sections using an air cool rather than the water quench required with the solution treatment of Ti-10V-2Fe-3Al. Ti-15V-3Cr-3Al-3Sn

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has been developed as a strip alloy due to its excellent rolling characteristics and high strength capabilities. It also has high formability for simple forming operations in the solution treated condition. It has been used extensively for clips and brackets to support the composite floor beams on the 777, and is used in the fabrication of the titanium-graphite laminates. Overall, titanium alloys are responsible for approximately 9% of the total weight of the Boeing 777 aircraft [19]. Most of the development funding for titanium alloys is being directed toward reducing cost. Much of this effort is directed toward more monolithic structures, resulting in significant cost savings through reduced part count, reduced machining, reduced tooling, reduced fastener count and the elimination of shims and sealants. Approaches studied include castings, larger forgings, welding, superplastic forming (SPF), and superplastic forming in conjunction with diffusion bonding. In addition, lower temperature SPF materials, offering significant impact on the economics of the process, have been developed and are being used on Boeing aircraft. Research in both powder metallurgy and the refinement of titanium ore is yielding substantial reductions in processing and raw material costs. Furthermore, welding technology (both solid and abrasion) for titanium is receiving considerable attention as a means of reducing cost via lower buy-to-fly ratios and reduced machining.

Fig. 5.13

Candidate titanium forgings for 777 main landing gear.

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5.2.7 EVOLUTION OF COMPOSITE MATERIALS Composite materials are used extensively on new airplanes due to their lightweight, desirable mechanical and thermal properties and their reduced maintenance. They offer the designer the ability to individually tailorable properties to meet the requirements of the component and/or structure. Advances in manufacturing and a growing confidence in their use has translated to a reduced cost and cycle time for producing and employing composite components.

5.2.7.1

COMPOSITE MATERIAL DEVELOPMENT

All of the early airplanes utilized various forms of composites from spruce spars to doped fabric skins. It took a major catastrophe to eliminate the last vestige of the original structural composites in airplanes. On March 31, 1931 a TWA wood and fabric Fokker tri-motor went down in Kansas killing all aboard including the legendary Notre Dame football coach Knute Rockne. Up to that time there had been a fierce debate over whether wood or metal was the better material for building airplanes. When wood rot was found in the wing of the Rockne airplane, the debate was over and all subsequent airplanes were made of metal. [20]

Since then, as the need for lighter weight solutions to metals evolved, the design community became more comfortable with composites and began to incorporate them into more complex structures. In addition, they were able to exploit unique characteristics of composites such as aeroelastic tailoring, demonstrated on the X-29 in the 1980s. Projections for the next decade show that the commercial aviation industry will consume 75% of the composites market for aerospace applications, with rotocraft and military fixed wing systems accounting for less than 25%. The future of composites for aerospace applications is growing, finding most of its use in commercial and general aviation aircraft where the need for efficiency is driving the transition. Composite raw materials used in the fabrication of structural parts consist of reinforcing fibers held together by a polymeric resin matrix whose main function is to transfer the load-induced shear stresses between the filaments of the reinforcing fiber. Chapter 2 gives an introduction to polymers and polymer matrix composites and their processing and properties. The polymeric resins used can be classified into two general classes of polymers: thermoplastic polymers, where thermal energy is needed to plastically deform the polymer, and thermoset polymers, where the polymer undergoes a curing reaction and is permanently set by thermal processing. The most highly used thermoset resins for aerospace structural applications are the epoxy resins, because they: .

Have the highest adhesive strength than any polymer

.

Bond to almost anything with a clean surface

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.

Offer versatile processing (e.g., room temperature, 1218C, 1778C, vacuum bag pressure or high autoclave pressure processing)

.

Have good resistance to chemicals as a result of strong covalent bonding

.

Provide good properties at a reasonable cost

Glass fiber is commonly used to provide the properties needed for structural applications. The main ingredient in glass fiber is SiO2 with different amounts of other oxides added to lower the melting point and melt viscosity to reduce manufacturing cost. E-glass (54%SiO2) is the most commonly used glass fiber with a modulus of 69 GPa modulus, a strength of 3,447 MPa, and strain-to-failure of 5%. S-glass fiber (65%SiO2) is used for higher strength applications and has a modulus of 83 GPa modulus, a strength of 4,516 MPa, and a strain-to-failure of 5.3%. Quartz glass (99.99 þ % SiO2), commonly used for radomes due to its radar transparency, is a more expensive fiber but has impressive strain-to-failure and strength values of 7.7% and 5,998 MPa, respectively. Carbon fibers are another widely used reinforcement material, with generally lower density and higher stiffness than glass. Produced from a variety of precursors, fibers based on polyacrylonitrile (PAN) are the most common. This thermoplastic polymer is spun into yarns containing 1000 (1K), 3000 (3K), 6000 (6K), etc. filaments which are subsequently stretched many times in length to further orient the molecular chains in the filaments in the direction of the fiber length axis (zero direction) . The thermoplastic PAN yarns are then converted into a thermoset, in a thermo-oxidative stabilization process, where the molecules interconnect and form a ladder structure, followed by the carbonization process at 1200–15008C to eliminate oxygen, nitrogen, and other noncarbon elements from the fiber and form carbon crystals. The carbonized fiber contains more than 95% carbon in the form of very small crystals oriented within 15 degrees of the fiber axis. Depending on the precursor, carbon fibers can be further heat treated to form graphite fibers with carbon in excess of 99%. Other commonly used precursors include, pitch, mesophase pitch, isotropic pitch and rayon. Carbon fibers are grouped by modulus as ultra-high (modulus .450 GPa), high (350–450 GPa), intermediate (200–350 GPa), and low (,100 GPa). The most controlled, stable, repeatable, and user-friendly form of composite raw material for the manufacture of aerospace composite parts is the prepreg (preimpregnated). For unidirectional (UD) tape prepreg, the fiber content is established by spreading the fibers pulled from a creel containing as many fiber spools as necessary for a desired fiber areal weight and the maximum UD prepreg width desired. Subsequent to impregnation, the UD tape is trimmed on line for clean straight edges. For fabric prepreg, the prepregger uses a continuous roll of fabric, which has already been woven and certified with a tightly controlled fiber areal weight, and impregnates it with a tightly controlled resin mixture to acquire the required resin content. The width of fabric prepreg is established during weaving of the fabric. The most widely used style of fabric is the 3K

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(i.e. 3,000 filaments per tow) plain weave fabric, with half the tows running in continuous length (08, Warp) direction and half in width (908, Fill) direction.

5.2.7.2

DEVELOPMENT TIMELINE

Composite materials have been attractive to aerospace industry due to their high specific strength and the remarkable capacity of tailoring their mechanical properties [21]. In particular, Boeing has a long tradition of using composites in its aircraft as demonstrated in Fig. 5.14. The first composites used were “wet lay-up” where dry fiber bundles are impregnated with polyester resin, similar to the construction of fiberglass boats. Wet layup required considerable skill at positioning the material within a short time before the resin would begin to cure and become unusable. The Model 377 Stratocruiser (first flight in 1947) achieved a 20% weight savings over metal ducting by using fiberglass composite. Supplier pre-impregnated fabrics (prepregs), which provide consistent resin content and eliminated the messy process of wet layup, were first used in 1961. The 727 used a first-generation fiberglass-reinforced 1218C cure epoxy composite for radomes and fairing panels. The 737 used both a first generation fiberglass reinforced 1778C cure fiberglass reinforced epoxy in the hot areas and a second-generation fiberglass reinforced 1218C cure epoxy (rubber toughened/self-adhesive) on radomes, fairings and control surface cover panels. These materials were mainly used with honeycomb core. The 747 used the same materials in the same locations, but on a much larger scale. The 747 rudder cover panels made with a second-generation

Fig. 5.14

Composite applications into commercial aircraft.

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Fig. 5.15

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NASA ACEE 737 carbon fiber stabilizer.

fiberglass reinforced 1218C cure epoxy, was the biggest composite part Boeing had flown up to that point. In 1972, NASA teamed with the commercial airplane industry (Boeing, McDonnell-Douglas and Lockheed) on the Aircraft Energy Efficiency Program (ACEE) to design and build carbon fiber reinforced polymer matrix (CFRP) parts [22]. The parts that Boeing initially built were 737 spoilers made from polyacrylonitrile (PAN) based standard modulus (220 GPa) carbon fiber reinforced 1218C and 1778C cure epoxy matrices. The design was full depth aluminum honeycomb with pre-cured skins secondarily bonded. The results showed that carbon fiber reinforced composites were viable. The next part Boeing built under the ACEE program was a 727 elevator. The basic design utilized a standard modulus carbon fiber with an un-toughened 1778C cure epoxy co-cured with aramid paper honeycomb core with adhesive film to make panelized skins, spars and ribs that were bolted together. The first primary structure that Boeing made for the ACEE program was a 737 horizontal stabilizer torque box (Fig. 5.15). The cover panels were co-cured skin and stringers. The spars were “I” sections made from pre-cured details secondarily bonded together. The material used was standard modulus carbon fiber with an untoughened 1778C cure epoxy, the same as the 727 elevators. The service experience of these parts as summarized in the final report was very good with two shipsets still in commercial service [23, 24]. The NASA ACEE contract provided Boeing the confidence necessary to commit to carbon fiber parts on the 767. The state of the art at the time utilized high resin content (42% by weight or 50% by volume) prepreg, which had more resin than required for weight efficient designs. The resin flow during the cure removes trapped air and/or volatiles and the excess resin is surface bled-off. In this manner, porosity is kept to a minimum

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and the cured laminate has a resultant fiber volume of approximately 64%. This concept did not work well, particularly for thick laminates. Boeing developed a “no bleed” process that used prepreg with the resin content that gave the desired final fiber volume without bleeding [25]. In addition to providing a more consistent product, it eliminated considerable amounts of expendable processing aids, which made this a less costly, more “green” process. The next breakthrough was in 1983 with the elimination of process control panels. These were extra panels made each autoclave run to verify the processing provided acceptable parts, similar to the prolongations used on forgings and castings. These panels were not really necessary since the critical processing information was recorded and did not need to be validated with test panels. This change allowed instantaneous approval of a cure cycle, rather than a one-to two-day wait while the test coupons were being tested. The first Boeing heritage airplane to use carbon fiber was the 767. The control surfaces (inboard ailerons, elevators and rudders) used the same form of material and design as the NASA ACEE 727 elevator. The spoilers and outboard ailerons used much of the design demonstrated by the NASA ACEE 737 spoilers. The doors and fairings used carbon fiber/aramid fiber reinforcement with a secondgeneration 1218C epoxy. Engine nacelles used a similar carbon fiber/aramid hybrid, but were impregnated with a higher temperature (1778C) epoxy resin. The 1778C cure resins used in prepregs at that time were designed to make laminates, not thin skinned, co-cured honeycomb structure. The majority of the composite parts on Boeing airplanes at that time were of the latter design. The 757 program, in conjunction with Hexcel, developed a material system specifically for co-curing with honeycomb. This material has become the de facto standard for secondary structure on nearly all Boeing models. The 777 program, in conjunction with Toray, developed an intermediate modulus carbon fiber prepreg for the primary structure, with modulus of 294 GPa, tensile strength of 5,860 MPa and a strain-to-failure of 2%. This material was used for the horizontal and vertical stabilizer torque box and the passenger deck floor beams. In addition to the higher modulus, this prepreg resulted in parts with significantly better impact resistance. Carbon fiber (T300) was also used on the 777 control surfaces and empennage ribs. Figure 5.16 captures the evolution of composite materials used in Boeing commercial aircraft. 5.2.7.3

CURRENT DEVELOPMENTS—COMMERCIAL APPLICATION

Trade studies conducted by Boeing and 787 partners showed composites superior to advanced metals for empennage, wing, and fuselage primary structures (Fig. 5.17). The choice to build the next all new Boeing jetliner from composite materials was driven largely by the great potential for reducing weight and life cycle costs. The 787 has set aggressive cost and weight targets for the airframe. The targets require substantial improvements in the processing as well as significant cost reductions for both composite and metallic structures.

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• Nanotechnology • Nonepoxy matrices • M5 fiber

Weight and cost Improvements

• Higher modulus (60 ms) carbon fiber • Higher fiber volume SM carbon fiber • Thermoplastic welding

• 15X lay-down rates • Non-woven stringer pre-forms • Systems integrated into structure

• Bonded assembly • Secondary bonded stingers • Eliminate safety fasteners

Mostly uncharted Known territory

• Mixture of SM and IM fibers • Industrial priced prepregs

• Ti-Gr (Ti foil producibility problems) • Vacuum bag cure 350F CGRP prepreg • Carbon/glass fabric reinforced thermoplastic laminate • PEKK CFRP UD tape for compression molding of structural thermoplastic parts

•Noncrimp fabric • BMS8-276 form 3 implemented (lower cost) • Automated fiber placement

• Hot drape forming/double diaphram forming/lash forming • Vacuum bag resin infusion/small batch processing technology • Self-adhesive co-curable standard modulus carbon fiber

• Pultrusion • 5X laydown rates • Toughened intermediate modulus fiber implemented

• Higher performance fiberglass allowables (class 4) • High speed stringer mills implemented • HAL cells (automated compaction cells)

• Optical ply layup templates • Autotape layup • Honeycomb co-curable standard modulus/carbon fiber epoxy • Untoughened standard modulus carbon fiber/epoxy

Time

Fig. 5.16

Advancing composite technology.

With the introduction of the Boeing 787 and the program decision to utilize high performance polymeric composites for the majority of the aircraft primary structure, a paradigm was created that was a first for the next generation of airliners. However, it was not an abrupt change. Figure 5.18 shows the increasing usage of composites in all aircraft over the decades as airplane performance

Fig. 5.17

Structural materials on Boeing 787 commercial aircraft.

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Fig. 5.18

Aerospace composite usage growth.

increased and prices reduced with increased markets and the advent of more sophisticated manufacturing facilities. The reason for this steady growth in composites application is clear when the advantages of composites are considered. Composites resistance to fatigue and corrosion can lead to reduction in maintenance down time. The higher strength-to-weight ratio translates to reductions in weight, increases in range and payload, and improvements to environmental performance. Use of composites also enables larger, more integrated structure with a corresponding increase in design options (e.g. one-piece barrel sections and larger windows). From a fabrication perspective, composites allow reduced manufacturing flows, reduced tooling and easier assembly. For the 787, the applications now include the most significant primary structure, the empennage, wings and the efficient one-piece fuselage sections (Fig. 5.19). In addition to the new technology on the 787 platform there is also

Fig. 5.19

One-piece fuselage barrel section (left), one-piece fuselage CAB section (right).

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a new philosophy of Open Architecture (i.e., new technology can be inserted after airplane #1 if the technology meets sufficient readiness and business case criteria). Together, these new paradigms come together to enable continuous consideration of new technology across a broad spectrum of opportunities. The development of the new aircraft brings opportunities for novel materials and structures. In particular the use of composite materials opens the door to the opportunity for multifunctional materials and structures, combining the capabilities for acoustic attenuation, fire retardation, lightning strike protection, power generation, and others, into the high performance light-weight anisotropic capabilities of the latest state-of-the-art composite. 5.2.7.4

CURRENT DEVELOPMENTS—MILITARY APPLICATION

Composites usage for military application has grown significantly from the early demonstrations on the F-14, F-15 and F-16. The F-14 and F-15 demonstrated the use boron-epoxy predominantly in the skins of the empennage. The F-16 employed graphite-epoxy for the same applications. In practice, the use of composites in military applications was limited and not used in highly-loaded primary structure simply due to a lack of understanding of the material’s failure mechanism. Since the early years of composites, the confidence in the material has grown, resulting in significant usage in primary and secondary structure. The vertical take-off V-22 Osprey boasts an airframe structure that is over 40% composite. Fighter aircraft are also seeing increased use of composites with the F/A-18E/F at nearly 20%, the F-22 at approximately 24%, and the F-35 projected at between 20 and 30%. The Air Force Research Laboratory’s (AFRL) Advanced Composite Cargo Aircraft (ACCA) [26] is the first X-plane focused on demonstrating advanced composite structures concepts, design methods and tools, and manufacturing processes. ACCA’s mission was to design, build and fly a militarily relevant aircraft constructed primarily from composite materials and assembled using large integrated and bonded structures. ACCA challenged current paradigms for flight demonstration of advanced concepts in order to reduce the cost and shorten the time to field advanced warfighting solutions. AFRL teamed with Lockheed Martin Aeronautics Company to modify a Dornier 328 regional jet (Fig. 5.20) and replace the fuselage and vertical tail with composite structures. The selection of a “donor” aircraft enabled the focus of the program to be on the structure, resulting in a low-cost quick demonstration of key composite technologies. No development of flight controls or aerodynamics was necessary. Lockheed’s design (Fig. 5.21) featured an MTM-45 out-of-autoclave allcomposite, bonded fuselage with approximately one tenth the parts of the donor aircraft. Additionally, it included a wider fuselage, a new composite cargo door, a vertical stabilizer with advanced fiber placement concepts for integral stiffening, and bonded assembly via pi joints developed under the Composite

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Fig. 5.20

ACCA—Modified Dornier 328.

Affordability Initiative (see Section 5.2.10.5). The new fuselage and vertical tail consist of 306 structural parts, an approximate 90% reduction from the original Dornier 328 fuselage and vertical tail. In addition, the new structure reduced the number of fasteners by 98%.

Fig. 5.21

ACCA structural layout—Composite fuselage, vertical tail, and fairings.

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Fig. 5.22

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ACCA fuselage skins-mated together (left), assembled with joint (right).

ACCA took advantage of new developments in composite processing to realize the part count reduction. Standard aerospace epoxies are cured in autoclaves at 1778C and at elevated pressures. This requires the use of hardened (metallic) tooling and it limits the size of part to what can fit into an autoclave. Since MTM-45 cures at 718C to 938C (with a 1778C post cure off the tool) and under vacuum pressure only, the use of inexpensive, short lead time tooling is enabled. In addition, part sizes are not limited to what can fit into an autoclave. Ovens can be built around the part. With this philosophy, the upper and lower fuselage skins were fabricated in two assemblies and up to that time were the largest MTM-45 parts fabricated. The main frames and floor supports were sandwich stiffened with HRH-10 Nomex honeycomb from Hexcel. The frames, floor supports, pressure bulkhead and cargo door substructure were bonded to the lower and upper skin with 3D woven “Pi” performs. The upper and lower fuselage assemblies were bonded together in a double lap shear configuration where overwrap plies were placed over the seam on the outer mold line and the interior of the aircraft, and the assembly then rolled into the oven to cure the joint (Fig. 5.22). Figure 5.23 shows the completed vertical tail assembly. The skins were precured along with the spars, ribs, and stringer webs, all of which were fabricated from aerospace standard IM7/977-3. The webs were cobonded into the assembly

Fig. 5.23

ACCA completed vertical tail.

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via “Pi” joints with IM7/977-3. The forward and rear spars were separate cobonded assemblies that were mechanically attached once the control surface mechanisms were installed inside the tail. The use of large integrated and bonded structures reduced labor hours needed to design, plan, source, schedule, fabricate, calibrate, purchase, store, and inventory tooling. The ACCA fuselage and vertical tail achieved a 90% and 80% reduction in part count respectively as compared with the original Dornier 328 aircraft. In addition, the 90% reduction in parts reduced the labor hours needed for designing, configuration controlled, planned, sourced, scheduled, tooled, kitted, fabricated, inspected, purchasing, packaging, shipping, handling, marking, and inventorying. It also reduces the need for tooling and fasteners, work instructions, and hole drilling and fastening operations as the parts are assembled. Composite bonded assemblies are forecast to cost 20–50% less than conventional fastened structures depending on the assembly. 5.2.7.5

FUTURE COMPOSITE POTENTIAL

There are material and processes improvements that could provide breakthroughs in composites. Several examples are high modulus organic fibers such as M-5TM that could provide significant structural performance improvements. Thick wide multi-axial reinforcements would allow laying the entire skin down in several passes instead of hundreds of passes. Another would be innovative processing such as E-Beam curing which would allow in-situ curing of large structure, eliminating the need for an autoclave and greatly reduce the residual stresses in a part after cure. The batch processing so common in aerospace will have to give way to a moving line approach to fabrication and assembly to reduce cycle time. Reduced cycle time allows for fewer sets of layup mandrels. Oven cure processes such as Vacuum Assisted Resin Transfer Molding (VaRTM), Bulk Resin Infusion (BRI), Resin Film Infusion, Controlled Atmospheric Pressure Resin Infusion (CAPRI) or oven cure prepregs will allow the use of tunnel ovens or part specific heating. The cure of the part can start immediately upon completion of layup, not when the autoclave or oven is filled with the full load. Development of new materials and large part fabrication techniques, along with improvements to design and analysis tools are needed to ensure robust transition of composites to aerospace structures. It has been demonstrated repeatedly that composites offer significant improvements to problems such as fatigue and corrosion, but further improvements are needed in understanding and designing for hot/wet compression and compression after impact. Figure 5.24 shows the developments of open-hole residual compression strength versus compression after impact strength for various generations of composite materials. The trend is moving in a positive direction for the expanded use of composites in stressing environments. The material utilization (i.e., buy to fly) for composite parts is significantly better than similar metal parts. Metal parts generally start off with oversize

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Fig. 5.24

337

Composite matrix improvements.

stock and trim down to the final part dimension. Composite parts are laid up one ply at a time and only where needed. Minimum weight designs require more intricate tailoring, which takes more time during design, programming and particularly fabrication. To provide the best balance between cost and weight requires consideration of the material cost, material utilization, as well as the fabrication time. As a general rule, a composite part is lighter than its aluminum counterpart, however the cost has been traditionally higher. One of the techniques that allow composites to be cost competitive with aluminum parts is through reduction of part count. This requires more monolithic structure, which combines many smaller parts and eliminates the fasteners. As a result, the 787 program developed an advanced fiber placement method to fabricate one piece fuselage barrel sections with co-cured hat-stringers in a single cure cycle. The drawback is that the complexity and cost of the part increases. However, if properly done, the savings in assembly more than offsets the cost increase in the detail part. As parts get more complex and costly, there is more incentive to validate the design and manufacturing process before beginning to design the part and tooling. Computer programs to predict part warpage and final dimensions will provide the confidence to design for no shims. These same programs will predict thermal uniformity and allow modification of the tool design, prior to tool fabrication, to provide a more uniform heat-up that will shorten the cure cycle.

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5.2.8 EVOLUTION OF FIBER METAL LAMINATES Fiber metal laminates are a relatively recent material concept that offer exceptional damage tolerant properties. Interest in these materials is growing as their unique properties are finding their way into solutions addressing design and repair of fuselage and wing skins. 5.2.8.1

ALUMINUM-BASED FIBER METAL LAMINATES: GLARE AND ARALL

Fiber metal laminates (FML) were developed at Delft University of Technology (TU Delft) as a family of hybrid materials for use in fuselage and wings of commercial airplanes [27]. These materials consist of alternating thin aluminum alloy layers (0.2–0.5 mm) and uniaxial or biaxial aramid or glass fibers prepreg (Fig. 5.25). Two grades of FML are commercially available: ARALL based on aramid fibers, and GLARE based on high strength S2 glass fibers. The laminate design creates a material with excellent properties. In addition to low density, the FML present high fatigue and corrosion resistance, excellent impact and damage tolerance. The bond lines with fibers (prepreg) act as barriers against corrosion of the inner metallic sheets, whereas the metal layers protect the fiber/epoxy layers from environmental exposure. The same configuration also reduces crack propagation, thereby leading to longer fatigue life. There are five different grades of GLARE [28] produced from two types of aluminum (7475-T761 and 2024-T3). Grades one and two are fabricated from unidirectional fiber glass laminates and 7475-T761 and 2024-T3, respectively. The remaining three grades are made from 2024-T4 aluminum sheets and bi-directional fiber glass laminates. ARALL laminates were developed primarily for wing applications, while biaxial GLARE-3 was created for fuselage applications. Under realistic loading conditions, GLARE laminates have exhibited crack growth rates 10–100 times slower than their monolithic 2024-T3 aluminum counterparts. Demonstrating the potential of using FML for aircraft structures, researchers at TU Delft [27] evaluated two GLARE configurations and compared those against Al 2024-T3. As GLARE was primarily developed as a fuselage material, longitudinal and circumferential joints of stiffened GLARE panels of the fuselage are usually made using rivets or hi-locks. In a fuselage structure these joints are the most likely locations for fatigue damage. The residual strength of 2024-T3 and GLARE

Fig. 5.25 Alternating layup of aluminum and glass-epoxy layers.

SUBSONIC AIRCRAFT MATERIALS DEVELOPMENT

Fig. 5.26

339

Residual strength of GLARE and Al 2024-T6 rivet joints.

riveted lap-joints after fatigue is shown in Fig. 5.26. The GLARE joint displays a better performance over the aluminum joint, merging high initial strength with slow strength reduction. For the FML, cracks initiate in one aluminum layer at the mating surface and remain limited to this layer for a long portion of the life. Conversely, the data for the 2024-T3 riveted joint demonstrates a rapidly decreasing strength once fatigue cracks initiate. For a fielded aluminum joint, a maintenance program would require inspections at relatively short intervals to prevent a situation in which an unstable crack growth leads to a catastrophic reduction of residual strength. The data presented for lap joints demonstrates that under realistic conditions, GLARE designs are unlikely to suffer from fatigue damage. The failure process of GLARE laminates is quite complex and can involve multi-fracture modes such as matrix cracks, fiber-matrix debonding, fiber fracture, fiber/matrix interfacial shear failure, and interdelamination of laminates [29]. For unidirectional GLARE specimens under longitudinal tensile load, fiber pull-out and interface-matrix shear mode are normally noticed in the fiber-epoxy layers, but global longitudinal splits are prevented by the aluminum layers. Under transverse tensile loading, matrix failure and fiber-matrix interface debonding/ fiber splitting are the main fracture modes in the fiber-epoxy layer of GLARE. Controlled delamination in GLARE and ARALL materials allows for the dissipation of energy through non-catastrophic mechanisms (such as opening of cracks) without failure of the fibers [30]. The laminates can be prepared and machined using the same technologies developed for aluminum alloys while still offering

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the high specific strength benefit of composite materials. FML can be stretched after curing in order to reverse the internal stress system in the material. Due to the difference in thermal expansion coefficients, the as-cured FML composite is under tensile stresses in the aluminum layers and compressive stresses in the fiber layers. During the stretching operation the aluminum layers deform plastically, while the fiber/epoxy layers remain elastic. Therefore, residual stresses are present in the laminate after unloading. After the post-stretching operation, the laminates have a favorable compressive (ideal for crack closing) stress in the aluminum sheets and a tensile stress in the fibers. Aircraft fuselages require special attention in transferring high shear loads around doors and window cutouts. To avoid stress concentrations at the corners, a complex set of doublers are generally employed around the cutout. In addition, care must be taken to limit deformations around the cutouts to eliminate interference with the doors. The unique anisotropic nature of GLARE and ARALL coupled with their lower stiffness make them ideal for application in these critical regions. Fabricated from ARALL, the C-17 cargo door demonstrates that these FMLs can overcome stress concentration conditions and be useful for aircraft designs [30]. Furthermore, hybrid composites such as ARALL and GLARE are suitable for high-impact regions (e.g. leading edges, inboard flaps) [31], and can greatly improve damage tolerant designs [32]. Significant weight savings of approximately 20% can be achieved using these hybrid composites in fatigue-prone regions such as pressured fuselage skins, stiffeners and low wing skins [31]. The use of GLARE in large portions of the A380 fuselage enabled Airbus to save approximately 800 kg [29]. Usage of ARALL in the Airbus A320 led to only 8% weight reduction on fuselage, but led to a much larger 30% weight saving for the wings [33]. GLARE provides an alternate material solution for patch repairs of aircraft structures due to its capacity for slowing crack propagation [34]. Although GLARE can be used for conventional repair methods like riveting and bolting, bonded patches can provide a uniform and efficient load transfer into the patch and reduce the risk of high stress concentrations caused by mechanically fastened repairs. The advantages of GLARE patches compared to other materials can be explained by a series of factors: 1) the small mismatch in the coefficient of thermal expansion with the aluminum skin, 2) GLARE’s excellent fatigue properties, and 3) GLARE’s mechanical properties, i.e. high strength, moderate extensional stiffness and high bending stiffness. Despite of these advantages, the complexity of damage propagation due to the large mismatch in mechanical and thermal properties between the different layers can lead to failure [35]. The difference in coefficient of thermal expansion (CTE) between patch material and adherent material plays a crucial role in patching effectiveness. When materials with different CTEs are bonded, both the cure cycle and operating temperature affect the thermal residual stresses in the patched area. The cure temperature will introduce permanent static stresses. In addition, there is also a thermal cycling effect caused by the operating temperatures of the aircraft. At

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cruising altitudes, the complete fuselage will cool down to temperatures as low as 2558C. When the patch CTE is lower than the parent material’s CTE, high tensile thermal stresses will develop in the crack, reducing the patch effectiveness. For instance, the CTE for aluminum alloys is typically around 23.4 ppm/8C, while for GLARE-2 it is 16.3 ppm/8C, and for boron/epoxy approximately 4.5 ppm/8C. Figure 5.27 [34] illustrates the influence of CTE between parent substrate and patch material for a Boeing 737–200 fuselage, cruising at an altitude of 10 km. The parent material is an Al 2024-T3 alloy under a hoop stress of 100 MPa and a longitudinal stress of 50 MPa with a crack diameter of 51 mm repaired by two different patch designs—one using boron/epoxy and the other using GLARE-2. For this case, the GLARE-2 repair with and without thermal effects presents a much higher stress intensity factor reduction. GLARE patch repairs have been successfully applied on USAF C-5A transport aircraft [36]. A problem in this class of aircraft is the formation of multiple cracks in the upper aft-crown section of the fuselage skin. These cracks are most likely caused by stress-corrosion of the 7079-T6 aluminum alloy. The crown section experiences significant longitudinal tensile bending in addition to biaxial tension due to internal pressurization. The cracks grow in the second rivet of a 5-rivet row joint. This is a clear indication that the crack is not an ordinary fatigue crack. The direction in which the crack grows indicates that the longitudinal stress is the dominant stress for crack growth. The complicated sub-structure (frame and stiffeners) affects the stress field around the crack. After a series of tests and numerical simulations using the finite element method, a patch repair was

Fig. 5.27

Stress intensity factor reductions for patch repairs.

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Fig. 5.28

Crack growth curves—unpatch and patched 7075 specimens.

designed and tested. The GLARE patch was made of GLARE-2-3/2-0.2 which is 3 layers of aluminum and 2 layers of glass-epoxy prepregs, with a total thickness of 1.1 mm. For bonding, the 3 M adhesive AF-163-2M was used. The unpatched panel demonstrated significantly higher crack-growth rate than the patched specimen (Fig. 5.28). In this case, GLARE repairs were very effective in slowing the crack growth under realistic crown fuselage spectrum and were able to withstand one full life cycle without a problem. Designers should be cognizant that in actual service, composite structures will be exposed to a variety of environmental conditions, including hot humid environments. This “hot-wet” combination can lead to hygrothermal forces and formation of residual stresses which can influence the behavior of the composite. Although the GLARE aluminum outer surface can act as a moisture barrier, water absorption can occurs on the free edges and rivet hole boundaries. Water absorption can lead to a reduction (between 17% and 32%) of the interlaminar shear strength (ILSS) due to micro crack formation. A typical reduction of ILSS (around 18%) has been reported when GLARE materials were submitted to thermal cycle fatigue [37]. 5.2.8.2

TITANIUM-BASED FIBER METAL LAMINATES: TiGr

TiGr (Titanium Graphite) belongs to the class of engineered materials known as Fiber-Metal Laminates (FML) and consists of layers of resin-coated titanium foil interleafed through the thickness of a carbon fiber reinforced plastic (CFRP)

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composite laminate. In the 1990’s, TiGr was developed during the NASA/Industry High Speed Civil Transport research program. It was considered for its resistance to environmental degradation and attractive high-temperature, high performance mechanical properties when used with carbon fiber/polyimide composites. In the 2000’s, TiGr was evaluated for use in conventional subsonic aircraft using a toughened epoxy carbon fiber composite material system. The use of TiGr fiber metal laminates as a design material for aircraft structure presents advantages over both traditional metallic and CFRP structures. Advantages over monolithic titanium include higher specific strength and stiffness, improved fatigue life, greater damage tolerance, and the ability to tailor the stiffness/strength properties to suit unique design constraints. Advantages over CFRP include improved compressive and notched-strength properties, better resistance to impact damage (improved visibility and residual strength), and additional tailorability due to the nearly isotropic nature of the included foil plys. Moreover, with higher percentages of titanium foil in the layup (still well below 50%), TiGr laminates can provide large improvements in bearing strength when compared to pure CFRP, approaching the bearing strength of pure titanium. Benefits have been demonstrated in other performance areas including solvent sensitivity and fluid resistance, lightning strike, and bonding and grounding among others that will enable a net reduction in structural gauge and/or systems weight for many components [38, 39]. Figure 5.29 shows a cross-section of a TiGr laminate with titanium plies selectively distributed to provide localized reinforcement. There has been a significant body of work conducted by The Boeing Company on the development and optimization of the constituent materials and processes used to fabricate TiGr laminates [39, 40]. One of the main building blocks of the TiGr material system is a prepared titanium foil, typically 125–300 mm thick. The thickness of the Ti foil plies has generally been chosen such that after processing, it will match the thickness of the surrounding composite plies, simplifying TiGr

Fig. 5.29

Titanium graphite (TiGr) material.

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laminate designs and layouts [40]. After extensive testing of many titanium alloys and heat treatments, a Ti 15-3-3-3 alloy in a solution treated and aged condition was selected based on its mechanical properties, cost, and suitability for rolling into a thin foil. Optimization of the Ti foil surface preparation process and the development of unique (and very thin) resin coatings for the foil resulted in an extremely robust and environmentally durable titanium-to-CFRP interface architecture. This surface treatment was tailored to the unique composite resin chemistry used in each of the aforementioned iterations of the TiGr material system. Significant research into the feasibility of production-related issues such as scale up of the Ti foil surface coating process, Ti foil auto-lamination and laydown over compound contours, Ti foil ply adhesion and trimming, and trimming and drilling of the cured TiGr laminates have all been addressed and optimized. Additionally, Non Destructive Inspection (NDI) and repair activities indicate that current CFRP methods can be applied to TiGr laminates with future optimization of parameters. The aforementioned developments in the area of TiGr laminates have made it a feasible lighter weight alternative to conventional aerospace materials for selected applications.

5.2.9 EVOLUTION OF STRUCTURAL REQUIREMENTS Discussion to this point has focused primarily on the evolution of material development and continuing attempts to improve properties. This section will attempt to relate evolution of basic material property data such as Fcy/Fty/Fsu to design values, which are used to design aircraft structures. The relationship between basic material and design properties is illustrated in Table 5.3, and applies to traditional designs utilizing current material solutions. 5.2.9.1

WING STRUCTURAL REQUIREMENTS

The commercial aerospace industry is at the beginning of the introduction of the next generation of highly efficient and quieter airplanes and is evaluating candidate lightweight structurally efficient materials and designs for use on wing structure. The Boeing 777 was a culmination of design solutions that over the years had pushed the static, fatigue, and damage tolerance performance of the wing structure to a point that maximized the capability of the 2324 skins, 2224 extruded stringers on the lower surface, and the 7055 skins and stringers on the wing upper surface. Table 5.4 identifies the material usage and rationale for its selection for the upper and lower surface of the wing. Currently, there is a great deal of experience with designing aluminum alloy wings. Based on known design drivers and the need for a useful balance of properties, new alloys are being proposed that have significant increases in static strength relative to the 7055, 2324 and 2224 alloys used on the 777. Specifically, the upper and lower surface materials have shown improvements in both compression yield strength (Fcy) and tension yield strength (Fty), respectively.

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TABLE 5.3 Design Property

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CRITICAL MATERIAL DESIGN PROPERTIES

Criteria Property

Critical Material Property

Property Evaluation

Fty, Ftu, Fbru OHT, FHT, NT

Fty – small hole out OHT – open hole tension Ftu – large hole out FHT – filled hole tension Fbru – Joint strength NT – notched tension

Fcy, Ec OHC, FHC, NC

Fcy – short columns Ec – long columns OHC – open hole compression FHC – filled hole compression NC – notched compression

Ftu, Fty, Fsu NC, NT

Ftu45, Fty45 – thin web Fsu – thick web NT – notched tension NC – notched compression

Fatigue strength, open hole, notched specimen, low load & high load transfer joint coupons

Low load and high load transfer joint coupons data most reliable for material evaluation For composite, cycling to validate no growth

K1scc, SCC threshold and exfoliation rating

Heavy reliance on service experience

Static strength Tension

Compression

Shear

Structure must remain elastic to limit load and carry Ultimate Load. For composite materials, manufacturing flaws and Barely Visible Impact Damage (BVID) must be included

Durability Fatigue

Design service objective with high level of reliability Corrosion

(Continued )

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TABLE 5.3 Design Property

CRITICAL MATERIAL DESIGN PROPERTIES (Continued )

Criteria Property

Critical Material Property

Property Evaluation

Crack Growth

Damage must be found before becoming critical. For composite material, structure must demonstrate no detrimental growth with visible flaw.

Fatigue crack growth characteristics CAI – compression after impact

Inspection intervals & methods

Residual Strength

Must carry limit load with large damage

Kapp, Fty, elongation H – Composite fracture toughness CAI

Kapp for low Toughness or wide panels Fty for high toughness narrow parts Hc for wide panels, CAI for local areas

Minimize within constraints

Density, material costs

Fabrication and maintenance costs must be accounted for

Damage Tolerance

Weight/ Cost

For the upper surface, the desire is to improve the static compression strength while not sacrificing fatigue performance, stress corrosion, or exfoliation corrosion susceptibility relative to today’s 7055-T7751. For the lower surface, the desire is for substantially improved tension strength, fracture toughness, fatigue performance and crack propagation rates. Aluminum suppliers understand how to utilize their aluminum alloy knowledge and material production methods to improve these properties, but guidance is required to achieve the correct balance. For example, if a high static strength is achieved, the question to answer is how high the fracture toughness and fatigue capability must be to take advantage of the improvement. This same philosophy is being used

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TABLE 5.4

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CRITICAL CONSIDERATIONS FOR WING DESIGN

a. Aluminum Wing

b. Carbon Fiber Wing Panels and Spars

Lower surface Skin (2324 Plate)

Stringer (2224 extrusion)

Residual strength, fatigue, damage tolerance, static strength (tension and shear)

Upper/lower surface Skin (CRFP tape)

Stringer (CFRP tape)

Upper surface Skin (7055 Plate)

Stringer (705 extrusion)

Static strength (compression and shear), residual strength, damage tolerance

Static strength, residual strength, damage tolerance (tension/ compression)

Spars Spar (CFRP tape)

Static strength, residual strength, damage tolerance (tension/ compression/ shear)

Spar Spar

Static strength, damage tolerance (tension/ compression/shear)

to balance properties of composite laminates between large notch compression, filled-hole compression, filled-hole tension, etc. Figure 5.30 shows the typical margins of safety versus wing station for various wing load conditions. Each of these margins is established using allowables from coupon, element, and subcomponent tests. It is essential to improve material properties of new candidate materials in a balanced fashion. Understanding the material property interrelationships is very important in achieving the requisite balance. A useful means of providing material suppliers with the correct balance is to plot these relationships against each other. In this manner, materials can be selected which will allow for structural weight reduction. The lower panel figure is considered differently from the upper wing panel where compression strength is the dominant property in terms of determining weight. For the lower panel, tension strength, residual strength, and fatigue all combined as key design drivers. Figure 5.31 gives the typical relationship between fracture toughness and tensile strength, fatigue stress to tensile strength,

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Fig. 5.30

Criticality of wing failure modes.

and crack growth rates to tensile strength. These relationships are absolutely necessary for deciding if a material is an acceptable candidate for the lower wing surface. The comparison of different material systems applicable for the lower wing panel is shown in Fig. 5.32. The graph includes specific properties for 2024, 2324, the best-case projection for advanced aluminum alloys, and carbon fiber materials. Figure 5.33 illustrates the evolution of upper panel strength, and includes the actual compression strength for 2024, 7075, 7055-T7751, the proposed strengths of next generation aluminum alloys, and the specific-strength of a representative aerospace quality graphite material.

Fig. 5.31

Lower wing panel material relationships.

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Fig. 5.32

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Lower wing panel material comparisons.

Approximately 10% lighter wing structure, compared to existing wing designs, can be achieved through the use of the new alloys and composites that are being developed. In addition to metallic wing structure, much emphasis is being placed on understanding the design, fabrication methods, systems integration and cost of carbon fiber structure. State-of-the-art carbon fiber designs are thought to save similar weight relative to the advanced aluminum alloys (on the order of 10% or so) when all systems integration features are considered in the complete design solution. One of the major challenges facing carbon fiber reinforced plastic materials is the cost of fabrication methods and materials. Strides are

Fig. 5.33

Upper wing panel strength evolution.

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being made in reducing raw material cost, improving material lay-down rates, minimization of material waste, and simplifying systems integration. These are coupled with techniques that eliminate the need for autoclaves. As these improvements are realized, carbon fiber composites will become more viable for wing primary structure. 5.2.9.2

FUSELAGE STRUCTURAL REQUIREMENTS

The 777 was a culmination of design evolution that over the years had pushed the static, fatigue and damage tolerance performance of the fuselage structure to a point that maximized the capability of the 2524 skins and 7150 extruded stringers. Table 5.5 identifies the traditional material usage and rational for their selection in the general monocoque structure. Figure 5.34 identifies the failure modes that are critical for typical fuselage structure. This chart typifies the balance in design between static strength, fatigue strength and damage tolerance. The new alloys and composite materials

TABLE 5.5

CRITICAL CONSIDERATIONS FOR FUSELAGE DESIGN

Aluminum fuselage Monocoque Skin (2524 Sheet) Stringer (7075 Sheet, 7150 Extr) Frames (7075 Sheet, Extr) Floors Beams (7075 Sheet, Extr) Seat Tracks (7150 Extr. Ti-6AI-4V Extr)

Fatigue, damage tolerance, corrosion Fatigue, static tension, compression strength Stiffness, fatigue, compression strength Corrosion resistance, static strength

Carbon fiber fuselage Monocoque Skin (CFRP tape and/or fabric) Stringer (CFRP tape and/or fabric) Frames (slit tape or braided textile) Floors Beams (CFRP tape, Reinforced Thermoplastic) Laminate (RTL), titanium Seat tracks (7150 Extr. Ti-6AI-4V Extr)

Static strength (compression), damage tolerance (tension/compression) Static compression strength (after impact), stiffness Stiffness, stability, static compression strength Corrosion resistance, static strength

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Fig. 5.34

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Aluminum monocoque critical failure modes.

being developed today have significantly improved static strength, while maintaining the toughness required for fatigue and damage tolerance. In particular, the new skin alloys have shown a potential to improve ultimate tensile strength (Ftu), tensile yield strength (Fty), and compression yield strength (Fcy) by as much as 20%. New stringer extrusions show static tensile strengths in excess of 690 MPa and compression yield strengths near 690 MPa. The challenge is to develop designs and assembly techniques that will fully utilize the new material capability while minimizing build costs. Reducing the weight of the fuselage implies operating at higher stresses given that the density of the aluminum is relatively unchanged. The 777 design operates at approximately 103 MPa hoop stress and an axial tension stress of 110 MPa under normal operating pressure. Operating at higher stresses allows for a reduction in basic fuselage gages. The effort to raise the general operating stresses has forced a focus on the tension yield strength, Fty, and fracture toughness, as defined by the material property Kapp, of the skin material. These material parameters determine the general residual strength of the skins with a circumferential crack. A well-balanced design requires that improvements in durability and crack growth are also achieved. The general operating tension stresses cannot be increased if there are no improvements in the durability, crack growth or residual strength capabilities of the structure. Past experience has shown that static sizing can be performed using axial stress limits to maintain a balanced design that satisfies durability and damage tolerance requirements. Monocoque designs have been limited to a static tension axial ultimate stress range of 345 to 379 MPa. This static tension limit is related to a fatigue axial operating stress of approximately 110 MPa. This approach led to well-balanced designs that are equally critical for static, fatigue and damage tolerance using existing aluminum alloy technologies. As future material yield strength and toughness increases, the general operating stress and likewise

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static stress limitations can be increased. This leads to general gage reductions and a corresponding weight savings. The structure must then be evaluated for durability and crack growth rates to ensure that the general design and material behavior can be supported by the maintenance and inspection procedures. The potential increased use of composites creates a competitive challenge for aluminum alloys. Figure 5.35 shows the relative axial stress capability for past composites and current toughened composites relative to conventional and future alloys. Weight savings on the order of 20% to 30% can theoretically be achieved with the use of composites when compared to current aluminum alloys. Alloy development for future aircraft is being challenged more than ever to improve performance as well as reduce overall life cycle costs of aircraft ownership for the airlines. To this point, the focus of discussion has been on evolution of aircraft structure with respect to design drivers and weight efficiency. Today, there is an integrated approach to material development, weight and life cycle cost. When examining the construction techniques of most commercial jet transports, it is apparent that these techniques are largely unchanged from those used in the 1950s. Wing panels consisting of skins and stringers, usually two spars and multiple ribs to support the panels and fuselages today are still built with similar skin, stringer and frame components as those used on the 707 and DC-8. Cost improvements have come in the form of improving the fabrication and assembly of the individual elements through automation of the machining and joining processes. There have been significant attempts to reduce the cost of major assemblies by machining them from a single piece of metal with the benefit being the

Fig. 5.35

Allowable operating stress based on axial residual strength.

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Automatic fiber placement

Fig. 5.36

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Completed fuselage sections

Composite fuselage barrel.

elimination of assembly time and labor. Specific examples include spars on some Airbus products, wing center section cover panels on the Lockheed C-130, and wing ribs on the 737NG airplanes. This is not always a clear cost savings as sometimes this practice causes a high degree of waste, and the increased raw material cost more than offsets the savings. In addition, limitations on ingot size and forming have generally prevented the large cover panels on the outboard wing from being integrally machined. Primary structure, fabricated from composite materials, tend to be more monolithic than traditional aluminum designs due to the raw material form and the method in which the parts are tailored, as well as the ability to co-cure or co-bond components together in a common cure cycle. For example, the 787 fuselage is fabricated using advanced filament placement processes as shown in Fig. 5.36. The fabrication is accomplished on barrel tooling allowing for loading of stringer hat sections before placing the skin. The entire assembly is cured in a single process. Looking forward however, increased ingot volume, combined with improved forming techniques may allow for the fuselage and wing skin cover panels being integrally machined from one piece of metal. If these hurdles can be overcome, the path is open to vastly reduced part count, assembly and non-recurring tooling cost. Wings then can be constructed from three unitized structure part families—spars, ribs, and panels. The cost savings from reduced part count and assembly are obvious. Less obvious is the reduction in non-recurring tooling cost—this stems largely from the ability to use determinate assembly methods to eliminate the need for large and expensive monument type tools. As a complement to the integrally machined part, there is a new interest and emphasis on welding techniques to essentially achieve the same benefits (i.e. reduced part count, minimize non-recurring and assembly cost). The latest techniques include friction stir welding (FSW) and laser beam welding. Time will tell if these turn out to be viable, cost-effective methods of construction for wing

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structure. All of these candidate cost reduction processes rely on development of materials to accommodate the desired manufacturing methods for both composite and metallic material choices.

5.2.10

MANUFACTURING REQUIREMENTS

Producibility of materials continues to be a significant design consideration. In order to meet design cost targets, it has become necessary to make sure new materials can be used cost effectively. The production processes used in traditional fuselage and wing construction today include stretch forming, shot peen forming, roll or brake forming, age creep forming of skin panels, chemical milling or machining of fuselage skins/stringers/frames, drilling and fastening of stringers and frames to skins, and heat treating or aging of selected skins/frames/stringers. 5.2.10.1

ROLL FORMING

Roll forming (Fig. 5.37) of stringers includes taper rolling to tailor stringer thickness during the process of forming the hat or zee section from the raw coil and strip stock. Each of these production processes has associated costs that vary with material types and forms. It is essential that the design engineers understand the effect of each process on material part production costs. Metallic material microstructure affects the ability to form radii on roll formed stringers and plays a significant role in the resulting stress corrosion resistance of the final stringer. Alloy producers have expended significant effort in developing material characteristics that would allow coil stock to be roll formed at relatively thick gages, as thick as 4.166 mm. Tailoring gages using the roll forming process is very efficient when compared to machining extruded sections. There is very little material discarded and the

Fig. 5.37

Fuselage roll forming process.

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Fig. 5.38

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Machining of large wing panels.

programmed rollers modify shape and gage in a continuous flow process. Machined stringers often require multiple machine passes or multiple set ups to completely machine the part. 5.2.10.2

DETERMINANT ASSEMBLY

With the increased use of determinant assembly techniques, dimensional stability of parts has become more critical than ever. Determinant assembly refers to the practice of designing parts that fit at predefined locations, precluding the need for setting gauges and other time consuming measurements or adjustment. The parts practically “snap together” while maintaining the necessary high positional tolerance and eliminating the need for assembly fixtures. The new assembly methods require full size coordinated pilot holes to align for the assembly of large monolithic parts. This concept reduces or eliminates fixed-base, monumentstyle tooling such as floor-mounted assembly jigs (FAJ) by using the airplane components as tools themselves. 5.2.10.3

HIGH-SPEED MACHINING

Fabrication costs for complex components can be reduced using high-speed machining techniques (Fig. 5.38). Per this method, machining with spindle speeds greater than 25,000 RPM and employing light-depth-of-cut practices can enable one-piece (or monolithic) fabrication of thin components. The F/A-18E/F fighter aircraft experienced a part count reduction of 42% simply by employing high-speed machining. Similarly the F-15 speed brake design was reduced from 500 assembled components to just one piece. To obtain the detail and surface finish required for these aerospace parts, light cutting is essential but quickly becomes economically infeasible at conventional feed rates. Increasing the spindle speeds to the point where the frequency of the cutting edge striking the material is the same as the resonance frequency of the milling machine, ultimately

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increases accuracy (enabling thin walled fabrication), surface finish, tool life and production rate. With an increasing use of high-speed machining and optimization of machine set-up, the aerospace industry has experienced a significant cost savings over traditional machining processes. 5.2.10.4

COMPOSITE PROCESS DEVELOPMENT

Automated tape layup was initially evaluated during the NASA ACEE 737 spoiler program in 1972. The concept was sound, but the equipment was very rudimentary. The Boeing Advance Composite Development Program (ACDP) investigated automated tape layup, pultrusion, filament winding, molding and resin film infusion in the early 1980s to reduce the cost of composites. The preferred materials at the time were prepreg unidirectional tapes and fabric. Prepregs are time and temperature sensitive materials, requiring freezer storage for long term storage and have very limited time at room temperature (out-time). Hand layup of prepreg, effective for small to moderate size parts, is not suitable for larger parts, due to the ergonomic limits, and layup time exceeding the material out-time limitations. The 737-300 was the first production program to use automated tape laying machines (ATLM). The size of the 737 parts elevators did not necessarily require the automated equipment, but this provided an opportunity to understand the capabilities of the equipment. The 777 was the first airplane to realize the full benefit from the ATLMs with the manufacture of carbon fiber/ epoxy stabilizer skins. Currently, contour tape laminating machines (CTLMs) and automated fiber placement (AFP) machines are the workhorses for moderately and highly contoured parts respectively (Fig. 5.39). Changes to hardware and software in conjunction with design for manufacturing as well as material form changes (wider and thicker) are allowing significant productivity improvements. An unanticipated benefit of the automated layup equipment was a major breakthrough in in-process inspection. Mounting laser pointers on the

Fig. 5.39 Automated layup methods—CTLM of stabilizer skin (left), advanced fiber placement (right).

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Fig. 5.40

Pultruded “I” stiffener.

tape laying head to indicate the ply edges eliminated the need for physical ply location inspection templates and greatly reduced the time to implement drawing changes. Use of freestanding optical layup templates, in conjunction with automated ply cutters even brought automation to hand laid up parts. Processes more commonly used in other industries, such as general aviation, boat building, and construction, have been looked at to determine the viability of these processes in commercial airplane applications. Pultrusion, used extensively in the manufacturing of beams and constant crosssection parts (e.g. ladder rails), has been investigated for stiffeners, stanchions and floor beams. Wide use of pultrusion for aerospace applications is not practical since, as stated previously, most axial members in aircraft require gage tailoring to be weight efficient. Pultrusion of standard “I” and “T” sections (Fig. 5.40) and honeycomb floor panels was looked at in the 1980s, but eventually dropped, partially because the need for constant cross-section parts is limited in an airplane. Processes such as resin transfer molding in conjunction with textile technology allow cost effective manufacture of complex geometry parts such as ribs, window frames, and fuselage frames. Chopped fiber systems are also being investigated for use in some discrete fittings and potentially the window frames for the fuselage. Composite laminated skins are produced in layers, done most efficiently in the form of continuous plies. The edge of panels where splicing or fastening occurs requires an increase in gage to provide adequate bearing strength. These local plies are typically smaller or narrower and force the tape laminate machines to make many short runs and many changes in direction. This creates an increase in machine time and slows part production. Also, co-bonded structure requires gentle ramp rates between interfacing parts unless specialized costly tooling is used. These gentle ramps add cost and weight to the parts in the form of excess material not required for strength. This parasitic weight penalizes the design. As a result, it is more efficient to build co-cured structure employing more aggressive ramps. Co-cured structure provides an added side benefit over co-bonded structure—it removes the need for complicated and sensitive surface preparation for co-bonding the pre-cured part to the uncured part. From an inspection and quality assurance perspective, it should be noted that a defect discovered in any component of a co-cured assembly is effectively a defect in the entire assembly.

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Aluminum designs tend to be easier to shape by machining. The base material is less expensive than advanced composites. As a result, it is important to utilize the composite material as efficiently as possible to minimize the amount of ply scrap when building parts. This is analogous to the aluminum “buy to fly” ratio. Although composite parts can be machined, it is costly and difficult to control the quality of the finished part. Many composite parts are profile or edge trimmed using routers or water jet cutters. Thickness machining is not performed except for rare circumstances. Sacrificial fiberglass plies are sometimes added to a part to provide a material that can be easily machined away. However, these materials will result in added weight to the base design. 5.2.10.5

COMPOSITES AFFORDABILITY INITIATIVE (CAI) [41]

In the mid 1990’s, the Air Force recognized that despite the potential of advanced composites to drastically reduce aircraft structural weights compared to conventional metal structures, the aircraft industry was reluctant to implement them in new aircraft. For example, despite projections early in the F-22 program of the airframe being 50% composite by weight, it settled back to 24%. Although composites were used on the F-15, F-16, and F-18 in small percentages, the perceived risk of this “new” technology was one reason cited in the F-22’s fallback from 50%. In addition as the F-22 faced cost scrutiny at the drawdown of the cold war, the high costs of composites compared to metals was another reason cited for this decrease. As a result, the Air Force Research Laboratory (AFRL) launched CAI to address these concerns. What resulted was a team of the AFRL, the Office of Naval Research, Bell Helicopter, Boeing, Lockheed Martin, and Northrop Grumman. State-of-the-art aircraft structures have thousands of parts and hundreds of thousands of fasteners. In addition, drilling holes and installing fasteners is a major source of labor and rework in aircraft structures. By the integration of parts and by bonding parts, structural assembly costs could be drastically reduced. As a result, CAI’s objective was to establish the confidence to fly large integrated and bonded structures with demonstrations performed for several aircraft, including the F-35, X-45, C-17, and X-32 (Boeing’s Joint Strike Fighter prototype). This required a multidisciplinary approach: maturation of materials and processes, an understanding of the structural behavior of bonded joints, quality assurance and non destructive evaluation to ensure bonded joints remain bonded throughout an aircraft’s service life, and the buy off of Department of Defense (DoD) aircraft certification authorities. This unusual breadth of functional areas for a research program represented a full spectrum of airframe design and manufacturing disciplines and was crucial to the success of this fundamental change in the way aircraft structures are built. The primary manufacturing technology pursued to fabricate large integrated structures was Vacuum Assisted Resin Transfer Molding (VARTM), a process used for making large yacht hulls. VARTM is a process that uses a lower than

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atmospheric pressure (typically full vacuum) to pull a liquid resin into a fiber bed. It was made famous in the boat building industry with the advent of the SCRIMP process (Seemann Composites Resin Infusion Molding Process). There are two key advantages of VARTM over conventional autoclave curing. First, is that an autoclave is not needed, resulting in reduced capital equipment costs and a much larger suppler base for part fabrication. Second is that the typical VARTM resins cure at a low enough temperature to enable the use of inexpensive tooling such as medium density fiberboard rather than the typical invar tooling used for 1778C (3508F) curing autoclave materials. While the aerospace industry had dabbled in VARTM over the years, CAI demonstrated its viability as a valid production method for large aerospace parts. CAI’s VARTM efforts resulted in fiber volumes and per ply thickness comparable to typical autoclave cured aerospace composite parts. In addition, the process worked with several resins, including EX1510, SI-ZG-5A, and VRM-34. Further use of the VARTM process would be enabled through the development of toughened resins with properties similar to 977-3 resin. CAI demonstrated VARTM on parts up to 4.6 m2. The process has shown its versatility as demonstrated by the broad range of parts fabricated as seen in Fig. 5.41. These demonstrations have shown VARTM can drastically reduce part and fastener counts through integrated structures and at much lower fabrication costs. For example, these parts displayed in Fig. 5.41 were originally subassemblies on the order of 20 parts and hundreds of fasteners and were converted into a single part using the VARTM process. 5.2.10.6

ASSEMBLY

Manufacturing processes also play an important role in the structural performance of the design. Improved assembly techniques must help to improve the static, durability or fatigue performance of the structure. For example, improving the quality of fastener installation can allow for higher operating stresses while providing similar fatigue reliability of current structures (Fig. 5.42). Conversely,

Fig. 5.41 VARTM demonstration parts—X-32 cockpit tub (left), C-17 nose barrel skin (center), C-17 nose landing gear door (right).

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Fig. 5.42

Automated splice fastening eliminates variability of hand installations.

poor fit-up of parts adds to assembly stresses that can lead to stress corrosion problems and fatigue problems of aluminum structure. Composite parts are sensitive to pull-off loads at flange radii such as for spar, rib and frame chords. Interlaminar tension stresses caused by bending must be controlled to avoid inter-laminar fracture and delamination. While bonded structural joints are currently in service on military aircraft including the F-18 and Global Hawk, there are legitimate concerns on the expanded use of bonded structures. The inability to discriminate between a good bond and a “kissing” bond (intimate contact between adhesive and structure without adhesion) has been the key roadblock to further use of bonded structures. To combat this, current certification policy requires a traveler coupon or a proof test with each bonded joint to ensure its structural integrity, each of which is expensive. In addition, adhesive aging and prolong “out time” will affect mechanical performance of the bonded joint. Despite this unease, bonded structures have tremendous potential for aircraft structures. If designed correctly, bonded aircraft structures can have greatly reduced part count and fastener count and also greatly reduced structural assembly times. CAI’s [41] bonded structures work centered on the “pi” joint (Fig. 5.43). This stiffener, shaped like the Greek letter p, can be co-cured or co-bonded to the skin. The pi joint has several advantages. First, it provides structural redundancy. The pi joint acts as a double lap shear joint, increasing the surface area for bonding. When used with EA 9394 adhesive, the pi joint takes advantage of the inherent properties of that adhesive, that being excellent shear properties. Component structural testing has shown that the joint is very robust and has predictable performance. Pi joints also paved the way for much reduced assembly times by eliminating the frequent checks needed to ensure the correct adhesive thickness tolerances, and reducing the “out time” concerns associated with faying surface bonds. Large component demonstrations (Fig. 5.44) resulted

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Fig. 5.43

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Cross-section of a Pi joint.

in drastically reduced assembly times. CAI studies showed that assembly times for bonding can be reduced from 50 to 80% over typical fastened structure depending on the article, translating to a cost savings of 20 to 50%. Part count has traditionally been a significant contributor to build cost. Figure 5.45 exemplifies traditional built up aluminum structure. Figures 5.46 shows the reduced part count expected due to the use of monolithic components for aluminum fuselage. The move to unitized structure helps to reduce part count and lower build costs. Unitized structure concepts integrate determinant assembly methods and demand tighter tolerance for detail components fabrication, but the reduction of tooling will dramatically reduce the entire product cost. Changes and improvements to the designs are easier to introduce since tooling modifications

Fig. 5.44 CAI demonstration articles—F-35 forward fuselage (left), X-45 wing carry-through (right).

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Fig. 5.45

Conventional 777 skin, stringer and frame construction.

are not required. Composite assemblies have taken advantage of more monolithic designs for some time. The philosophy has been slow to develop in metal designs for the reasons previously stated.

5.2.11

IN-SERVICE REQUIREMENTS

Material performance while in-service is a reflection of the design and impacts the life cycle cost of the airframe. Environmental exposure to hail, lightning strike, electromagnetic effects from electrical bonding and grounding of systems, proximity to corrosive materials, moisture and extreme temperatures can impact the choice of material systems. Wiring and power systems require adequate current return paths to protect for system surges and lightning strikes. The relative ease of repair techniques for incidental damage must also be taken into account. Repair methods play a significant role in detail design and fabrication method analysis to ensure that the operators can maintain fleet dispatch reliability by facilitating rapid, structurally efficient repair procedures.

Fig. 5.46

Future monolithic frame and shear tie construction.

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5.2.11.1

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HAIL AND INCIDENTAL IMPACT

In the vehicle design and manufacturing development phases of a project, consideration must be given to impact situations such as hail or bird strike. Resistance to impact due to hail is a design requirement that establishes minimum wing and fuselage skin gages. These gages are resistant to damage from all, but not to the most severe hail storms. The aluminum structure in service today can be as thin as 0.91 mm. Hail impact with sufficient size and energy tends to dent thin gage aluminum. Fortunately, the damage can be repaired or filled depending on the criticality of the impact size and location. In addition to hail, consideration must be given to bird impact and impacts experienced during the aircraft assembly process. For composites, hail impact requirements often establishes minimum gage for composite structure. Composite materials will not dent like aluminum sheet, but will suffer from inter-laminar damage if the impact has sufficient energy. Risk of delamination after being exposed to ground hail can be costly for the airlines requiring extensive non-destructive inspection if it is not part of the original design requirement. Composite design philosophy today is based on a “no detrimental” damage growth approach. In order to achieve this, the component allowables are developed from damaged samples with impacts of sufficiently high energy levels to create barely visible and visible damage and tested under expected cyclic loads to establish that the damage does not grow. Based on these allowables, component gages are determined that would be sufficient to preclude delaminations due to most hail and other impacts, and create operating strain levels such that delaminations would not negatively impact the strength of the structure or grow to a detrimental state. For co-bonded structure subject to impact damage, care must be taken to ensure that bond line failure due to impact does not lead to catastrophic failure of a major component. Fuselage and wing structure are also exposed to incidental impact damage due to service vehicles and galley carts while being serviced on the ground. Additionally, there are occasional on-ground airplane-to-airplane contacts that require repair before operation. Incidental damage to conventional fuselage and wing structure can range from small dings, nicks, gouges to the more severe skin penetrations or severed stringer and skin elements. Small dents, scratches and gouges can often be blended out of aluminum structure. Composite structure does not suffer from fatigue concerns for small incidental damage, but delaminations must be identified to ensure they are not of a size that would degrade the static strength below regulatory requirements. 5.2.11.2

LIGHTNING AND GROUNDING

Airplane system requirements affect the way material systems can be used in a fuselage. The fuselage contains a number of electrical systems that support avionics, in-flight entertainment, galleys, lavatories and antennas. Aluminum

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structure provides a natural current return path if systems are appropriately attached, whereas composite structure does not provide an adequate current path due to the high electrical resistance of the epoxy resin system. Grounding provisions are needed within the fuselage to provide a current return path. Outer skins made from composites require shielding to help dissipate current over the fuselage surface. This has been addressed by adding either expanded foil or bronze/copper mesh over the exterior surface where threat levels are severe. The mesh can be woven into the carbon or fiberglass fabrics currently used for aerospace applications. The lightning strike and electrical grounding protection schemes in a composite wing are very similar to those employed in the fuselage. The additional concern is the presence of fuel and fuel systems. Wire mesh, buss bars, sufficient metallic substructure and special fastener installation techniques are available to meet the lightning threats. The design features required in an all-aluminum wing are: 1) high-interference fasteners that are cap sealed to prevent corrosion from deteriorating the intimate connection for the grounding path, and 2) electrical bonding jumpers where necessary. 5.2.11.3

EXTREME TEMPERATURE AND MOISTURE

A material must be able to withstand the extreme environments to which commercial, military and general aviation aircraft are exposed. Temperature extremes in the aircraft can be caused by exposure to harsh external environments or the high temperatures generated by internal equipment such as hydraulics, electrical power units, high voltage electrical boxes, or auxiliary power units. Internal equipment can create local temperatures in excess of 1208C. If a material system is temperature limited, it may require cooling techniques to keep the structural temperature within acceptable limits. These cooling systems could add additional cost, weight, and maintenance requirements to the aircraft. Designs that use materials of dissimilar thermal coefficients of expansion must account for additional loading induced by a thermal growth mismatch. For example, a thermal growth mismatch may create a more fatigue critical environment for the aluminum structure when rigidly fastened to parts made from carbon fiber materials. In addition, high ambient temperatures and moisture absorption in composite materials can lead to degradation of material properties, and necessitates the evaluation of hot-wet material characteristics. 5.2.11.4

CORROSION

The fuselage may be exposed to a number of corrosive fluids, (i.e., spills from galleys and lavatories) that can collect on floor beams and lower skin panel assemblies. Moisture from condensation can collect on the crown skins and migrate towards the belly. These fluids can create a corrosive environment for aluminum alloys. Material selection and development must accommodate the need for

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increased corrosion resistance. In-service experience has helped to drive development of material forms and heat treatments that are less susceptible to corrosion. Corrosion is still a significant maintenance item for current aircraft. Composite materials are corrosion resistant, however use of composite materials in contact with aluminum creates a galvanic corrosion path that can accelerate corrosion of the aluminum. Proper care must be taken in the design development to ensure that the materials are adequately protected. 5.2.11.5

REPAIR

Once damage is inflicted, it is critical that the materials and design architecture be repairable. Some materials and damage scenarios might require bonded repairs while other material systems and locations can utilize bolted repairs. Repair options also vary depending on whether the damage is introduced prior to delivery or during service with an airline. The basic goal of repairing composite parts is to restore structural integrity in as short a time span and at the least cost. From a structural stand point, generally what an engineer strives to do is to match stiffness and strength of the repair to the parent material while keeping weight gain down to a minimum. There are of course many other factors to consider as well, and those are often part of the trade-offs in designing such repairs. Factors such as durability (fatigue, corrosion, and environmental effects), aerodynamic smoothness, weight & balance, service requirements (temperature, fluid, moisture effects), fuel system sealing, lightening protection, cost, schedule (downtime, facilities, equipment, skill level of personnel available to perform repairs, materials, etc.) must be considered for an effective repair. It should be noted that the repair must restore the capability of the part for design ultimate load and full service life in order for it to be considered permanent. Damage to composite parts can come from a variety of sources. In manufacturing damage can come from improper processing resulting in delaminations or higher than acceptable void levels, mis-drilled holes, oversized holes, fiber breakout on drilled holes, etc. There is damage that comes from mishandling such as dropped tools, service trucks impacting composite structure, and other maintenance inducted damage. Then there is environmental damage due to such elements as hail, lightning strikes, debris, and bird strikes. Damage to composite structure is typically categories into three levels. The lowest level is “negligible damage”. Generally this is damage that can be permitted to exist “as is” without any negative effect on the part’s structural properties or limits. It may be repaired with simple smoothing of nicks and/or scratches to prevent any degradation due to environmental effects. The next level is “repairable damage”. Parts/structures categorized at this level can be restored to design ultimate loads. Parts/structures that cannot be repaired so as to have adequate margins of safety (i.e. are beyond repairable limits) fall into the category of “replacement”.

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For parts that can be repaired (i.e. are “repairable”), there are basically two different types of repairs, bolted and bonded. Bolted repairs have many advantages; they are not affected by the environment, there is less surface preparation necessary, they are generally easier to perform (i.e. less skill is required, no refrigeration or heat blanket required), they are suitable for field or line station repair and they are easier to inspect. The cons are that they add weight to the structure, require holes in the parent structure (weakens it, source of fatigue and stress concentration), mis-drilling can make structure susceptible to delamination, are generally not suited for honeycomb structure or thin laminates, and it results in nonuniform load introduction into the parent structure. The patch/ doubler material can be metallic or composite, but if metallic attention must be paid to the potential for galvanic corrosion. Bonded repairs inherently overcome one of the most negative attributes of bolted repairs, namely nonuniform load introduction into the structure. This is due to the load being evenly distributed over a much wider area (i.e. the entire bondline not just through bolts/fasteners) and this makes it well suited for repairing thin laminates and honeycomb structure. The negatives though include degradation in service (via temperature/humidity), significant surface preparation steps (following the process is paramount), difficulty in inspecting the bondline (i.e. lack of being able to detect the “kissing” bond), they are more difficult to design, they are more difficult to perform (i.e. requires technicians with higher skills, heat blankets, vacuum equipment and hotbonders required to cure the laminate (wet layup or prepreg)), and they are more time consuming. For thicker laminate structure scarf (tapered) repairs are the most efficient for uniform load transfer but it expensive to perform and requires extra time and skill to perform. Repairs for honeycomb structures, which tend to be lightly loaded, can be more difficult to assess than for laminate structure. This is because the damage to the core beneath the skin may remain undetectable until the skin is removed. Permanent repairs of sandwich structures with damage greater than the allowable damage level require elevated temperature cures (hot bond repairs). Hot bond repairs are preferable to bolted repairs due to the fragility of the minimum gage honeycomb structure. If the core is damaged, it should be removed and replaced. The replacement core must be thoroughly dried, otherwise the trapped moisture will turn to steam during the heated repair and will blow out the core causing more damage than originally present. For this reason heating rates and soak temperatures are important parameters to watch during the repair. Honeycomb laminates can be part of the flight control surfaces for aircraft and how much additional weight and where it is located on the part can be significant factors as they relate to weight and balance and flutter control. Bonded repairs also include injection repair of delaminations in laminates or for skin to core disbonds in honeycomb structures. This has the advantage of repairing damage without removing the damaged materials, and can be done easily in the field. A two part low viscosity resin/adhesive is injected into the

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damaged area at an edge, through drilled holes, or set of holes at one edge of the delamination and then sucked through to another hole at the opposite end of the delamination. The disadvantage is that the condition of the delamination is unknown (i.e. is contamination in the hole from hydraulic oil or grease) and hence the efficacy and permanence of this type of repair is also unknown. The type of repair (that is the skill level of the technician and the facilities required) dictates where and at what level it can be performed. For instance, for bolted repairs only minimal skill levels and facilities are necessary, hence this type of repair can be conducted at the field level (military) or line station (commercial) at facilities with minimal capabilities (forward operating base for the military; airport for the commercial sector). For more technically challenging repairs, for instance large area scarf repairs to thick laminate structure, it would be more suitable to be performed at depot level (military) or maintenance base (commercial), since these locations have better facilities, higher skilled technicians and the appropriate environments to affect these repairs. The use of composites in the design and fabrication on aircraft structure is continuing to increase, and affecting repairs to these structures will be crucial to keeping them economical. This will necessitate the continued development of composite repair methods that are structurally adequate as well as economically practical.

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Keith B. Bowman‡ Wright-Patterson Air Force Base, Ohio

5.3 ROTORCRAFT The major difference between rotorcraft and fixed-wing aircraft of the previous sections is that lift is generated using rotating blades instead of wings, giving the rotorcraft the ability to hover. The general design rule is that the blade tips need to stay subsonic, which limits the forward speed of the rotorcraft before the retreating blade stalls, thereby losing lift. Helicopters are relatively short range slow vehicles for lower altitudes with the capability of taking off and landing anywhere. Advanced rotorcraft, such as tilt rotors, attempt to keep the hover capability of helicopters with the benefit of higher forward speeds of fixed-wing aircraft. In each case, rotor blade design is a challenge in an operational environment where high rotational speeds combined with long thin flexible structures leads to millions of cycles where vibration and flutter dominate. These considerations all drive material selection.

5.3.1 ROTORCRAFT PRELIMINARY CONSIDERATIONS Rotorcraft with their ability to use any available open space as a potential landing site have become ever more useful and relied upon in modern society. The ability to operate off any surface poses unique material challenges to the design of the system and in particular to the design of the rotor blades. While most rotorcraft materials are used and qualified according to industry standard requirements outlined elsewhere in this book, there are a couple of unique applications that are worth mentioning. The emphasis of this section is on design aspects unique to rotorcraft and not often considered in the design of fixed-wing systems. The design and manufacturing of composite rotor blades with operational considerations such as field repairs and rotor blade erosion coatings are critical aspects of rotorcraft design that require special material considerations.

5.3.2 ROTOR BLADE OVERVIEW While some legacy designs still use metal blade spars, newer designs are made from composites. This poses numerous material challenges during design and qualification. Not only do rotor blades generate extreme numbers of fatigue cycles, but due to their extreme flexibility, dynamic stability also drives the material selection. Material strength, stiffness and mass distribution all must be traded to achieve a viable blade design. A micron-level change in thickness of a ply or adhesive over the entire length of a blade can result in an unacceptable mass distribution. While many aircraft designs search for lighter weight materials, ‡

Air Force Research Laboratory.

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Fig. 5.47

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“D” spar example from a composite rotor blade.

rotor blades in many cases cannot reduce or add weight without affecting the dynamic stability. Material weight is very tightly controlled and blades that weigh several hundred pounds must be within ounces of each other to balance all of the blades on the rotor head. Composite rotor blades that are still flying on Army rotorcraft started design and development in the early 1970’s with all fiberglass designs. The spar consists of a heel, leading edge and upper and lower spar walls which are referred to as a “D” spar due to the shape (Fig. 5.47). The fairings aft of the spar are sandwich structure consisting of fiberglass skins over non-metallic honeycomb. The spar, leading edge, heel and fairing are laid-up and cured as separate details before being secondarily bonded together. Blades still use a metal leading edge, generally titanium or steel, which protects the leading edge and provides chord stiffness in some designs. The outboard portion of the leading edge is a replaceable nickel strip. Tungsten weights are used for dynamic stability and tuning the blade for track and balance. Metal trim tabs are common and since fairings are fiberglass, aluminum lightning mesh is also used. Other metals are used to provide a lightning path from the blade into the rotor hub. Some designs use metal fittings to mount the root end to the rotor hub that are either mechanically attached to the blade or adhesively bonded to the composite. Other blades use composite pin wrap designs to attach to the rotor hub. As a result of the blade flexibility with composites, metal components sometimes cannot be used due to high blade strains or must use a layer of soft film adhesive in order to not drive loads into the metals. For example, metal trim tabs are being replaced by rubber wedges because of the high strains. Newer design blades incorporate carbon fiber into the blade spars, although most of the fairings are still fiberglass sandwich. Military rotor blades include fiberglass with the carbon in the spar for ballistic tolerance. The high strain-to-failure of glass is beneficial to meet this requirement. More advanced

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aerodynamic designs with swept, tapered, anhedral and dihedral blade tips put more material aft of the quarter chord and require careful consideration to keep the blade stable. These new designs consequently complicate the manufacturing of blades, specifically the tooling for composites, and require materials to react the out-of-plane loads generated by not keeping everything straight. The primary loading on a rotor blade is due to the centrifugal force (Cf). While flap bending, chord bending and torsion are also predominate load cases, the Cf load dominates all of them. Except for ground loads, compression loading on the blade is rare. Since the predominant load is Cf, rotor blades spars use unidirectional prepreg running spanwise. Composite plies running +458 are common as a wrap to react the torsion loads. Blade fairings commonly use either +458 or +608 plies to address torsion loads. Earlier blades also used co-cured skins on the fairings. Due to the lower consolidation pressure for co-cured skins (so as not to crush the core), as well the core cell walls being the only place where the external pressure is reacted, microcracking is common and provides opportunities for moisture to wick into the structure. Between microcracking and the X-ray tracers for radiography not bonding well with the resin, water ingression is problematic. This results in the blade eventually failing track and balance. Blades must be dried under controlled temperature and vacuum to draw water out without further damage to the blade. For some larger blades, the amount of water that is removed can total several gallons. Newer blades use precured skins that are secondarily bonded to the honeycomb to reduce the consequences associated with unintended microcracking. Most composite blade spars and composite rotor hub components are thick structures with hundreds of plies. Maintaining ply orientations and preventing wrinkling and marcelling during manufacturing is a major challenge that needs to be addressed, as is the subsequent inspection of these thick laminates to discern component quality and integrity. For some spars, ultrasonic inspection is not feasible and newer technologies such as thermography cannot drive enough thermal energy into the spar to return usable results. Radiography is especially useful in finding wrinkles and is commonly used when other nondestructive inspection (NDI) techniques are not suitable.

5.3.3 ROTOR BLADE ALLOWABLES Static design allowables for composite rotor blades follow industry practices set out in MIL-HDBK-17/CMH-17. A-basis allowables are typically used due to the criticality of the structure and the single-load path aspect of a typical blade. Some blades use 1778C cure materials while others use 1218C cures, especially in press cures. Lower cure temperatures (e.g. 1218C) can be problematic in providing a glass transition temperature high enough for the 828C wet environment typically required for static allowables. In some cases, detailed thermal analyses using finite element models are used to justify a lower temperature. While sitting on the ground, blades can reach temperatures exceeding 938C, but can

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quickly cool to ambient temperatures once the blade starts spinning. The qualification temperature used in the design depends on the rate that the blade cools and the time to reach the flight condition where a maximum load maneuver is executed. In addition, all blade laminates are assumed to be saturated. Some argue that there is not enough humidity to saturate blades operating in extreme desert heat, however it is not uncommon to transport blades from the high humidity of Guam to rotorcraft in the desert in less than 24 hours. Due to the laminate thickness, it is not likely that all of the plies will reach saturation, but the outermost plies can reach critical saturation levels and may be the most highly loaded plies in the laminate. To generate fatigue allowable curves, the environments that are required are the maximum ambient operating temperature, typically 49–548C, at wet conditions and the minimum ambient operating temperature, usually 2408C. It is not feasible to expect the rotorcraft to be at the extreme environmental conditions for its entire lifetime. In reality the blades will never see these environmental conditions over the million of fatigue cycles in their lifetime, but designing to these extremes provide adequate conservatism. For other extreme load cases such as crash or ballistics, imposing environmental knockdowns given the vast uncertainties in the events is not justified. These analyses use room temperature allowables given that crashes are extremely rare events and that the parts are allowed to fail, though not to the point they compromise the cockpit and cabin. While load enhancement factors are traditionally used for accounting for the environmental knockdowns during room temperature qualification tests, in some cases the structure must be environmentally conditioned. In one case for a bonded blade attachment fitting, the adhesive was critical at cold dry conditions while the laminate was critical at hot wet conditions. A single load enhancement factor was not feasible. This required conditioning a full scale blade section in a humidity chamber for over six months.

5.3.4 FATIGUE Fatigue is the key design driver for rotor blades, and is the reason that pushes the design toward composite materials rather than metals. Cracks were common in earlier metallic blades, resulting in degradation of integrity and service life. Unlike metals, composites fail in such a way that a single defect is unlikely to cause a catastrophic failure of the component. In addition, the fatigue life for a composite material generally levels-out to a relatively flat endurance limit. Keeping the component strains below this value is the goal of the designer. One of the most important decisions in developing fatigue curves for rotor blades is the selection of the curve shape. The curve shape will have the appropriate knockdowns applied and then used for the life calculations for the blade. The particular curve shapes are determined according to each manufacturer’s fatigue

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methodology, which are company-specific based on lessons learned. The airworthiness authority must agree to the fatigue methodology and the resulting advantages and limitations. Unlike fixed wing aircraft that are generally designed to a damage tolerant or fail-safe structural methodology, rotor blades are designed to the safe life methodology. That is that no component should ever crack in its lifetime. Usually the structural life is considered to be 10,000 hours. When considering that the blades revolve at 200–300 revolutions per minute it is easy to see how hundreds of millions (if not billions) of fatigue cycles drive the design. Under the safe life methodology, six full-scale fatigue specimens are tested for each blade section. In rare cases, four specimens will be used with substantial statistical knockdowns to their lives. One of the biggest design drivers for military rotor blades is ballistic tolerance. An incredible amount of damage can be generated by a large caliber high explosive round. The blades must be able to survive long enough, typically 30 minutes, for the rotorcraft to get out of the hostile area and be able to land. Equally impressive is the amount of damage a composite blade spar can tolerate and still meet these requirements. As long as the fibers in the spar are not severed too badly, the Cf can hold the pieces straight and keep it flying. Once the blade stops, the damage may act like a piano hinge. Figure 5.48 illustrates the amount of damage that a blade can sustain and still be able to function long enough to safely land. Other blade damage requirements are similar in nature and include tree strike and lightning strike. Rotor blades are a typical pathway for lightning strikes on rotorcraft. Main rotor blades are the entry point, and tail rotor blades may be the exit. Tree strike requirements account for the possibility of severing branches 5 to 10 cm in diameter. Like ballistic damage, the intent of both requirements is to

Fig. 5.48

Rotor blade damaged by hostile fire.

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insure that the blade does not instantaneously fail, giving the pilot enough time to land.

5.3.5 COMPOSITE FIELD REPAIR Composite rotorcraft structures pose a myriad of repair challenges. Rotorcraft can be operated out of austere environments with little to no support structure, and can be easily damaged by landing on rocks, trees, or fence posts, or for military systems be damaged by enemy ground fire. No hangars or other buildings may be available to provide a clean environment to repair the composite or adhesively bonded structures. These locations do not have freezers, autoclaves or ovens to use prepregs or film adhesives. They may not even have power to run a hot bonder or vacuum pump and may only be reached by another rotorcraft. The tools required to repair the structure must be self-powered and fit inside of a suitcase. The Army is buying battery operated hot bonders and vacuum pumps to meet this need. The speed of repair is extremely important. The ability to fix the rotorcraft and get it back flying, especially for military rotorcraft, may be an issue of life and death. Some composite repairs in the Army must be accomplished in less than an hour which negates using standard composite repair techniques. Non-removable fuselage damage must be repaired with mechanically fastened metallic patches. Blades and removable fuselage structure are often swapped-out in the field. Due to the criticality of rotor blade spars, most damage to blade spars requires scrapping the blades. New research efforts are trying to qualify blade spar repairs for small arms damage. Since any repair to the blade spars requires full scale fatigue testing, it is easy to understand why spar repairs remain unusual. Just a single full scale fatigue test requires months to complete and numerous tests are usually required to establish the reliability of the repairs. Blade repairs involving fairing damage are typically limited in size due to the weight increase. Typically wet lay-up resins and paste adhesives result in heavier repairs. Beyond a certain size or number of repairs on a blade fairing, the blade can no longer meet the track and balance requirements. Either the blade must be returned to the factory to have a new fairing installed or most likely scrapped. Similarly, leading edges are an integral portion of the spar and cannot be removed and replaced. Some blades have supplemental doublers that can be bonded onto the leading edge in order to return the blade to service.

5.3.6 ROTOR BLADE EROSION COATINGS The world’s deserts are an extremely difficult environment to operate a rotorcraft. Figure 5.49 shows an OH-58 and UH-60 completely covered with sand. Rotor downwash kicks up tremendous amounts of sand and dust, resulting in erosion of rotor blades, pitting of the cockpit transparencies, damage to engine components, and a general accumulation throughout the rotorcraft. Rotorcraft hovering over any unimproved areas will kick up sand, rocks, and other debris that can

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impact the blade and result in damage. Since most leading edges (other than the leading edge nickel strips) are not repairable, rotor blade erosion can be a significant cost. Military operations in Southwest Asia have identified the critical need for better rotor blade erosion materials. A brief history of polymer erosion coatings back to the days of Estane in the 1970’s can be found in [42]. Most current erosion protection coatings that have been applied to Army blades [43] consist of polyurethane tapes, boots or coatings. Current blade tapes use pressure sensitive adhesives (PSA) which make them easy to install and remove. Some polyurethane tapes are not compatible with PSA and require a two-part epoxy adhesive to bond the tape to the blade. While these adhere well, removal and replacement of the tape must be accomplished at the depot. Newer coatings include sprayable polyurethanes as well as moldable boots [44]. The latest research is focused on even longer life coatings using ceramics and metals that can survive at least 1,000 hours in a sand erosion environment [45]. One important lesson that has been learned in recent erosion qualifications is to reduce or eliminate the involvement of the maintainer’s role in applying these protective coatings. Current tape materials applied in the field have widely varying lives due to the installation. One operational unit, with a technician dedicated to applying blade tape and paying particular attention to the installation, can get several hundred hours from existing tapes, whereas another unit may have tape sling -off at start-up even after waiting the 24 hour cure time. As a result, in-field repairs using tape materials have been banned due to this variability. By applying the material in the factory with appropriate engineering controls and then providing a small repair kit for field maintainers, the consistency of the coating performance greatly improves. Unlike with swept blade tips which have been in operation for many years, conforming the tape material around more complex shape blades that incorporate anhedral/dihedral tips for improved aerodynamic performance is proving challenging. The ultimate goal is a coating that performs well in both sand and rain. Existing metal leading edges perform well in rain, but poorly in sand. Polyurethane

Fig. 5.49

Rotorcraft in Southwest Asia.

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Fig. 5.50

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Titanium sparking due to sand erosion.

materials perform well in sand and poorly in rain. A single material that meets both requirements as well as can be integrated onto the rotor blade is desired. There is some interest in developing a military specification for rotor blade erosion materials. Since there is so much tailoring of the qualification requirements to address the rotorcraft mission and the type of material, coming up with a general specification becomes impractical. Another problem with rotorcraft operation in sandy environments is what is called the “Halo effect.” Many blade leading edges are titanium which when experiencing sand erosion results in the blade glowing. This is extremely detrimental to military rotorcraft operating at night since it makes the aircraft clearly visible. Figure 5.50 shows a titanium sample in the sand erosion rig giving off sparks from the sand particle impingement. Rotor blade erosion has shown to be a serious problem with rotorcraft operating out of Southwest Asia. Extensive erosion research has been conducted since the early 2000’s, building on research conducted after the first Gulf War in 1991. The Department of Defense Rotor Blade Erosion Working Group [46] was created to leverage numerous research and development programs across DoD and standardize test methods for rotor blade erosion. The Working Group includes representatives from the Army, Navy, Air Force, and Coast Guard, and works collaboratively with industry to ensure the lessons learned are available to the larger rotorcraft community. 5.3.6.1

SAND EROSION

Existing metal leading edges are extremely susceptible to sand erosion. Figure 5.51 shows a tip cap that has been completely eroded through after only 10 hours in an

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Fig. 5.51

Eroded main rotor blade tip cap.

erosive environment. Most blades integrate a replaceable nickel strip on the outboard portion of the leading edge. However, erosion just inboard or aft of the nickel strip can result in excessive damage and require scrapping the rotor blade. The operational impact of sand erosion can be substantial [47, 48]. Blades may last only 10 hours before excessive erosion leads to scrapping of the blade. This is especially evident when pilots are training for operations and landing in dust clouds (called brown-out) prior to deployment. With the cost of a single main rotor blade in the hundreds of thousands of dollars, just one rotorcraft can cost a million dollars or more in blade replacements after a few hours of operation. In addition, the operational impacts of rotorcraft down for blade replacement are significant. It can take a day or more to remove and replace a blade, and several days to receive a new blade from the supply system if spares are not on hand. Typically AR 70-38, MIL-HDBK-310 and MIL-STD-810 [49, 50, 51] are used to define the environmental conditions for military rotorcraft. For rotor blades, these resources do not adequately provide the detail necessary to evaluate advanced materials for sand erosion. AR 70-38, MIL-HDBK-310 and MILSTD-810 provide general guidance on the sand composition and primarily focus on the sand concentration and particle size distribution. In addition, it has been shown that the particles’ chemical composition (differing amounts of quartz, feldspars, pyroxenes, black oxide materials and a large amount of silts and clays) will result in different densities and particle momentums [52]. In 2010, a new military standard, MIL-STD-3033 [53], was provided that detailed testing procedures for rotor blade erosion coatings and provided a more detailed sand definition. MIL-STD-810 uses the Krumbein number of 0.5 to 0.7 for roundness and sphericity to define the particle shape. Research has shown [52] that the Krumbein scale for particle shape does not adequately define the more aggressive sands from Southwest Asia. This research, based on a geological analysis of various sands, uses a relationship between the sand particle aspect ratio and the ratio of an apparent

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circumferential radius (RC) to the apparent area radius (RA) of the particle. The ratio of RC to RA can be described as: The measurements of greatest concern for this study were the ratio of the apparent circumferential radius (RC ) to the apparent area radius (RA), as well as the aspect ratio of the grain. A large circumferential to area radius ratio value (RC/RA) declares that the average roughness or angularity of a grain is very large meaning that the grain is pointy. [52]

Figures 5.52 to 5.55 [52] show that compared with the Krumbein scale, sands in Southwest Asia characterized as “geologically fresh” are significantly more fragmented with sharp edges and points. Sands typical of the Krumbein scale are characterized as “potatoes” in comparison. Figure 5.53 shows a commercially available golf course sand that better represents the Kuwait sand in Fig. 5.55. Figure 5.54 shows a foundry sand that was previously used for erosion testing that does not compare as well to golf course sand as a test media. Beside the morphology of the sand, other parameters such as what impact angle to subject the coatings to, and whether to test the blades to static or dynamic conditions become key in the testing of coatings. Experience has shown that polyurethane materials erode more at low impact angles (approximately 308), whereas metals erode more significantly at 458 [54]. Static tests (as discussed in MIL-STD-3033) do not simulate the centrifugal loads experienced by both the blades and coating. Material failure may occur earlier when combining centrifugal loads with sand exposure for some materials. Weigel [55] describes

Fig. 5.52

Plot of particle angularity vs aspect ratio for the Krumbein Scale.

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Fig. 5.53

Fig. 5.54

Plot of particle angularity vs aspect ratio for golf course sand.

Plot of particle angularity vs aspect ratio for foundry sand.

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Plot of particle angularity vs aspect ratio for Kuwait sand.

using a whirling arm rig for sand erosion testing. The whirling arm rig exposes the material to a sand environment at a set spin rate in an attempt to better simulate the operational environment of the rotor blade. However, while static testing applies no load to the specimen, the whirling arm rig may apply too much. A whirling arm rig may apply up to 2000 g’s on the specimen when the actual rotor blade only experiences about 500 g’s. The advantage of a static test is that up to 16 test specimens can be tested in the time that only a few can be evaluated using the whirling arm rig. The definition of failure for a rotor blade coating is also a requirement for qualification. Mass Loss is the easiest method and simply involves comparing the pre- and post-test mass of the material. Generally, the material will lose mass, but polymer coatings misleadingly can gain weight when the sand becomes imbedded in the coating [54]. To complicate matters, the weight gain was found to vary as a function of impact angle due to the elastic versus plastic response of the polymer. The measurement accuracy also becomes an issue when looking at very small mass losses. Nickel substrates can require measurements to five significant digits (based on a 2.54 cm  2.54 cm specimen) to detect any mass loss. Volume Loss is a better measure of the amount of erosion, but is more difficult to quantify due to wear non-uniformity, erosion wear exposure area accuracy, and uneven sample wear overlap. MIL-STD-3033 describes various methods for determining the performance of coatings based

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on the specific material type. Extremely thin coatings, such as diamond, may not be measureable by either mass loss or volume loss. In this case, penetration to the substrate may be the evaluation method employed, and includes assessing the cracking, pitting, debonding, or wrinkling of the coating or substrate. Another parameter of sand erosion testing that must considered is velocity. MIL-STD-3033 currently lists 222 m/s (730 ft/s) for the speed of the erosion jet, which correlates to the conditions at the rotor blade tip. Depending on the size and mass of the particle, there is a range of velocities that may exist in the plume. Progressing inboard along the blade, the velocity will decrease and materials being considered near the root end of the blade must be evaluated at an equivalent velocity. 5.3.6.2

RAIN EROSION AND ENVIRONMENTAL EXPOSURE

A rain drop hitting the tip of a rotor blade is a highly dynamic event, imparting a shear wave into the erosion coating material [55]. The wave is extremely transient and results in large strains in polymeric coatings and ultimately failure in the bond line. The impact energies are similar to a .22 caliber bullet. The combined effect of both rain and sand exposure is even more detrimental since the damage to the coating surface from sand makes it more susceptible to rain erosion [44]. The test standard used to assess rain erosion is ASTM G 73-98 [56]. Figure 5.56 shows an image of a whirling arm rig that is frequently used. In this configuration, the specimen is placed at the end of the arm and circulated through a line of raindrops of a certain diameter and rate. To experience rain erosion, the rotorcraft does not necessarily have to fly through rain [48]. Missions that require hovering over bodies of water

Fig. 5.56

Air Force Research Laboratory whirling arm rain erosion rig.

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(e.g. search and rescue, fire fighting, etc.) can experience rain erosion damage as water droplets are kicked up by the rotor wash. 5.3.6.3

OTHER OPERATIONAL AND ENVIRONMENTAL CONSIDERATIONS

In addition to sand and rain, other environmental conditions must be considered in the qualification of coating and rotor materials. Hydrolysis of polyurethane coatings was discovered to be a problem in the early days of development [42, 55]. In humid environments in the southeastern United States, Panama and Guam, the coatings would hydrolyze and literally drip off the blades. Hydrolysis testing should be a standard test requirement in any qualification program for polymers. Polyurethanes also can experience a propensity for fungal growth, and should undergo fungus testing as noted in MIL-HDBK-454 [57] and MIL-STD-810 [51]. Other environmental parameters such as ultraviolet solar radiation effects (for polymeric coatings) and salt fog (for metallic coatings) should also be considered as part of the qualification process. The key is that the entire operating environment of the rotorcraft must be accounted for during the qualification phase. Downwash from a rotorcraft landing in unimproved areas can kick up rocks, sticks and other debris, posing a potential harmful situation to the blade’s coating materials [47, 48]. The impact of this debris on the coating can cause damage sites that can further degrade from subsequent erosion mechanisms. Typically rotor blades were qualified for tree strike using hardwood dowels, but these methods do not cover debris impact caused by the rotor wash. Though recent efforts have looked at understanding the effects of debris impact [45], there currently are no known test standards to asses and characterize this form of damage.

5.3.7 DESIGN AND INTEGRATION CONSIDERATIONS FOR EROSION COATINGS There are numerous considerations that must be addressed to integrate an erosion coating onto a rotorcraft. The success or failure of a rotor blade erosion coating in many respects has little to do with its erosion performance but entirely to the integration of the coating on an aircraft. Many of these requirements are not even levied by materials engineers. 5.3.7.1

ADHESION

Any erosion coating must undergo adhesion testing to ensure it will not separate from the blade throughout the operational envelop. The adhesion of the coating to each substrate material (usually titanium or steel, nickel and composite) must be verified by standard lap shear and peel coupon tests. Some polyurethanes get extremely soft at wet-elevated temperature conditions, but the failure mode is still cohesive and is not a concern. Testing at temperature is a critical aspect in verifying coating performance since rotor blades can easily experience upper surface temperatures in excess of 938C when ambient temperatures are

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49–558C like in the deserts of Southwest Asia. In addition, an aerodynamic analysis of the loads on the coating must accompany the adhesion testing to provide some reasonable basis that the testing is simulating the operational environment. 5.3.7.2

SUBSTRATE COMPATIBILITY

In addition to the adhesion to the substrate, the compatibility of the coating with the blade substrate must be taken into consideration. The coverage area of a coating is typically confined to a small portion of the leading edge near the blade root, but extends across the entire blade at the tip. Polyurethanes generally do not have any issues with being applied to either metal leading edges or composite surfaces. However, coatings applied by High Velocity Oxygen Fuel (HVOF) or High Velocity Air Fuel (HVAF) have application temperatures in the thousands of degrees. This makes them applicable only to the metal leading edges, and in many cases, requires application before the leading edge is included in blade assembly. This can leave the leading edge susceptible to damage during subsequent production operations. For coatings applied by HVOF or HVAF, another coating must also be included to provide the necessary coverage of the composite blade fairings at the tip. 5.3.7.3

STRAIN COMPATIBILITY

As previously noted, rotor blades can be considered very flexible structures. As a result, erosion coatings must be compatible with the blade strains. Level flight can generate vibratory strains in excess of 1000 m1 [58], and spanwise and edgewise strains at around 3000 m1 [45]. This strain compatibility is a reason very brittle materials fall out of consideration as coating materials. As an example, chemical vapor deposited diamond is extremely resistant to sand erosion and results in little external profile impact at 25 mm thick, but is not compatible with more flexible substrates. However, polymers have no problems meeting the strain requirements of the blade, but require special consideration to meet erosion demands. 5.3.7.4

WEIGHT AND SPAN MOMENT

While most coatings are relatively thin, 0.508 to 1.524 mm, the large surface area results in a significant weight gain. For instance, in the case of the V-22, the erosion protection of just one blade with 0.508 mm of polyurethane can add an estimated 6.5 kg per blade [59]. Not only is this added weight a major concern, but how the weight is distributed along the blade needs to be addressed. Since higher velocities and greater erosion are found at the blade tips, more material is applied there. This increases the span moment on the blade, which may negatively affect the fatigue lives of both the blade and rotor hub. Just a single pound of extra weight out at the blade tip can increase the span moment enough to

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noticeably reduce the fatigue life of a blade that would normally experience millions of cycles. A rule of thumb is if the coating increases the span moment less than 1%, then the additional weight can be accommodated with simply changing the track and balance weights. Though the span moment concern is especially a problem with heavier metallic coatings, most applications of polyurethanes increase the span moment more than 1%, requiring a thorough fatigue analysis to assess the impact. Operation experience has shown that even after 15 months in Iraq, blades with a polyurethane coating still met their span moment requirements even with repairs since the additional applied coatings quickly eroded in the harsh environment. In fact, four blades came within 18 inch-pounds of their previous values [48]. One drawback from existing sprayable coatings is the variation experienced in the coating thickness and weight along the blade. As previously mentioned, blade track and balance is a critical parameter. With existing hand spray methods, the coating weight varies enough that a static balance to rebalance the blade is ineffective, instead requiring a dynamic balance on the whirl stand. This greatly increases the costs of coating application and may delay the production schedule since whirl towers are in high demand and can be choke points in production. Robotic spray application is being considered to ensure coating uniformity along the blade length. 5.3.7.5

AERODYNAMIC FACTORS

Integrating the coating into the blade design requires that the blade be requalified through full scale fatigue testing. Because of the time and expense of this, most erosion coatings are aftermarket applications which consequently can alter the aerodynamic profile of the blade. Depending on where the coating ends chordwise, the aerodynamic performance could be affected. Computational fluid dynamic models of the coated V-22 blade showed minimal effect of the coating on aerodynamic performance [59]. However, each blade needed to be analyzed since the blade contours could have varied and the chordwise location of the coating lip may have resulted in different aerodynamic parameters than initially expected. In similar cases, a high altitude hover flight test may be necessary in order to characterize the performance loss, especially in hover out-of-ground effect. An instrumented rotor blade and hub might be required to determine if the component loads increase, thus reducing the fatigue lives. To minimize the coating effects on aerodynamic performance, robotic spray application technologies that can feather the coating edge may mitigate detrimental aerodynamic performance issues. 5.3.7.6

REPAIR

As noted in Section 5.3.6, applying coating materials in a facility with the appropriate engineering controls and then providing a small repair kit for maintainers

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Fig. 5.57

On-aircraft repairs of polyurethane erosion coatings.

in the field has resulted in much greater consistency of the coating performance. However, this necessitates the need for a repair method that is compatible with both the rotor coating system and substrate. Fielded sprayable coatings require regular attention with the top coat usually being repaired every 40 hours and the basecoat every 120 [47, 48]. Top coat repairs can be accomplished on aircraft with a small aerosol spray gun as shown in Fig. 5.57, while basecoat repairs require removal of the blades and subsequent repair in a hangar. The operational tempo of rotorcraft and the vulnerability of the top coat require that repairs are necessary every couple of days, which may impose constraints on meeting mission objectives. Besides time-between-repair, the coating application method is another consideration for the maintainer in the field. Field repair of coatings applied by HVOF or HVAF require high temperature guns that can lead to unacceptable temperatures, possibly damaging the composite structure underneath. However, with proper inspection and maintenance, blades can last indefinitely without removal and reapplication of the coating at the depot or factory. 5.3.7.7

COATING REMOVAL

In the normal life of a blade, erosion coatings will require removal and reapplication. This is especially true for blades that have replaceable leading edges that contain deice heater blankets. The challenge is that existing sprayable polymeric coatings are designed for abrasion resistance and have good adhesion strength to the blade, hence making the coating removal problematic. Sprayable polyurethanes have been removed by a combination of plastic media blast (PMB), power orbital sanders, and hand sanding. The PMB can be used successfully on the metal leading edge while less aggressive methods are necessary to remove the coating from composites to avoid damaging the laminate. Removing an erosion resistant coating from a single blade can take several man-days of work

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to accomplish [48]. Figure 5.58 shows a blade after the polyurethane coating was successfully stripped. 5.3.7.8

ENVIRONMENTAL DESIGN CONSIDERATIONS—ICING, LIGHTNING, AND P-STATIC

Flying in weather can lead to icing conditions that detrimentally affect the aerodynamic performance of the rotor blades. To counteract this problem, de-icing systems are often employed. Any use of erosion coatings must ensure compatibility with the de-icing system. Existing blade tapes [43] require that the de-ice system be disconnected when flying using blade tapes. This can adversely affect operational flexibility. For the V-22 where the operational altitudes can lead to icing conditions down to 2208C, erosion polyurethane coatings with fillers that improve thermal conductivity experience an unintended reduction in erosion capability [59]. Research into newer ice phobic coatings to meet this requirement have been largely unsuccessful, showing little if any improvement in preventing ice accumulation while trading off erosion performance [58]. Rotor blades are a typical pathway for lightning strikes on rotorcraft where the main blades are the entry point, and tail rotor blades can be the exit. In addition, when considering the high speed rotation of the blades through air with snow, dust or rain, charging via precipitation-static (P-static) must be a design and operation consideration. Polyurethane coatings can add an insulative layer to the top of the metallic lightning strike material, resulting in additional damage to the rotorcraft if a lightning strike occurs. For the V-22 blade, the coating profile initially called for a uniform 0.508 mm thick coating, but subsequent studies demonstrated that a thickness of 0.152 mm was required to give acceptable

Fig. 5.58

Main rotor blade stripped of coating.

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lightning strike performance [59]. Understanding the charging environment and impact of the coating material and profile is essential to mitigating the detrimental effects of lightning and P-static effects. 5.3.7.9

ELECTROMAGNETIC INTERFERENCE (EMI)

One final consideration for erosion coatings is any interference with the radios in the rotorcraft. Radio signals from fuselage mounted antennas under the rotors can be bounced or absorbed by the coatings leading to undesirable radio performance. A qualitative electromagnetic compatibility check during initial flight tests can verify any interference.

5.4 ENABLING MATERIALS, STRUCTURES, AND MANUFACTURING PROCESSES Current design development is using past experiences coupled with aggressive weight and cost targets to develop future materials and structural configurations. Weight is a key performance parameter influencing the range, efficiency and payload capabilities of the aircraft. Both non-recurring and recurring costs directly influence the purchase price and profit potential for the manufacturer. In-service performance influences the operating costs as seen by the operator. Multiple business case scenarios are required to establish development and production plans. Design evaluations encompass the total life cycle costs including in-service maintenance and repair, which is dramatically influenced by the material selection and system design. Figure 5.59 depicts the relative material

Fig. 5.59

Relative structural cost and weight for material systems.

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performance for different material systems. New materials and processes are the key to achieving the target, enabling highly efficient products at a reasonable cost.

5.4.1 NANOTECHNOLOGY Nanotechnology is the creation of uniquely designed materials through the control on the nanometer-length scale and the exploitation of novel properties and phenomena developed at that scale. Nanotechnology is not just very small, but it exists in the scale where continuum mechanics and quantum mechanics meet. The consequences of these nano-scales are: .

Powerful effects for very small volume fractions: Nanoscale structures such as nanoparticles and monolayers have very high surface-to-volume ratios, making them ideal for use in polymeric materials.

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The promise of multi-functionality: By creating nanometer (nm) scale structures, it is possible to control fundamental properties of materials like melting temperature, magnetic properties, charge capacity, mechanical properties and even color, without changing the materials’ chemical composition.

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The promise of true engineered materials: Utilizing this potential will lead to new, high-performance products and technologies that were not previously possible, by circumventing traditional performance tradeoffs.

In Fig. 5.60, the concept of step increases in material performance and design flexibility is shown from isotropic metals, to anisotropic composites with complex resin morphology and tailored load carrying capability with coupled deformations, to the latest paradigm, nanocomposites in which mechanical performance can be linked to improvements in a wide array of electrical, thermal, optical and chemical behavior. In terms of mechanical performance, the possibility is there for eliminating the classical reciprocal relationships between toughness and strength and get improvements in both, as well as increased stiffness. But the challenges have been affordability, processability, and from the business side, building supplier value-chains to deliver technology to the airplane. There are dozens of types of nanoparticles/nanomaterials ranging from single walled nanotubes with aspect ratios in the hundreds, to multi-walled nanotubes, nano-gold particles, nano-oxides, nano-clays, nanosheets, and hybrids. The large majority of them capable of being dispersed into conventional composites creating the so-called nanocomposites. As mentioned by different authors, nanocomposites are the new generation of composites materials where nanoparticles dispersion into polymeric systems allowed formation of nanostructures inside the matrices changing the composites overall behavior. The first nanostructured composites (polyamide 6/organophilic montmorillonite clay) were synthesized by the Toyota Research Group in 1990’s [60]. When the net properties of polyamide 6 (PA6) were compared against the PA6/nanoclay, the advantages in terms of stiffness were evident. Generally,

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ing eas ls r c in teria rd wa f ma o t ty o nd tre pabil g ca uin ntin ering o Anisotropic C ine composite eng materials

Isotropic metal materials • Mechanical performance enhancements through alloying and heat treatments

Fig. 5.60

Nanomodified composite materials • Broad mechanical property improvements • Flammability and solvent resistance enhancements • Conductivity and CTE tailoring • Inherent color optical qualities, etc. • Multifunctionality

• Fiber orientation optimizing effects on strength and stiffness • Laminate tailoring for coupled deformations • Independent toughness improvement through polymer alloying and controlling phase morphology

The promise of true multi-functional engineered materials.

three methods for the formation of polymer/clay nanocomposites are used: 1) intercalation of monomers followed by in situ polymerization, 2) direct intercalation of polymer chains from solution, and 3) polymer melt intercalation [61]. Each of these methods has its advantages and disadvantages. The direct melt intercalation into clay layers is a preferred process for the formation of hybrid materials because this method is solvent free and it is compatible with conventional industrial processes, such as extrusion and injection molding and other polymer processing techniques [62]. The melt intercalation approach (also known as direct melt intercalation) involves annealing (statically) or melt-compounding (under shear flow) a mixture of the polymer and organically modified clay (organoclay) above the softening or melting point of the polymer [62]. Under these circumstances and if the layer surfaces are adequately compatible with the chosen polymer, the polymer chains can crawl into the interlayer galleries and form either an intercalated or an exfoliated nanocomposites according to the degree of penetration. Furthermore, the melt intercalation method allows the use of polymers which were previously not suitable for in situ polymerization or solution intercalation. Melt intercalation is the most efficient method for dispersing nanoparticles into thermoplastics, as has been successfully demonstrated by dispersing surface modified montmorillonite clays into polyamides (PA11 and PA66) [61, 62]. For thermoplastics, the molecular weight affects the exfoliation rate during the melt intercalation process. It has been shown that composites formed from high molecular weight (HMW) and mid molecular weight (MMW) PA6 were

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well-exfoliated, consisting largely of individually dispersed platelets; whereas, those based on the Low molecular weight (LMW) matrix revealed a mixed morphology, containing both agglomerates of multi-layered stacks and individually dispersed platelets [63]. The superior morphology of the MMW and HMW-based nanocomposites translated into better reinforcement, demonstrated by higher composite moduli and yield strengths. The good performance of PA-nanoclay nanocomposites led to the development of different classes of nanocomposites, including epoxy-based systems. Research has shown that nanoclays such as organically modified montmorillonite (Cloisite 30B) can be dispersed into epoxy systems without major efforts [64]. In many cases, the complete exfoliation or a fully disperse intercalated structures formation is virtually impossible, but still contributes to significant improvements in stiffness and strength. In one study [65], the amount of Cloisite 30B was varied by weight through shear mixing from 1% to 10% in an epoxy system, and an increase up to 80% was noted in the elastic moduli. Another study [66] demonstrated that both stiffness and toughness can be improved by using nanocomposites. For this case, the epoxy system (resin—diglycidyl ether of bisphenol A and cure agent—triethylenetetramine) demonstrated a maximum impact strength when 1% by weight of montmorillonite was added through direct mixing. Several researchers have demonstrated that an octadeyl amine modified montmorillonite (e.g. nanomer I30E from Nanocor) is more suitable for dispersion in epoxy systems and can lead to better results [67, 68]. Another inorganic layered silicate in its nano-dimension form (i.e. fluorohectorite) has demonstrated significant improvements to epoxy systems. Like montmorillonite, the fluorohectorite has tremendous ability to exchange ions, favoring the penetration of the polymer or polymer precursors between these layers and consequently the formation of an exfoliated nanocomposite. In one study [69], it was reported that a 54% increase of the Young modulus was observed when only 10 wt% of layered silicate was added to the epoxy matrix. However, the tensile strength was reduced by 36% and the elongation at break was also affected. When the nanomodified epoxy matrix was added to glass fibers, the results were quite different. A remarkable 27% increase in flexure strength was observed while the Young’s modulus enhanced only by 6%. Two hypotheses can explain this increase on flexural strength. First, the improvement on the compressive strength of the epoxy matrix is due to the presence of the layered silicate, increasing the bending strength of the corresponding fiber reinforced composite. The second is the presence of the silicate layers at the surface of the glass fibers, which may improve the interfacial properties between the matrix and the fibers. There are four commonly accepted advantages offered by nanocomposites: 1) improves fire retardancy, 2) increases heat distortion temperatures, 3) improves flexural modulus, and 4) promotes a decrease in permeability [70]. Improvements in flammability have been observed with reductions in peak heat release rate (PHRR) of 40 to 65% at very low loadings of layered silicate in select polymers.

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Demonstrations using different types of nanoclays (Cloisite 15A, 20A, 30B, 93A) have shown that by varying the nanoclay content up to 10 wt% and employing a mass loss calorimeter for flammability tests, a reduction in the peak heat release rate is possible over the baseline polymer with an increase of nanoclay content. In particular, 10 wt% of Cloisite 30B added to a polyamide resulted in a 73% reduction on PHRR and a reduction on mean CO yield when exposed to a 50 KW/m2 heat flux for 180 seconds. The addition of carbon nanotubes will yield similar results, with demonstrations showing that a 5 wt% addition can result in a 60% reduction of PHRR. The use of nanosilicates up to 10% has been shown to yield no effect on polymer flammability [60]. Graphite nanosheets offer another approach to improving fire retardancy, material stiffness and strength [71]. Graphite nanosheets (GNSs) are formed from expandable graphite when submitted to a thermal gradient of approximately 10008C/sec. The large thermal gradient causes the rapid loss of the polymeric matrix placed between graphite layers, resulting in a large expansion of the graphite layers. The graphite layers thickness approaches the nanoscale as the volume is held constant. To make the nanographite layers capable of establishing covalent bonds with polymeric systems, nanographite layers (also called nanosheets) must be submitted to a chemical oxidation in the presence of concentrated H2SO4 and HNO3 acid [72]. The large elastic modulus of approximately 1.0 TPa, makes GNSs reinforcement a viable option for improving the stiffness and strength of composites [73] while providing additional multifunctional capability. Graphene and graphene-derived materials also have been evaluated for their ability to toughen epoxy resins, demonstrating significant improvements with as little as 0.1% graphene by weight added to the polymer [74]. Several researchers have studied the effects on electrical conductivity due to the addition of nanosheets to polymeric matrices [75–77]. GNSs tends to align in the polymer matrix under shear-force, resulting in a polymer-based nanocomposite with unique piezo-resistivity [78]. For instance, natural flake graphite is a layered filler with a carbon axis lattice constant of 0.66 nm. The crystalline lattice of graphite consists of graphene layers formed by sp2 hybridized carbon atoms bounded together by weak van der Waals forces. This configuration translates to an electrical conductivity in the range of 104 Sm21 at room temperature. Carbon nanotubes (CNTs) are promising reinforcement options for composite materials [79]. Possessing a unique honeycomb lattice (a graphene sheet) rolled into a cylindrical configuration, single-walled CNTs have been the center of much research over the years due to their intriguing dimensional scale and their remarkable electro-mechanical properties. In general, a carbon nanotube has a diameter in the nanometer range with a length greater than 1 mm, offering a dimensional scale much smaller than most advanced semiconductors devices developed so date [80]. Moreover, single-walled nanotubes (SWNTs) have a predicted specific strength of approximately 600 times larger than steel [81]. Other important benefits of CNTs are their remarkable thermal and electrical properties.

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CNTs are thermally stable up to 28008C (in vacuum), demonstrate a thermal conductivity about twice as high as diamond, and may exhibit a capacity to carry electric current a thousand times better than copper wires. With such promising properties, CNTs are natural candidates for reinforcement of advanced composites. However, their cost is still a limiting factor, with one single gram of SWNT of 99.99% purity costing approximate $500 USD (in 2015) [60]. Vapor grown carbon nanofibers (VGNF) are significantly larger than CNTs, but offer a more affordable reinforcement solution while still providing improved mechanical, electrical and thermal performances. Nanotechnology can be applied also to the reinforcing fiber of a composite laminate. For many decades, advanced fiber-reinforced composites have been used as viable primary load-bearing structures. Although the in-plane loading and stresses have been handled by various configurations of fiber architectures, such as 1D (i.e., unidirectional tapes) and 2D (i.e., woven fabrics), the intralaminar and interlaminar stresses have remained major issues, resulting in relatively weak interlaminar fracture toughness and fatigue resistance. Toughening the matrix can resolve some of the issues related to delamination, but to further impede interlaminar crack initiation and growth, a nanofibrous nanostructure in between the layers can provide additional benefits. This has been demonstrated using unique bottom-up multiscale hierarchical manufacturing of carbonnanotube forests grown on fibers present in adjacent plies to give 3D multifunctional nanocomposites [82, 83]. Multi-walled carbon nanotubes (MWCNTs) were grown uniformly on all of the exposed fibers on the surface of the fiber cloth. The nanotube-grown fabrics were infiltrated with a high-temperature epoxy, then stacked to yield a nanoforest ‘sandwich’ structure, and the laminated structure was cured in an autoclave. As a result, Mode I and Mode II fracture toughness improved by about 300% and 50%, respectively. Flexural modulus, strength, and toughness improved by about 5%, 100%, and 300%, respectively. Structural Damping improved by 400–500%. Through-the-thickness thermal and electrical conductivities increased by 50% and 50,000%, respectively, and through-thethickness CTE decreased by about 60% making the developed 3D hierarchical nanocomposite a true multifunctional composite. The much-improved through-the-thickness electrical conductivity observed in the 3D nanocomposites would potentially impart to these structures an electrical sensing capability for structural health monitoring during delaminations (i.e., crack initiations and/or propagations). The vertical arrays of nanoforest nanotubes in the thickness direction of the composite improve the mechanical properties without compromising the in-plane properties, and alleviate the problem of agglomeration when nanotubes are randomly introduced into the matrix to produce nanocomposites. Such 3D nanocomposites could pave the way for the application of CNTs in structural composite. The future looks bright for nanotechnology, and ongoing research is continually discovering new applications for this once little understood material. Beside the structural, electrical and fire retardancy benefits already mentioned, nanoscale

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technologies are envisioned to support advancements in aerospace applications such as: .

Acoustically enhanced composite materials

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Enhanced adhesives and bonded repair methods

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Hydrophobic/ice phobic coatings

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Transparent nanocomposite materials

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Multi-functional structure-power combinations yielding high power densities

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Nanoscale power sources and embedded conductors

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Nanoscale sensors and electro-active coatings for integrated system health monitoring

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Thermal management heat pipes, radiators and thermal planes

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Lightweight non-permeable hydrogen storage systems

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Lightweight air purification filters

There are also risks and concerns associated with the maturation and transition of nano technology. From a processing perspective, there are issues regarding uniform dispersion of nano particles in resin systems due to their natural tendency to aggregate. In addition, nano-additives can increase viscosity and decrease flow, thereby inhibiting resin infusion processing methods. Cost is another major issue that needs to be addressed. The high price of additives can be as high as $250,000 USD/lb for some of the most promising materials. This not only inhibits transition to commercial application, but stifles the fundamental research needed to mature these materials. From a health and environmental perspective, the effects and risks of nano materials have not been fully characterized and will need to be clearly defined before widespread application can be expected. This assessment will need to be made based on the specific application and intended operational use. Once the viability and the cost benefits become clear for nano materials, scaling production to meet demand will become a primary challenge. Since nano scale materials come in many different forms, the nanotech suppliers will need to work closely with their many customers to meet their requirements while balancing the economics of operating possibly several different production lines. For the industrial customer the hope is that application of nano materials to new products can fit well into their current manufacturing methods, requiring minimal retooling. Regardless of the application, the first step is to get users and the public familiar with the benefits and limitations of nanotechnology. Education is the key to moving forward with this unique class of materials.

5.4.2 ADDITIVE MANUFACTURING Additive manufacturing (AM), also known as 3D printing, has been used for prototyping for over twenty-five years and is now emerging as a mainstream

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manufacturing technology for direct part production. The benefits of AM include shorter lead times, mass customization, and the ability to produce parts with complex geometries that are unable to be manufactured using traditional methods. The term “additive manufacturing” describes a suite of processes that “join materials to make objects from 3D model data, usually layer upon layer, as opposed to subtractive manufacturing methodologies, such as traditional machining” (ISO/ASTM 52900:2015). AM is not considered to be a new technology, as foundational patents were filed for stereolithography in 1986 and for material extrusion in 1989. However, AM exhibited a “boom” in the mid to late 2000s when key patents expired for these processes resulting in a deluge of desktop 3D printers and also a growing interest in metallic 3D printed structures. The major process types, as defined by ISO/ASTM 52900:2015, materials and markets are in Fig. 5.61 [84]. Materials for polymeric processes may be UV-cured photopolymers, thermoplastic powders for powder bed fusion, or thermoplastic filament or pellets for material extrusion. Some of the higher performance thermoplastic materials include those in the polyamide (nylon), polyaryletherketone (PAEK) and polyetherimide families. Materials for metallic processes span a wider selection such as nickel alloys 625 and 718, CoCr, Ti-6Al-4V, AlSi10Mg, stainless steel 316L and many others. Each process has advantages and disadvantages and the suitability of a process is dependent on the requirements of the application. AM is not a “push button” technology and considerations of the full manufacturing process such as design, material selection, processing, post treatment/processing, and testing/evaluation are very relevant. Additive manufactured parts are very process-sensitive, similar to welded materials or composite materials, in that the resulting part properties

Fig. 5.61

Additive manufacturing ASTM process types.

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are highly dependent on the processing conditions. For final part production of aerospace parts, the most promising process types are material extrusion, powder bed fusion, and directed energy deposition. Material Extrusion systems may range from very large robotic gantry systems such as the “Big Area Additive Manufacturing (BAAM)” system, industrial-grade systems within thermally stable environments such as Fused Deposition Modeling from Stratasys, and small desktop (consumer-grade) equipment of which there are many options available. Aerospace applications include direct part fabrication for nonstructural applications such as environmental control ducting. Figure 5.62 shows an additively manufactured composite processing tool for a carbon fiber epoxy compact inlet duct, able to withstand a 1778C cure. The tool was built in eight days using material extrusion equipment and a polyetherimide feedstock (ULTEM 9085), at a fraction of the cost of a metallic tool. After the preimpregnated fiber was laid up on the tool and cured, the tool was dissolved quickly using a small amount of solvent. Polyetherimide tools are being used for sheet metal forming and other manufacturing processes. AM tooling, fixtures, jigs, and surface treatment masks are just a few of the various manufacturing aides under development to support maintenance and repair operations at U.S. Air Force and other military depots. Similar structures as shown in the Fig. 5.62 are being used for aircraft environmental control ducting greatly reducing weight and assembly costs. Thin-walled ducts are particularly advantageous for weight reduction and are used throughout many commercial aircraft platforms. Powder Bed Fusion is a process where thermal energy (typically from a laser or electron beam) selectively fuses powder feedstock within a bed. The powder may be metallic or polymeric and this process may produce some of the most geometrically complex parts. A well-known aerospace application is the Leading Edge Aviation Propulsion (LEAP) engine fuel nozzle by General Electric, which received Federal Aviation Administration (FAA) approval in 2015. Previously

Fig. 5.62

Additively manufactured composite processing tool.

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made up of 19 parts, the fuel nozzle was redesigned and manufactured as one part using AM, decreasing weight by 25% and increasing the durability of the nozzle by 5X. Due to the redesign, the entire LEAP engine is projected to be 15% more efficient [85]. Directed energy deposition involves applying thermal energy selectively to fuse materials as they are being deposited onto a substrate. In this process, the thickness of each layer and the area over which the energy is applied is typically much greater than that in powder bed fusion, so a rougher surface with visible layer lines is produced. Typically directed energy deposition parts need significant post processing and many highly complex geometries cannot be produced with this method. The thermal energy may be laser, electron beam or plasma, with the process either blowing powder feedstock material or introducing a wire feedstock material into a thermal energy zone. Additive manufacturing allows for 1) the design and production of parts not able to be made through other methods, 2) more efficient design of structures, 3) reduced manufacturing times through the reduction of manufacturing steps including assembly, 4) reduced material waste through improvement of “buy-to-fly” ratio via net or near net shape parts, and 5) reduced machining costs for hard-to-machine materials such as titanium and Inconel. This process is ideal for applications that require advanced geometric complexity and/or customization. Geometric complexity is far more achievable using additive manufacturing, with multiple design iterations being accomplished much more quickly before arriving at a final part design. In addition, AM enables more efficient use of resources, fewer processing steps, net-shape production, less assembly (unitization of structures/parts), less post processing, and less wasted material and energy. It also enables the reverse engineering for out-of-production parts, a pressing need for military legacy aircraft. Another unique aspect of AM is that it requires a relatively small footprint, allowing for local or remote fabrication of parts. This last point is a huge advantage in situations where supply chain logistics are challenging such as on oil rigs, a military forward deployed area, or even the International Space Station. As promising as AM is to the future of manufacturing, it is not without challenges and is not suitable for all applications. Some of the disadvantages inherent to the processes are issues with dimensional accuracy and surface finish due to the layering process. Many additive manufactured parts are somewhat anisotropic, with diminished properties perpendicular to the layering direction (Z-direction). This reduction in properties may be as much as 30–50% depending on the AM method and must be considered when designing a part. Variations in the thermal environment during a build may result in residual stresses and distortion. Directed energy deposition processes may impart significant residual stresses. This distortion can be mitigated somewhat with build path planning and the use of supports to thermally and mechanically stabilize parts. Aerospace structural parts, particularly those that are safety-critical or fracture-critical, have high certification barriers. Five critical factors governing the design and use of

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aerospace structures are 1) Stabilized material and/or material processes, 2) Producibility, 3) Characterized mechanical properties, 4) Predictability of structural performance, and 5) Demonstrated supportability [86]. Considering these factors for AM, concerns arise with demonstrating a stabilized process, as many processes exhibit significant build-to-build and machine-to-machine variability. AM processes have tens to hundreds of controllable process parameters that interact with complex part geometries, build layouts, and scan strategies to produce unique and varying local processing states. Even with identical process parameters, build path variations, part type and orientation may result in variation in mechanical properties and the presence of anomalies such as porosity or unfused feedstock material. Drastic differences in dimensional accuracy and surface finish can be seen with changes as simple as the orientation of a part in the build. In order to qualify AM materials and processes and certify AM parts for aerospace applications, the processes must be controlled, with inherent variability both minimized and predictable. Process variability may be reduced through process sensing and controls. Once variability is reduced, designers will need statistically-relevant mechanical property data. In addition, nondestructive evaluation techniques need to be established for quality assurance. Greater understanding is needed also on the effect of powder feedstock properties such as particle size distribution, powder spreadability, and machine decontamination after switching feedstocks. The effects of post processing (e.g. hot isostatic pressing, machining) on the microstructure and material properties need to be understood and communicated to the designer early in the design phase to accommodate any potential knock-downs in mechanical or thermal performance. Additive manufacturing is a promising suite of technologies for many aerospace applications with the promise to significantly unitize and lightweight structures. Collaboration through regional clusters, industry consortia and publicprivate partnerships have been beneficial for the development of additive manufacturing. To realize the full benefits of AM, process stability and the understanding of relationship between process conditions and performance is required. This knowledge will be realized through advanced process control, nondestructive evaluation methodologies, and standards development. The development of standards for AM is extremely important to facilitate widespread adoption, especially for high value industries such as aerospace. In 2016, America Makes and the American National Standards Institute (ANSI) launched the AM Standardization Collaborative to coordinate and accelerate the development of industry-wide AM standards and specifications, with the purpose of facilitating the growth of the AM industry. The major standards development organizations involved continue to work with industry to address these gaps including collaborating on the R&D needed to feed standard development activities. The Federal Aviation Administration remains highly involved with developing policy and guidance towards the use of AM within new aircraft as well as for maintenance, repair and overhaul applications.

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5.5 CLOSING COMMENTS Material development and selection is of paramount importance in designing a new aircraft. Safety, affordability, and performance, are all affected by the choice of materials. Over the years, the commercial and military aerospace industry have evolved an integrated approach taking into account the design requirements, critical failure modes, controlling material properties and manufacturing processes. This is a process that requires extensive and detailed trade studies looking at all these critical factors as well as what materials can be developed in a given time frame. The material development and selection for the Boeing 787 and 777X are the culmination of the development and refinement of this approach over earlier commercial airplane programs. Future programs will benefit from breakthroughs that are anticipated in all areas of structural materials.

ACKNOWLEDGEMENTS The authors and editor would like to recognize and thank the many people that supported the writing of this chapter. From The Boeing Company, Dr Moe Soleiman provided valuable assistance in reviewing the chapter and contributing the section on composite materials, Dr Rod Boyer for contributing the titanium section, Dr James Cotton for contributing the aluminum section, Mr Paul Dufour for review of the additive technology section, and William Grace was a valuable resource for information on the titanium-graphite fiber metal laminate material called TiGr. From the Air Force Research Laboratory (AFRL), Dr John Russell, Mr Barth Shenk and Mr Richard Holzwarth contributed significantly to the summary of ACCA and CAI, Dr Jennifer Fielding prepared the overview on additive manufacturing, and Dr Brett Bolan provided valuable insight to the section on aircraft repair. In support of the rotorcraft sections, the authors and editor would like to acknowledge the contributions of Ms Lynne Pfledderer from AFRL, Andrew Phelps from the University of Dayton Research Institute, Marc Pepi and Rich Squillacioti from the Army Research Lab, Shek Hong from Hontek, and Lou Centolanza from the Army Aviation Applied Technology Directorate. The authors would also like to thank Professor Mehrdad Ghasemi Nejad of the University of Hawaii for reviewing and contributing to the Nanotechnology section.

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CHAPTER 6

Materials for Spacecraft Miria M. Finckenor NASA Marshall Space Flight Center, Huntsville, Alabama

6.1 INTRODUCTION The general knowledge in this chapter is intended for a broad variety of spacecraft: manned or unmanned, low Earth to geosynchronous orbit, cislunar, lunar, planetary, or deep space exploration. Materials for launch vehicles are covered in Chapter 7. Materials used in the fabrication of spacecraft hardware should be selected by considering the operational requirements for the particular application and the design engineering properties of the candidate materials. The information provided in this chapter is not intended to replace an in-depth materials study but rather to make the spacecraft designer aware of the challenges for various types of materials and some lessons learned from more than 50 years of spaceflight. This chapter discusses the damaging effects of the space environment on various materials, materials that have been successfully used in the past, and materials that may be used for a more robust design. The material categories covered are structural materials, thermal control for on-orbit and reentry, shields against radiation and meteoroid/space debris impact, optics, solar arrays, lubricants, seals, and adhesives. Spacecraft components not directly exposed to space must still meet certain requirements, particularly for manned spacecraft where toxicity and flammability are concerns. Requirements such as fracture control and contamination control are examined, with additional suggestions for manufacturability. It is important to remember that the actual hardware must be tested to understand the real, “as-built” performance, as it could vary from the design intent. Early material trades can overestimate benefits and underestimate costs. An example of this was using graphite/epoxy composite in the International Space Station (ISS) science racks to save weight. By the time the requirements for vibration isolation, space shuttle frequencies, and experiment operations were included, the weight savings had evaporated [1].  Materials

Engineer, Materials and Processes Laboratory, Environmental Effects Group, Associate Fellow, AIAA.

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

403

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M. M. FINCKENOR

6.2 SPACE ENVIRONMENT Hardware exposed to space must withstand all aspects of the space environment. This includes vacuum, thermal cycling, charged particle radiation, ultraviolet radiation, and, in some environments, plasma effects and atomic oxygen. Micrometeoroids and space debris particles may impact at high velocities; shielding is discussed in Sec. 6.9. All of these may have significant effects on material properties either alone or in synergism. The hard vacuum of space, with pressures below 1024 Pa (1026 torr), causes some materials to outgas, which in turn affects any spacecraft component with a line of sight (LOS) to the emitting material. This is discussed in more detail in Sec. 6.4. Thermal cycling occurs as the spacecraft moves through sunlight and shadow while in orbit or conducts a “rotisserie” maneuver to keep solar exposure even. Thermal cycling temperatures are dependent on the spacecraft component thermo-optical properties (i.e., solar absorptance as, or how much solar energy the material absorbs, and infrared emittance 1IR, or how much thermal energy can be emitted to space). The lower the ratio of a to 1, the cooler the temperature of the spacecraft surface. Thermal cycling can cause cracking, crazing, delamination, and other mechanical problems, particularly in assemblies where there is mismatch in the coefficient of thermal expansion. Charged particle radiation includes protons and electrons with a wide range of energies. Spacecraft operating in or outside the Van Allen belts are exposed to much greater radiation levels than those in low Earth orbit. Charged particle radiation along with ultraviolet radiation can cause cross-linking (hardening) and chain scission (weakening) of polymers, darkening and color center formation in windows and optics, and single-event upsets (SEUs) in electronics. Radiation shielding is discussed in more detail in Sec. 6.8. Plasma refers to the ionized molecules in the upper atmosphere that have been excited by interaction with ultraviolet radiation and are affected by the Earth’s magnetic field lines. The plasma environment varies with altitude, latitude, time of day, and solar activity. Interaction with plasma and charged particles in the space environment contributes to the buildup of surface charge, especially in higher-voltage systems. This surface charge can damage electronics, produce SEUs, trigger arcs in solar arrays or power systems, and cause dielectric breakdown of structure of surface coatings. NASA-HDBK-4006, Low Earth Orbit Spacecraft Charging Design Handbook, covers plasma interactions and mitigation techniques for high-voltage space power systems (.55 V). Atomic oxygen (AO) is produced when ultraviolet radiation reacts with molecular oxygen in the upper atmosphere. Currently only found in low Earth orbit between 100 and 1000 km altitudes, AO oxidizes metals, especially silver and osmium. AO reacts strongly with any material containing carbon, nitrogen, sulfur, and hydrogen bonds of 5 eV bond energy or less, meaning that most polymers react and erode away. Polymers containing fluorine, such as Teflon, react

MATERIALS FOR SPACECRAFT

405

synergistically, where the reactivity to AO increases with longer exposure to ultraviolet radiation [2]. Some materials, such as ceramic coatings, can be bleached by exposure to AO.

6.3 SPACECRAFT DESIGN AND MATERIALS REQUIREMENTS An old joke about the aerospace field is that when your paperwork weighs as much as your rocket, you are ready to launch. Good recordkeeping is vital to mission success. Not only a parts list, but also complete part identification and traceability (test reports, inspection records, application, and location of an individual part or material) are essential. Materials used in critical applications, such as life-limited materials, safety- and fracture-critical parts, liquid oxygen/gaseous oxygen (LOX/ GOX) batch-sensitive materials, or materials requiring treatment to prevent hydrogen embrittlement should be traceable by lot through all critical processing steps and the end-item application. Vendors should understand the requirements for certifying fracture-critical hardware. In one notable case, a NASA contractor was convicted and fined heavily for improperly heat-treating, aging, and falsifying quality testing on aerospace hardware that was used in the space shuttle and space station programs, as well as commercial and military aircraft and missile programs over a period of 16 years [3]. Fracture control requires that all space vehicle parts must be assessed to determine if structural failure in a space vehicle system would result in catastrophic failure. If the assessment determines that failure of the part would result a catastrophic failure, then that part must be subjected to full fracture control, including nondestructive evaluation (NDE). NDE methods for flaw or crack detection include eddy current, fluorescent penetrant, magnetic particle, radiography, and ultrasonics. Components that are exempt from fracture control are those that are clearly nonstructural and not susceptible to failure as a result of crack propagation (e.g., insulation blankets, electrical wire bundles, and elastomeric seals). Designs should be developed using concurrent engineering practices, including manufacturability. Manufacturability is extremely important in spacecraft component designs to provide the best benefit to cost and schedule. Manufacturability factors to take into account are listed in Table 6.1. As much manufacturing development as possible must be performed at full scale, including demonstration articles and mock-ups. When materials and processes are scaled-up from small laboratory specimens to full-scale spacecraft components, a multitude of unforeseen problems can arise. Material properties can vary, machining and forming issues can arise, and, when welding, the increased heat sink of large components can cause weld property variations that often require modifications to tooling and/or weld process parameters [4]. Shelf life of components must be considered during manufacturing. Organicbased materials have a limited shelf life and limited static age life (i.e., time in a

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M. M. FINCKENOR

TABLE 6.1

FACTORS TO CONSIDER FOR MANUFACTURABILITY

Factor

Consideration

Drawings

Use geometric dimensioning and tolerancing. Avoid double dimensioning. Choose dimensions close to standard stock. Use common angles if possible, 458 as opposed to 408. Limit decimal places to the specificity required. Create a separate drawing for finishing if complex masking or several processes are used on one part.

Tolerances

Use realistic tolerance levels. Be aware of tolerance stackup. Consider inspection and tool access to areas for verification.

Tapped holes

Only design holes tapped 1.5 (or less) the diameter. For blind holes, consider thread relief or do not tap to the bottom of the hole to prevent burr buildup.

Internal radii

Specify largest radii possible. Use same radii when practical.

Edges/thickness

Reduce sharp corners or points, which are breakable. Avoid thin webs and walls and deep holes to minimize distortion.

Part holding

Provide extra stock on all sides for clamping or chucking the workpiece.

Assembly

Design for disassembly. Provide clearances for wrenches. Include access holes where necessary.

Materials

Select materials that are readily available to manufacturing. Be aware that some materials are not available domestically, and their certifications can be difficult to get or unreliable. Select materials with shortest processing time, machining, heat treating, etc. Select materials with the simplest storage requirements.

Processes

Select processes that have been verified and are available to manufacturing.

Composite materials

Ensure selection of material system for which manufacturing has experience and verified processes. (Continued )

MATERIALS FOR SPACECRAFT

TABLE 6.1

407

FACTORS TO CONSIDER FOR MANUFACTURABILITY (Continued )

Factor

Consideration

Surface finishes

Specify minimum finishes.

Coatings

Use processes available and established in manufacturing. Involve coating specialist, manufacturing, and engineering before deciding on the best practice. Consider effects of coating processes on part dimensions, optical properties, etc. Consider problems with coating holes, blind holes, and difficult masking requirements. Use coating-specific drawings for complicated masking.

Heat treat

Consider using precipitation hardening alloys such as 17-4PH, 15-5MO, or 12-8MO, which only require a soak at the relatively low temperature of 480–6208C (900–11508F) from 1–4 h with an air cool, in place of the common alloys like 4340 or 4130 steels, which require an austenitizing soak at 815–8438C (1500–15508F) with a quick quench into oil followed by a tempering soak at 480–6008C (900–11008F). There is significant oxidation and scaling with the latter type of heat treat. On aluminum weldments, increase weld size or add gusseting rather than use heat treat to get the weldment back to a T6 condition, particularly with close tolerances or surface finish. This requires a solution treat at near melting temperature and a quick quench in water or other approved quenchant, followed by aging. Warping of the part and rupturing of enclosed areas can occur.

Welding

Use American Welding Society Standard Welding Procedure Specifications where possible. Keep weld length at a minimum. Select a joint with least amount of filler. Do not overweld. Use square vs round tubing for structural applications. Design for inspectability and accessibility. Allow for shrinkage and distortion. When using structural I and H beams, be aware of the irregular dimensions of the mill-supplied beams and call out your tolerances accordingly. It is hard to meet a þ/2 .030-in. tolerance when the beams may differ .250 in. from center line out to end of flange. (Continued )

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M. M. FINCKENOR

TABLE 6.1

FACTORS TO CONSIDER FOR MANUFACTURABILITY (Continued )

Factor

Consideration

Painting

Ensure documented, verified processes are available. Provide proper level of surface cleanliness. Consider the ability to get a paint gun perpendicular to the area being painted.

Shop capability

Size and weight of component Forklift/crane limits Documented/verified processes Welding processes Sheet metal capability Surface treatment capability Heat treat sizes Cleaning/painting size limitations

Electrical/electronic components

Consider manufacturing capability and lead time requirements.

Packaging and storage requirements

Size of component. Environmental controls required. Available space and equipment to meet storage requirements. Packaging material requirements, particularly for electrostatic-sensitive parts.

nonoperating mode in an ambient environment). The properties of polymeric resins, catalysts, some lubricants, thin polymer films, sealants, adhesives, elastomers, and other materials may change slowly with time, even if sealed. Most manufacturers specify the shelf life, but often storage conditions are not specified. Generally, lower storage temperature extends the shelf life, as does minimizing exposure to light (both sunlight and fluorescent lighting). Another factor that influences the shelf life is the length of time that the product has been exposed to air and the constituents in it, such as oxygen, moisture, and other active agents.

6.4 FLAMMABILITY, TOXICITY, AND OFFGASSING CONSIDERATIONS All spacecraft and ground support equipment materials must meet the criteria of NASA-STD-6001 (formerly NHB 8060.1), Flammability, Offgassing, and Compatibility Requirements and Test Procedures, for flammability, odor, offgassing, and fluid compatibility, depending on the environment to which the materials are exposed. Applicable environments are habitable environments,

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409

liquid oxygen and gaseous oxygen systems, breathing gases, and reactive fluids. Table 6.2 is from NASA-STD-6001 and lists each test required depending on the environment for use. NASA maintains the Materials and Processes Technical Information System (MAPTIS) online database of materials test results. NASA-STD-6001 Test 1 is upward flame propagation, or basic flammability. The control of flammability hazards on space hardware is extremely important because of crew safety. Two major lessons learned from the Apollo 1 fire investigations in 1967 were that ignition sources can never be eliminated and that propagation paths must be eliminated. If major propagation paths are eliminated, then, should a fire develop, it would be small and localized and would self-extinguish with little harm to the crew or vehicle systems. Limiting the total amount of combustible material is also important. Guidelines to control flammability can be found in NSTS 22648, Flammability Configuration Analysis for Spacecraft Applications, and MSFC-PROC-1301, Guidelines for the Implementation of Required Materials Control Procedures. Material behavior in oxygen-rich environments is discussed in Chapter 2.9 of this book. NASA-STD-6001 Test 4 is electrical wire insulation flammability. An electrical fire during STS-61-A Spacelab D-1 mission in 1985 led to a better awareness of insulation fraying and abrading. During this incident, when the insulation was worn through, the wire short-circuited and caught fire but the breaker did not trip. Electrical breakers need to be readily accessible to cut the power to equipment on fire. The lessons learned from the Russian oxygen generator fire onboard the Mir space station in 1997 led to the addition of a containment shield and stricter quality control for oxygen canisters on the ISS [5]. NASA-STD-6001 Tests 7 and 12 determine offgassing. Offgassing or toxicity testing of materials to minimize trace contaminant gases ensures air quality in habitable areas. Nonmetallic materials including coatings, adhesives, and potting compounds may contain irritants such as formaldehyde, n-butanol, and aliphatic aldehydes. Although activated carbon filters are effective at removing some of these contaminant gases, they should not be relied on for long-duration manned missions. Bakeout of offgassing material is recommended to drive off the volatiles, usually for 48 h at 508C (1228F). Outgassing still refers to the volatiles released from a material but is a separate test from offgassing for materials used in a vacuum environment. All external spacecraft materials must be tested at a minimum, ASTM-E-595, Standard Test Method for Total Mass Loss and Collected Volatile Condensable Materials from Outgassing in a Vacuum Environment, with not more than 1.0% total mass loss (TML) and 0.1% collected volatile condensable material (CVCM) recommended. Depending on the spacecraft contamination control plan, ASTM-E-1559, Standard Test Method for Contamination Outgassing Characteristics of Spacecraft Materials, testing may be required. Rather than just TML and CVCM, ASTM-E-1559 provides outgassing rates over time and at different temperatures for better dynamic modeling of any contaminants.

410

TABLE 6.2

M. M. FINCKENOR

NASA-STD-6001 REQUIRED AND SUPPLEMENTAL TESTS FOR EACH MATERIAL USE

Environment

Test Number

Type1

Title

Habitable flight compartments

1

R

Upward flame propagation

Other areas5

LOX and GOX environments

2

2

R /S

Heat and visible smoke release rates

3

S

Flash point of liquids

4

R

Electrical wire insulation flammability

6

R3

Odor assessment

7

3

R

Determination of offgassed products

8

S

Flammability test for materials in vented or sealed containers

10

S

Simulated panel or major assembly flammability

12

S

Total spacecraft offgassing

18

R

Arc-tracking

1

R

Upward flame propagation

2

2

R /S

Heat and visible smoke release rates

3

S

Flash point of liquids

4

R

Electrical wire insulation flammability

8

S

Flammability test for materials in vented or sealed containers

18

R

6

Arc-tracking

3

Odor assessment

3

R

7

R

Determination of offgassed products

13A

R

Mechanical impact for materials in ambient pressure LOX

13B

R

Mechanical impact for materials in variable pressure LOX and GOX

14

S

Pressurized gaseous oxygen pneumatic impact for nonmetals

17

R4

Upward Flammability of Materials in GOX (Continued )

MATERIALS FOR SPACECRAFT

TABLE 6.2

411

NASA-STD-6001 REQUIRED AND SUPPLEMENTAL TESTS FOR EACH MATERIAL USE (Continued )

Environment

Test Number

Type1

Title

Breathing gases

1

R

Upward flame propagation

6

R

Odor assessment

7

R

Determination of offgassed products

13A

R

Mechanical impact for materials in ambient pressure LOX

13B

R

Mechanical impact for materials in variable pressure LOX and GOX

15

R

Reactivity of materials in aerospace fluids

Reactive fluids 1

R – Required test; S – Supplement test. Required test only for surface areas greater than 4 ft2 (0.37 m2 per each use). Not required for materials inside hermetically sealed containers (see Section 2.1.3). 4 Not required for materials if they pass the criteria of Test 1 in that environment. 5 Includes all areas outside the habitable flight compartment. 2 3

Materials in LOS to sensitive optics may be tested per MSFC-SPEC-1443, Outgassing Test for Non-Metallic Materials Associated with Sensitive Optical Surfaces in a Space Environment. This test was developed during the Hubble Space Telescope program to ensure that materials that passed ASTM-E-595 did not evolve enough volatiles to impact optical performance in the ultraviolet wavelengths. Reflectance measurements are made on a magnesium fluoride/aluminum mirror before and after exposure in vacuum to the candidate material, with a change of more than 3% being cause for rejection of the material for the proposed application. Outgassing, or molecular contamination, can severely impact spacecraft performance. For example, the navigation camera on the Stardust spacecraft was blurred by contaminant deposition and required several heating cycles to improve image quality [6]. The heating cycles did not remove all of the contamination, and so there was limited resolution in the images Stardust took of the Wild 2 comet nucleus during flyby. One of the lessons learned from the Shuttle–Mir Program that benefited the ISS was the impact of contamination on radiator coatings and solar arrays. Samples of Z-93 white ceramic thermal control coating flown on the Passive Optical Sample Assembly (POSA)–I experiment were degraded to nearly end-of-life solar absorptance after only 18 months in space. Electron spectroscopy for chemical analysis (ESCA) and ellipsometry analysis of nearby optical samples

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Fig. 6.1

Localized contamination on ISS.

˚ of contamination had been deposited on one side of the indicated 5000–10,000 A experiment. These samples were LOS to a stowed solar array with large amounts of unbaked silicone [7]. The ISS program took steps to minimize molecular contamination by both materials selection and thermal vacuum bakeout, with a requirement of less ˚ contamination per year. Samples of Z-93, AZ93, and other white than 130 A ceramic thermal control coatings on the Materials on International Space Station Experiment (MISSE) indicate the success of the contamination control plan for ISS. Localized contamination can still be found on ISS (Fig. 6.1), but overall the performance of the radiators and solar arrays has not been impacted by molecular contamination at time of publication. Outgassing may be detrimental to materials even in the internal spacecraft environment. Silicone outgassing hampered the performance of beds used for carbon dioxide removal in the Environmental Control and Life Support System (ECLSS) on ISS. These beds were vented through the station’s vacuum system, promoting the outgassing. Thermal vacuum bake of the beds before flight significantly improved the performance.

6.5 STRUCTURAL MATERIALS Strength-to-weight ratio is usually the critical factor in choosing structural materials for spacecraft (see also Chapter 3). Static and dynamic loads must be considered, along with thermal performance, corrosion protection, manufacturability, reparability, and cost. Values for allowable properties of structural

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materials in their design environments should be taken from Metallic Materials Properties Development and Standardization (MMPDS, formerly MILHDBK-5); MIL-HDBK-17, Plastics for Flight Vehicles; MIL-HDBK-23, Structural Sandwich Composites; or other approved sources. Use-dependent properties must also be considered, such as dielectric constant for radomes or gas permeability for tanks. High-strength alloys of aluminum, titanium, and stainless steel have been in common use for decades. However, the 5000-series aluminum alloys containing more than 3% magnesium shall not be used in applications where the temperature exceeds 668C (1508F) because grain boundary precipitation above this temperature can create stress-corrosion sensitivity or exfoliation. This includes 5083-H32, 5083-H38, 5086-H34, 5086-H38, 5456-H32, and 5456-H38. For the same reason, the 300-series corrosion-resistant (CRES) stainless steel should not be used at temperatures above 3708C (7008F) for extended periods of time. With their higher chromium and nickel content, austenitic stainless steels are more resistant to stress-corrosion cracking than ferritic and duplex stainless steels. In general, titanium alloys and high nickel content alloys are resistant to stresscorrosion cracking. MSFC-STD-3029, Guidelines for the Selection of Metallic Materials for Stress Corrosion Cracking Resistance in Sodium Chloride Environments, states, “Many copper alloys containing more than 20 percent zinc are susceptible to stress corrosion cracking, even in the presence of alloying additions which would normally impart resistance to stress corrosion.” The standard also warns that protective coatings may only delay onset of stress corrosion, not entirely prevent it. Surface treatments such as carburizing or nitriding may increase susceptibility to stress-corrosion cracking. Aluminum–lithium alloys have 10% or more weight savings over standard aerospace aluminum alloys. For example, aluminum–lithium alloy was used in the super-lightweight tank (SLWT) for the space shuttle, with a weight savings of 7000 lbs. over the original external tank. Friction stir welding has been used on aluminum–lithium alloy as well as other aluminum alloys previously considered unweldable. Friction stir welding requires neither inert shielding gas nor filler material, reduces the number of weld defects, and has higher weld joint strength over fusion joining processes where the metal is melted. Space structures that demand tight tolerances in coefficient of thermal expansion, such as telescope optical benches, are usually made of composite materials. A wide variety of fibers, including graphite, boron, fiberglass, aramids, and carbon are available, as are many polymer resin systems, including epoxy, phenolic, polyimide, and polysulfone. The fiber may be in tow, tape, sheet, or woven form for traditional polymer-matrix composites. Metal-matrix composites (MMCs) and ceramic-matrix composites (CMCs) may have particulate or fiber reinforcement, where the fiber can be continuous or discontinuous (chopped fibers or whiskers). These are used where high toughness is needed.

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Fig. 6.2

X-33 cryogenic tank after failure.

For nonmetallic materials, atomic oxygen erosion may be a concern if used in low Earth orbit, as is chain scission or cross-linking of polymer chains in ultraviolet and particle radiation environments. These are usually surface effects, but a long exposure to atomic oxygen may compromise the strength of thin composites. For example, composites on the leading edge of the Long Duration Exposure Facility (LDEF) satellite lost a full ply to atomic oxygen over 5.8 years [8]. A high-radiation environment may result in strength loss and embrittlement for some polymeric materials. Honeycomb structures have also been used for their excellent stiffness. These can be with either composite or metallic facesheets and cores. A lesson learned from the X-33 honeycomb composite tank is the need for closed cell cores (dependent on application) and redundant permeation barriers in cryogenic applications to eliminate cryopumping. Cryopumping is “the influx of gas into an unclosed volume resulting from the vacuum generated when cryogenic temperatures liquefy and condense the gas on the cryogenic boundaries of that volume” [9]. When the X-33 tank was filled during testing, liquid hydrogen leaked through microcracks in the inner facesheet, and the nitrogen purge gas was pulled through microcracks in the outer facesheet to fill the core. When the tank was drained and started warming, the core pressure increased to the point of failure, when the outer facesheet and core separated and peeled away from the inner facesheet (Fig. 6.2; also see Chapter 7).

6.6 THERMAL CONTROL MATERIALS Passive thermal control of a spacecraft can be by a surface treatment, application of a coating, or covering with a multilayer insulation blanket. Surface treatment of

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metals is generally required to prevent corrosion before launch. Environmental protection directives have led to increasing restrictions on hexavalent chromate, which is commonly used in chemical conversion coatings. Newer conversion coatings may contain trivalent chromium or no chromium at all. Corrosion and space environmental effects testing of these coatings are continuing. Although chemical conversion coatings per MIL-C-5541 provide more than adequate corrosion protection, they are usually very low in thermal emittance and may not meet temperature requirements on orbit. Manned spacecraft that may require extravehicular activities (EVAs, also known as spacewalks) will have more stringent thermal control requirements for touch temperatures, generally 2118 to þ1138C (2180 to þ2358F). Better passive thermal control in terms of absorptance (a) and emittance (1) may be achieved with anodizing per MIL-A-8625. MIL-A-8625 calls out three types of anodize: Type I chromic acid, Type II sulfuric acid, and Type III hard anodize. Reduction of chromate use has virtually eliminated Type I chromic acid anodize. Type II sulfuric acid anodize that will be exposed to space should be processed with a hot water seal, not nickel acetate, as the residue from the nickel acetate seal yellows after a brief exposure to ultraviolet radiation. Type III hard anodize is also performed with sulfuric acid but usually at lower temperatures to get a thicker oxide layer, when needed for wear resistance. However, hard anodize should be carefully evaluated before use on parts subjected to fatigue. Boric/sulfuric acid and phosphoric acid anodize have been tested in space and may be useful when the desired thermal properties cannot be met by sulfuric acid anodized aluminum (generally a ¼ 0.45/1 ¼ 0.80) without compromising corrosion protection. When a lower a/1 ratio is needed, passive thermal control coatings or paints are used. Those qualified for use on spacecraft generally have one of four types of binder: polyurethane, epoxy, silicone, or potassium silicate. Acrylic-based paints have very poor performance in space and should not be considered. Polyurethane and epoxy-based coatings have limited life in low Earth orbit, because atomic oxygen will erode the binder, leaving only pigment particles behind [10]. Silicone, usually a low-outgassing type, is used when some UV darkening can be allowed or if the geometry is complicated enough where coating adhesion is a concern. Coatings with silicone binders should be used with caution when in LOS to sensitive optics. Potassium silicate coatings can be difficult to apply and are sensitive to contamination, but they have the advantage of proven durability in space for long periods of time and are generally considered safest around optics [11]. Manufacturer-recommended cure times for coatings should be followed (e.g., putting Z-93 potassium silicate coating in vacuum before the seven-day cure is complete will cause the coating to crack and debond). Manufacturers’ cure conditions may not be adequate to ensure the coating does not become a contamination source to sensitive optical surfaces. Materials may require thermal vacuum bakeout and should be tested under conditions equivalent to on-orbit extremes.

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Passive thermal control coatings and anodizes may be electrically insulative enough to build up surface charge in the space environment. Sufficiently high surface charge can even break down the anodized layer on aluminum (Fig. 6.3) or result in operational anomalies. NASA RP-1390 details a number of satellites that had system failures due to the plasma environment [12]. Some of these spacecraft charging events caused loss of control of the satellite. These events sometimes required station-keeping fuel to stabilize, therefore shortening the life of the satellite. Static-dissipative materials or conductive coatings can bleed off this surface charge before damage occurs. Indium tin oxide–coated thin films have been used for many years, but special care must be taken not to crack the coating. Fiberglass cloth with conductive thread has also been used. Conductive or static-dissipative thermal control coatings usually have conductivity on the order of 105 to 108 ohms/square. Coatings loaded with carbon for conductivity (e.g., Electrodag 501) should not be used in low Earth orbit, as the carbon will be eroded by atomic oxygen, but they may be acceptable for other environments. It is of utmost importance that the coating demonstrate conductivity in high vacuum after adequately conditioning (at least 24 h). In the past, some coatings were advertised as conductive when they were actually dependent on water content for their conductivity. As the water was removed by vacuum, these coatings became insulative. Multilayer insulation (MLI) blankets use multiple reflectors, usually thin polymer films with vapor-deposited metal on one or both sides, separated by lightweight, low-thermal conductivity materials. These thin films are fragile, as evidenced by the damage by purge on the Huygens probe (Fig. 6.4); therefore, a durable outer cover is usually added to the MLI blanket, and an inner cover

Fig. 6.3

Breakdown of anodized aluminum in plasma environment.

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Fig. 6.4 Worker repairing multilayer insulation on the Huygens probe after an abnormally high-pressure purge ripped the delicate material. may be added as well. Common MLI blanket materials and their properties can be found in NASA TP-1999-209263, “Multilayer Insulation Material Guidelines.” It should be noted that MLI blankets composed of Mylar reflector layers should be kept below 1208C (2508F) because of shrinkage and loss of tensile strength. MLI blankets require atmospheric pressure of less than 1025 torr to prevent convection and gas conduction between the reflector layers. The reflector layers may have perforations or porolations to allow better venting. MLI blankets should not be pulled taut; pinch points, heat shorts, and cutouts should be minimized. Organic threads should not be used in low Earth orbit unless they are adequately protected from atomic oxygen erosion. Hook-andloop tape fastener (e.g., Velcro or Aplix) seams should have some blanket overhang, preferably 1 in., to prevent atomic oxygen erosion of the polymeric hooks and loops down to stubs. It is recommended that unmated hook-andloop tape fasteners not be exposed to more than 1  1020 atoms/cm2 of atomic oxygen. Electrical grounding required for MLI depends on the size of the blanket and the environment. Proper grounding requires that the spacer netting be cut away and grounding tape, such as aluminum tape with conductive adhesive, applied in a continuous accordion between the reflector layers of the blanket (Fig. 6.5). A hole is punched through all of the layers, and a grounding bolt with washers, eyelet terminal, and locking nut is installed. A single conductor wire is generally used between the eyelet terminal and ground. In addition to MLI blankets, the large dewar on Gravity Probe-B used vaporcooled metallic shields. The dewar contained superfluid helium cooled to near absolute zero. As the helium was slowly vented during spacecraft operations, it was circulated through the shields. This design was successful in that the helium onboard stayed liquid for 16 months.

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Fig. 6.5

Electrical grounding design for multilayer insulation blanket [13].

6.7 THERMAL PROTECTION MATERIALS Thermal protection materials are separate from thermal control materials in that thermal control materials are used to moderate on-orbit temperatures, and thermal protection materials are generally for higher temperatures, such as around engine exhaust or for reentry. These temperatures may reach 28008C (50708F). Heat shields may be made from reusable materials, such as tiles or

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CMCs, or one-time use materials, such as ablatives. Heat shield materials selection is dependent on the peak heat flux and stagnation pressure during reentry (Fig. 6.6), as well as mechanical strength, density, entry angle, and the shape of the heat shield (i.e., blunt-body, sphere-cone, biconic, or nonaxisymmetric). Determining the thickness of the heat shield must balance weight vs the uncertainty of performance. For example, the ablation modeling of the Galileo entry predicted that the nose of the shield would ablate considerably more than the shoulder region. However, the data from the ablation sensors revealed that this was not the case. The TPS of Galileo recessed about 44 mm in the shoulder region, compared with the predicted ablation distance of about 33 mm, leaving a margin of only 10 mm [14, 15]. The stagnation point recession model predicted that the nose would recede about 88 mm, but the measured recession value was 41 mm. Silica ceramic tiles, or high-temperature reusable surface insulation (HRSI), were developed for use on the space shuttle. They are lightweight, can be formed in various densities, and can handle reentry temperatures up to 12608C (23008F), but they are fragile and easily damaged. They also require a coating to be waterproof. Toughened unipiece fibrous insulation (TUFI) tiles are stronger and tougher. The X-37B (Fig. 6.7) orbital test vehicle uses similar heat shielding as the space shuttle, with lightweight tiles on the belly and flexible insulation blankets for lower

Fig. 6.6 Peak heat vs stagnation pressure for various missions [14]. MER is Mars Exploration Rover; MPF is Mars Pathfinder.

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Fig. 6.7

X-37B after landing at Vandenberg Air Force Base.

temperature areas. Quilted blankets of woven silica fiber, silica batting, and aluminoborosilicate fiber were used on the space shuttle where reentry temperature stayed below 6498C (12008F). Reinforced carbon–carbon (RCC) was used in the nose cap and wing leading edges of the space shuttle. It can be used where reentry temperature exceeds 1,2608C (2,3008F). Carbon fibers in a silicon carbide matrix (C/SiC) were used in the nose cap, leading edges, and steering flaps for the X-38 vehicle, although this was only ground tested and not flown. Multilayer high-temperature ceramics such as zirconium boride and silicon carbide [16, 17] or nanocomposites [18] can be used to protect carbon–carbon composites against oxidation. Ablative heat shields are usually a honeycomb material with a resin or polymer injected into each cell. Materials include the Avcoat used on the Apollo capsules, phenolic-impregnated carbon ablator (PICA) used by the Stardust sample return capsule, and SLA-561V used on the Viking landers.

6.8 RADIATION SHIELDING Space radiation can be from galactic cosmic rays, trapped radiation, auroral radiation (polar orbits only), and solar flare. Most of the radiation is in the form of protons, electrons, and alpha particles (helium ions), but heavier ions do impact spacecraft. In addition, the reaction of radiation with higher Z (higher atomic number) materials creates Bremsstrahlung radiation. Low Z materials, such as polyethylene, are good protection from beta particles (electrons) and Bremsstrahlung radiation. A composite of low and high Z materials, particularly

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where the high Z material is sandwiched between two layers of low Z materials, can improve shielding per areal weight performance for shielding electronics. Adequate radiation shielding for long-term manned missions still needs to be developed. Astronauts may take cover in a “storm cellar” with extra radiation shielding during a solar flare. Liquid hydrogen and other fuel tanks as well as water tanks have been suggested for multipurpose radiation shielding of astronauts, as has incorporating lunar or Martian regolith into a polymer binder for in situ resource utilization [19]. Research is continuing in this field, particularly with hydrogenated graphite nanofiber and other nanocomposites. Active shielding generated by either magnetic or electrostatic fields has also been proposed but may be impractical due to interference with avionics and communications.

6.9 METEOROID/ORBITAL DEBRIS SHIELDING Hypervelocity impacts by meteoroids and orbital debris are a significant hazard for spacecraft. Average velocity for orbital debris is 10 km/sec, whereas meteoroids can travel as fast as 60 km/s. For comparison, a high-speed rifle fires at 1 km/s. The Russian satellite Kosmos-1275 was at an altitude known to be populated with space debris (977 km) and in a high-inclination orbit that increased relative velocities to that debris. In July 1981, the satellite broke up into more than 200 trackable fragments, likely due to collision with space debris. Although large pieces of orbital debris, usually 10 cm and up, can be tracked and avoided, smaller particles are still capable of damaging hardware, particularly windows. Over the 30-year life of the space shuttle, at least 45 windows had to be replaced because of impacts (Fig. 6.8). Windows on the ISS are

Fig. 6.8

Space debris impact on shuttle window.

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Fig. 6.9

ISS Cupola with open shutters. (Photo courtesy of NASA.)

triple-pane and, in the case of the Cupola (Fig. 6.9), have shutters to minimize the likelihood of impacts. The fundamental method of protecting a spacecraft from hypervelocity impact is the Whipple shield. The sacrificial bumper breaks up the impacting particle into smaller pieces or, in case of extremely high impact velocities, vapor. The number of bumpers and the spacing between them and the pressure wall or other critical hardware will be determined by desired probability of no penetration (PNP) and spacecraft weight and volume constraints. The most common Whipple shield material is aluminum. However, in the case of the Deep Impact mission, the impactor had five sheets of copper to differentiate it from any comet material. Various composites and honeycomb materials have also been used for bumpers. Thermal protection and hypervelocity impact protection can be combined in an MLI blanket. The ISS used a “stuffed Whipple” design, which includes a blanket of Kevlar and Nextel between the bumper and the pressure wall [20]. Other aramid materials such as Nomex, Twaron, and Aracon (which is coated Kevlar) may also be used. It should be noted that an MLI blanket may provide better protection when spaced evenly between the bumper and the pressure wall. For manned spacecraft, the Whipple shield design will also depend on what material is being used for the pressure wall. A study of the aluminum alloys 2219-T87 and 5456-H116 indicated increased petaling (where the aluminum peels back like flower petals, which can then detach and cause damage) and spallation in the 5456-H116 dual-wall at impact velocities higher than 7 km/s [21]. Recommended reading on this topic is the Handbook for Designing MMOD Protection [22].

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6.10 OPTICAL MATERIALS Window materials, such as quartz, fused silica, and sapphire, should have low impurity content. Space radiation reacts with the impurities to form color centers, degrading the transmission. Coating the window with cerium oxide will also improve the durability in space. Magnesium fluoride has been flown in both window form and as a layer over aluminum for a mirror. Workmanship is important, as some magnesium fluoride samples have endured the space environment quite well, whereas the Hubble Space Telescope’s original wide field/planetary camera flood mirror suffered blistering and peeling. Magnesium fluoride does react with atomic oxygen to form an oxide layer, as do lithium fluoride, calcium fluoride, and barium fluoride. Zinc selenide, which is used in some infrared optic applications, also reacts with atomic oxygen to form a thin oxide layer. This oxide layer increases absorptance. Most telescope designs do not directly expose the mirror(s) to the space environment, but there are space environmental effects data for gold, platinum, iridium, and nickel optical coatings [23]. Nickel reacts with atomic oxygen to form a thin oxide layer, with an accompanying drop in reflectance. Osmium should not be used in low Earth orbit applications as it reacts strongly with atomic oxygen and erodes away. An optical solar reflector (OSR) is a second-surface mirror where the metal, usually aluminum or silver, is protected by a layer of quartz. OSRs usually have low solar absorptance and high infrared emittance and can be used for thermal control in a high-radiation environment where metalized Teflon is contraindicated due to the increase in emittance with exposure.

6.11 SOLAR ARRAY MATERIALS The ISS uses silicon solar cells with ceria-doped borosilicate cover glass for power generation. Performance monitoring over the years indicates a 1.5%–3.5% power loss/year in a fairly low-radiation environment [24]. The Solar and Heliospheric Observatory (SOHO) is orbiting the L1 Lagrangian point, with a much higher radiation environment, but about the same degradation per year of 2% [25]. SOHO’s solar arrays are back surface reflector silicon solar cells and ceria-doped microsheet cover glass. Degradation due to proton radiation was observed, particularly during solar eruptions in July 2000 and November 2001. These solar eruptions with accompanying protons cause displacement damage, where radiation interacts with the solar cell lattice, producing defects that reduce the solar cell’s output voltage and current. One of the solar arrays flown on the Mir space station for more than 10 years was returned to Earth and was found to have 58% degradation, or 5.8% per year [26]. This is the same low Earth orbital environment as the ISS. A combination of factors led to this higher rate of degradation, including electrical arcing,

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Fig. 6.10 SCARLET solar array with curved silicone Fresnel lenses above multijunction solar cells.

meteoroid/debris impacts, degradation of interconnects, and heavy silicone contamination of the solar array. Although the behavior of silicon solar cells is well known, silicon has been replaced with higher efficiency multijunction solar cells on recent spacecraft, including the Mars Exploration Rovers Spirit and Opportunity. Multijunction solar cells typically have layers of germanium, gallium arsenide, and gallium indium phosphide. The Deep Space 1 probe included the solar concentrator array with refractive linear element technology (SCARLET) solar arrays with silicone linear Fresnel lenses for light concentration deployed over GaInP2/GaAs/ Ge solar cells (Fig. 6.10) [27]. Multijunction solar cells have the same challenges in a radiation environment as silicon solar cells with displacement damage. Data from Spectrolab, a leading manufacturer of solar cells, indicates approximately 1% per year performance loss in space for their high-efficiency triple-junction GaInP/GaAs/Ge solar cells over 15 years, more than likely with ceria-doped borosilicate cover glass to increase the durability [28].

6.12 LUBRICANTS The vacuum of space presents a challenge for keeping mechanisms lubricated. Hermetically sealing moving parts to keep the lubrication intact may be possible in some cases but is usually considered impractical. When choosing a liquid lubricant (oil, fluid, or grease), the lower the volatility, the better. Also, the component design must include provisions to retard creep, evaporation, and draining by gravity. The temperature extremes of space must be considered also, as liquid lubricants have a limited temperature range of operation. Some commonly used liquid lubricants include perfluoropolyethers, multiple alkylated cyclopentane, and polyalphaolefins.

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A better choice for moving mechanical assemblies exposed to the space environment is a solid lubricant. Solid lubricants include molybdenum disulfide, tungsten disulfide, niobium diselenide, graphite powder, silver, Teflon (polytetrafluoroethylene), and nylon. Unbonded solid lubricants include loose powder, brush-on lubricants, or spray-on lubricants. Burnishing, or applying loose powder onto a component with pressure, is another application method but is subject to wide variability in thickness and wear performance. This type of lubricant is not recommended for space-exposed applications. Bonded solid lubricants may be in a resin binder or inorganic binder. McMurtrey [29] notes that the wear behavior of resin-bonded lubricant films is different from other solid lubricants because of initial wear-in, loss of loose material on the bearing surface, and compaction. Inorganic-bonded solid lubricants in general do not perform as well as resin-bonded lubricants at room temperature but may be more suitable for high temperatures. Sputtered molybdenum disulfide may be difficult to apply for certain parts but has excellent coefficient of friction. Graphite-based lubricants will not work in space due to the lack of moisture and will become abrasive. The molybdenum disulfide lubricants are the better choice for space use, especially in sliding applications [30]. Lubricants containing chlorofluorocarbon constituents should not be used with aluminum or magnesium if high shear stresses can be imposed. Temperature limits should also be considered, such as for graphite, which loses its absorbed water at higher temperatures and is not recommended for use above 10008F (5388C). Molybdenum disulfide oxidizes at 7508F (3998C) and may also be affected by high humidity. Storage in an inert atmosphere is recommended for sensitive parts lubricated with molybdenum disulfide to decrease exposure to humidity. The Solar Alpha Rotary Joint (SARJ) on the ISS is an example of insufficient understanding of tribological interaction and therefore a weak lubrication design, which resulted in severe surface distress on orbit. The joint allows the solar arrays to rotate to maximize incident sunlight. It was noted during flight operations that more power was required to rotate the starboard SARJ than the port side one and that there was more vibration. Inspection of the starboard SARJ during a spacewalk revealed a large quantity of metal particles (Fig. 6.11) that was a spalled nitride layer of the rolling surface where the trundle-bearing assemblies contacted and rolled on the steel truss structure or race-ring joint [31]. These trundle bearings that formed the structural connection between the large rotating and nonrotating truss sections also included tapered roller assemblies that allowed the “alpha” rotation of the arrays to track the sun. The trundle bearings, similar to track follower bearings that hold an amusement roller coaster cart to a track, were designed for ostensibly rolling contact for low friction. But the on-orbit operation included just enough detrimental slip that the solid film lubricant scheme, a thin gold coating on the rollers, could not reduce the total slip/roll interface friction to a tolerable level. This combined friction ultimately exceeded the capability of the hardened nitride rolling surface

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Fig. 6.11

Damage in the SARJ on the ISS.

to withstand the subsurface stresses, resulting in a unique, massive spalling of the nitride layer. The damage required three more spacewalks to remove the trundle bearings; clean out the loose, hard metallic spalling debris; replace the trundle bearings; and lubricate the system, this time with vacuum-compatible grease that also included a molybdenum disulfide additive. The starboard SARJ, as of this writing, is operating fine with the repair, although the rolling surface is much rougher and softer than the intended design with the nitride layer intact. Although a rigorous ground test program was implemented to predict the life of the lubricant fix, ultimately time will tell if the repair was effective. Also, there are valid concerns over long-term outgassing of openly greased surfaces of such magnitude that are in LOS to contamination-sensitive surfaces on the ISS. Meanwhile, the port side SARJ, which was also lubricated with the perfluorinated MoS2 grease, continues to operate with no evidence of spalling damage. This SARJ anomaly example does not promote the use of a specific grease lubricant over other solid films, greases, or their base oils. Rather, the anomaly highlights the need to fully understand the system-level tribological behavior and desired lubrication regime (boundary to fully separating lubricant films) to properly design mechanisms that will perform with acceptable friction and tolerable wear. All too often a lubricant is called upon to correct weak tribological designs, much like treating a symptom rather than the responsible disease. Such “lube fixes” are often not even considered a principal design change as there is no perceived change to the “hardware.” Incorporation of a better tribological awareness in all mechanisms in the concept and design phase can save a lot of major headaches later on. Even so, all boundary lubrication regime tribological contacts (solid film lubricants, starved fluid lubricant films, slow moving high load, etc.) or other questionable tribological couples should be tested at the

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highest system and fidelity level practical. Modeling or analytically predicting behavior of boundary lubricated contacts requires understanding the interaction of more variables than can usually be adequately known.

6.13 SEAL MATERIALS Seal materials discussed in this section are limited to those used to maintain vehicle pressurization, pneumatics, and hydraulics. These may not be suitable for applications related to propulsion systems (see Chapters 10–13). Any seal material used in hydraulic systems should be checked for fluid compatibility. In general, seals are made of metal or elastomer. Metal seals can be made of soft aluminum, copper, or stainless steel, although nickel, Monel, and Inconel have been used in the past. A wide variety of elastomeric seal materials is available, including butyl rubber, silicone, Viton, Teflon, Kel-F fluoropolymer, and Neoprene. Seal materials to be used on windows should be thermal vacuum baked before flight to minimize outgassing.

6.14 ADHESIVES Two classes of adhesives will be covered in this section: structural adhesives, such as those used in honeycomb laminate manufacture, and nonstructural adhesives, such as the pressure-sensitive adhesive used for thermal control tapes. The coefficients of thermal expansion (CTE) of adhesives and substrates should be evaluated to ensure that a CTE mismatch does not lead to problems. When using an adhesive on two substrates with different CTE, the adhesive should have an adequate amount of elasticity (rubber or polyurethane) to compensate. Surface preparation requirements must be specified and rigorously enforced to maintain bond characteristics. Failures in large propulsion components have been attributed to inadequate control of the bonding process, usually either in the surface preparation steps or in the storage and handling of the adhesive material, including mixing and application of the two-part epoxies. Structural adhesives can be two-part epoxies, modified phenolics, or thermosetting resins. Cytec FM-300 and 3M AF-191 film adhesives have been used for honeycomb core/face bonds, and Hysol EA9394 paste for external splices. The epoxy family of thermosetting adhesives exhibits good solvent resistance and good elevated temperature properties to 3508F. Disadvantages are twocomponent mixing requirements, limited pot life, exothermic reactions, and deterioration of properties in hot and wet environments. Adhesives that use epoxy-amine or amide polymer systems will react with copper and result in blue to green corrosion products. Large propulsion components that use adhesively bonded components generally require long pot and working life and

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tailored cured properties. These materials typically are not available off the shelf; therefore, the adhesive manufacturers should be involved in identification or development of an adhesive that meets the design requirements. Pressure-sensitive adhesives are usually either acrylic or silicone based. 3M 966 (sometimes referred to as Y966) is commonly used in tapes and can be used on either metal or composite surfaces. 3M 9406PC is also acrylic and outgasses less than the 966. 3M 9703 is an electrically conductive acrylic adhesive. Arclad 8026, NuSil CV-1144, and DC93-500 are common silicone adhesives for spacecraft applications. Cleanliness and proper preparation of the surface before tape or adhesive application is essential. Other adhesive materials include polyurethanes, cyanoacrylates, and polyimides. The polyurethane adhesive family has excellent low-temperature flexibility. Disadvantages are complex mixing and application procedures, short pot life, moisture sensitivity in cured and uncured states, and poor elevated temperature performance. The cyanoacrylate family has good strength, fast setting, and ease of use and exhibits good adhesion to metal substrates. Disadvantages are high cost, poor durability on some surfaces, and limited solvent and elevated temperature resistance. Polyimides have excellent high-temperature capability (greater than 3708C/7008F) with high strength. Disadvantages are high cost, low peel strength, poor processing characteristics, low tack, and elevated temperature cure of 3438C/6508F.

6.15 LESSONS LEARNED Mac Louthan [32] teaches how catastrophic failures in materials engineering can be traced to six root causes that are applicable to spacecraft design, development, manufacture and assembly. Perhaps keeping these causes in mind can help to prevent future failures.

6.15.1

DEFICIENCIES IN DESIGN

The meteoroid shield and one of the solar panels on Skylab were damaged during launch because venting had not been adequately addressed. Sixty-three seconds after launch, internal pressure built up behind the meteoroid shield until the shield moved into the supersonic airstream. The shield was then torn from the orbital workshop part of Skylab, damaging one solar panel and blocking a second solar panel from fully deploying. This left the space station with little power or thermal control. Two EVAs were required, the first to deploy a temporary sunshade and the second to free the blocked solar array. Venting analysis had been performed but assumed a sealed aft end on the meteoroid shield. The need to keep this seal was not adequately communicated to aerodynamics, structural design, or manufacturing personnel [33].

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6.15.2

429

IMPROPER MATERIALS SELECTION

During the period of May through July 1998, one or more of the redundant spacecraft control processors (SCPs) failed in each of three Hughes HS-601 commercial communications satellites. This resulted in one of the satellites being removed from service. Three SCP failures were attributed to intermittent or continuous short circuits caused by the growth of conductive filaments, known as “tin whiskers,” from the tin-plated surface of an electronic assembly or its cover [34]. Tin plating is used on many electronic devices and on space hardware for corrosion protection and ease of soldering. Pure tin plating is subject to a crystal instability problem. The normal cubic crystal form of pure tin is unstable and tends to change to tetragonal at temperatures below 108C (þ508F). This instability increases with decreasing temperatures to a maximum at 2408C (2408F). This crystal transformation results in an expansion of the crystal lattice. This expansion causes the tin plating to crack and spall off the substrate. This phenomenon is called “tin pest” or “tin disease.” It can be prevented by adding small amounts of lead, bismuth, or antimony to the tin [35]. Tin whiskers create a potential electrical problem whenever printed wiring boards with tin plating are used for extended duration, regardless of the environment and whether the boards are conformal coated. Tin whisker formation has been widespread, and numerous incidents of electrical equipment failures have been attributed to the formation; therefore, copper conductors and circuit paths on printed wiring boards should not be electroplated with pure tin. Potential tin whisker growth can be eliminated by incorporating a minimum of 1.5% lead into the electrodeposited tin and reflowing the tin alloy.

6.15.3

DEFECTS IN MATERIAL

In the 1980s, Martin Marietta received a lot of two-part primer for the external tank in which the wrong solvent reducer had been shipped, even though the can was labeled with the correct material and the shipping paper showed that the correct material had been shipped. This primer was flight critical for bonding the urethane foam to the aluminum tank substrate. The error in labeling was not caught until after a technician sprayed a 1000-sq-ft tank dome, noted that the material did not behave normally, and reported it. The dome area had to be cleaned by hand sanding and scrubbing with Scotch Brite and a solvent [36]. Not only does this example point to the need for experienced personnel during manufacture and inspection, but it also shows the need for a culture where problems can be identified and taken care of, rather than ignored, covered up, or glossed over. Often material performance can be affected when manufacturers change material formulations or manufacturing process steps without the knowledge of the material user. Numerous occurrences were noted during the space shuttle

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program when unreported manufacturing changes resulted in unexpected material performance during flight. What manufacturers or raw materials suppliers may consider insignificant changes can require expensive hardware rework or disposition rationale development when discovered downstream in the manufacturing or launch preparation flow.

6.15.4

IMPROPER PROCESSING

During the Apollo 13 mission, the number 2 oxygen tank exploded during a “cryo stir” procedure. The tank held a “slush” of liquid oxygen with a fill line and heater running down the center, and the cryo stir procedure turned on an internal fan to keep the slush from stratifying. The original specifications for the tank and heater assembly were for 28 V dc power. This was later changed to 65 V dc; however, the thermostatic switches for the heater were missed in the changeover. These switches could accommodate the higher voltage during tank pressurization because they normally remained cool and closed, but they could not open without damage. During assembly, the tank was accidentally dropped, loosening the fill tube assembly and allowing gas leakage. Because the tank could no longer drain properly, an improvised detanking procedure was developed at Kennedy Space Center. During detanking, as the switches started to open at the upper temperature limit, they were welded permanently closed by the too-high voltage. Failure of the switches led to the heater being on for much longer and at much higher temperatures than expected, melting the fan motor wire insulation. The higher tank temperature was not noticed because the temperature gauges only measured to 808F. In flight, when the crew performed the cryo stir, the wire short-circuited, arced, and ignited its insulation, triggering the explosion in the oxygen tank [37].

6.15.5

INAPPROPRIATE ASSEMBLY

The first test of the tethered satellite system during STS-46 in 1992 was halted when the reel mechanism jammed. Only 260 m (853 ft) of the 20 km (12.43 mi) tether was deployed. An investigation found a protruding bolt from a latestage modification of the tether reel system.

6.15.6

INADEQUATE SERVICE

The high-gain antenna for the Galileo spacecraft did not deploy properly after launch. The reason for this is not precisely known, but it is likely that the three years in storage while waiting for launch after the Challenger disaster led to the jammed mechanism. The most likely culprit is evaporation of the lubricant, but damage could have occurred while in storage or during one of the unplanned trips between the Jet Propulsion Laboratory in California and Kennedy Space Center in Florida during the launch delay. The backup low-gain antenna saved

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the science mission. Because of this incident, NASA communications satellites with similar antenna designs have the pins replaced just before launch, and new lubricant is added [38].

6.16 CONCLUDING REMARKS As with any engineering task, the requirements must be properly understood, and the materials must be selected to meet those requirements with appropriate safety margins. Design of a spacecraft should take into account the interaction of all the components: structures; thermal; propulsion; avionics; guidance, navigation, and control; and so on. The full life cycle of the design should be considered, including safety, verification, reliability, operability, and maintainability. Decisions on materials and manufacturing should not be “stovepiped.” Spacecraft materials should be selected based on the environment and the time of exposure, with either durability in that environment or known degradation for appropriate endurance and end-of-life properties. They should be correctly assembled and maintained, with proper compatibility with surrounding materials and corrosion prevention. Attention to detail is never wasted. Good communication is essential. Communication with manufacturers and vendors can prevent problems with changes in processing or discontinued products. Interaction with other disciplines can contribute to an understanding of interfaces, identify deficiencies, and improve requirements. Ongoing dialogue can keep a small change from becoming a big problem. To confine our attention to terrestrial matters would be to limit the human spirit. —Stephen Hawking

ACKNOWLEDGMENTS The author gratefully acknowledges the assistance of numerous reviewers of this chapter, especially Mike Prince, Chip Moore, Tammy Hampton, Tim Vaughn, DeWitt Burns, Gary Pippin, Alan F. Stewart, and Jeff Finckenor.

REFERENCES [1] Blair, J., Ryan, R., and Schutzenhofer, L., “Lessons Learned in Engineering,” Lessons Learned Workshop, October 2007, p. 201. [2] De Groh, K., and Smith, D., “Analysis of Metallized Teflon Thin-Film Materials Performance on Satellites,” Journal of Spacecraft and Rockets, Vol. 41, No. 3, 2004, pp. 322–325.

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[3] NASA Office of Inspector General, “Semiannual Report to Congress, April 1, 2000–September 30, 2000,” http://oig.nasa.gov/SAR/sar0900.pdf [retrieved 8 Jan. 2018]. [4] Vaughn, T., “Manufacturing Development Challenges for Exploration,” Presentation to the Ares Program Manager Review, July 17, 2009. [5] Thieme, T., “ISS Learns from Mir’s Flame-Up,” Popular Science, January 24, 2003, http://www.popsci.com/military-aviation-space/article/2003-01/ iss-learns-mirs-flame [retrieved 8 Jan. 2018]. [6] NASA Jet Propulsion Laboratory, “Stardust Vision Nearly Restored,” January 11, 2001, http://stardust.jpl.nasa.gov/news/vision.html [retrieved 8 Jan. 2018]. [7] Zwiener, J. M., Kamenetzky, R. R., Vaughn, J. A., and Finckenor, M. M., “Contamination Observed on the Passive Optical Sample Assembly (POSA)-I Experiment,” SPIE Paper No. 3427A-20, SPIE International Symposium on Optical Science, Engineering, and Instrumentation, San Diego, CA, July 1998. [8] Whitaker, A. F., and Young, L., “An Overview of the First Results on the Solar Array Materials Passive LDEF Experiment (SAMPLE), AO171,” LDEF First Post-Retrieval Symposium, NASA CP-3134, June 1991. [9] “Final Report of the X-33 Liquid Hydrogen Tank Test Investigation Team,” NASA Marshall Space Flight Center, Huntsville, AL, May 2000. [10] Stein, B., “LDEF Materials: An Overview of the Interim Findings,” LDEF Materials Workshop, NASA CP-3162, Nov. 1991. [11] Finckenor, M., Pippin, H. G., and Frey, G., “MISSE Thermal Control Materials with Comparison to Previous Flight Experiments,” 9th International Space Conference: Protection of Materials and Structures from the LEO Space Environment, Toronto, Canada, 20–23 May 2008. [12] Bedingfield, K. L., Leach, R. D., and Alexander, M. B. (eds), “Spacecraft System Failures and Anomalies Attributed to the Natural Space Environment,” NASA Reference Publication 1390, August 1996. [13] Finckenor, M. M., and Dooling, D., “Multilayer Insulation Material Guidelines,” NASA TP-1999-209263, April 1999. [14] Laub, B., and Venkatapathy, E., “Thermal Protection System Technology and Facility Needs for Demanding Future Planetary Missions,” International Workshop on Planetary Probe Atmospheric Entry and Descent Trajectory Analysis and Science, Lisbon, Portugal, 6–9 October 2003. [15] Milos, F. S., “Galileo Probe Heat Shield Ablation Experiment,” Journal of Spacecraft and Rockets, Vol. 34, No. 6, 1997, pp. 705–713. [16] Corral, E. L., and Loehman, R. E., “Ultra-High-Temperature Ceramic Coatings for Oxidation Protection of Carbon-Carbon Composites,” Journal of the American Ceramic Society, Vol. 91, No. 5, 2008, pp. 1495–1502. [17] Tului, M., Lionetti, S., Pulci, G., Rocca, E., Valente, T., and Marino, G., “Effects of Heat Treatments on Oxidation Resistance and Mechanical Properties of Ultra High Temperature Ceramic Coatings,” Surface & Coatings Technology, Vol. 202, 2008, pp. 4394–4398. [18] Koo, J. H., Lee, J., Lao, S., Jor, H., Pilato, L. A., Wissler, G., and Luo, Z. P., “Nanomodified Carbon/Carbon Composites: Further Thermo-Oxidative Studies,” Proceedings of International SAMPE 2007 ISTC, SAMPE, Covina, CA, 2007.

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[19] Tripathi, R. K., and Nealy, J. E., “Mars Radiation Risk Assessment and Shielding Design for Long-Term Exposure to Ionizing Space Radiation,” IEEEAC Paper No. 1291, Version 4, IEEE Publications, Piscataway, NJ, 23 Nov. 2007. [20] 3M, “3M Nextel Ceramic Fibers and Textiles: Technical Reference Guide,” 3M Advanced Materials Division, St. Paul, MN, 2016, http://www.3m.com/market/ industrial/ceramics/pdfs/CeramicFabric.pdf [retrieved 8 Jan. 2018]. [21] Schonberg, W. P., “Aluminum 2219-T87 and 5456-H116: A Comparative Study of Spacecraft Wall Materials in Dual-Wall Structures Under Hypervelocity Impact,” Acta Astronautica, Vol. 26, No. 11, 1992, pp. 799–812. [22] Christiansen, E. L., “Handbook for Designing MMOD Protection,” NASA JSC-64399, Jan. 2009. [23] Vaughn, J., Linton, R., Finckenor, M., and Kamenetzky, R., “Evaluation of Atomic Oxygen Effects on Metals and Optical Thin Films on EOIM-3,” AIAA Space Programs and Technologies Conference, AIAA 93-4104, Sept. 1993. [24] Garza, I., Munoz, J., and Martinez, A., “SAW Power Performance Assessment,” Report to International Space Station Vehicle Control Board, Aug. 2011. [25] Rumler, P., Schweitzer, H., and Evans, H., “SOHO Solar Array: A Performance Evaluation of 2.5 Years in Orbit and Capabilities for a Mission Extension of 5 Years,” Proceedings of the Fifth European Space Power Conference (ESPC), Tarragona, Spain, 21–25 Sept. 1998. [26] Visentine, J., Kinard, W., Brinker, D., Scheiman, D., Banks, B., Albyn, K., Hornung, S., and See, T., “MIR Solar Array Return Experiment: Power Performance Measurements & Molecular Contamination Analysis Results,” AIAA-2001-0684, Jan. 2001. [27] Murphy, D. M., “The Scarlet Solar Array: Technology Validation and Flight Results,” Proceedings of the Deep Space 1 Technology Validation Symposium, Pasadena, CA, 2000. [28] Bailey, S., and Raffaelle, R., “Space Solar Cells and Arrays,” Handbook of Photovoltaic Science and Engineering, edited by A. Luque, and S. Hegedus, 2nd ed., Wiley, New York, 2011, pp. 365–401. [29] McMurtrey, E. L., “Lubrication Handbook for the Space Industry,” NASA TM-86556, Dec. 1985. [30] Fusaro, R. L., “NASA Space Mechanisms Handbook,” NASA TP-1999-206988, July 1999, Chaps. 15 and 22. [31] Elliot, H., McFatter, J., Sweeney, D., Enriquez, C., Taylor, D., and McCann, D., “The International Space Station Solar Alpha Rotary Joint Anomaly Investigation,” 40th Aerospace Mechanisms Symposium, May 2010, NASA/CP-2010-216272, pp. 187–206. [32] Louthan, M. R., “Life Lessons of a Failure Analyst”, ASM International, March 31, 2016, ISBN: 978-1-62708-110-8 [33] “NASA Investigation Board Report on the Initial Flight Anomalies of Skylab 1 on May 14, 1973,” NASA, 13 July 1973, https://history.nasa.gov/skylabrep/SRcover. htm [retrieved 1 Jan. 2018]. [34] Jet Propulsion Laboratory Standard “Flight Project Practices, Rev. 7,” JPL Doc ID 58032, September 30, 2008, Paragraph 7.3.6. [35] Babecki, A. J., “Problems with Pure Tin Coatings,” NASA Materials Engineering Branch TPI No. 021, Sept. 2002.

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[36] Morring, F., “Lessons: Forty years of space flight experience are a guide to human space flight today,” Aviation Week & Space Technology, December 6, 2010, pp. 68–70. [37] “Report of Apollo 13 Review Board,” NASA N70-76479, June 1970. [38] “Galileo’s antenna: pinning the problem,” Science News, Vol. 140, No. 10, Sept. 7, 1991.

CHAPTER 7

Materials for Launch Vehicle Structures Grant Henson Invariant Laboratories LLC, Westlake, Ohio

7.1 INTRODUCTION This chapter concerns materials for expendable and reusable launch vehicle (LV) structures. An emphasis is placed on applications and design requirements, and how these requirements are met by the optimum choice of materials. Structural analysis and qualification strategies, which cannot be separated from the materials selection process, are described. An LV is an airborne system that delivers a payload from the ground to suborbital, orbital, or interplanetary space. The payload is usually housed in a space vehicle or satellite that is not considered part of the LV. When it is not important to distinguish the payload from the space vehicle, both may be referred to as the payload. Modern LVs are designed with a particular type of payload in mind (astronauts, earth-orbiting instruments, interplanetary probes, etc.), but at the dawn of the Space Age, vehicles performed multiple duties. For example, the Atlas, Titan, and Thor/Delta vehicles all began as long-range weapons and were later adapted for orbital delivery. Sounding rockets such as Aerobee (historical) and Black Brant can leave the atmosphere but do not enter orbit. For the purpose of this chapter, shorter-range missiles that never leave the atmosphere are not considered LVs. Most LVs, including Atlas, Delta, Ariane, and Proton are expendable. Expendable vehicles are flown only once; the upper stages may be disposed of through a controlled reentry or may be left in orbit as “space junk,” while the first stage or booster falls to Earth in a cleared area. The term booster usually means the first stage of a multistage LV and will be used in that sense here. Reusable systems may incorporate a single vehicle that both launches the payload and houses it while in space, the prime example being the Space Shuttle Orbiter. The Orbiter, and the similar Soviet Buran vehicle, are here considered LVs rather than space vehicles because they must sustain atmospheric flight loads and environments similar to those sustained by expendable boosters.

Copyright # 2018 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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Therefore, the materials selection aspects are much the same as for expendable LVs. Proponents of reusable vehicles assert that they can be cheaper and more reliable than expendables. On the other hand, recovery and refurbishment are costly, and a failure of a vehicle intended for reuse is more damaging to schedules and budgets than a failure of an expendable vehicle. The envisioned benefits of reusability have led to recent investment, both public and private, in reusable vehicle development. One source [1] claims that a reusable variant of the Aerobee sounding rocket was flown; if so, it was the first reusable vehicle. Notable reusable orbital LV programs that never demonstrated powered flight were the Sea Dragon, X-33, X-34, and K-1. The first stage of the Soviet/Russian Energia vehicle, developed to lift the Buran Orbiter as well as other heavy orbital payloads, was designed to be reusable for at least 10 flights [2]. However, it has never actually been recovered and reused. The DC-X/-XA was an early demonstration of reusable rocket flight within the atmosphere. SpaceShipOne reached suborbital space in 2004, landed, and repeated the feat. However, neither of these systems led to a sustained record of operations. In 2015, a New Shepard vehicle, including both the booster stage and the space vehicle, was recovered from suborbital flight and then successfully reflown 61 days later. Also in 2015, the booster stage of a Falcon 9 was recovered by powered descent onto a land-based pad after having launched a payload to orbit; since then several attempts to descend onto a seagoing platform have been successful. Today, LVs are considered, along with aircraft, part of a single endeavor we call aerospace. But various dictionaries date this term only back to the late 1950s, at least a decade after the guided missile, for better or worse the archetype of the modern LV, was developed. In most nations, the initial authority for developing guided missiles rested with the artillery or ordnance corps, not the air corps. The relevance of this observation is that although LV materials and structures technologies have much in common with those of aircraft, the degree of commonality is perhaps less than one might think. Investment in LV development and operation is now a small part of the overall aerospace economy. However, for several decades, political and military imperatives drove high expenditures on LV development, leading to significant advances. New materials and structures had to be developed in parallel with other vehicle systems in “crash” programs, under high risk of technological failure, in order to satisfy aggressive performance requirements within the desired time frame. Although the pace of innovation was slow for decades, increased emphasis on cost reduction and improved reliability continue to drive incremental advances  For

example, the U.S. Census Bureau reported about $23 billion in deliveries of “guided missile and space vehicle manufacturing,” “guided missile and space vehicle propulsion unit and propulsion unit parts manufacturing,” and “other guided missile and space vehicle parts and auxiliary equipment manufacturing” in 2005, which surely includes many billions spent on non-launch-vehicle hardware such as antiaircraft missiles. Compare this to $114 billion in deliveries of aircraft and related items [3]. Considering that many countries manufacture aircraft but not LVs, LVs probably constitute less than 10% of the global aerospace economy.

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in materials and structures technology. Also, large, qualitative improvements in computing capabilities and newly available precursor materials have provided a technology push to encourage further advances in LV materials and structures. Because materials selection for LVs is affected by laws and regulations that vary from country to country, it is important to note where LVs are built and used. Until the 1970s, the United States and the Soviet Union (Russia and Ukraine) dominated LV production. More recently, France and China have developed and operated a significant number of LVs. Within the last few years, India, South Korea, North Korea, and Iran have also developed LVs. The French and Ukrainian vehicles are launched from different countries than the ones they are produced in. Also, many vehicles contain major substructures or engines built in several different countries. Table 7.1 shows orbital launches broken down by country of final factory assembly. Chapters 11 and 12 of this book are dedicated to materials for the solid rocket motors and liquid rocket engines, respectively, that propel LVs. Propulsion materials and structures are mainly affected by the loads and environments generated within the engine or motor itself, such as thrust chamber pressure. TABLE 7.1

ORBITAL VEHICLES LAUNCHED OVER TWO RECENT PERIODS, GROUPED BY COUNTRY OF PRODUCTION [4, 5]

Period

Country of Production

1990 to 1998

United States

39%

Russia

32%

France

13%

Ukraine

9%

China

5%

Japan

2%

Israel, India 2007 to mid-2009

Share

Russia

,1% 30%

United States

26%

China

14%

Ukraine

13%

France

8%

India

4%

Japan Iran, Israel, North Korea, South Korea

2% ,1%

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However, a section is provided in this chapter on large solid rocket motor cases because they can form a significant part of the load-bearing capability of the vehicle as a whole. The structural failure of a large strap-on solid rocket motor on an Ariane 5 or the Space Shuttle, or the solid rocket boost stage of the Ares I, would doom the vehicle structure rather than just the propulsion system. Inclusion of solid rocket motor cases with the structural system rather than the propulsion system follows the precedent set in Chang’s work [6]. Also, propulsion support systems such as propellant feedlines are included in this chapter because they are usually the responsibility of the LV contractor.

7.2 LAUNCH VEHICLE STRUCTURES The typical missile-derived expendable LV may be thought of as a stack of tanks with an engine at one end and a payload at the other. The fuel and oxidizer are contained in separate tanks. In more detail, the engines are mounted to the aft end of the tanks and exert thrust through a reinforced structure. The tanks are connected with thin-walled cylinders called skirts or intertanks. Complete stages are connected to one another through cylindrical shells called interstages or adapters. When the connected stages are of different diameters, the adapter has the shape of a truncated cone, which may have its smaller diameter forward or aft. When the smaller diameter is aft, the structure may be referred to as a boattail. The forward end of the vehicle is formed by a tapered shell that also encloses the payload. This structure is referred to as the payload fairing, payload shroud, nose fairing, or nose cone. Inside the nose cone, and attached to the forward end of the upper stage, is the payload. The payload is attached through a payload adapter or payload fitting. Therefore, at the forward end of the vehicle, there are two primary load paths: the payload fairing or outer branch and the payload attach fitting or inner branch. Usually, but not always, the tank walls themselves carry the primary loads. Occasionally, if a stage is much smaller in diameter than the payload compartment or the booster, the entire stage may be contained in a non-load-bearing aeroshell or aerofairing. The major substructures are attached using bolted flanges. The connections may be made with the vehicle in either the horizontal or the vertical position, in a factory or at the launch site. The final placement of the payload onto the vehicle frequently takes place with the vehicle actually sitting on the launch pad. Figure 7.1, a cutaway view of the Saturn V LV used to launch astronauts to the Moon, shows the location of the tanks, engines, and payload. The Apollo payload was unusually large and bulky and resided within a complex fairing topped by an escape rocket. A nearly cylindrical interstage can be seen joining the booster to the second stage, and a conical one can be seen joining the second and third stages. Some internal structures and stiffeners in the tanks are visible. The booster fuel and oxidizer tanks are joined by a cylindrical intertank, whereas the secondand third-stage tanks have common bulkheads to save weight and volume.

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Fig. 7.1 Cutaway view of the Saturn V launch vehicle with the Apollo payload, showing major substructures. Anonymous, “Illustration of Saturn V Launch Vehicle,” NASA image MSFC-0100979, 1967.

The outer mold line is the outermost surface of the cylindrical structure, visible from the outside, whereas the inner mold line is the inner surface. These terms, common in composite molding processes, are used even if there was no molding involved in building the structure. LV shell structures may completely lack internal bracing or stiffening, may have stiffeners integrally machined into the wall, or may have mechanically attached stiffeners or braces. Extensive internal framing is rarely used in LVs except in thrust structures. The term membrane is used to refer to the part of a shell structure far from attachments or other discontinuities, in which only in-plane loading is significant. This same area may be called acreage, especially when discussing thermal protection systems. In contrast, flanges, door seals, bolt lines, and the like may be called details or closeouts; closeouts especially refer to small items or fasteners that are the last to be installed when building the vehicle. Reusable designs with winged launch and reentry vehicles do not conform to the description just given. The Space Shuttle is functionally split into the reusable

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Orbiter, the partially reusable Solid Rocket Boosters, and the expendable External Tank (ET). Many different concepts, from single-stage-to-orbit to staged systems comprising a winged vehicle piggybacked on a more conventional missile-like booster, have been proposed. Wilhite [7], in the context of a trade study, discusses some of the materials selection aspects of advanced fully reusable designs. It is telling that only rather exotic materials (a metal matrix composite with silicon carbide fibers and monolithic titanium aluminide) were considered feasible for the two-stage-to-orbit systems he explored.

7.3 BASIC MATERIAL CHARACTERISTICS As with other aerospace applications, the most important characteristics of LV materials are material strength, based on any applicable failure criteria; material stiffness, as quantified by the elastic modulus or moduli; mass density; nature of the failure modes (gradual or sudden); ability to tolerate small-scale damage; and mechanical and chemical compatibility with nearby materials. Long-term damage resistance or durability are not as important in expendable LVs as in reusable ones and are much less important than in aircraft. In many LV applications, the foregoing must remain favorable at very high or low temperatures and in the presence of humid, corrosive, or other degrading environments. Because most LVs use cryogenic propellants, properties at very low temperatures are important; high-temperature properties can also be important because of the aerodynamic heating encountered in the high-speed atmospheric part of the trajectory. Knowledge of material characteristics must be quantitative in order to play a direct role in structural system trade studies. The stiffness and density of most materials are consistent enough to be treated as deterministic values for a particular material at a given temperature. However, material strength displays sample-to-sample variation that must be taken into account in both design and analysis; design values based on tenth- or first-percentile strength are more important than average strength. Further, if the factors tending to cause variations in strength are poorly understood, high safety factors must be used to preserve reliability, leading to heavier structures. Equally important is manufacturability. Without the ability to shape or assemble a material into an efficient structure, the material’s intrinsic advantages become meaningless. For instance, a single carbon nanotube is extremely strong, but until a carbon nanotube structure of useful size can be manufactured while preserving this extreme strength, that material will not play a significant economic role. Aspects of manufacturability that are especially relevant to LV applications include weldability, machinability, ease of making a composite laminate and formability or “drape” of plies, and ease of assembly using fasteners, co-curing, adhesives, locking features, and so on.

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Thermal properties may also be important; in particular, it is desirable to have thermal expansion characteristics that are predictable and compatible with adjacent materials, including tooling. These general characteristics must be associated with relevant, measurable material properties, or at least be translated into standardized tests. A good summary of the properties and tests most relevant to structural design can be found by reviewing the data tables in the universally referenced Metallic Materials Properties Development and Standardization (MMPDS) published by the Federal Aviation Administration (FAA) [8]. This reference was formerly known as MIL-HDBK-5. In this work we find data on material strength, including typical values and statistically derived lower-bound design allowables for tensile yield and rupture (“ultimate”), compressive yield, shear rupture, and bearing yield and rupture; elongation to break; tensile and compressive Young’s modulus; shear modulus; Poisson’s ratio; density; and thermal conductivity, heat capacity, and coefficient of thermal expansion (CTE). These properties are reported for a wide range of tempers of commonly used metals. They are usually given for various thicknesses because heat or age treatment affects metals differently depending on the thickness. Also, they may be given at elevated or cryogenic temperatures for various exposure times, or plots of temperature adjustment factors may be provided. In some cases, full-range stress-strain curves are provided. These are required in order to perform stress analysis in the plastic range. Finally, S-N (fatigue) diagrams and Paris-region crack growth curves are provided for many alloys. Metal properties at cryogenic temperatures depend strongly on the crystal structure. Face-centered cubic metals such as aluminum and the austenitic stainless steels experience a rise in ultimate strength but a lesser increase in yield strength, which preserves their ductility. Body- centered cubic metals such as the ferritic steels tend to experience a greater increase in yield strength than in ultimate strength, which results in more brittle behavior. For composite materials, which are generally not isotropic, more extensive (and expensive) testing may be required for full characterization. To take full advantage of the directional stiffness and strength properties of composites, directional material properties must be available. Composite properties are not as readily available as metal properties because of the proprietary constituents and processes that are used, and hence they are not widely applicable. However, one frequently consulted reference that may be used for initial design calculations is the Composite Materials Handbook [9], formerly sponsored by the Department of Defense as MIL-HDBK-17. In this handbook we find data on strength, modulus, and elongation to break for fiber, tapes, prepreg cloth, and laminae under various temperature and moisture conditions. This information, in combination with thickness and ply angles for laminate designs, may be used to build up the full laminate stiffness matrix. Much of these data are labeled by fiber volume fraction, ply thickness, and other processing parameters, but these parameters may vary so much in practice that it may be difficult to find directly applicable handbook data.

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MMPDS defines the A-, B- and S-values as statistical minimums for design use. Roughly speaking, the A- or S-values are suitable for nonredundant structure and the B-values are suitable for redundant structure. The A-value is the value that 99% of all samples are expected to exceed, at the 95% confidence level. The B-value is the value that 90% of all samples are expected to exceed, at the 95% confidence level. The S-value is not a statistically derived value but rather a specification minimum. S-values may be substituted for A-values provided the material is screened to ensure the S-value is met. Although every materials and structures engineer should be thoroughly familiar with these definitions, their significance should not be exaggerated. It has been said that “typically, less than 1 percent of composite structures on large aircraft is actually governed by unnotched laminate strengths” [10]. Although this may be overstating the case, clearly the familiar uniaxial tensile strengths are not the last word in material characteristics. Hart-Smith [10] states that “joints, damage tolerance, and stiffness” govern the choice of the rest of the materials. These may be regarded as a minimum set of properties needed to produce a credible preliminary design. However, many other properties, in particular strength properties under flight-like combinations of loads and including stress raisers, are important. Even with the widespread availability of finite element analysis, it is still important to characterize material strength in realistic regimes through careful testing. A detailed, nonlinear, validated finite element analysis may well prove more expensive and less reliable than a well-planned test to determine, for example, the fatigue life of a bonded joint. Some examples of strength testing from the literature are biaxial strength [11], cryogenic fracture toughness and fracture toughness ratio [12], hardness, tangent modulus, impact, notched fatigue, weld coupons, and creep-rupture [13]. In addition to numerical property data, MMPDS and the Composite Materials Handbook also include information on applications, materials processing, corrosion resistance, maximum service temperatures, and other information relevant to the designer. A comprehensive handbook on materials selection for launch vehicles (and space systems in general) that is more oriented toward physical/chemical properties and compatibility is MSFC-HDBK-527, Materials Selection List for Space Hardware Systems, published by NASA Marshall Space Flight Center [14]. This handbook provides an extensive summary of knowledge concerning the corrosion, stress-corrosion cracking, propellant and working fluid compatibility, flammability, toxicity, and thermal vacuum stability properties of aerospace materials, both metallic and nonmetallic. Another excellent reference is the Aerospace Structural Metals Handbook [15]. This work, which was originally sponsored by the U.S. Air Force Research Laboratory, contains not only extensive tables of data, but also a cross-reference so that the same alloy may be located under names that may vary from producer to producer or country to country. Data are usually typical properties rather than statistical minimum design values. The book is now available as an online database.

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Per-piece raw material cost is usually small compared with tooling and labor costs at the low production rates typical of LVs. Therefore, the cost of the material in its unprocessed form is rarely an important consideration in materials selection. If a material is commercially available in the required sizes and quantities and on the needed schedule, it is a candidate for use in an LV structure, practically regardless of cost. Historically, space programs would even specify custom materials having no existing commercial applications and therefore being subject to unknown cost and production fluctuations; for example, Rocketdyne developed NARloy-Z specifically for use in the linear aerospike engine and Space Shuttle Main Engine [16]. But lately this high-risk, high-reward approach has been discouraged. Although much of the effort to develop requirements and materials for reusable vehicles stemmed from the Space Shuttle Orbiter program, the same questions had to be addressed by the designers of the Soviet/Russian Buran Orbiter [17]. Much of this development had to take place independently because of the political situation. Unlike the Space Shuttle Orbiter, Buran did not have booster engines, only orbital maneuvering engines; launch was solely by means of external boosters. The Buran designers found that riveting was not compatible with graphite-epoxy composites, due to inadequate impact strength. They also reported that due to galvanic corrosion, it was not possible to use aluminum fittings with composites, so titanium was used instead. This problem was largely solved on the Space Shuttle by careful material compatibility studies. As in the West, the Buran designers noted that the strength and stiffness properties of composites tend to vary more than those of metals. Finally, the Buran designers identified fastening and joining as the key challenge in designing with composites, a finding that many composites designers will agree with.

7.3.1 DURABILITY AND REUSABILITY Fatigue, fracture, and aging characteristics are less important for expendable LVs than for aircraft or reusable LVs. However, when long delays between manufacture, testing, and operation must be accommodated, thermal and chemical aging as well as ambient moisture uptake should be considered in materials selection. Repeated ground tests can consume some of the fatigue life. Material characteristics that are particularly important in reusable vehicles are resistance to fracture and the propagation of cracks under fluctuating loads, ductility, resistance to stress corrosion, the ease with which damage can be found and characterized, and chemical and electrochemical compatibility with other materials or contained fluids. Structures in a reusable vehicle will obviously experience more loading cycles than if the vehicle were expended, but airliner-style operations in which thousands of flights may be accumulated are not yet possible for LVs. For example, the Space Shuttle Orbiter airframes had a design lifetime of 100 missions. For metal primary structure not exposed to high load fluctuations and designed to withstand flight

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loads without macroscopic yielding, 100 missions will not consume a significant amount of the high-cycle fatigue life. However, undetectable preexisting cracks on highly loaded structures or near-stress raisers may grow to dangerous lengths within 100 flights. Failures due to fracture may pose a risk to nearby components if a moving part is liberated. Also, low-cycle fatigue, which by definition requires significant plastic deformation, can be important on expendable LVs. In the present context, it is sufficient to understand that the fracture failure mode occurs when a fatigue crack grows to its critical size (the size at which unstable, catastrophic propagation of the crack occurs). Predicting the initiation of a crack is outside the normal scope of the fracture analysis; the analysis assumes the existence of the largest undetectable crack at the worst-case location at the time of inspection. The fracture or “safe-life” analysis predicts the growth of the crack under the expected spectrum of fluctuating loads. It predicts how long the loads may be sustained before the crack reaches its critical length. Safe-life analysis† may be defined as the understanding and quantification of life estimates. Safe-life-critical structures are likely to be included in the LOLI (limited operating life item) listing of the vehicle. LOLI hardware can be lifelimited due to corrosion life, battery life, time of operation, thermal cycles, and so on, but here we focus on the safe-life fracture analysis. A LOLI definition is provided for the vehicle that includes the “zero time” and how cycles are to be counted. A quality control group tracks the cycles for each vehicle. As far as safe life is concerned, LOLI counts are counts of stress excursions beyond a defined level, and the zero time is the time at which flaw inspection was done. Reinspecting the structure is a way to reset the zero time and gain additional life. For an expendable vehicle, the service life is an assumed, fixed number of load cycles high enough to allow checkout and multiple launch attempts, each involving a load cycle due to tank prelaunch pressurization. As more and more vehicles of a particular type are launched, fewer launch attempts should be needed per actual launch, so the assumed service life may decrease. For life-limited structures, the shorter assumed service life can lead to higher life margins, greater tolerance for manufacturing discrepancies or found flaws, and lighter-weight structure in case there is an opportunity for design changes. Figure 7.2 shows an idealized crack growth curve for a metal under fluctuating stresses. This is commonly referred to as a da/dN curve, where a is the crack length and N is the number of cycles. The many factors influencing this curve, such as stress ratio and frequency, are discussed in detail in preceding chapters. Because of the short life of an expendable LV, crack growth concerns are frequently in Region 3 of the da/dN curve. Being unstable in nature, Region 3 predictions can be unreliable. When the metal is ductile, much of this Region 3 crack growth is of a tearing nature. In situations where production discrepancies or damage during prelaunch operations †

This material on fracture-based safe life and fracture control was contributed by John Hilgendorf, structural analysis lead for Delta II, United Launch Alliance.

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Fig. 7.2 Idealized plot of crack growth as a function of stress cycles for a metal. (By J. Hilgendorf, United Launch Alliance.) occur, it is sometimes necessary to remove conservatism to adequately assess the risk associated with the damage. In these cases, elastic-plastic fracture mechanics or other less conservative theories may be used. When sustained loading is part of the load spectrum, stress corrosion of the potential flaw needs to be considered. Keac (or KIssc) is a truncated value that toughness can degrade to, under sustained loads. The stress-corrosion resistance may need to be taken into account for pressure vessels storing fluids used to pressurize pneumatic, hydraulic, or ullage pressure systems. The time at load can be as short as a few hours. For vehicles considered to be at risk of failure due to crack propagation, a formal fracture control program may be implemented. Information describing how to write a fracture control plan may be found in NASA Standard 5003 [18]. A fracture control program classifies parts as fracture critical if they exceed a certain mass; are uncontained, non-fail-safe, or part of a pressurized system; or meet other criteria that suggest serious consequences in case of failure. For fracture-critical components, the fracture control program applies special analysis, testing, and inspection requirements to reduce the chance of a harmful fracture. These vary from program to program but generally amount to an analytical determination of the smallest crack that could grow to critical size before the next regular inspection and an inspection plan that will detect a large percentage of cracks larger than that critical size. In addition, the fracture control program places restrictions on the materials that may be used and specifies

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the documentation needed to ensure that the correct material has been used, that it has been processed in a way to discourage the initiation of cracks, and that the proper inspections have been performed. It also specifies a factor to cover analysis uncertainty: typically, a fracture-critical part may be used for one-fourth of the life predicted by the safe-life analysis before it must be reinspected. Because they involve inspection, fracture control programs are most commonly seen in aircraft and in reusable LVs such as the Space Shuttle. Expendable vehicles cannot be inspected after use unless they are recovered, and then they will not be flown again anyway. However, expendable vehicles must undergo ground tests that consume some of the safe life of the parts, and inspection is possible after ground tests. Therefore, fracture control may be applied in expendable vehicle programs to a limited extent.

7.3.2 SPECIALIZED MATERIALS Most of the foregoing discussion applies to metals and composites, which are by far the most important materials used in LV structures. Their useful regime is linear elastic, and the effects of temperature and other environments on their behavior is small enough that it may usually be accounted for with adjustment factors. If a metal structure does yield, the amount of yielding is small enough that deformation plasticity in the form of an isotropic Mises yield function, followed by a Ramberg–Osgood description of plastic flow, is usually sufficient. For more complex materials such as elastomers, foam, and adhesives, materials testing becomes even more expensive and time consuming, and good property data accordingly are harder to come by. Fortunately, these materials are often used in applications where very accurate mechanical property data are not vital. Many of these materials display time-dependent behaviors, such as relaxation and creep, and have strong temperature dependence. They may also have nonlinear stress-strain curves or such a large strain during operation that they must be treated with one of the many nonlinear theories of mechanics. For materials that are not linear elastic, the distinction between phenomena and properties becomes important. Phenomena are behaviors such as elasticity, creep, and relaxation that can be observed and measured without assuming a particular material model. Observing material phenomena can be useful for screening or lot acceptance and can suggest an appropriate material model, but phenomena are usually insufficient inputs for accurate simulation of structural response. To conduct accurate analyses and simulations, a material model (constitutive equation) must be assumed, and only then can the properties defined in the model be measured. For instance, some type of stiffness may be measured for all elastic materials, but once one is forced to consider large strains of a compressible material, a large-strain model containing three properties may be necessary. A conventional uniaxial tension test will not suffice to determine the three properties; multiple specialized tests are needed. A less desirable but nevertheless

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common approach is to adjust the properties until analysis agrees with a variety of measured responses that are similar to the actual application of the material.

7.3.3 RATIONAL METHODS OF MATERIALS SELECTION Materials selection is a part of structural design optimization, whether the optimization is done intuitively by an experienced designer working on a minor variation of an existing design or quantitatively through the use of a large material properties database and algorithms for adjusting hundreds of design variables. The classical approach to optimum design, including materials selection, was comprehensively reviewed by Gerard [19]. It involves the definition of a design index based on a requirement. For example, to optimize a thin-walled column, equations relating external load to the critical stress for two failure modes (column instability and local buckling) are derived, and by requiring that the margin of safety for both failure modes be minimized, a design index is determined. In this example, the index is a function of Young’s modulus, some sort of plastic modulus, the load, and the length of the column. Given a set of values for some of these parameters, the others can be chosen to optimize the design index. The design index approach is only tractable for problems involving a few key parameters. The ability to determine ahead of time which parameters are key is an aspect of engineering genius that not everyone enjoys. But by computerizing the process, the number of variables can be greatly enlarged, so an intuitive ability to narrow down the design space is less important. One such approach was documented by Mukhopadhyay [20]. Chapter 3 of the present book discusses materials selection in greater detail.

7.4 STRUCTURAL DESIGN AND REQUIREMENTS Materials selection is as much a part of the design process as sizing. In fact, the two cannot be separated. Therefore, the requirements and criteria that impinge on the structural sizing process also impinge on materials selection. Development practices in LV materials and structures are an interesting combination of extreme conservatism and bold risk-taking. Modern LV development programs typically budget for zero or one test flight before an expensive payload is launched. Differences in payloads and trajectories tend to limit the amount of knowledge that can be carried from one flight to the next. When a military service decides to launch a billion-dollar, one-of-a-kind payload critical to national security on an expendable LV in a configuration that may never have been flown before, the materials selection, structural sizing process, and testing are held to standards that owe more to custom than science. The launch decision itself is a major, irrevocable commitment of resources based on a significant extrapolation of experience. Therefore, the extrapolation process must be as rational as possible.

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The following discussion is necessarily general because program- specific policies are usually trade secret and/or export-controlled. This section does not purport to review unusual or innovative structural qualification methods, or specific reliability requirements, that are not documented in the public domain.

7.4.1 CONTRACTUAL REQUIREMENTS By far the most significant requirements are those imposed by the procuring agency or, in the case of commercial operations, by the payload client. In some cases, these requirements are actually drawn up by the LV contractor itself, subject to revision and approval by the procuring agency. Requirements exist in a hierarchy that is managed by systems engineers primarily to ensure that the LV delivers a functioning space vehicle to the desired orbit and secondarily to minimize the cost, development time, danger to the public, and other factors. The structural system, propulsion system, guidance and navigation system, and other systems are considered subsystems of the LV system as a whole. Blair and Ryan [21] provide a good overview of requirements and standards and how detailed design criteria are derived from them. A set of top-level functional requirements for the structural system that could well apply to many different LVs is to support and protect the other vehicle systems and the space vehicle such that they can function properly; to contain and deliver working fluids to the propulsion, guidance, and other systems; to maintain an aerodynamically acceptable shape; and to do these things in a way consistent with the functioning of the other vehicle systems—for example, by allowing electrical grounding. Top-level requirements may specify not only the performance goals to be met, but also the likelihood that the design will meet them. It may be required that the vehicle be 98% likely to meet all requirements—that is, to place an intact payload into the proper orbit 49 out of 50 times on average. It is difficult or impossible to predict whether a complex machine like an airplane or an LV will satisfy such a requirement simply based on the design. There are too many interacting failure modes. For aircraft, the large number of repeated operations makes it possible to develop some empirical rules of thumb. But even an empirical approach is usually not possible for LVs because of the low numbers of identical vehicles and operations. Some researchers have attempted to use a Bayesian statistical approach to circumvent the lack of data [22]. An alternative might be to break down the vehicle into a few standard subsystems and try to reuse those standard designs on many different vehicles, thus providing a significant experience base. But for LVs this is the exception rather than the rule. Top-level reliability requirements are best interpreted as a general statement positioning the desired reliability relative to similar systems. It is healthy to realize that perfect reliability is neither possible nor desirable. For example, the Japanese space development agency set a reliability goal of 96% for their

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H-2A vehicle, stating honestly that they would not be “aiming for the ultimate in design” [23].

7.4.2 LAWS AND REGULATIONS The preceding discussion covered requirements imposed by the procuring agency or self-imposed by LV contractors. Another class of requirements is that imposed by laws and regulations. These seek to minimize overflight and environmental hazards to the public. In the United States, the FAA regulates commercial space operations but not operations carried out by, or on behalf of, the federal government. This excludes the majority of launches (and reentries) from FAA scrutiny. Also, Title 14 of the Code of Federal Regulations does not impose the same very detailed structural requirements on LVs as it does on aircraft. It is mostly concerned with hazards from expended stages, reentering payloads, and mishaps. The FAA’s relationship with the private space launch industry is still evolving, but it appears that private launches will not be regulated as closely as passenger aircraft. Therefore, vehicle safety laws and regulations do not significantly constrain materials selection for LV structures. However, environmental regulations have had a significant and ongoing impact on materials selection for LVs, particularly in the area of coatings and insulation. Heavy metals such as cadmium, mercury, and lead were once commonly used in metals processing and plating, but as it has become widely known that these substances are poisonous, regulations have greatly reduced their use. Beryllium has important aerospace structures applications due to its thermal stability, but beryllium dust is toxic and must be handled carefully. Also, the use of asbestos insulation and chlorofluorocarbon blowing agents for foam insulation has been greatly reduced by environmental regulations.

7.4.3 RANGE SAFETY The other major class of requirements is that imposed by operators of launch ranges to minimize the risk of injury to personnel and damage to ground equipment. Military, government nonmilitary, and commercial organizations alike must adhere to range safety rules. The vast majority of LVs are operated out of the ranges listed in Table 7.2. For many years, the governing range safety document for the Eastern and Western Ranges of the United States was EWR 127-1, Eastern and Western Range Safety Policies and Procedures [24]. Although EWR 127-1 states that it is “applicable to all organizations, agencies, companies and programs conducting or supporting operations on the ER and WR,” it now only governs programs introduced at the Ranges before 2004. Since 2004, Air Force Space Command has issued the manuals AFSPCMAN 91-710, Range Safety User Requirements Manual [25], and AFSPCMAN 91-711, Launch Safety Requirements for Air Force Space Command Organizations [26], as replacements for EWR 127-1. The

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TABLE 7.2 Name

MAJOR SPACE LAUNCH RANGES

Launch Location(s)

Notes

Eastern Range

Cape Canaveral Air Force Station, Florida

Mainly low-inclination orbital vehicles on a southeastward ground track

Western Range

Vandenberg Air Force Base, California

High-inclination orbital vehicles on a southward ground track and suborbital vehicles westward toward Kwajalein Atoll

Wallops Research Range

Wallops Island, Virginia Small suborbital and orbital vehicles in eastward to south-ward directions

Guiana Space Centre

Kourou, French Guiana Orbital vehicles to a wide range of inclinations

Baikonur Cosmodrome

Tyura-Tam, Kazakhstan Orbital vehicles along a corridor extending northeastward over Russian territory

Plesetsk Cosmodrome

Arkhangelskt Oblast, Russia

Northward into high-inclination and polar orbits

Sea Launch

Equatorial Pacific Ocean

Low-inclination orbital launches

former is binding on all range users, but the latter is binding only on Air Force space programs. EWR 127-1 sets as a general goal that the risk of injury or damage to the public due to space launches should be no greater than that normally accepted in day-to-day activities, including the risk due to airplane overflights. It uses language such as “all reasonable precautions shall be taken” and “lowest risk possible.” Section 3.12 of EWR 127-1 contains detailed requirements for testing and analysis of pressurized systems and structures on LVs.‡ It requires that materials be compatible with working fluids, seals, lubricants, and so on, from the standpoint of flammability, ignition and combustion, toxicity, and corrosion and requires the range user to supply evidence in the form of a report. It specifies that material compatibility should be based on T.O. 00-25-223, Integrated Pressure Systems and Components (Portable and Installed), Chemical Propulsion ‡

As defined by EWR 127-1, a pressurized system is a system such as a helium storage bottle that is primarily designed to contain internal pressure, whereas a pressurized structure is a system such as a main propellant tank that carries both internal pressure and significant external loads.

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Information Agency Publication 394 [27], MSFC-HDBK-527 [14], or independent testing. EWR 127-1 also specifies qualification, acceptance, hydrostatic proof, and leak testing requirements for pressure vessels and pressurized systems. It requires quite specific design solutions to reduce risk, such as the location of drains and vents, design of interconnects, and the like. It addresses graphite-epoxy composite overwrapped pressure vessels (COPVs) in a separate appendix, which requires demonstration of a leak-before-burst (LBB) failure mode for metallined COPVs, nondestructive evaluation of the composite overwrap, special fluid compatibility testing, and design/test/pedigree record-keeping in accordance with MIL-STD-1522 [28]. These requirements are for the safety of ground personnel and the public. For small-diameter lines in particular, static design factors may be as high as 4.0, and required safe life may be as long as four expected service lives. The very detailed and prescriptive regulations in EWR 127-1 were consciously relaxed in the new AFSPC manuals, not necessarily with the intention of raising risk, but rather to change the approach from risk avoidance to risk management. Some specific materials selection rules in EWR 127-1 have been deleted from the new manuals. The thinking behind this is outlined in a National Academy of Engineering study [29]. Quantitative requirements have replaced the “all reasonable precautions” language, and the range user is given more discretion in implementation. This initiative was partly driven by the desire to reduce the cost of range safety and make the ranges more attractive to commercial users.

7.4.4 VERIFICATION AND QUALIFICATION A vehicle can meet all design requirements but still fail to deliver the payload to orbit. Further, because of randomness in material properties, dimensions, and loads, one successful flight of a system does not guarantee future flights will also succeed. Even in the case of a reusable vehicle, 49 successful flights do not verify the requirements are met if the design lifetime is 50 flights. The vital question, and one that the materials and structures engineers must help answer, is whether the next flight will be successful. Analysis and review of ground test and previous flight data are necessary, bearing in mind that predictions of future flight performance are at best a rational extrapolation of experience. The benchmarks determining whether the system is ready for the next flight are set cooperatively by the materials and structures engineers, the systems engineers, and others. Some engineers, notably Sarafin [30] in reference to satellite structures, refer to these benchmarks as verification criteria rather than requirements. The distinction is made in order to discourage blind adherence to rules, because, after all, those criteria only represent an educated guess as to the best way to build confidence in system reliability. The overall means of qualifying LV structural hardware for flight may be a contractual mandate, a company policy, or simply tradition, but the preferred

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method of qualifying LV primary structure will always be a single-article test to limit load times a factor. Other verifications such as proof testing or analysis are adjuncts to this basic approach. Requirements can be so narrowly written that they are really prescribed designs that hold back the state of the art. It would not be desirable, for instance, to require propellant tanks to be designed to a one-size-fits-all specification such as that used for rivets. But although excessively narrow requirements and standards may have been imposed in areas such as avionics, this was not the case in materials and structures. The U.S. Department of Defense abandoned military standards and even prohibited contracts from citing them as requirements for a time in the 1990s [31]. This was part of a government-wide political initiative that affected NASA (“better, faster, cheaper”) as well [32]. After a series of highprofile failures in the late 1990s, procuring agencies concluded that wholesale abandonment of standards was too extreme, and systems engineering processes began to reintroduce them [33].

7.4.5 STRUCTURAL QUALIFICATION In this section, the most commonly used concepts in structural qualification are introduced. Although terminology varies, these concepts appear in most government standards concerning structures, and knowing their meaning is a prerequisite to understanding the various qualification strategies. Design limit load is the maximum expected in-service load. Programs may be very precise; a common definition is that limit load is the 99.7 percentile of a distribution of loads that may be generated by analysis, flight measurements, or both. Such loads are usually generated from a finite number of samples, so it is often stated additionally that the 99.7 percentile load must be determined to a confidence level of 90%. Design factors are factors by which limit load is multiplied to determine the no-yield condition (the load at which the structure must not suffer detrimental deformation), the proof condition (a load used for acceptance testing), the no-break condition (the load at which a structure must not lose its load-carrying capacity, through breakage or instability), and other hypothetical load levels used in analysis. Design factors are chosen by, or subject to the approval of, the procuring agency. Test factors are analogous to design factors but are used to factor up the limit load for testing purposes, as opposed to design purposes. They are usually equal to the corresponding design factors, but they do not have to be. For example, if limit load is 10 tons, and the design ultimate factor is 1.25, analysis must show that the structure will withstand a load of 12.5 tons. Most likely an ultimate load test would also specify a load of 12.5 tons, but it could specify 14 tons or some other factored-up value. Because limit load already takes quantifiable uncertainties into account, design and test factors can be viewed as insurance against “unknown unknowns.”

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Capability is a lower bound on the ability of a structure to resist detrimental deformation and to maintain its load-carrying capacity. It is determined by analysis using material yield and ultimate strengths (which are lower-bound values) and the least favorable dimensions allowable in built hardware. Margin of safety, or simply margin, is the fraction by which the capability exceeds the no-yield or no-break conditions. Thus, continuing the previous example, if the structure is predicted to buckle or break at a load of 15 tons, the ultimate margin would be 15=(1:25  10)  1 ¼ þ22%

(7:1)

The sign is customarily shown on a margin even if it is positive. Using this definition, the capability may be viewed as the load at which the margin of safety is zero. The demonstrated load is the load by which the test factors were multiplied in generating loads during a successful test. Generally, there are two tiers of design factors: a lower set of values, meant for use on structures that have been tested, and a higher set, meant for use on structures that have not been tested. To be entitled to use the lower, “tested” set of design factors, a structure cannot be exposed to flight loads in excess of the demonstrated test load. In such situations, the demonstrated load becomes the allowable load for the structure. Even if the margin is positive at the allowable load, flight loads must not exceed it; otherwise, the lower design factor is no longer justified. The demonstrated load is sometimes known as the limit test load, and the demonstrated load times the ultimate test factor is sometimes known as the ultimate test load. However, these should not be confused with design limit and ultimate conditions. The test loads are fixed once the test has been completed, but the design conditions may vary as knowledge is gained about the LV. For an untested structure, the allowable load is the load at which the margin of safety is zero. In other words, for an untested structure, the allowable load equals the full capability. In contrast, large test articles are not usually tested to full capability or to destruction, only to design limit load or less, thus constraining the flight article to an allowable load at which ample margin may exist. The “hidden margin” between the allowable load and the capability of a tested structure is an important fact to consider when comparing the relative risk of testing versus not testing a structure. Testing can uncover a dangerous condition that analysis alone might miss, even when higher safety factors are used to compensate for the lack of testing. The relationship between the various design conditions, the test and analysis results, and the design factors and margins is illustrated in Fig. 7.3. This figure shows the predicted flight loads and predicted failure loads in the form of histograms, which could be generated by Monte Carlo simulations or from an assumed distribution. For instance, an individual failure load might be calculated Monte Carlo–style from random draws of material strength and dimensions from

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Fig. 7.3 Distributions of predicted loads and failure loads, and the definitions of key load levels and design quantities. distributions consistent with sampled test and dimensional data. Or, more commonly, it may simply be a Gaussian distribution fit to a mean and variance. Flight loads are more likely than failure loads to be built up from random underlying contributors, but in principle both can be done that way. The figure shows the capability as a lower limit on predicted failure loads and the design limit load as an upper bound on predicted flight loads. The illustration shows the typical circumstance in which flight load predictions are more scattered than failure load predictions. This arises from greater underlying uncertainty in wind statistics, trajectories, and other inputs to the loads analysis, as well as uncertainty in the analytical model itself. It also shows that the capability and limit load do not enclose every single predicted load, and, in that sense, they are not truly bounding values, although we call them that for convenience. The demonstrated limit load is a single value, shown in gray on the figure. It is typically close to the design limit load. The intent is usually to test the structure to exactly the limit load, but limit load can change as new knowledge is gained. Finally, the figure shows that the design factor provides separation between limit load and the no-fail condition, and the separation between limit load and the capability is a function of both the design factor and the margin of safety.

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One may hear a statement like the following: “The test article was loaded to 140% of the no-yield condition, so a tested margin of 40% has been established.” This is not a correct use of the term margin, because the test was of a single article that could have been stronger than average. Margins are based on lower-bound strength, not averages. It would, however, be correct to say, “The test load was 90% of capability, so there was a 10% margin of safety during the test.” The capability represents the lower-bound strength, and the test load is known, so there is no need to account for uncertainty in the load. Therefore, the stated margin of 10% is meaningful. Stiffness is as important as strength in LVs. The thin-walled construction, combined with the strength and stiffness properties of typical materials, tends to render the buckling margins about the same as the strength margins, and both are always checked. From a material properties standpoint, stiffness is less variable than strength and is less affected by temperature and moisture. Therefore, nominal modulus values are often sufficient, especially for metals. In composites, a lower-bound stiffness may be obtained by testing “hot-wet” samples—that is, coupons saturated with moisture and held at the maximum expected service temperature. But composite stiffness properties that are truly applicable at the scale of a full structure can be challenging to measure. Specially laid up and cured coupons may have different microstructure than the full-scale component. Coupons cut out of a full structure may have damaged edges. However, analysis for stiffness is less exact, and therefore more conservative, than analysis for strength. The buckling failure mode is the one most influenced by stiffness. Because the buckling load of a thin-walled shell is strongly affected by slight geometric imperfections and edge constraint, adjustment or “knockdown” factors derived from experiments on subscale specimens are applied. These factors may lead to a reduction in the predicted buckling strength of 50% or more, as compared with the theoretical value for a geometrically perfect shell. Factors documented in a NASA monograph [34] were originally developed from experiments on small plastic cylinders. Bushnell comprehensively reviewed the state of the art in shell buckling analysis through 1980 [35]. Recently, recognizing the major role played by buckling knockdown factors in vehicle design, NASA conducted a Shell Buckling Knockdown Factors research program that was the most significant work in the field in decades and that experimentally supported a significant refinement and reduction in the factors [36].

7.4.6 PITFALLS, CONTROVERSIES, AND ENGINEERING JUDGMENT Stated requirements, and the strictness with which they are applied, vary between programs. Knowing what to require in a particular situation depends largely on factors specific to each program. Such factors are neither public nor readily transferable to new situations, so this discussion is limited to the pros and cons rather than advocacy of particular solutions. Because primary structure must be qualification tested to the no-break condition, if predicted loads increase, for instance

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due to payload weight growth, analysis refinements, correction of mistakes, and so on, the structure must be retested. However, an expensive and time-consuming retest will only be contemplated if the increase in loads is “significant.” There may be special provisions for allowing higher loads on a structure than what it was tested to, possibly using a sliding scale of design factors. Also, it is sometimes not easy to determine the range of applicability of a structural test. If a material must be slightly changed from that used in the test article, is the design with the new material still qualified, or must it be retested? From this scenario comes the idea of qualification by similarity. This refers to a formal process of demonstrating that a design may be considered test qualified even though it is not identical to the test article. A detailed comparison of material, geometry, and manufacturing differences is necessary, as defined in MIL-HDBK-340 [37]. An example of qualification by similarity occurs when a propellant tank must be enlarged to meet new mission requirements. The course usually followed is to “stretch” an existing, qualified design. Often, the stretched design may be considered test qualified, even though it is longer than the original test article. The guiding requirement in such cases is that the new design must have the same failure modes as the original, with equal or higher margins of safety. There is controversy in the definitions of primary and secondary structure and their implications for testing. The fundamental divergence may be illustrated by considering two structures, A and B. Suppose Structure A was successfully qualification tested and has zero margin of safety using tested design factors. Structure B was not qualification tested but has zero margin of safety using higher, no-test design factors. May the two structures be considered equally acceptable under all circumstances? One school of thought says that the reliability added by using the higher, no-test design factors completely compensates for the lack of testing. Using typical values, consider that a structure with zero ultimate margin using a design factor of 1.60 would have a margin of 60% if a design factor of 1.00 were used. From this perspective, a program may elect not to test some primary structures. The distinction between primary and secondary is then made mostly on the basis of size: the vehicle can tolerate “fat” designs of small structures needed to accommodate no-test design factors, but cannot tolerate fat designs of larger structures. Therefore, larger structures are tested only to enable the use of lower, tested design factors. This less conservative viewpoint is characteristic of programs without heavy involvement of a procuring government agency. The other school of thought posits that higher safety factors can never completely compensate for the risk of an analysis shortcoming that would only be revealed by testing. Therefore, primary (critical, nonredundant) structure must be qualification tested, whether or not it has positive margins using no-test design factors. Also important is the “hidden margin” discussed previously. There is a long history of success in operating structures qualified by test, but those structures were usually neither tested nor flown to their full capabilities. Structures qualified

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solely by analysis, using a typical no-test design yield factor of 1.60, have allowable loads 1 2 (1/1.60) ¼ 38% lower than capability. But successfully flown, tested structures were most likely limited to loads 20% lower than capability simply because they were not tested to full capability and were limited in flight to the testdemonstrated load. Therefore, allowable loads for nontested structures would be not 38% lower than the experience base, but rather only about 18% lower. The “pad” provided by the no-test factors of safety does not appear quite so comfortable when viewed this way. Structures that are nearly always considered primary are fairings; payload fittings and adapters; main propellant tanks; interstages, intertanks, skirts, and transition sections; and engine thrust structures.

7.4.7 OUTLOOK The level of conservatism that ultimately proves more cost effective is different in every case and is what makes structures engineering more than just a calculation process. It is not surprising that the organization that bears the cost of testing tends to take a less conservative approach, whereas the organization that bears the cost of a failed mission tends to be more conservative. When the same organization bears the costs of both testing and flight failures, a rational ordering of priorities is forced. But often, the responsibilities are separated, and the negotiated level of conservatism is determined by a political process, not an objective technical one. Current flight rates are too low to conclusively prove which approaches are superior. The structural subsystem itself, and especially any single structure, must have a very remote chance of failure in order for the vehicle as a whole to have a reasonably small (say, one in a hundred) chance of failure. It is not uncommon for the required probability of failure for a particular structure to be on the order of one in a million. Even if a less conservative approach leads to double the chance of failure (say, 2  1026) for a single structure, this will not be empirically distinguishable from a more conservative approach over the life of a program. The danger is carrying this thinking over, by inattentive systems engineering or lax verification of requirements, to every structure. Then, of course, the vehicle as a whole will have twice the risk of failure. It has been noted that a truly reusable LV would allow requirements to be made more rational, as they are in aeronautics, by generating a large performance database for the same flight article. A look at LV failure statistics shows that the overall demonstrated reliability of LVs worldwide was 96% for the period 1984–1994 [6]. Of the failures, the propulsion system was by far the leading cause (27 out of 43 failures).§ Just five out of 43 failures were attributable to primary structure in that period: a payload fairing failure on a Chinese CZ-2E, a Centaur liquid oxygen tank failure, and three §

The author has counted solid rocket motor case failures as structures failures.

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solid rocket motor case failures, including the well-known Challenger disaster. Many failures cause the vehicle structure to be destroyed, but these are usually due to primary failures in other systems, leading to loads in excess of those the structure was designed to sustain. In such scenarios, the structure is not considered the root cause of failure. A probabilistic approach to structural integrity would dispense with the question of primary versus secondary structure. Instead of using design factors, in a probabilistic approach, each component would be assigned a probability of failure considering all sources of uncertainty.

7.5 PRESSURIZED STRUCTURE The majority of material in a LV is found in integral load-bearing propellant tanks. This section is mostly confined to discussion of materials for the tank shells; tanks also have small parts such as sumps, lids, and outlets that are subject to different requirements than the shells. Propellant tanks function as pressure vessels, containing fluids under moderate pressure and often at cryogenic temperatures. However, unlike stationary pressure vessels, propellant tanks must sustain large, highly variable primary flight loads. This has been the case since the early days of rocketry, when for reasons of weight, external load-bearing shells protecting tanks from flight loads (as in the V-2) were replaced by integral load-bearing tanks. Also, the need to reduce mass has required that propellant tanks be much more lightly constructed, with far smaller design factors than stationary pressure vessels. Finally, propellant tanks in expendable vehicles are operated for only a short time, so longterm, time-dependent processes such as creep and corrosion are less relevant. Flynn, in a book covering all aspects of cryogenic engineering, devotes some discussion of propellant tanks as compared to other applications of cryogenic technology [38]. He also provides a useful discussion of cryogenic insulation, which will be discussed later in this chapter. Government standards such as range safety requirements consider the main propellant tanks to be “pressurized structures” rather than pressure vessels (refer to EWR 127-1 [24] for one formal definition), reserving the designation of pressure vessel for smaller tanks such as propulsion system pressurization tanks that do not bear significant external loads. Factors of safety and other requirements are much different for pressurized structures as opposed to pressure vessels. Propellant tanks are of three basic designs. The commonest is the stiffened metal shell, structurally stable under the load of its own weight when empty and unpressurized. Stiffening is generally by integrally machined stiffeners in an isogrid or orthogrid pattern rather than by mechanically fastened stringers. Such designs are constructed of aluminum alloys. The next most common is the “steel balloon” design, which is very thin walled and not structurally stable under the load of its own weight unless pressurized or stretched. Its stability

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before fill and pressurization is maintained by pressurization with an inert gas or by mechanical tension applied by a holding cradle. This design was most famously applied in the Atlas missile. Both the stiffened and balloon-style metal designs may be of a single tank space, containing either fuel or oxidizer, or combined fuel and oxidizer tanks separated by a common, dome-shaped internal bulkhead. The common-bulkhead tank offers mass and size savings over separated fuel and oxidizer tanks and has been used in such high-performance upper stages as the Saturn S-II and S-IVB [39] and the Centaur. A drawback of this design is the need for the common bulkhead to control heat flow between two propellants that may be at vastly different temperatures. The third type of design is the composite tank. Whereas non-cylindrical shapes would be very difficult to achieve in a mass-efficient manner with metallic shell designs, such shapes are less troublesome with composites. Also, composite tanks offer potentially significant mass savings through higher material specific strength and the ability to orient the primary load-carrying direction of a composite laminate along the expected loading direction. Composites also offer better resistance to fatigue and flaw propagation, because microscopic flaws tend to be blunted and stopped by the fibrous microstructure, although accumulated fatigue damage can result in increased permeation of propellant. With all these advantages, much effort has been expended on realizing an operational composite propellant tank, but to date, successes have been few. All tank designs must perform the basic function of containing the liquid propellants during testing, fueling, and flight. Propellants vary from RP-1, a highly refined kerosene, to cryogenic liquid oxygen (LOX) and liquid hydrogen (LH2), to storable but often toxic combinations such as hydrazine and nitrogen tetroxide. All have properties that constrain the designer’s choice of propellant tank materials, and cryogenic propellants require that the tank be insulated to minimize boil-off. In almost all cases, tanks must sustain aerodynamic and inertial flight loads, which for the typical long, cylindrical tank means a combination of axial compression and bending. The Space Shuttle external LOX tank is a special case in that it receives axial aerodynamic loading directly due to its position at the forward end of the tank assembly. Inside the tank, various baffles and propellant management devices must be supported. Finally, depending on the tank’s location in the vehicle, main propellant feedlines and electrical tunnels must be supported, either as an external appendage or through centerline tunnels as in the Saturn S-IC stage [39]. The tank contents must be fed to the engines under pressure. For a pressure-fed propulsion system, propellants are forced directly into the combustion chamber by ullage pressure. The ullage is the unfilled space at the forward end of the tank. For pump-fed engines, moderate pressure is still necessary to prevent cavitation in feedlines. Just before launch, large tanks are pressurized using a ground supply of gas; once the booster engines have been started, the gas

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supply may be provided by the engines through a repressurization system. For smaller stages, an onboard supply of inert pressurant is often used. It is worth recalling the basic relationship between load and internal forces for a pressurized thin-walled cylinder subject to an external compressive force P0 and bending moment M0 at the tank bottom. (Shear force is usually not significant when considering the overall section forces acting on an LV.) For increased generality, suppose the cylinder contains a quantity of liquid of density r and that it is accelerating forward at a rate a. The tank and its contents together do not form a continuous elastic body, so they must be analyzed separately. The pressure at a distance z below the free surface of the liquid is p(z) ¼ rz(a þ g) þ pull

(7:2)

where pull is the ullage pressure and g is the acceleration of gravity. A separate free-body diagram shows that the axial compressive force in the tank shell at location z is P(z) ¼ P0  m(z)a  pR2 rh(a þ g)  pR2 pull

(7:3)

where h is the total height of the liquid in the tank, R is the tank radius and m(z) is the mass of the tank aft of location z. Part or all of the force P ¼ P0 2 m(z)a 2 pR 2rh(a þ g) may be provided by a separate loads analysis. It may also include vibratory effects and other terms not shown in this simple analysis. Consider the typical case where the force is given in the form P(z) ¼ P  pR2 pull

(7:4)

The bending moment at all locations, assuming for simplicity no lateral forces or angular acceleration, is M ¼ M0. Bending stresses due to the moment load M are calculated as though the tank were a slender, hollow beam of wall thickness t. The longitudinal stress has its maximum (highest tensile) value at one of the two points on the cross section farthest from the bending axis and its minimum (highest compressive) value at the other such point. The largest longitudinal compressive stress is P(z) M  2pRt pR2 t P M p R ¼  þ ull 2pRt pR2 t 2t

sz, comp (z) ¼ 

(7:5) (7:6)

and the largest longitudinal tensile stress is

sz, tens (z) ¼ 

P M p R þ 2 þ ull 2pRt pR t 2t

(7:7)

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The hoop stress is

su (z) ¼

[rz(a þ g) þ pull ]R t

(7:8)

The quantities 2M 2M and p (7:9) eq ¼ P  R R are called equivalent axial loads [40], and in terms of them the longitudinal stresses are þ ¼Pþ Peq

sz, comp (z) ¼ 

þ Peq

2pRt  Peq

sz, tens (z) ¼ 

2pRt

þ

pull R 2t

(7:10)

þ

pull R 2t

(7:11)

In the preceding equations, s represents the average stress over the wall thickness. Often, a local analysis that considers the variation of stress between the skin and the stringers or the core and facesheets of a built-up wall is needed. In such cases it is useful to work in terms of q, the integral of stress over the wall thickness: þ Peq

þ

pull R 2

(7:12)

þ

pull R 2

(7:13)

qu ; [rz(a þ g) þ pull ]R

(7:14)

q(z, comp) ;  q(z, tens) ; 

2pR  Peq

2pR

The quantity q is called the line load or tensile flux. Note that in the previous development, axial force is taken as positive in compression. These equations apply to large tanks and cylindrical adapters except where local irregularities or constraints render the underlying assumptions invalid. For a structure such as an adapter or interstage that contains no liquid, the terms containing density may be deleted. However, internal pressure in such structures may be important. Consider that an adapter with a radius of 100 in. and a wall thickness of 0.2 in. will experience a longitudinal wall stress of 0.25 ksi for every psi of internal pressure. From Eqs. (7.8), (7.10), and (7.11), we see that in the absence of external load and static head, the state of stress in the membrane is biaxial with a hoop-to-longitudinal ratio of two. External loads will cause this ratio to vary significantly from two. Conventionally, material strength is determined from uniaxial tensile tests, and then a combined-stress yield theory such as the Mises theory is used to calculate a scalar effective stress from the actual biaxial state of stress in the application. Although a large amount of experimental effort has been directed

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toward gaining a more sophisticated understanding of metal yielding and rupture under biaxial stresses (e.g., see Bert et al. [41]), the results seem to be little used today. The use of the maximum principal stress failure criterion for metals is near universal, but consider that a ductile material has a higher ultimate stress than its strength at rupture. In fact, for some high-strength steels, the stress is higher at the offset yield point than at any subsequent time [41]. Although maximum principal stress correlates very well to rupture strength, it is possible that ultimate stress, which is the material property customarily used to indicate failure, might be predicted better by alternative criteria. The foregoing discussion only addresses strength. Tanks may also fail by global or local buckling, or by the fracture of a flaw at far-field stresses below yield. In practice, the margin of safety tends to be about the same for strength and buckling failures. The fracture failure mode, which is managed by controlling the initial flaw size, may not be close to the others in criticality. Proof pressure testing is usually required, if not by the procuring agency, then by the range safety organization. Pressure testing at cryogenic temperatures is very expensive, so proof testing is usually done with room-temperature nitrogen gas or water. The ratio of yield to ultimate strength, and the fracture toughness, of many materials is different at room temperature than at the service temperature. Thus, it is not a trivial problem to devise a room-temperature proof test that exercises all failure modes of a cryogenic propellant tank adequately. Designing for light weight requires that the structure be quite thin walled. Thicknesses (or effective thicknesses, in the case of stiffened structure) can be on the order of one tenth of an inch for a section 200 in. in diameter (R/t ¼ 2000). For comparison, a soda-pop can has R/t  1000. Methods of flaw screening over large areas are usually sensitive enough to allow very small initial flaws to be assumed in the safe-life analysis and thus to provide ample safe life.} Automation of flaw screening can be developed during production planning. Years ago, flaw screening was provided via proof test; a flaw that could survive the proof test without catastrophic propagation was considered very likely to survive flight as well. This was usually performed on pressure vessels and pressurized structures. A more rigorous screen may (depending on the material) be provided by a proof test at cryogenic temperatures. For many materials, at colder temperatures the yield strength increases, permitting testing to a higher pressure, and the fracture toughness decreases, reducing the margin against catastrophic flaw growth. Methods of flaw detection include dye penetrants, ultrasound, x-ray, magnetic particles, and eddy current inspection. The inspection method is chosen based on cost, the required sensitivity, the accessibility of the area to be inspected, surface finish and coating/plating, and the material. MSFC-STD-1249 [42] is an oft-cited standard covering inspection methods. }

This discussion is provided by John Hilgendorf, structural analysis lead for Delta II, United Launch Alliance.

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Some materials have high fracture toughness relative to yield strength, so a larger flaw can be tolerated. The ratio of toughness to yield strength is significant due to the need to restrict stress levels below yield strength. Conversely, a highyield-strength material with low fracture toughness will need to be screened for very small flaws, which is the case with some high strength steels with low ductility. In some cases, when hardware is received, it is found to have been inadequately inspected, or the results of the inspection may show that the design intent was not met. It may prove faster and cheaper to conduct additional analysis, inspection, and testing to accept the discrepant hardware than to scrap the structure and manufacture a new one.

7.5.1 LEAK-BEFORE-BURST CRITERION Usually, pressure vessels and pressurized structures must satisfy the leak-beforeburst (LBB) criterion. The LBB concept is found in all industries that use pressure vessels, including aerospace, energy, and ground transportation. A definition given in a commonly cited military standard [28] is “a fracture mechanics concept in which it is shown that any initial flaw will grow through the wall of a pressure vessel and cause leakage rather than burst.” The purpose is to prevent catastrophic or explosive failures of pressure vessels or pressurized structures that may damage nearby flight hardware and launch facilities or injure personnel. In flight, a tank that has the LBB property may fail gradually enough that the mission can still be completed. It also provides time to depressurize or safe the system once a detectable leak has occurred. If all pressure vessels in a system are held to the LBB standard, safety rules and nearby systems need not be designed to withstand an explosive, catastrophic failure. This saves money. The LBB property may be verified by testing, analysis, or a combination of the two. A burst test that results in gradual leakage rather than sudden rupture is a demonstration of LBB. However, a test of a single article is of limited use unless it can be shown that an initial flaw not obviously detectable existed in a critical location. Analysis is necessary to determine the worst-case location and orientation. Flaws may be intentionally introduced into the test article to cause leakage to occur first at a location of interest. Analytically, to demonstrate LBB, it must be shown that the vessel can withstand the expected operating pressure when a leaking (through-wall) flaw exists. Said differently, a crack growth analysis must show that the critical flaw size is larger than the wall thickness. 7.5.2

STABLE METAL TANKS

Structurally stable metal tanks are the most common design. Historically, 2000series aluminum alloy has been by far the most popular material in this application, although recently, lighter aluminum–lithium (Al-Li) alloys have been

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used. For relatively slender tanks, the cylindrical tank barrel may be formed as a single ring if small enough, but more commonly it is built up from panels. The end domes are usually spun and may be of a different temper from the barrel. The barrel and domes are joined by welding. Squat tanks such as the S-II stage LOX tank have been laid out completely as domes welded from gores, with no cylindrical section. Large domes may be produced by explosive forming, as in the S-II stage [39]. Mynors and Zhang [43] discussed the widespread use of explosive forming in the 1970s, detailed the advantages and disadvantages, and described a research program exploring potential modern applications. Small end closures may be present at the apex of the domes, and these are usually bolted on so that they may be removed if necessary. Barrel panels are stiffened either with extruded stringers or with integrally machined stiffeners. The integrally machined designs demand that plate be available in fairly thick gauges (1 in. or thicker). Stiffeners may be created by machining or chemically milling pockets into a thick plate. The machining process leaves thickened weld lands, which are necessary because welds are not as strong as the as-machined metal. Machining of stiffeners is conducted when the panel is still flat, as a rule. Once machined, the panels are bump-formed or brake-formed into cylindrical arcs and then welded into a barrel of circular cross section. To avoid local buckling of ribs during forming, the machined pockets may be filled with a thermoplastic compound that is allowed to cool and harden before rounding the panel. The hardened compound provides stability to the thin ribs. After forming, the compound is melted out [44]. Because of the large amount of material that is removed, integrally machining the stiffeners may result in a scrap ratio of as much as 80%. This can be a significant cost for the more expensive alloys, and has been a motivation to attempt to produce Al-Li panels with extruded rather than machined stiffeners [45]. The isogrid pattern [46], in which the integral stiffeners are a network of equilateral triangles, is by far the most popular of the integrally stiffened tank wall designs. It offers the stiffness and mass efficiency of other stiffener patterns but preserves the large-scale isotropic behavior of the panels, so that they may be modeled as shells with “equivalent isotropic” properties. Although the simplifications made possible by isotropic behavior may not appear to be very advantageous in detailed stress analysis, when rapid iterations must be done in design trade studies, isotropic behavior is a significant benefit. Meyer et al. [46] provided the definitive work on isogrid design and stress analysis. The stiffening of tank walls by integrally machined stiffeners increases both the extensional and the bending stiffness of the walls. This improves the buckling (particularly the local buckling) resistance and the ability to withstand concentrated loads perpendicular to the shell surface at openings or attachments. The same principle is followed for structures other than tanks, where integral ribs or mechanically attached stringers, corrugation, or sandwich construction may be used. Propellant tank barrels and domes are invariably joined by welding, but welding is challenging in this application because of the relatively thin material

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and the tapered thicknesses that are used to save weight. Weld schedule development is time consuming, and external support is usually necessary to avoid distorting the shell due to the required heat input. Mendez and Eagar [47] provide an overview of the state of the art in aerospace welding technology; a more detailed discussion is presented in the section on manufacturing later in this chapter. The 2000 series of aluminum alloys has historically been the material of choice for stable tank designs and remains dominant, although in the last 10–15 years, the Al-Li alloys have also become significant. Chapter 2 covers aluminum alloys in detail. The 2000-series alloys are aluminum–copper alloys with the percent of copper varying from 0.9% to 6.3%. In these alloys, the intermetallic compound CuAl2 serves as the primary strengthening ingredient. Silicon and lithium are added to allow room-temperature age hardening, as well as improve the forgeability and strength. Trace amounts of manganese, magnesium, and titanium are present to refine the grain and inhibit stress corrosion [48]. Alloys for tank applications must be weldable, so that large barrels can be built up from smaller panels, and their strengths must be insensitive to notching at cryogenic temperatures. The 2000-series alloys were the highest-strength weldable alloys available for many years. Higher-strength alloys such as the 7000 series are available, but their poor weldability and cryogenic notch toughness relegates them to use in interstages, where they are assembled using fasteners and not subject to extremely low temperatures [49]. A very popular tank material is alloy 2219, a high-strength, weldable aluminum alloy whose principal alloying element is copper (6.3%) [50]. It has been the primary tank structural material in the Saturn S-IC stage [49], and the standard-weight and lightweight (LWT) Space Shuttle External Tank designs [51]. Alloy 2219 is a wrought, heat-treatable, precipitation-hardening alloy developed by Alcoa in 1954 for high-temperature structural applications [50]. However, its excellent properties at cryogenic temperatures are what makes it attractive for LV tanks. Its full strength is developed by solution heat treatment followed by aging. Cold work may be applied before aging to further enhance the precipitation-hardening process. Reheat of clad grades (not commonly used in LVs) may reduce the alloy’s resistance to stress corrosion. The most widely used temper of 2219 in LV tankage is T87. In this grade, in-plane A-basis ultimate tensile strengths are 63–64 ksi, with B-basis strengths only about 1 ksi lower, indicating very good control of strength variability. Yield strengths are around 51 ksi. Elongation to break is 6%–7% for the thinner gauges of plate. As with all aluminum alloys, the elastic modulus is around 10.5  106 psi, one-third that of steel, so significant springback often occurs in cold-formed parts. Very thick alloy 2219 shapes have lower yield and ultimate strengths than thinner ones. Thickness at the time of solution heat treatment, not the final machined thickness, should be taken into account when establishing design allowables. The tensile strength of aluminum alloys is increased by cryogenic temperatures. For example, at LOX temperature (22978F), 2219-T87’s ultimate and

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yield strengths are 20% higher than at room temperature. At LH2 temperature (24238F), the strengths are more than 30% higher. This increase in strength is frequently taken into account in design margin calculations. However, large, thinwalled tanks may buckle at a lower compressive load than that necessary to cause a failure in strength. In such cases, it is the cryogenic elastic modulus, not the cryogenic strength, that determines the compression capability of the tank. The increase in modulus is not as impressive as the increase in strength; for 2219-T87, it is only about 10% at LH2 temperature [8]. One problem associated with the use of alloy 2219 has been the difficulty of chem-milling in the T3 temper. This problem was encountered with the hydraulic bulge-formed and chem-milled dome gores of the SIC, and ultimately led the designers of the Shuttle External Tank to abandon chem-milling and adopt the more capital-intensive, but easier to control, stretch forming process [51]. Alloy 2219 is also subject to surface corrosion, especially in the clad grades. It was found that foam insulation on a 2219 substrate resulted in collection of a chloride-rich liquid in the salt air environment of the southern United States, which caused extensive corrosion after exposure of many months [51]. The other workhorse aluminum alloy for stable tank designs is 2014. Alloy 2014 has copper as a principal alloying element (4.4%) but at a lower level than 2219 (6.3%) [52]. It was developed in 1928 primarily for use in aircraft structures as forgings and extrusions; for LV tanks, the sheet or plate forms are used. Alloy 2014 generally has higher strength than 2219: in the T6 temper, its A-basis tensile strength is 64–67 ksi, a few percent stronger than 2219 [8]. Alloy 2014 is a precipitation-hardening alloy. Unlike the widely used 2219-T87 grade, commercial tempers of 2014 are not cold worked. As with 2219, considerable springback may occur after cold forming, and this is typically corrected by “overforming” [52]. Both 2219 and 2014 are easily machinable, which is important in designs with integrally machined stiffeners. Alloy 2014 has been used in the Titan II booster, the Saturn S-II stage, and the Saturn S-IVB stage [49]. The Saturn I, designed in the late 1950s, used the Al-Mg alloys 5456 and 5083, but these are rarely considered now due to their lower cold notch toughness and greater susceptibility to corrosion. However, they are more weldable than the 2000-series alloys. That is, they lose proportionately less strength and ductility in the welded condition [49]. Both 5456 and 2014 appear to have been early candidates for the S-IC stage [53], but 2219 was ultimately selected. Another aluminum alloy, 6061, was used on the Agena tanks [54]; although this alloy still has some applications in other vehicle structures, it is no longer used for tanks. Welding processes for tanks have an influence on materials selection. Historically, most tanks have been fusion welded. The S-IC stage used gas tungsten-arc welding (GTAW) to join 2219 panels [51], a practice that continues to be popular. More recently, plasma arc welding has been implemented. Variable-polarity plasma arc (VPPA) welding, in which the arc polarity is periodically changed to reduce the accumulation of dross, was successfully implemented on the Shuttle

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ET and Delta IV programs and has also been used to join Al-Li alloys [55]. The large Soviet/Russian Energia booster used electron-beam welding in its tanks [2]. Within the last 20 years, the development of friction stir welding (FSW) has been a major advance in tank manufacturing. FSW was first developed in the 1990s and is now used in production on several LVs. In this process, a rapidly rotating pin moves along the weld lands, mixing clean base metal, which welds spontaneously. It produces a higher-strength and higher-ductility joint than fusion welding because the material is never melted [56]. FSW is particularly attractive for aluminum alloys because of their low hardness. FSW was introduced into production on the Delta II program in 1997 [55]. But FSW is more sensitive to weld land alignment deviations than fusion welding. Aside from 2000-series aluminum alloys, the material with the widest current application to propellant tanks is the Al-Li series of alloys. These alloys contain only a small amount of Li by weight (about 1%), less than their Cu content of 2%–4%, but they are known as Al-Li alloys to contrast them with non-Li-containing alloys. An Al-Li alloy was developed specifically for aerospace applications as early as the 1950s, but problems with fatigue, fracture, and weldability precluded its widespread use in the United States until the 1990s [51]. Although all wrought alloys are anisotropic in strength and stiffness to some degree, Al-Li is anisotropic enough that it must be structurally analyzed as such. One study found that 2195 Al-Li extrusions [57] had direct and off-axis strengths differing by as much as 20%, depending on the depth through the section. In the early 1990s, funding became available for a major redesign of the Shuttle ET with the primary goal of reducing weight. Weight reduction became necessary when it was decided that the International Space Station (ISS) would be put into a high-inclination orbit accessible to Russian launchers; the Shuttle then had to reduce its empty weight to be able to reach the ISS. A series of weldable Al-Li alloys under the Weldalite trade name was available to Lockheed Martin, prime contractor for the ET. The redesigned tank was given the abbreviation SLWT, for super-lightweight tank. The Al-Li alloy 2195 ultimately selected for parts of the SLWT is lighter than the formerly used alloy 2219, but has yield strength about 20% higher at both ambient and cryogenic temperatures [12]. It is also about 8% stiffer than 2219. However, 2195 is less formable in the T3 condition than 2219, so an early attempt to simply drop it in as a replacement for 2219 resulted in damaged forming equipment. The remedy was to solution treat and quench the 2195 into the T0 condition, then stretch form and shape to the T3 condition, and finally age to the T8 condition [51]. Alloy 2195 is also less ductile than the 2000-series aluminum alloys. Ultimately, all of the ET tank barrels as well as the intertank thrust panels were changed to Al-Li. It was also found that fusion welds on Al-Li were more susceptible to hot cracking than on 2219 and that the subsequent repairs were more difficult. Process changes involving a smaller heat load, a backside inert gas purge, and weld bead planishing were necessary to enable the needed repairs [51].

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However, weld quality concerns led the Marshall Space Flight Center to investigate FSW for the Al-Li tank components. FSW was implemented on the ET starting in 2002. Other applications of Al-Li have been the DC-XA and X-33 research vehicles. In both cases, composites were used for the LH2 tanks, but Al-Li was used for the LOX tanks. Composite LOX tanks require a protective liner, typically a halogenated polymer, to reduce the chance of ignition [58]. The DC-XA LOX tank was built in Russia from Al-Li alloy 1460 [59]. The Ares I upper stage was a structurally stable, common-bulkhead propellant tank design with friction stir-welded Al-Li 2195 tank barrels and domes. The common bulkhead was a sandwich construction consisting of 2014 facesheets enclosing a phenolic honeycomb core. The bulkhead was to be joined to the barrels by a 2219 Y-ring [60]. In a pump-fed stage, the propellant is held under low pressure in the tanks, then pumped to the injection pressure after it has left the tank. The tanks therefore may be constructed lightly, and stresses due to external flight loading are comparable to those due to internal pressure. In contrast, pressure-fed stages do not have pumps; the propellants are forced into the engine by holding them under high pressure in the tanks. This type of design is used when simplicity and reliability are paramount. Injection pressures for pump-fed engines may be several thousand psi, which would require inordinately heavy tankage. But pressure-fed systems are designed to require only moderate injection pressures. Although this is higher than the tank pressure in a pump-fed stage, it is low enough that the tank can be flight worthy at an acceptable weight. Stresses in tanks for pressure-fed stages are dominated by internal pressure loads. Some pressure-fed designs have used internal bladders to expel propellant from the tank rather than externally supplied gas. Many basic design and materials selection aspects are discussed in Wagner [54]. Pope and Penner [61] described testing of multilayered bladder materials consisting of various arrangements of polyethylene terephthalate (PET) film, composite balloon film, aramid film, and polyimide film. They found through subscale testing that a PET-balloon film fabric provided good performance under cryogenic conditions, with the lowest permeability. Gleich and L’Hommedieu [62] performed similar studies on wire-reinforced metallic bladders of annealed austenitic stainless steel. Calabro et al. [63], in the course of system studies for an advanced pressure-fed cryogenic upper stage, proposed combining a 2219 aluminum LOX tank with a filament-wound graphite-epoxy LH2 tank in a common-bulkhead design. The LH2 tank used an internal aluminum foil liner. The working pressure was 270 psi. Thermal insulation was provided by externally applied polyurethane foam. Many LVs use hydrogen as a propellant in the booster, the upper stages, or both, so the compatibility of materials with hydrogen must be thoroughly understood. Cataldo [64] summarized the findings of several research programs investigating hydrogen embrittlement in high- pressure storage tanks, fasteners, and weldments. Although the focus was on titanium alloys and Inconel 718, useful

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information is provided on a wide variety of aerospace metals. High pressure was not always a necessary condition for problems with hydrogen compatibility. Hydrogen embrittlement of metallic materials is discussed in Chapter 2.

7.5.3

BALLOON TANKS

The Atlas vehicle designed by K. E. Bossart at Convair Division of General Dynamics in the early 1950s is exemplary of this type of design. The other notable application is the Centaur upper stage, also developed by General Dynamics. The Atlas maintained the balloon tank design through several intercontinental ballistic missile (ICBM) variants, the early Atlas E and F space LVs, and the Atlas I, II, and III commercial space launchers (Fig. 7.4). The Centaur stage still uses the balloon tank design. Balloon tanks require either mechanical tension (“stretch”) or internal pressure to keep them from collapsing under their own weight before operation. In operation, the pressure required for propellant feed is sufficient to keep the tank stable under flight loads. The following information is taken primarily from the review by Martin [65]. Balloon tanks have very thin walls (as thin as 0.01 in., thinner than three sheets of copier paper) and are built from corrosion-resistant steel. In the Atlas and Centaur, most of the tank skins are made from stainless steel alloy 301 in the extra full-hard (EFH) grade. Skins that must be formed into a shape other than a circular cylinder, such as conical transitions or domes, are made from 1/2 and 3/4 hard grades, for improved formability. Because the tank walls are so thin, machined Fig. 7.4 Atlas launch vehicle carrying John Glenn to orbit. The balloon propellant tanks can be seen; the LOX tank is forward and covered with frost, whereas the fuel tank is aft and its shiny stainless-steel skin is clearly visible. (Public domain photograph by NASA, 1962.)

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TABLE 7.3 Alloy

COMPARISON OF ALLOYS 2219-T87 VS 301 EFH A-basis Yield, ksi

Density, lb/in3

Specific Strength, ksi/(lb/in3)

2219-T87

60

0.103

583

301 EFH

200

0.286

699

reinforcing rings must be placed at locations where external hardware such as feedlines, electrical tunnels, or strap-on booster rockets must be attached. These rings are made from 321 stainless steel, because it is more machinable than 301. Both 301 and 321 are austenitic stainless steels, whose primary alloying elements are chromium and nickel. In the very early phase of ICBM development, a vehicle was designed using the balloon tank concept but with aluminum instead of steel as the material. However, comparing the specific strength of 2219 aluminum and 301 EFH stainless steel at LOX temperature, as in Table 7.3, 301 stainless steel offers an advantage, especially in the EFH condition and at cryogenic temperatures. Also, aerodynamic heating of the skin must be considered. The Atlas missile was designed as an ICBM and had to be able to withstand a depressed trajectory that resulted in skin temperatures as high as 7008F. At this temperature, the stainless steel loses only 17% of its room-temperature strength, whereas the aluminum loses more than 80% of its strength. Aluminum could only be used if it were highly insulated, at an inert mass penalty. The 10 foot diameter Atlas balloon tank barrels were constructed from stubby bands 32 in. high. The bands were “stovepiped” together (i.e., inserted into one another a short distance), resistance seam-welded, and then spot-welded on both sides of the seam weld for added strength. The longitudinal welds in the bands and dome gores were resistance butt-welded, and then a doubler was applied with several rows of spot welds. No filler material was used in the resistance welds, although it was found that placing nickel foil between the workpieces produced stronger spot welds. 7.5.4

COMPOSITE TANKS

Although light weight is always a major goal in the design of aerospace structures, it is especially important in LV stages that ultimately will be propelled to orbit. In staged vehicles, the inert weight of boosters is jettisoned once the booster’s fuel supply is exhausted. However, the orbital stage is not jettisoned, so there is a very high motivation to keep its inert mass to the absolute minimum. Every pound of inert mass on an orbited stage is one less pound of payload that can be carried. The vision of a reusable, single-stage-to-orbit (SSTO) vehicle with airliner-like operations has existed since the earliest speculations about space travel. Such a vehicle would have no jettisonable boosters, with all of its inert mass propelled

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into orbit, reentering the atmosphere, and returning to Earth to land. Therefore, structural mass efficiency is paramount. Barring unforeseen developments in propulsion technology, any SSTO vehicle must have a structure that is at the absolute maximum efficiency possible with known materials. The imperative to minimize inert mass has been one of the major reasons so much research effort has been directed toward composites, the other major reason being the ability to fabricate complex cross-sectional shapes with inexpensive tooling and processes. The tensile strength-to-weight ratio of graphite fibers is many times that of the aluminum alloys and steels typically used in propellant tanks. But the raw tensile strength-to-weight value that is so favorable for graphite fibers can be misleading. To produce a useful structure, the fibers must be incorporated into a matrix; this decreases the tensile strength by about 50% and adds the weight of the matrix, which carries little load. Also, unlike a true pressure vessel, the skin of a pressurized structure will not always be in tension. Compression loading raises the possibility of buckling. Although composites with elastic moduli several times that of an equivalent-weight metal design may be produced, it is difficult to control the geometric imperfections that are so damaging to buckling resistance. The polymeric matrix of conventional composites places an upper limit on the service temperature. Conventional graphite-epoxy composites lose strength and stiffness rapidly when temperatures reach 2008F to 3008F, due to softening of the matrix. Thus, composite tanks must be insulated or protected from skin heating by trajectory limitations. This is especially constraining to the design when the trajectory includes reentry, as it does for a reusable vehicle. Improvements in both thermoplastic and thermoset matrix materials are potentially a means of raising the temperature limit. Also, especially for tanks of complex shape, reinforced joints are necessary. The need to reinforce these joints and to insulate a composite tank against aerodynamic heating tends to erode the weight advantage over a metal tank. It has been stated that a composite tank can represent a 20%–40% weight saving compared with an equivalent metal tank [66, 67]. In the specific case of the DC-XA, NASA claimed in a press release that the composite LH2 tank was 37% lighter than the metal tank used on its predecessor, the DC-X [68]. The vast majority of the composite experience base has been with laminates; that is, panels built up from several layers of material manufactured in a previous process. The challenges of joining laminated panels, and their poor interlaminar strength, has led to an interest in braided, woven, and knitted textile preforms manufactured by resin transfer molding (RTM) and resin film infusion [69]. These preforms offer a way to join laminated panels without a subsequent bonding process or discrete fasteners. They can also be used to fabricate braces and bulkheads that are not panel-like in geometry. As stated in the previous section, composites were first applied to LH2 tankage. Since that time, composite tanks compatible with LOX have been developed, but a

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protective liner separating the composite walls from the LOX was necessary to reduce the potential for ignition [58]. LOX may also chemically degrade the matrix through oxidation. The National Aero-Space Plane (NASP) or X-30, a SSTO system contemplated in the 1980s, sought to use composite liquid hydrogen tanks. Hartunian [70] recounted something that often occurs in high-risk developments: despite plans laid by knowledgeable people, significant technical challenges do not come to light until some work is actually done. In the case of the NASP tank, scaling up the concept from laboratory scale to production scale introduced some difficulties. The IM7/PEEK composite initially identified as the one with the best resistance to microcracking could not be scaled up. The cure temperature and pressure, and the required cooling rates, could be achieved at small scale but not at production scale. After the failure to cure the production-scale tank, the engineers changed the PEEK matrix to 8551-7A epoxy. The epoxy matrix design was successfully fill/drain cycled, but the program was canceled for other reasons. Two more recent programs intended to advance the state of the SSTO art were the DC-XA and the X-33 suborbital technology demonstrators. These programs used composite cryogenic propellant tanks. The DC-XA vehicle flew twice, with the composite tank performing satisfactorily [71], whereas the X-33 never flew, largely due to development difficulties with its composite LH2 propellant tank, including a major test failure. Most of the interest in composites for propellant tank applications has centered around graphite-epoxy. Both the DC-XA and the X-33 used graphite-epoxy tanks, and the DC-XA also used a composite LH2 feedline. In addition, the DC-XA used a composite intertank structure. The composite structures on the DC-XA were developed with the aid of rapid prototyping methods [69]. The X-33 tank was a sandwich design with graphite-epoxy facesheets and an aramid-reinforced phenolic honeycomb core. The core contained empty spaces that were not vented. The X-33’s development difficulties and 1999 test failure have strongly influenced research in the field since that time. An overview of that design and failure is now presented as a way to introduce the key materials and structures issues involved in composite cryogenic tanks. Aerodynamics forced the X-33 tank to be structurally much more complex than typical LV tankage. It consisted of a lobed outer barrel constructed from composite sandwich and monolithic composite internal stiffening frames (Fig. 7.5). In addition, bulkheads and thrust tubes were attached to support primary structural, landing gear, and control surface loads. The X-33 tank could almost be considered a composite fuselage filled with LH2. The X-33 tank was in the process of being qualified in a protoflight program. This entails testing the actual flight article to load levels higher than the maximum expected flight loads, but not as high as a single dedicated test article would be subjected to. The tank had been cryogenically cycled three times, subjected to proof pressure while filled with LH2, and then subjected to one external test

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Fig. 7.5 X-33 liquid hydrogen tank on a test stand at NASA Plum Brook Station, Sandusky, Ohio. Note the complex, lobed shape of the tank. (Public domain photograph by NASA photo C-1998-1295, 1998.) load case while filled. A few minutes after the tank had been drained, it suffered a catastrophic delamination. It was found that cold gaseous hydrogen had entered the sandwich core from the inner volume of the tank by permeating the inner facesheet. At the same time, ambient nitrogen gas was drawn into the core through the outer facesheet. The permeation processes were abetted due to the strain induced by the test pressure and loads, combined with the low temperatures, which caused leak paths to develop. As the tank cooled to LH2 temperature as it was filled, the trapped gases condensed into liquids, creating a partial vacuum that drew additional gases into the core. Upon draining, the tank began to warm to room temperature, and the pressure in the core rose as the liquefied hydrogen and nitrogen warmed up and began to evaporate. The pressure resulted in a sudden debond of the entire area of the inner facesheet. A preexisting bondline flaw, in the form of a piece of slippery tape found between the core and facesheet, probably contributed to the failure. This failure mode is called cryopumping. Generally, in the context of aerospace structures, cryopumping refers to the condensation of gas in a void and the drawing in and condensation of additional gas due to the lowered pressure in the void, followed by the possibly destructive rapid venting of the gas upon reheating. In cellular insulations such as polymeric foam, cryopumping occurs when the insulation is cooled by contact with a tank filled with cryogenic propellant, then heated as the vehicle ascends, the tank empties, and aerodynamic friction heats the insulation. Liquid air condensed in voids in the foam is vaporized and will blow a hole in the foam if it cannot gradually vent. Cryopumping was a known condition that the X-33 design was supposed to accommodate, and the failed core in fact had a measured cryopumping pressure

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that was lower than the design value, but local, unobservable peaks in the pressure may have exceeded the bondline capability. Despite ultrasonic nondestructive inspection, the PTFE tape, as well as other debonded areas, were not detected before testing. They were only observed after the test article had failed. The possibility of manufacturing flaws difficult to screen by inspection or proof testing has always been a disadvantage of composites, especially in sandwich constructions. It is a mistake to conclude that the X-33 failure proves composite tanks can never work, because that application was much more demanding than conventional applications. It is known that thermomechanical cycling, which is much more severe on a reusable vehicle like the X-33 than on an expendable, is the primary driver of permeation and leaking. After all, composite filament-wound, monocoque solid rocket motor cases have been successfully used for years, and mechanically they are similar to liquid propellant tanks. However, composites are not as clean a solution as they might appear to be from a naive conception of their raw material properties. In particular, the need to characterize and control permeability without the use of a liner has been the thrust of much recent research in composite tanks. During and after the X-33 program, several research projects have sought to improve the performance of composites in cryotank applications. Heydenreich [72] described system studies carried on in Europe to establish which tankage applications could most benefit from the use of composites. He pointed out the need for a mechanically strong yet thermally insulating design, suggesting that a liner would be necessary to prevent permeation. He also recognized the fact that composites do not exhibit plastic behavior, which requires a different design philosophy than for metal tanks. Sankar et al. [66] conducted a multiyear research program aimed at developing improved analytical models of gas permeation through composite panels at cryogenic temperatures and under complex, fluctuating stress states. In particular, they examined the effect of interacting distributions of oriented cracks in the different layers of a laminate. Transverse microcracking due primarily to thermal stress is known to contribute to permeation. A fracture-mechanicsbased approach was used to predict crack densities and permeation rates. They also performed testing that showed cryogenic cycling caused a degradation in the resistance of panels to permeation due to the opening and propagating of cracks. The testing showed that textile (woven) composites had less permeation than laminated composites after cycling; this was attributed to the lack of propagation of transverse cracks. Morino et al. [67] carried out preliminary tests using a subscale tank with a liner, focusing on the Y-joint at the dome-barrel intersection. They noted the difficulty of maintaining a quality laminate in such locations and aimed their testing at this area. They observed matrix microcracking at low stress levels when the matrix was cold. Graf et al. [73], noting the need for leakproof adhesively bonded joints in cryotank applications, tested a double-lap joint design. They showed that the lack of a

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peel-ply surface preparation, as well as the use of an adhesive primer, reduced the bond strength. Overall, they found, as in other investigations, that cryogenic temperatures reduced the strength of their components by 50% or more. They showed a size effect; that is, the larger the bonded surface, the lower the supported shear stress. Such effects are usually attributable to the greater likelihood of bondline defects as the bonded area increases. Miller and Meador [74] found that clay-based layered silicate nanocomposites, dispersed in the epoxy matrix, significantly reduced thermal expansion and gas permeability in the resin both before and after cryogenic cycling. The degree of reduction was directly related to the weight percent of nanocomposite. They also found that, although the nanocomposite matrix led to a laminate with lower flexural strength than plain epoxy resin, the nanocomposite retained its strength after thermal cycling. It appears that after cycling, the nanocomposite laminate had strength comparable to the plain resin laminate. However, these encouraging results did not translate to decreased permeability when the nanocomposite matrix was used in a subscale test bottle. Pavlick et al. [75] investigated the strength of advanced matrix materials. The resins were tested in the form of tensile and fracture samples machined from neat plaques. Tensile strength, modulus, elongation to break, toughness, and fracture properties were measured at temperatures ranging from þ3208F to 23108F. It was found that cryogenic temperatures tended to increase the strength and decrease the elongation to break of the matrix materials. Trends in fracture properties were unclear. A candidate liquid crystal polymer matrix material was found to be generally more brittle and less tough than the three other resins, all polyimides. Black [76] discussed recent advances in research on composite tanks for cryogenic fluids. An unlined composite LOX tank for the since-canceled X-34 reusable vehicle was successfully tested for fill/drain cycling and impact resistance. The ability of composite tanks to incorporate more complex shapes than those of metal tanks has been enhanced by in situ fiber placement, which can produce thick, curved structures that do not wrinkle during cure, and can eliminate the need for debulking. Another new manufacturing method that eliminates the need for debulking is to lay up a panel by ultrasonically bonding thin layers of prepreg tape. Linerless tanks may be possible if toughened, advanced matrix materials are used. Even composite tanks still must use heavy metal bosses for fluid connections. However, composite bosses manufactured by RTM have been tested. 7.5.5

SOLID ROCKET MOTOR CASES

Large solid rocket motor cases are discussed in this chapter because of the significant flight loads (in addition to self-generated internal pressure) they carry. Although they usually are “strapped on” and therefore are not in the primary load path, in one vehicle, Ares I, a motor case did form the bulk of the booster primary structure. Solid motors also provide primary structural support in

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solid-fueled missiles. Because of their size and rigidity, solid motor cases are attractive locations for the attachment of auxiliary flight systems, and they also must support strap-on booster nose cones and aft fairings. In this respect, they have more in common with liquid main propellant tanks than with the combustion chambers of liquid rocket engines. However, they must withstand pressures that can exceed 1000 psi, far higher than the pressures in propellant tanks. A solid motor case is composed of a barrel section, a forward dome and closure, and an aft dome with provisions for mounting a nozzle. Smaller motors such as the GEM-40, -46, and -60 strap-ons for the Delta II, III, and IV, and the Atlas V solid rocket motor can be produced as a single, monolithic unit. Very large motors, including the Titan III and IV strap-ons and the Space Shuttle Solid Rocket Boosters, must be manufactured in segments in order to be transportable over the road. Motor cases and segments are permanently loaded with propellant by the manufacturer, and therefore must be handled carefully as they are transported to the launch site, where they are assembled or “stacked.” A “case-bonded” (as opposed to cartridge-loaded) motor typical of those used for LVs consists of an outer shell closed forward and aft by domes, and the assembled pressure vessel is lined with an insulating material that both protects the case from the heat of combustion and facilitates the bonding of the propellant to the case. The propellant is then cast directly into the lined case and cures to a rubbery consistency. Neither the propellant nor the insulation provides significant strength or stiffness to the motor as a whole, so they are not discussed further here. Additional details are given in Chapter 11. The pressurized envelope of a motor case is capped by a forward closure, which usually houses the igniter, and an aft closure that must provide an attachment for the nozzle. Also, forward and aft skirts are usually provided for attachment to other vehicle structures. These are integral with the motor case. Except for the very largest first-stage boosters, solid rocket motor cases are designed based on the pressure stress plus flight loads amounting to some fraction of the pressure stress. As with main liquid propellant tanks, cyclic loading during proof testing may cause flaws to propagate. But solid rocket motor cases are also subject to pressure oscillations at frequencies up to 1000 Hz during the motor burn [77]. Therefore, nondestructive inspection methods of similar type and significance as those previously discussed for liquid propellant tanks also apply to solid rocket motor cases. Motor cases are generally constructed of high-strength steels, titanium, or filament-wound graphite epoxy. Pressure stresses usually preclude the use of aluminum except for very small motors. Metal cases may be built from rolled and welded sheet or by seamless methods such as drawing or spinning. The presence of a welded seam lowers the strength of the nearby material and requires heat treatment and careful inspection. Steels that are commonly used are D6AC, the 18% nickel maraging steels, and 4130 alloy [77]. Steels requiring postfabrication heat treatment may pose a problem because of the very large diameter of the

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finished product. There is a limit to how large the structure can be before exceeding the capacity of commonly available heat treatment facilities. Solid rocket motor cases were one of the earliest applications of filamentwound composites technology. Peters [78] states that motor cases “were primarily responsible for accelerating filament winding from a laboratory curiosity to the major industry it is today.” As with propellant tanks, a major reason composites are attractive as a material for motor cases is the ability to orient the strong direction of the material along the direction of highest loading. This leads to greater structural efficiency than is possible with an isotropic material. In motor cases, more so than other structures, it can be stated with high confidence that the state of stress is close to biaxial, with the axial stress about half of the hoop stress. Flight loading is small compared with internal pressure and will not significantly alter this ratio. The titanium alloys and high-strength steels commonly used for motor cases have specific strengths of about 850 ksi/(lb/in3), whereas composites can achieve three to five times this value. Other reasons to use composites include lower-cost and more adaptable tooling, relatively low-cost raw materials, and imperviousness to corrosion. The thermal environment for motor cases is not significantly different from that of noncryogenic primary structure. Although combustion temperatures are as much as 4000 K, this extreme temperature does not have time to penetrate through the very poorly conducting solid propellant and insulation to the case. Several programs, including Titan and Space Shuttle, have developed composite filament-wound replacements for motor cases that were initially metal. Not all of these new designs were put into production. In the case of the Space Shuttle, the filament-wound motor offered a definite mass fraction advantage over the existing design, but the extra capability was only needed for polar orbit launches from the Western Range, which were canceled after the Challenger failure [79]. The Delta II uses up to nine large strap-on GEM-40 solid rocket motors. The GEM-40, -46, and -60 have graphite-epoxy filament-wound cases. Filamentwound cases have even been able to meet the very stringent mass efficiency requirements of upper stages. The inertial upper stage (IUS) developed as an upper stage for both Titan and Shuttle, incorporated two aramid-epoxy filament-wound motors. Gargiulo et al. [80] compared failure envelopes generated by several commonly used composite failure criteria to test data for pressurized filament-wound tubes. Bert and Hyler [81] and Alexander and Fournier [82] provided two early studies of materials selection for solid rocket motor cases. Pionke and Garland [83] compared D6AC and 18-Ni maraging steel from the standpoint of subcritical crack growth behavior in motor case applications. This research was conducted in the course of early Space Shuttle system studies. They found that D6AC had inferior corrosion and stress-corrosion resistance and also experienced a decrease in cycle life when exposed to temperatures needed during refurbishment operations.

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7.6 FEEDLINES, SMALL LINES, AND PRESSURE VESSELS Many tubes and pipes are necessary to supply fluids to the propulsion and guidance systems. These components range from small tubes less than an inch in diameter to main propellant feedlines, which can be 18 in. or more in diameter. The larger lines frequently must have gimbals or flexible sections so that thermal and mechanical stresses do not build up, especially where the lines connect a strap-on to a main booster that may experience large relative motions. Also, lines may connect to the inlet valve on a gimballed engine that undergoes large motions.

7.6.1 FEEDLINES Feedlines are different from other pipes and tubes due to their large size, higher criticality, and high flow rates. Operating pressures are similar to those in the tanks. Some lines are downstream of pumps and the pressure can be several thousand psi, but pipes downstream of feed pumps are usually considered part of the propulsion system and therefore fall outside the scope of this chapter. Either the fuel or the oxidizer tank may be in the forward position. The feedline from the forward tank has a downcomer that may run along the side of the aft tank or may penetrate the tank. The downcomer can be more than 50 ft long. Feedlines are usually constructed of 321 corrosion-resistant steel (CRES), although 347 CRES, Inconel 718, Hastelloy, and A-286 have also been used [84]. Inconel 718 and Hastelloy are especially suited to areas experiencing fluctuating loads and corrosive environments. Feedlines can experience a high fluctuating load component relative to the mean load because of dynamic excitation and flow-induced vibration. They may also vibrate during pogo, which is an undesirable resonant interaction between the motion and pressure of the fluid and the structural modes of the feedlines or adjacent hardware. The DC-XA included a composite LH2 feedline among the technologies it demonstrated [85]. Metals for feedlines must have high ductility because of the need to form elbows and bends. They must be formable, weldable, and compatible with common lubricants. They must also have adequate performance at low temperatures, when cryogenic fluids are involved. They must be chemically compatible with the working fluid. A particular problem is hydrogen embrittlement (see Chapter 2); Inconel 718 is incompatible with high-pressure hydrogen for this reason. A corrosive or chemically active environment can significantly lower the fracture toughness of materials. Also, some fuels undergo rapid or even explosive reactions when they contact certain metals. For example, the breakdown of certain hypergolic propellants is catalyzed by some of the trace alloying elements present in many metals. Cryogenic lines may require insulation, whether they are inside or outside the vehicle shell. Insulation is required to minimize boil-off, maintain the fluid within the required temperature and pressure, and prevent geysering. Geysering occurs

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when the fluid in a vertically oriented line partly vaporizes, and the bubbles rise and rapidly exit the top of the line. Insulation on feedlines uses much the same technology as the lightweight thermal protection systems for vehicle primary structure. Both large feedlines and smaller tubes may be subject to safety factors and testing requirements that are quite different from primary vehicle structure. Lines that are small and can be pressurized when personnel are nearby may be held to safety factors as high as 4.0. When EWR-127 applies, many safety precautions are required. Proof pressure testing is almost always mandatory, and the many system functional and leak checks that are carried out can consume a significant portion of the safe life of a small line.

7.6.2 PRESSURE VESSELS LVs need to store small quantities of hydraulic fluid, secondary propulsion or reaction control propellants, helium for system pressurization, and the like. High pressures may need to be withstood. The classic design for this application is a Ti-6Al-4V welded sphere. A more mass-efficient design, widely used today, is the COPV. This design uses a very thin metal shell only as a leak liner. The membrane strength is provided chiefly by a filament-wound composite layer on top of the metal liner, usually graphite epoxy or aramid epoxy. The liners may be titanium alloys or Inconel. The two-layer construction allows the liner to be placed in a state of residual compression by initially pressurizing the tank beyond the yield point of the liner. This process is called autofrettage or sizing. When the pressure is removed, the overwrap elastically recovers, imposing a compressive stress on the liner. In subsequent pressure cycles, the liner will not go into tension until the sizing pressure is exceeded. This process greatly improves the pressure and fatigue capability of the liner. Obviously, if autofrettage is to be done, the material selected for the liner must have a stress-strain relation that permits it. Low variability in the yield strength and draw properties is needed in order to keep the results of the autofrettage operation within control. The inspection and safe-life analysis of COPVs have been extensively studied, and specialized standards exist [86–88]. However, with the liner strongly compressed when the vessel is empty, liner buckling must be prevented. A good, continuous bond of the liner to the overwrap is necessary. Unbonded areas due to inadequate adhesion or protruding weld beads on the liner can cause the liner to buckle. The LBB requirement is not entirely straightforward to apply to COPVs because of the separate liner and overwrap.

7.7 UNPRESSURIZED STRUCTURE Here, unpressurized structure means passively vented structure that experiences low pressure differentials, no more than a few psi. For these structures, pressure

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is not a driving factor in design. Examples are fairings, nose cones, skirts, adapters, thrust structures, wings, and control surfaces. Usually, at launch, a mixture of gases, primarily air, exists at near-atmospheric pressure in the interior spaces of these structures. These gases may be very cold if near a cryogenic tank and may contain gaseous propellants or oxygen due to prelaunch venting operations. After launch, as the vehicle ascends through the atmosphere, the internal gases escape the structure through vents or natural leak paths. Unpressurized structures may need to maintain a controlled interior temperature and humidity environment, as with a payload fairing, or there may be no control of the interior environment, as is usually the case with intertanks and thrust sections. As with all airborne structure, the strength-to-weight ratio is the most important design characteristic, and when liquid propellants need not be contained, there is more freedom to optimize the materials and structure for light weight. Therefore, unpressurized structures have seen greater use of composites, and the stronger grades of aluminum, whose lower fracture toughness is less of a disadvantage than it would be in structure that sees pressure cycling, may be considered. Lighter designs can result. The 7000-series aluminum alloys are often used in unpressurized structure. These alloys have zinc as their major alloying element, and have a much higher static strength than the 2000-series alloys used in propellant tanks. However, the 7000-series alloys are not as resistant to damage from repeated loading as the 2000-series alloys and have less favorable cryogenic properties.

7.7.1 INTERTANKS, SKIRTS, ADAPTERS, AND MORE A space LV is, functionally, a number of tanks connected in series, with an engine at the aft end and a payload at the forward end. The structures used to connect the primary functional pieces are known variously as intertanks, interstages, engine sections, skirts, and adapters. The generic term adapter will be used to refer to any of these types of structures. Adapters may be simple cylindrical shells providing a space for the end dome of a tank, or they may support feedlines, pneumatic and hydraulic lines, wire harnesses, and other items on internal brackets or shelves. Often the umbilical connections that supply the vehicle with ground electrical power and provide propellant fill and drain capabilities are located in adapter structures. Because of the available internal space, guidance and navigation hardware, telemetry equipment, inert gas tanks, and hydraulic pumps are often located in adapters. Thus, an adapter may have an outer shell that is primary structure and inner shelves or brackets that are secondary structure. Armstrong et al. [89] examined the use of a beryllium–aluminum alloy for use in lightweight stiffened cylindrical barrels, particularly from the standpoint of cost. Both integrally machined orthogrid designs and bilayer corrugated-smooth designs were considered. They concluded that the beryllium alloy would be 50%

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lighter than an equivalent-performance aluminum design, but as discussed earlier, beryllium dust is toxic and the expensive safety measures required in manufacturing tend to cut into its inherent advantages. Composites are used to a much greater degree where there is no need to contain a liquid. Therefore, they have many applications in unpressurized LV structures. These applications are similar in requirements and performance as the use of composites in aircraft, the pros and cons of which (weight savings, part count reduction, ability to fabricate complex one-piece shapes, etc.) have been addressed in other chapters. Large composite structures pose design, manufacturing, and maintenance challenges that are different from those for metals. Vosteen and Hadcock [90] surveyed industry experts and concluded that using composites requires a period of materials development before product development begins; that scale-up to production can be challenging; that bonded and fastened joints require more precision than in metal structures; and that tooling must be adaptable to allow design changes, control dimensions, and adjust for springback. LVs generally experience a greater temperature range than aircraft. Composite structures on an LV may be close to cryogenic propellant tanks; conduction through the structure and cold vapors emitted during fueling can result in extremely low temperatures. During atmospheric flight, an LV proceeds through hypersonic speeds, and without some means of insulation, heating due to aerodynamic friction can raise the temperature of composites well beyond the softening point of the matrix. Therefore, the low- and high-temperature behavior of composites is relatively more important in LVs than in aircraft. Adhesively bonded joints, as well as adhesive bonds of core materials to composite facesheets, are especially susceptible to strength reduction at extreme temperatures. It is expensive enough to adequately characterize a bonded joint at room temperature, but when large temperature and humidity ranges must be considered, the task becomes that much more involved. Kobayashi et al. [91] discussed the development of a composite interstage for the H-2A vehicle. The interstage shell was a foam-core, graphite-epoxy facesheet sandwich manufactured by co-curing. The role of geometric imperfections in the buckling capability was investigated. A good description of the structural qualification test is given, in which cryogenic temperatures were imposed at the aft end of the interstage to simulate the in-flight conditions due to an adjacent propellant tank. Such approaches are often necessary in structural qualification tests for LVs.

7.7.2 PAYLOAD FAIRINGS AND NOSE CONES A conical or tapered shell is used to provide a low-drag shape for the forward end of the vehicle and to protect enclosed payloads during ground handling and atmospheric flight. When a payload is enclosed, the structure is known as a payload fairing or shroud; when no payload is inside (as at the forward end of a strap-on booster), it is called a nose cone.

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Nose cones are permanently attached to the strap-on booster and go with the jettisoned boosters when they have completed their burn. Payload fairings are jettisoned once the vehicle has ascended out of the atmosphere and air drag has ceased. Because a nose cone does not need to protect a payload, the functional demands placed on it are less stringent. Another application is the nose cone of a vehicle that undergoes a head-first atmospheric reentry. This type of nose cone must be able to resist the extreme heat and pressure of reentry, and must be constructed of heavy heat-sink and shielding materials. Therefore, it is a quite different structure from a nose cone that must function only during ascent. A very early study of materials for this type of nose cone is given by Stalder [92]. Even during ascent, nose cones are subject to high heat fluxes and therefore must incorporate heat-resistant materials, especially at the apex. The Space Shuttle Orbiter nose cone is made of reinforced carbon–carbon, which can withstand temperatures exceeding 30008F. Carbon–carbon is a fibrous composite consisting of graphite fibers in a pyrolytic graphite matrix. Expendable vehicles may use superalloys or other heat-resistant metals at the nose cap. Payload fairings, being at the extreme forward end of the vehicle, do not need to sustain as much axial load as other structures. Therefore, stiffness is relatively more important than strength for a fairing. The fairing must maintain the shape of the payload compartment so that there is no danger of contact or interference between the payload and the fairing. It must be able to resist the very highintensity sound waves (160 dB or higher) that reverberate around the launch pad after engine ignition but before the vehicle has risen above the surrounding terrain. These sound waves can be intense enough to excite panel vibrations on the fairing. The fairing may be required to attenuate the liftoff acoustics to protect the payload. The fairing must also be stiff enough so that it does not grossly deform during jettison; the motions and deformations should be linear and easily predictable. A payload fairing design used on the first Atlas–Centaur launches was made of fiberglass [93]. However, increasingly stringent payload protection requirements and the need to reduce weight whenever possible led to the use of more advanced materials, in sandwich or stiffened shell designs as a rule. The core and facesheets of sandwich shells are often composed of different materials, such as laminated composite facesheets over a phenolic or aluminum honeycomb core, or a foam core. Such constructions require the joining of dissimilar materials, usually by adhesive bonding or co-curing, and may suffer from corrosion or stresses induced by differential thermal expansion. These problems may be solved by using the same material for both the core and facesheets. The Ariane 4 fairing, a conventional design that is 20 years old but can still be considered state-of-the-art, is described by Butcher [94] in the context of a separation test. The fairing shell is largely made of a sandwich of graphite-epoxy facesheets with an aluminum honeycomb core. The forward end of the fairing is an aluminum skin-stringer design that has a layer of cork insulation. This is a less

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expensive and possibly lighter approach than using a superalloy nose cap. The fairing-vehicle separation system consists of a tension belt or clampband that secures the aft end of the fairing to the rest of the vehicle under tension provided by two steel bolts. The Russian Soyuz LV has payload fairings whose shells are sandwich structures composed of an aluminum skin with aluminum honeycomb core [95]. Schwingel et al. have described an experimental structure composed of an aluminum foam core with aluminum facesheets [96]. The sandwich layup was manufactured by rolling the facesheets over a layer of mixed aluminum and gas-generating material. In a subsequent foaming process, the sandwich was heated until the gas-generating material was activated, causing bubbles to expand in the core and increase the thickness of the sheet without an increase in mass. The large foam cells produced by this process were about as big a honeycomb cell. A prototype conical adapter was built using this process and successfully tested to about half of the limit loads applicable to the conventional structure it was meant to replace. Homogeneous core/facesheet sandwich structures such as these overcome the problems of material incompatibility but cannot be tailored as precisely as sandwiches with differing core and facesheet materials. Lane et al. [97] investigated a fairing design composed of tubes joined into a sheet, subsequently formed into a cylindrical barrel. The tubes were then punctured on the inside of the barrel to reduce the acoustic levels inside the barrel. This design, known as the chamber core fairing, is intended to provide acceptably low sound levels inside the fairing without the need for the usual nonstructural acoustic blankets. They built a laboratory-scale specimen and measured noise reduction equal to that provided by blankets for low-frequency noise. The specimen was constructed of inner and outer filament-wound facesheets with composite tubes between them. There may be difficulties in integrating the cylindrical chamber-core barrel with the required conical shape at the forward end of the nose. Ochinero et al. [98] described the design optimization and subscale wind tunnel testing of an unconventional large asymmetric payload fairing intended to accommodate very bulky payloads. They discussed an optimization procedure that resulted in the selection of carbon fiber–reinforced facesheets and a Rohacell foam core. This design was governed strongly by buckling rather than strength, which is typical for payload fairings. Consideration was given to buckling behavior beyond the elastic stability limit (postbuckling), which has been applied in practice to balloon propellant tanks but is not usual for other structures. The use of a Rohacell foam core highlights important considerations, discussed in more detail in the following section, related to core materials. A primary reason for using Rohacell for this application was the relatively low  Material

on the Large Asymmetric Payload Fairing and the subsequent section on core materials and inserts were contributed by Tomoya Ochinero and Eric Roulo, Structural Mechanics Corporation.

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knockdown factor imposed by the program. Program requirements dictated the use of core material-specific knockdown factors on the strength of the sandwich panel to account for separation between the core and the facesheets due to such variables as manufacturing imperfections, microbuckling, and moisture entrapment. A more traditional honeycomb core has a tendency to entrap moisture in the cells of the core material. The entrapped moisture can evaporate as the payload fairing is subjected to the high temperatures and low pressures of ascent. Without adequate venting features to relieve the subsequent pressure rise within the core materials, the facesheets can become separated from the core material. For this application, the program dictated a significantly larger knockdown factor for an aluminum honeycomb core than for a Rohacell core. It is academically interesting to note that the fairing would have been lighter if it had been designed using aluminum honeycomb core if only the knockdown factors were equal. Another interesting note that highlights the struggle between idealized design optimization and the realities of manufacturing and operational requirements on this application is the uniform thickness of the core. The optimization analysis showed that a significant weight savings was achievable by tailoring the core thickness to vary with respect to location on the fairing. With Rohacell, it is easier to continuously vary the thickness of the core than with an aluminum honeycomb core. However, the manufacturing constraints on this program required a uniform thickness continuous core. This resulted in a compromise where the core is thicker in many regions where a thinner core would have sufficed. It is notable that despite these design constraints, a fairing with twice the volume of the standard fairing was achieved with only a 33% weight increase.

7.7.3 CORE MATERIALS Core material is used to separate thin composite facesheets and increase the structural efficiency in bending. The purpose of this core material is to tie the facesheets together in shear, thus allowing them to work together in bending. For this reason, when modeling, the properties of the core must be properly taken into consideration. One often overlooked core property is the in-plane modulus of aluminum honeycomb cores. Facesheet-stabilized aluminum honeycomb has a significant in-plane modulus that must be accounted for when conducting thermal analysis or thermal distortion analysis of sandwich parts with thin facesheets. A good reference to estimate the modulus in the absence of test data is Scolamiero’s work [99]. Incomplete bonding between core edges and facesheets is often the source of many manufacturing-induced flaws that cause disbonds and panel failures. These can be mitigated by using reticulation to premelt one layer of film adhesive on the bare core. This increases the bond fillet between the honeycomb and facesheet and has been shown to dramatically increase sandwich panel integrity. The downside is the increase in processing time and the use of twice the number of film adhesive

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layers, which increases the mass of the panel and adds more high-CTE adhesive into the panel. For large panels and complex sections, core splices are required. The need to use separate core sections and then bond them together with foaming adhesive adds another design detail with challenging analysis requirements. For most aerospace applications, the foaming adhesive has stronger shear strength than the core, so if the dimensions of the splices are controlled to ensure proper adhesion, the core splice is stronger than the base materials. Splices should be designed to be away from any load introduction points and as far away from highly loaded regions of the panel as practically possible. With sandwich structures that ascend to outer space or have rapid depressurization requirements, vented core is required to prevent the facesheets from blowing off. An approach to compute this failure mode is presented by Kaplan [100]. The vapor needs to have a pathway to ambient, requiring edge closeouts to also be vented. Mylar closeout tapes come perforated for such applications. The core out-of-plane shear strength is used to introduce out-of-plane loads via potted inserts. Potted inserts are placed in sandwich panels to connect ancillary components, such as equipment boxes, to the panel and provide load paths for panel-to-panel connections. Most companies have proprietary insert designs, but off-the-shelf designs are sold commercially. The analysis of these joints is complicated and is described in great detail in the Insert Design Handbook [101]. Test data for these joints is required to validate the design before production. Attention should be paid to the potting compound for this style of insert. With extreme thermal environments, the out-of-plane CTE difference can cause the potting compound to either shear the core or force the failure of the core-to- facesheet bond. Potting compound weight can also become a major driver to otherwise highly efficient sandwich panel construction, as hundreds or thousands of inserts may be added to a lightweight panel to provide fixation to all the equipment that must be placed on it. With the proper attention to the additional design and analysis details of sandwich panel construction, weight savings can be realized over traditional structures. The designer must be vigilant to ensure that the additional failure modes and behavior of the structure are well understood to prevent failures.

7.8 THERMAL PROTECTION AND INSULATION Thermal environments are a significant factor in materials selection for LVs. Most liquid-fueled vehicles use cryogenic LOX at 22978F as the oxidizer, and some use LH2 fuel at 24238F. Even though cryogens are loaded only a few hours before launch, there is ample time for the tank walls, domes, and adjacent hardware to become extremely cold. Venting and leakage of boiled-off propellants create plumes of cold gas that may surround vehicle structures and cause cooling of areas not in direct contact with liquid propellants. Insulation, typically in the

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form of closed-cell polymer foams sprayed on or bonded on as precured panels, is used to protect hardware from extreme cold and to manage the boil-off of loaded propellants before and during launch. All LVs must ascend through the atmosphere, typically for two minutes or so. The competing effects of decreasing air temperature with altitude and increasing frictional heating with acceleration can cause structural skin temperatures to decrease or increase. Insulation serves to moderate the temperature of the structure during this period. Thus, the insulation applied to a cryogenic propellant tank needs to retain acceptable mechanical and thermal properties at temperatures ranging from as low as 24238F to plus several hundred degrees Fahrenheit. For the two commonest structural materials, aluminum and graphite epoxy, temperatures must be kept below about 2008F in order for the structure to retain sufficient strength and stiffness. Aluminum is more tolerant of heating than graphite epoxy but still weakens appreciably when temperatures exceed 2008F. High-strength steel is less affected by high temperatures. In some areas, such as the forward end of a nose cone or an area subject to a standing shock wave, temperatures can be high enough to require high-temperature (refractory) alloys or carbon–carbon. The leading edges of the Space Shuttle wings and its nose cap are made of carbon–carbon with a silicon carbide coating to prevent oxidation. The nose caps of expendable vehicles may be made of beryllium alloys or high-temperature superalloys. However, exposure to high temperatures is brief, so time at temperature is usually not a consideration except after many flights of a reusable vehicle. The most widely used material for expendable LV thermal protection systems (TPSs) is polyurethane foam. These foams can be sprayed on, poured into molds placed over vehicle features, or bonded on in the form of precured sheets. Foams suitable for use in TPS applications are relatively rigid. Their microstructure consists of packed bubbles or closed cells with polyurethane walls. The polyurethane itself is created by the catalyzed reaction of a polyisocyanate with a polyol. During the casting process, which takes place either directly on the structure when foam is sprayed on or poured in place or in a factory where precured sheets are made, two parts are mixed. One part is the polyisocyanate, and the other part is the polyol, catalyst, blowing agent, and surfactant. The cells are generated when the blowing agent, suspended in the liquid mixture, expands. When the mixture cools, rigid-walled cells remain, initially containing mainly the blowing agent. As time passes, the blowing agent gradually diffuses out of the cells, and air diffuses in. By the time the foam is put into service as an insulator, the cells may still contain mostly blowing agent or a mixture of blowing agent and air. The insulating characteristics of the foam can thus change with time because the thermal properties of the changing cell contents play a significant role. Until the early 1990s, the most common blowing agents in TPS foams were the chlorofluorocarbon (CFC) refrigerants CFC-11 and CFC-12. These agents, although nonflammable and nontoxic, were recognized as damaging to the

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ozone layer and were gradually banned in some countries.†† Manufacturers no longer able to obtain CFCs sought substitutes such as hydrochlorofluorocarbons (e.g., HCFC-141b), but these too were eventually banned. Changes to the blowing agent require the foam to be requalified for its intended use. Different blowing agents generate different cell sizes and shapes and affect the thermal properties of the insulation. Requalification tests may indicate that process changes are needed to maintain the foam’s performance. By varying processing parameters such as flow rate, temperature, and ambient curing conditions, a variety of foams can be generated. A surfactant may be used to control the size of the cells. The stiffness of the cell walls themselves is a function of the precursor compounds and the ratio and conditions under which they are mixed. The stiffness and strength of the foam is a strong function of the cell size: smaller cells mean a denser, stiffer, and stronger foam. Over smooth, featureless areas, sprayed-on or bonded-on foams are usually used. Sprayed-on foam is applied in several passes; in the time between passes, the exposed surface of the previous pass can partially cure. A “knit-line” may then form at the boundary surface between two passes, consisting of two adjacent layers of aligned cell walls that appear as a thickened solid wall running through a field of randomly oriented cells. As the foam rises, the forces of gravity, surface tension, and internal pressure create cells of dispersed size that tend to be oblong, with the long axes aligned in what is called the rise direction. Noever et al. [103] showed microphotographic studies of the effect of gravity on the cell size, shape, and void frequency of foams. Their control sample was manufactured in zero gravity during a sounding rocket flight. The existence of a distinct rise direction has to do with the fact that the liquid foam has to be constrained into the desired shape by the structural surface, a partially cured previous pass, and/or a mold. The rise direction and the knit-lines result in anisotropic mechanical and thermal properties. When complex shapes such as flanges or fastener heads must be insulated, foam is usually poured into molds so that it can closely conform to the underlying surface. Whether foam is poured or sprayed, when the structure has a complicated shape, voids may occur due to the inability of the foam to conform exactly to the surface. Voids may also occur between spray passes and simply as enlarged cells, which will develop in scattered locations due to the slightly incomplete mixing of components. Also, knit-lines will exist wherever a poured area meets a separately poured area or a sprayed area. Machining or shaving may be necessary to achieve a low-drag outer profile for both poured and sprayed foams. The various failure modes of foam insulation received intense study following the Columbia accident. Stresses sufficient to fail foam may be caused by cryopumping, thermal expansion, flexing and stretching of the structural substrate, ††

With regard to the major LV manufacturing countries, the United States and France banned CFCs by 1996. Russia and Ukraine were attempting to eliminate the substances but having some difficulties achieving full compliance. CFCs were still available to Chinese manufacturers as of this writing [102].

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thermal cycling, preexisting flaws, voids and unbonds, and probably several other failure modes that have yet to be conceived. Bednarcyk et al. [104] provided a discussion of the failure modes from the micromechanics viewpoint along with an analytical framework for predicting failure under complex combinations of stress, temperature, and pressure histories. The Space Shuttle contains both major types of TPS: a low-strength, lightweight layer of foam on the expendable External Tank, and more capable, reusable insulation on the Orbiter. The Orbiter is not only reused, it also must withstand the rigors of atmospheric reentry, which are a far more challenging thermal environment than launch. Figure 7.6 illustrates the location of the different types of TPS on the Space Shuttle Orbiter. Reentry TPS technology for reusable LVs has its roots in the (primarily ablative) TPS designs for the early expendable capsules. A summary of the state of the art in ablative heat shield materials for reentry vehicles was given by Bauer and Kummer [105]. They described the design of a low-density, filled silicone ablative material cast into a nonmetallic honeycomb reinforcement, bonded to a plastic sandwich structure, as applied to the Gemini spacecraft. This was an advance over the Mercury heat shield, which was a glass-phenolic, and a step in the direction of the Apollo Command Module heat shield, which was silica fiber-epoxy

Fig. 7.6 Thermal protection system of the Space Shuttle Orbiter. (Excerpted from RL Dotts, DM Curry and DJ Tillian, “Orbiter Thermal Protection System,” NASA Johnson Space Center, Houston, Texas.)

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resin again cast into a nonmetallic honeycomb support structure. These early ablative systems were extremely heavy. The Apollo shield made up almost a third of the total weight of the Command Module. A reusable TPS with a great deal of operational experience is the ceramic tiles covering most of the Space Shuttle Orbiter. The development of these tiles was a major pacing item in the Shuttle program as a whole. There are four different types of tiles, with differing capabilities, used in different areas. All of the tiles are composed of amorphous silica fibers with a 0.015-in.-thick reaction-cured borosilicate glass coating on the side facing the atmosphere. The system is tiled, rather than a continuous sheet, in order to allow thermal expansion of the substrate and individual replacement of damaged tiles. In low-temperature areas (7508F to 12008F), the tiles are colored white, whereas in high-temperature areas (up to 23008F) the tiles are colored black to improve radiative heat transfer [106]. The rest of the Orbiter acreage is insulated with flexible blankets. Carbon–carbon can endure higher temperatures than any other aerospace structural material, up to 30008F. It is relatively expensive, difficult to work with, and subject to oxidation. Titanium and the nickel superalloys are the next most expensive. As metals, they are strong and can be worked with conventional tooling, but they are also heavy. More advanced concepts have involved nonmetallic, felt, or ceramic blankets and tiles [107]. The never-completed X-33 and its envisioned full-scale successor VentureStar were to have used an advanced metallic combination TPS/aeroshell. It was to be constructed of titanium and Inconel. This represented a departure from the ceramic tile “acreage” TPS of the Space Shuttle and was meant to improve the durability of the vehicle. The metallic TPS was intended to be rain-proof, resistant to impact damage, and easily serviceable by replacing panels. However, the hottest areas of the structure, such as the nose and leading edges, were still planned to have carbon–carbon or carbon–silicon carbide panels [108]. In operation it was found that the Shuttle TPS was easily damaged and required careful observation and maintenance. This was known long before the Columbia failure, which can be seen as involving two separate TPS structural failures: one when foam insulation came loose from the External Tank and another when the loose piece of insulation struck the carbon–carbon leading edge of the Orbiter’s wing, fatally damaging the ability of the wing to withstand the reentry thermal environment. When the X-33 was developed, much effort was directed toward developing a more robust TPS. TPS’s are usually considered nonstructural and are simply attached to the outer moldline of the structure. However, recent research has been done on loadcarrying TPSs, called integrated TPS’s. Gogu et al. [109] compared materials for a corrugated core sandwich panel integrated TPS. They considered Ti-6Al-4V, zirconia, and an aluminosilicate/ ceramic oxide fiber composite as web materials; aluminum, graphite epoxy, and vacuum hot-pressed beryllium as bottom facesheet materials; and Inconel 718, aluminosilicate/fiber, and carbon/carbon as top facesheet materials. They

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concluded that using the aluminosilicate/fiber for the web and top facesheet and beryllium as the bottom facesheet led to a design only one-third the mass of the heaviest design. Lindell et al. [110] described analysis and testing of an inflatable reentry vehicle incorporating a flexible fabric-type TPS consisting of layered polyimide and woven ceramic fabric. Because the structure was inflatable, it could be much larger than conventional reentry vehicles (60–90 ft diameter). A large surface area-to-weight ratio leads to lower heating and therefore less stringent requirements on the TPS than would exist for other concepts carrying the same mass. Rakow and Waas [111] investigated an integral TPS consisting of actively cooled metal foam sandwich panels. The panels were composed of metal facesheets brazed to an open-cell metal foam core, with a coolant fluid circulated through the open-cell core structure. They discussed the advantages of this concept over previously considered actively cooled honeycomb core panels, which required separate coolant tubes to be built into the structure. The tubes do not permit as even a cooling effect as the metal foam. Henson [112] developed a class of continuum models for materials with small fluid-filled passages as may be used for active cooling. Fesmire [113, 114] discussed the testing and potential applications of aerogel materials in LV TPSs. Gels are materials that are mostly liquid by weight but have a crosslinked network that contributes enough rigidity that the material can support stress without flowing. An aerogel is a gel in which the liquid part has been replaced by a gas, resulting in a very low-density, porous material. Fesmire showed that aerogels are less prone to cryopumping than conventional foams because of their high and finely dispersed porosity. Also, they are hydrophobic and therefore do not permit frost and ice to accumulate as do some other insulating materials. Yao et al. [115] described the design and fabrication of a nickel-based superalloy honeycomb nonstructural TPS for reusable applications. They measured the strength and thermal properties of the panel and developed an oxidation-resistant coating containing a high-emittance layer for improved thermal performance.

7.9 MANUFACTURING CONSIDERATIONS This section discusses LV manufacturing considerations based on several resources [6, 69, 116–122]. Manufacturing of LVs is a process that transforms raw materials into a space LV.‡‡ This includes tanks, engines, structure, and necessary subsystems for full operations. This process has three phases: Fabrication, assembly, and checkout. Fabrication involves processing raw materials into the basic components for an LV. Examples of these components include commercially available metal plates and bars, fasteners, composite ‡‡

This section was contributed by Clyde S. Jones III, NASA Marshall Space Flight Center.

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materials, adhesives, coatings, tubing, castings, and forged metal. Assembly is the process by which components are collected from suppliers and assembled into complete systems. Most LV factories are primarily assembly facilities. Checkout is the process of verifying that the vehicle is ready for delivery. It is usually distributed during assembly, so that defects can be detected before too much value is added. A final checkout is performed as a last step before delivery to the launch site, typically along with a functional or operational check.

7.9.1 MANUFACTURING PLANNING AND EXECUTION Planning for manufacturing of space LVs is similar in many ways to that in other industries. The size of components and types of materials are comparable to commercial aircraft, and quality standards share common approaches. Unique issues in LV manufacturing are primarily related to their low production rate and high cost. Even the most popular LVs rarely exceed a production rate of one unit per month, and most are produced at less than half that rate. In contrast, the Boeing 747, for instance, with a similar size and complexity, is produced at a rate of one to six per month. With other commercial manufacturing, the comparison is even more pronounced. The automobile industry may produce one thousand vehicles in a shift, and each vehicle is far less valuable. The significance of this difference in production rate is manifested in several ways. If the production process for a particular component or assembly is only performed a few times in a year, there will be a stronger reliance on written procedures to assure that the part is produced correctly. A space LV has a greater cost per component, and so each processing step is financially riskier than in mass production. With large, expensive components and precise fit-up tolerances, tooling to position the components can be very complex. Manufacturing simulation computer systems are used to help optimize the flow of large assemblies through the factory. As the cost of computing power declines, simulation systems are an economical way to analyze different manufacturing scenarios and iterate an optimum flow. These systems can then use design information to program robots, machine tools, and welding systems for very complex assemblies. Simulation systems can adjust the programs of large complex machines to fit unique model configurations and even compensate for some types of geometric imperfections in components. Fabrication of an aluminum-phenolic sandwich structure for the common bulkhead on the Ares Upper Stage demonstrated how manufacturing simulation systems could match two welded aluminum domes with their phenolic sandwich material. Although the welded domes had small areas that did not match the design within the tolerance required to complete the adhesive joints, computer systems match-machined the phenolic to fit the imperfect parts and successfully completed the adhesive bonded assembly. A successful manufacturing planning system will provide for tracking the use of different materials and components to allow traceability in the case of defects. If

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the certification of any particular lot of parts or material used in manufacturing comes into question at any time, the manufacturing planning system can determine where the questionable parts were used on any vehicle, allowing replacement, or even acceptance by further testing or analysis. In such situations, accurate information on the pedigree of any part or material used on the vehicle can be invaluable. Nonconformances, meaning processes that were carried out differently than the design intent, are bound to occur, and so a process for disposition is necessary. Some nonconformances are acceptable. A Material Review Board develops and documents the disposition of a nonconforming part or process. A typical process consists of discovery of the nonconformance, documentation of the technical details and application, determination as to whether corrective action is needed, and, if necessary, development of a corrective action.

7.9.2 WELDING Welding is the primary assembly method for large cryogenic tanks. A pressurized tank using welded joints can reduce dry mass and part count compared with a mechanical joint and is less likely to leak over a wider range of operating conditions. Disadvantages include the requirement of a high skill level for production workers and the cost of nondestructive inspection processes to screen for cracks or related defects. Historically, welding has been a critical technical and schedule driver in production of LVs [39]. Welding aluminum for LV tanks and structures has been a well-proven process since the 1960s. Aluminum alloys commonly used include 2219, 2014, 2024, and 2195. These alloys have the advantage of high specific strength and good fracture toughness at cryogenic temperatures. An important feature of aerospace aluminum alloys used for manufacturing LVs is that they have better fracture properties at cryogenic temperatures, so that a less expensive room-temperature acceptance test is sufficient. Aerospace aluminum alloys exhibit lower mechanical properties in the weld joint than areas unheated by welding, due to oxide trapped as the metal solidifies and cracking as the metal cools. Strength reduction can be mitigated by adding extra thickness at the weld joint. The weld process is usually developed to concentrate the heat as much as practical, allowing higher welding speeds. This reduces the heat-affected zone, minimizing heat effects on the base material temper. Welds on an LV structural element are usually made automatically rather than manually. This results in more consistent heat input along the weld joint. This consistency makes the weld properties more predictable and reduces distortion. Over the years, advances in computing hardware and software have made automatic welding systems more consistent over a wider variety of production conditions. In the 1960s, and during the first few builds of the Space Shuttle External Tank, electronic servocontrols with operational logic provided by relays

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were the norm for welding automation. By the mid-1980s, digital computers were commonplace for automation, improving the operator interface and providing more accurate adjustment of all weld parameters that affect the quality of the process. A very important improvement by digital control systems was detailed recording of parameters as the weld progressed. Computers have allowed for precise programming of each parameter before welding starts, allowing the welding engineer to build a successful scheme for each joint. As welding progresses, the computer records each parameter multiple times each second. The result is detailed data on each weld, which can be compared with previous attempts to finetune the procedure. Robotic welding was introduced for LV applications in the late 1980s. Robots apply the consistency of welding automation to joints with complex curvature. The programmable path of the robot can reduce the cost of motion control compared with a specially designed system for a specific geometry. Robots using the GTAW process were able to join a wide variety of components previously welded manually on the Space Shuttle main engine in 1989 and are still in use today. A robot using the plasma arc process was used for a saddle joint on the docking nodes of the ISS in the mid-1990s. The robot used eight axes of motion to position both the component and the weld torch in the ideal orientation for a successful weld. Because the robot could be programmed for multiple paths, it was also used for other welds on the ISS structure, eliminating the need for additional welding systems. Currently, the Orion crew vehicle uses one robot to perform FSW for every weld joint on the vehicle, including circumferential and linear geometries. Using a robot to bring the welding process to multiple fixtures and weld stations reduces the overall floor space that would have been required for conventional welding. The universal programmability feature inherent in the robot is ideal for the low production rate of LVs, providing a cost-effective approach to design changes and different model configurations. Many different welding processes have been used successfully in a production system on operational LVs, including gas metal arc, resistance, GTAW, plasma arc, electron beam (EB), and FSW processes. Gas metal arc has been phasing out since the 1960s because it is prone to porosity and oxide inclusions when welding aluminum. When used on the Saturn vehicles, the process required significant rework compared with the welding processes used today [39]. Resistance welding processes have been used extensively on LVs. It worked well with the 301 and 321 stainless steels used in the Atlas family and is still used for the Centaur upper stage. This process has not found similar success in aluminum structures, primarily due to inconsistent quality. This is likely due to the high resistance of aluminum oxide that quickly forms on the surface of aluminum, affecting the current flowing between the electrodes. The overlap design of a resistance-welded joint also leads to difficulties in applying nondestructive inspection techniques. Other applications for this welding process include structural covers for insulation systems.

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GTAW is still commonly used on aluminum welds for LVs. These processes have higher energy density than the gas metal arc process, resulting in a smaller heat-affected zone and thus higher mechanical properties. GTAW and plasma welding processes can be operated in alternating polarity, which provides a cathodic cleaning action to aluminum during welding. This reduces the presence of oxides, minimizing the chance for oxide inclusions in the weld zone, and improves flow of the molten pool. Oxides are further discouraged in the weld zone by abrading the joining surfaces, through draw filing, wire brushing, or other mechanical means. Because aluminum will develop a surface oxide quickly, usually a time limit is established between completion of surface cleaning and when welding starts. If this limit is exceeded, additional cleaning is required before welding. EB welding uses a high voltage to accelerate electrons, which are focused using magnetic fields to melt metals for welding. This welding process has the advantage of very high energy density, which can penetrate and join thick parts with minimal distortion and minimal effect on the temper of adjacent material. It is used extensively on LVs to assemble engine components, hermetically seal batteries, and join thick materials used in heavily loaded structural parts. The process takes place in a vacuum, so metals that oxidize at elevated temperatures, such as titanium, can be welded with minimal risk of included oxides. Because the process must take place in a vacuum chamber, there is a practical limit to the size of components that can be EB welded. It is also limited to metals that are nonmagnetic, which will not deflect the beam during welding. FSW has been adopted by LV manufacturers rapidly since its invention in the early 1990s. FSW is ideally suited for aluminum because it is relatively soft at elevated temperatures. This allows commonly available tool steels to be used for the pin that applies friction to the part. It also reduces the forces that must be reacted by the weld tooling. Although titanium and ferrous alloys have been welded with the FSW process, aluminum alloys are the most common application. The first application of this process in a production environment was in Europe, fabricating aluminum structures for shipbuilding in the mid-1990s, applied to a 6000-series alloy. The first LV application was by Boeing on a Delta II variant that first flew in 2001, which applied the process to the 2024 alloy on longitudinal welds. Lockheed Martin and NASA developed a more complex application for longitudinal barrel welds on the Space Shuttle External Tank in the early 2000s. The External Tank used Al-Li 2195 for these parts, and the weld joints tapered in thickness from almost 16 mm at the LH2 tank aft dome down to 8 mm at the LH2 tank forward dome. This application required a more sophisticated pin tool that could adjust its extension as the weld traveled along the joint. An automated method to control the pin extension was developed to maintain the proper depth and stir the weld completely through the weld joint thickness. The last five External Tanks produced took advantage of this new technology. The Delta IV LV was designed with FSW in mind. All the longitudinal welds were joined using FSW, whereas circumferential welds used a version of variable polarity

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plasma arc. The design of the LOX and LH2 tanks eliminated some circumferential welds by increasing the number of barrel panels and longitudinal welds. A common problem in all welding processes is distortion. A distorted component is more difficult to join to adjacent structure and has higher residual stress, both of which reduce structural efficiency. Distortion resulting from the weld process comes primarily from shrinkage in the weld zone but can also result from the interaction of residual stresses in each component and how they change after welding heats the parts. Because highstrength materials are often used in LVs, distortion is exacerbated, as localized shrinkage in the weld area is not distributed across a larger area by yielding. There are a variety of mitigation techniques for weld distortion and fit-up issues. Welldesigned fixtures position the parts precisely, and pneumatic actuators restrain the parts during application of heat. (Hydraulic actuation is rarely used for welding fixtures to avoid contamination by leaking fluid.) Alignment is measured before welding, and extra pressure is applied, or trimming operations are used, to bring the fit-up within specifications. Tack welds can be used to restrain the parts and maintain alignment as heat is applied. Spacing, depth of penetration, and the sequence of application are all important parameters in tack welding. Weld processes with low energy density and a less concentrated heat source, such as gas metal arc, are usually more prone to distortion. Areas with thinner material, or areas that require more heat passes, exhibit more distortion. Weld repair areas are more prone to excessive distortion because the part is subjected to multiple weld passes and solidification shrinkage in repair areas. The additional heat reduces the strength of the base metal by changing any previous tempering processes. Multiple welds in the same area will also act on any residual stresses in the components being joined, producing additional distortion and residual stress. High-energy-density weld processes such as EB and laser welding result in less distortion. Resistance welding, plasma arc, and GTAW fit between these two extremes. This is primarily due to a smaller molten pool along the weld seam, which reduces metal solidification shrinkage. FSW produces less shrinkage because it does not melt the material. After welding is completed, procedures typically require measurement of the joint geometry to verify that reinforcement, peaking, and offset (or mismatch) are within specification. If corrective measures are warranted, planishing can be used to compress the weld reinforcement and correct geometry problems. In rare cases, additional welding passes can be used to shrink certain areas to bring the geometry into compliance. This approach is less often used because of the risk of distortion.

7.9.3 MECHANICAL ASSEMBLY PROCESSES Mechanical fastening systems are well-developed for use on LVs. Major structural elements are joined using bolts and related fasteners with precision, accuracy, and

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predictable mechanical properties. Although the pressurized components of LV tanks are more typically welded, mechanically fastened components are used for propellant feedline attachments, venting components, personnel access covers, and instrumentation feed-throughs. Bolted connections allow disassembly and reassembly. Keys to success with bolted joints include good fit and adequate fastener torqueing. Success is verified by measuring torque on the fastener and a leak test. If fasteners are to be threaded to blind holes in an aluminum structure, a threaded insert is normally used. In aerospace applications, threaded fasteners require at least one locking device to prevent loss of preload, and lock wire is typical for this application. Thread-locking compounds are not commonly used due to temperature extremes experienced on LVs, but thread-sealing compound has been used on the Space Shuttle external rank to reduce infiltration of liquid nitrogen behind thermal protection foam. Riveting has been used in unpressurized structures of LVs such as the intertank subassembly of the Space Shuttle External Tank.

7.10 SUMMARY, TRENDS, AND OUTLOOK Preparing for the launch of an expensive, specialized payload on an expendable vehicle involves “good practice” processes that do not always have a firm scientific basis. Low flight rates make it difficult to rationally assess the costs and benefits of analysis, testing, and quality control. The verification criteria, qualification strategies, and analysis methods that have matured over the past few decades have been described here. Space LVs use many of the same materials as aircraft: the 2000-, 6000- and 7000-series aluminum alloys; laminated and filament-wound composites; highstrength steels; and titanium alloys. The need for mass efficiency is the primary driver for both aircraft and LVs. However, the frequent use of cryogenic propellants, as well as high aerodynamic heating environments, impose challenging thermal conditions on LVs. On the other hand, the short lifetime of expendable launchers reduces the importance of fatigue, fracture, corrosion resistance, and other properties governing long-term material behavior. For reusable vehicles, fatigue and fracture can be just as important as in aircraft, and the design of a robust, reusable TPS for atmospheric reentry requires all materials and structures technology to be brought to bear. Most of the material processing and joining technologies used in aircraft are also used in LVs. Welding is a key technology in LV structures. FSW is arguably the most significant advancement in the state of the art of materials and structures since the development of composites. Aluminum–lithium alloys, now introduced on a large scale in the Space Shuttle External Tank, represent a significant improvement in strength-to-weight ratio over conventional aluminum alloys. Composite propellant tanks can offer further gains in mass efficiency with judicious design, but the need for robust

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joints and minimization of permeation after fatigue remain significant roadblocks to the use of composites in pressurized structure. However, filament-wound composite solid rocket motor cases are a mature and widespread technology. Composites continue to be an active area of research. Bolted and bonded composite-to-composite or composite-to-metal joints present challenges to both design and analysis. Textile preforms, new methods of curing, and new matrix materials are all pathways to meeting these challenges. New materials such as aerogels, metal foam, and nanocomposites can be fabricated and tested at the laboratory scale; these materials may soon find applications in production. Another technology enabling the wider use of composites is rigorous methods for predicting gradual progression of damage and assessing residual strength. Looking further into the future, nanostructured materials such as carbon nanotubes and graphene sheets appear to hold great promise. These materials have interesting electrical and thermal properties as well as extremely high specific strength and stiffness. Current research seeks to reduce the cost of producing such materials and to assemble them in quantities usable for structural applications. Modifying current materials such as polymeric matrix materials for composites by the addition of nanostructured materials may be a significant first step in their more widespread use (see Chapter 4 for details.) A system study predicted a factor of two improvement in weight if conventional carbon fiber composites were used throughout a structure, but a factor of 10 improvement if projected properties of carbon fiber nanotube-reinforced materials could be realized [123]. Advanced materials identified in work by Harris et al. [123] and potentially applicable to LV structures included the following: titanium–aluminum alloy; alumina fiber/aluminum matrix composite; aluminum and titanium alloy foam as core materials for sandwich structures; aluminum–beryllium alloys; silicon carbide fiber/beryllium matrix composite; carbon nanotube fiber/aluminum matrix composite; single-crystal metals, nanotube-reinforced alloys, and new superalloys for high-temperature applications; and ceramic matrix composites. Bionics or biomimetics [124] is another material and structural concept that is a current topic of research. It has long been realized that if a structure were capable of large-scale adaptations, it could be optimized for two or more very different environments and therefore be much more efficient than a one-size-fits-all design. Flaps, slats, and trim tabs may be regarded as first steps down this path. Swing-wings and deployable space structures display yet more adaptation, but these continue to use conventional materials. A bionic structure would incorporate flexible skin materials capable of large strains, as well as internal bracing akin to a skeleton and actuators akin to muscles. Integral fluid passages could provide both thermal control and the ability to change the shape or stiffness of the structure by changing pressures or flow rates. Such concepts could answer requirements for extremely efficient, adaptable, robust LV structures in future reusable, SSTO systems.

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[107] Rasky, D. J., Milos, F. S., and Squire, T. H., “TPS Materials and Costs for Future Reusable Launch Vehicles,” RTOP 242-35-10, NASA Ames Research Center, Moffett Field, California, 2000. [108] Sumrall, J., Lane, C., and Cusic, R., “VentureStar . . . Reaping the Benefits of the X-33 Program,” Acta Astronautica, Vol. 44, No. 7–12, 1999. [109] Gogu, C., Bapanapalli, S. K., Haftka, R. T., and Sankar, B. V., “Comparison of Materials for an Integrated Thermal Protection System for Spacecraft Reentry,” Journal of Spacecraft and Rockets, Vol. 46, No. 3, 2009, pp. 501–503. [110] Lindell, M. C., Hughes, S.J., Dixon, M., and Willey, C. E., “Structural Analysis and Testing of the Inflatable Re-entry Vehicle Experiment (IRVE),” Proceedings of the 47th AIAA/ASME/ASCE/AHS/ASC Structures, Dynamics and Materials Conference, Newport, RI, 2006. [111] Rakow, J. A., and Waas, A. M., “Response of Actively Cooled Metal Foam Sandwich Panels Exposed to Thermal Loading,” AIAA Journal, Vol. 45, No. 2, Feb. 2007, p. 329. [112] Henson, G. M., “Engineering Models for Synthetic Microvascular Materials with Interphase Mass, Momentum and Energy Transfer,” International Journal of Solids and Structures, Vol. 50, No. 14–15, 2013, pp. 2371–2382. [113] Fesmire, J. E., “Aerogel Insulation Systems for Space Launch Applications,” Cryogenics, Vol. 46, No. 2–3, Feb.-March 2006, pp. 111–117. [114] Fesmire, J. E., and Sass, J. P., “Aerogel Insulation Applications for Liquid Hydrogen Launch Vehicle Tanks,” Cryogenics, Vol. 48, No. 5–6, May–June 2008, pp. 223–231. [115] Caogen, Y., Hongjun, L., Zhonghua, J., Xinchao, J., Yan, L., and Haigang, L., “A Study on Metallic Thermal Protection System Panel for Reusable Launch Vehicle,” Acta Astronautics, Vol. 63, March 2008, pp. 280–284. [116] Johnson, M. R., “Friction Stir Welding Takes Off at Boeing,” Welding Journal, Vol. 78, No. 2, 1999, pp. 35–39. [117] Cohen, H., “Space Reliability Technology: A Historical Perspective,” IEEE Transactions on Reliability, Vol. R-33, No. 1, 1984, pp. 36–40. [118] Hoffman, E., Pecquet, R., and Arbegast, W., “Compression Buck-ling Behavior of Large-Sale Friction Stir Welded and Riveted 2090-T83 Al-Li Alloy Skin-Stiffener Panels,” NASA TM-2002-211770, NASA, 2002. [119] Nemeth, M., “Effects of Welding-Induced Imperfections on Behavior of Space Shuttle Superlightweight Tank,” Journal Spacecraft and Rockets, Vol. 36, No. 6, 1999, pp. 812–819. [120] Davis, D., and McArthur, J., “NASA Ares I Crew Launch Vehicle Upper Stage Overview,” Proceedings of the 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, AIAA, Reston, VA, 2008. [121] Chang, I.-S., “Space Launch Vehicle Reliability,” Crosslink, Vol. 6, No. 2 (Spring 2005), pp. 22–32. [122] Ryan, R., and Townsend, J., “Fundamentals and Issues in Launch Vehicle Design,” Journal of Spacecraft and Rockets, Vol. 34, No. 2, 1997, pp. 192–198. [123] Harris, C. E., Shuart, M. J., and Gray, H. R., “A Survey of Emerging Materials for Revolutionary Aerospace Vehicle Structures and Propulsion Systems,” NASA TM-2002-211664, NASA, Hampton, VA, 2002. [124] Bar-Cohen, Y. (ed.), Biomimetics: Biologically Inspired Technologies, CRC Press, Boca Raton, FL, 2006.

CHAPTER 8

Materials for Exploration Systems Peter A. Curreri NASA Marshall Space Flight Center, Huntsville, Alabama

8.1 INTRODUCTION Progress has been made on materials to construct spacecraft and aerospace structures for rocketlaunches from Earth for exploration missions. These materials have enabled the development of a substantial satellite industry and the space probes beyond Earth orbit. However, once we leave the sphere of influence of Earth or attempt to construct objects in space more massive than a space station, Earth-based materials and processes become less important than space-based materials and processes. The reason is simply that the Earth is physically a large planet with a large associated gravity relative to the propulsion energy that can be attained from oxidation of chemical fuels. Thus, although the technology of the chemical rocket engine has reached maturity (the space shuttle main engine was the most efficient engine ever built and the space shuttle used reusable technology), the costs of space launch in the last 30 years have not decreased appreciably [1]. If we are to pursue the ambitious development of Earth–moon space for space solar power, permanent human occupation of the moon, and sending humans to Mars, then the basic laws of physics and economics dictate that we develop and use materials and processes in space to the exclusion, as much as possible, of Earth launch and resupply. Because the costs of space infrastructure for a single mission to extract and process in situ materials can be prohibitive, the temptation is to continue to design single missions based on the Apollo model of “planting a flag and footprints,” doing a survey, collecting samples, and returning to Earth. However, this model has prohibitive recurring costs because it creates no in-space infrastructure and has proved not to be sustainable even for human exploration. For the more ambitious (and arguably more important) goals of creating a space solar power industry and the human settlement of space, the Earth launch model does not scale up effectively [2] to enable a substantial program. Lunar and asteroid materials must be used.  Ph.D.

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Similar to our experience with industry on Earth, one must first invest in the mining and manufacturing infrastructure in space to attain a sustainable program of human exploration and development and eventual human habitation of space. Analysis has shown [3] that for the large-scale economic development of space, three things are necessary. We must manufacture in space, we must use in situ space resources, and we must have a workforce that is living in space.

8.2 DEVELOPMENT OF A TECHNICAL BASIS FOR MATERIALS PROCESSING IN SPACE Much of the materials processing technology developed for use on Earth is applicable to materials processing in space, but there are fundamental differences that the materials community must address. Chief among these differences is the difference in gravity, which can have profound effects on materials processes and the resulting materials properties [4]. A fundamental scientific basis has been developed over the last 40 years by the low-gravity materials science experiments on free-fall platforms near Earth and on various spacecraft and laboratories in Earth orbit [5].

8.2.1 MATERIALS SCIENCE AND PROCESSING IN SPACE The science of materials processing in microgravity provides a technical basis for the use of space resources for the exploration and development of space [6]. The Apollo flights to the moon and their robotic precursors emphasized the study of lunar geology (Fig. 8.1a). Sample returns from lunar human and robotic missions have enabled extensive study [7] of lunar mineralogy. The near elimination of gravity during flights to the moon enabled carry-on experiments to study important materials processing phenomena such as surface-driven convection. It was realized that the free-fall conditions in low Earth orbit enabled, for the first time, long-duration materials science experiments in continuous microgravity. After the human lunar flights, excess hardware was used to loft the first large space station, Skylab (Fig. 8.1b). Extensive materials science and processing experiments were performed during the months of human Skylab operations. The Skylab experiments included the microgravity solidification of metals and semiconductors. These experiments included the first results [8] showing that solidification in microgravity could reduce defect formation in semiconductor crystals. After Skylab, research in microgravity solidification continued, first using free-fall drop towers and parabolic aircraft flights (Figs 8.1c and 8.1d) and with the advent of the space shuttle and spacelab experiments using both pressurized [9] and telescience manned orbital laboratories [10] in the space shuttle cargo bay (Fig. 8.1e). Results from the materials science and processing in space studies were a main justification for the building and utilization of the International Space Station (ISS) laboratories in which this science continues.

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a)

c)

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b)

d)

e)

Fig. 8.1 a) Apollo lunar dust experiment. (Courtesy NASA.) b) Skylab orbital laboratory. (Courtesy NASA.) c) Low-gravity experiments during parabolic aircraft flight. (Courtesy NASA.) d) Sounding rocket payload. (Courtesy NASA.) e) USMP Spacelab furnace in shuttle cargo bay. (Courtesy NASA.)

Since the early 2000s, NASA has reemphasized plans to again send humans beyond Earth orbit to the moon, asteroids, and Mars. Thus, in addition to the study of microgravity materials processing in the ISS, the study of processing of lunar, Mars, and asteroid resources to enable human exploration is also emphasized. Thus, there is a renewed interest in lunar, Mars, and asteroid mineralogy, not only for astrogeology but also for understanding its potential for in situ processing to enable human presence and economic growth in space. Studies of the moon have found [7] that the lunar geology differs significantly from Earth’s. The difference includes the relative abundance of elements. For example, relative to the Earth’s surface, hydrogen, carbon, and nitrogen are scarce on the lunar surface. Although it is believed that the original source of

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the lunar materials was from the Earth’s crust, the lack of atmosphere on the moon has changed this composition, for example, to enable iron to exist in the reduced elemental state. This must be considered, for example, when designing extractive metallurgical processes [11] for lunar materials. Extensive studies of solidification and processing in low gravity [4] have shown that the elimination of gravity-induced convection and sedimentation can significantly affect microstructure and resulting material properties. Orbital laboratories have enabled materials processes to be studied under long-duration microgravity conditions; however, materials processes under the partial gravities of the moon (0.167 g) or Mars (0.369 g) have only been studied for durations of tens of seconds during aircraft parabolic flights. Many phenomena, for example, reaching a diffusion-controlled steady state during metal alloy solidification, require minutes or hours in reduced gravity. Another aspect important to materials selection in reduced gravity is the hierarchy of materials in terms of specific mechanical properties [12], which is discussed next.

8.2.2 IMPORTANT CONSIDERATIONS FOR USING NONTERRESTRIAL MATERIALS The discussion of “hierarchy of materials” in the context of nonterrestrial materials presented in the following sections was first presented by Stefanescu, Grugel, and Curreri [12]. Materials processing in low gravity, however, can differ quite significantly from the same processes on Earth. The absence of buoyancy-driven flow, for example, changes the solidification processes that are fundamental to most manufactured goods. In space, solidification can yield lower defects and more homogeneous crystals, which could yield better semiconductors for computer chips or, conversely, could increase the grain size and change phase composition of metal alloys, which could yield poorer mechanical properties. Thus, just as terrestrial materials science has been essential for technological advance on Earth, it is expected that materials science in low gravity will enable cheaper, more robust methods for manufacturing and hence extended human presence beyond Earth orbit. In this section, the use of in situ resources to produce metal alloys for use on moon or Mars bases is discussed. The particular focus is candidate metallurgical extractive processes, new hierarchy of materialsspecific properties, and the effects of reduced gravity on microstructure and material properties.

8.2.3 HIERARCHY OF MATERIALS Use of materials for specific applications is based on their mechanical properties. An example of such a hierarchy is given in Fig. 8.2, where tensile strength is used as the main criterion. ADI is austempered ductile iron. Based on this criterion, the best material is low-alloy steel and the poorest is magnesium. Other criteria, such as yield strength, elasticity modulus, elongation, or combinations of these, can be also selected. This kind of hierarchization is acceptable when the weight of the part

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Fig. 8.2 Hierarchy of materials based on tensile strength. is not an issue. However, if weight becomes an issue, as in aerospace applications or even in today’s automotive applications, other criteria that take density or weight into account may be used (Table 8.1). Such a criterion may be the ratio tensile strength/ density (specific strength) or maximum load/unit weight. The latter is preferable for this analysis because it includes the role of gravitational acceleration. Furthermore, it is nondimensional. The order will be changed, with titanium becoming the best and gray iron the least desirable material (Fig. 8.3). Magnesium becomes more competitive. The planetary bodies of interest for this discussion have significantly different gravitational accelerations than the planet Earth. The gravity level on the moon is 0.167 g, whereas on Mars it is 0.369 g. Because of the change in the g level, the numbers will change again. However, changing the g level alone will not alter the hierarchy but only modify the numbers by a factor proportional with the gravitational acceleration. Further significant change in hierarchy will be brought about if cost is included in the criteria used. As illustrated in Fig. 8.4, where the cost was assumed to be the processing cost on Earth, titanium becomes the least desirable choice. Because weight is included in the evaluation criteria, the gravity level will affect the numbers. Although, as indicated before, the relative hierarchy will not

TABLE 8.1

QUALITY CRITERIA USED TO ESTABLISH HIERARCHY OF MATERIALS

Quality Criterion

Symbol

Units

Tensile strength

TS

MPa

TS/r

m2/s2

TS/(r . g) $/(TS/r . g)

—

Tensile strength/density a

Load/unit weight

Cost/load/unit weight a

For unit length.

$

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Fig. 8.3 Hierarchy of materials based on maximum load/unit weight.

be altered, the difference between the various materials will change as a function of gravity level, as shown in Fig. 8.5. A clear compression of material properties is shown as the gravity level decreases. What this means is that on the moon, the decision to select one material over another may be based mostly on the availability of the material because the differences on the cost/ load/unit weight criteria are minimal. However, processing costs may be widely different than those on Earth, an issue that will be addressed in the next section.

8.2.4 MATERIALS AVAILABILITY AND EXTRACTION Naturally occurring “native” gold, silver, and copper were undoubtedly the first metals used by mankind. In a practical sense these elements were malleable and could be plastically deformed. Gold and silver took on ornamental roles, perhaps because of scarcity but more likely due to their inherent softness; copper, on the other hand, work hardened and could be hammered into useful, durable tools. Unfortunately, most metals prefer to be combined with other elements such as oxygen, sulfur, or chlorine and are thus designated as ores. Fig. 8.4 Hierarchy of materials based on cost/load/unit weight.

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Fig. 8.5 Influence of gravity level on the cost/load/unit weight criteria. Subsequently, our prehistoric ancestors discovered smelting, the process by which an ore could be reduced in a wood fire to its base metal. This technically innovative process provided them with lower melting point metals such as lead, tin, and zinc; combining the latter two ores with copper and/or copper ore produced bronze and brass, alloys with improved properties. The Bronze Age succumbed to the Iron Age, which then shaped civilization, set the stage for the Industrial Revolution, and arguably continues to this day. Relatively tiny amounts of native iron were either formed under unusual conditions [13] or in some meteorites [14]. Iron, however, prefers to be combined with oxygen and its ores [e.g., hematite (Fe2O3) and magnetite (Fe3O4)]. These ores are well represented, although not uniformly distributed, in the Earth’s crust. Although the high temperatures required to reduce these ores precluded early man from producing pure metal, a spongy mass consisting of iron and slag was formed that could be hot-worked to useful shapes. The blast furnace eventually evolved, in which the combination of ore, flux (limestone), coke (distilled coal), and air produced a high carbon “pig” iron. Today, in view of economy and properties, iron and its alloys are by far the most used metals. The intent of this briefest introduction to iron production is to convey a sense of its long history and the associated trials and tribulations encountered and overcome in developing the science and technology to what it is today. Similar convoluted developments characterize production of other metals of interest, including aluminum, magnesium, and titanium. Earlier and recent space expeditions have found metal-bearing rocks and soils (i.e., resources) on two of our nearest planetary neighbors, the moon and Mars. A comparison of typical soil analysis [15] for the three celestial bodies of interest is given in Table 8.2. According to past surveys, the moon has soils particularly rich in Al in its Highland regions, and in Ti in its Mare regions. Both the moon and Mars have significantly higher levels of Fe than Earth. Thus, if the composition of the soil is any indication of the availability of these metals for extractive processes, and based on the analysis of the hierarchy of materials presented in Figs 8.2–8.4,

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TABLE 8.2

TYPICAL SOIL ANALYSIS (WT %) OF CELESTIAL BODIES OF INTEREST

Body

O

Mg

Al

Si

S

Ti

Mn

Fe

Earth

47

2.3

8

27

0.04

0.5

0.1

5.1

Mars

42–45

2–5.5

4.2–6.6

20–26

0.9–2.5

0.4–0.7

0.4–0.7

10–15

Moon

40–45

4.9–6.8

5.8–14

19–22

0.06–0.1

0.3–5.6

0.05–0.2

4–15

it might be anticipated that iron and titanium will play a major role in the material competition to build the structures needed for extension of human civilization to the moon and Mars. Because the moon lacks an atmosphere, its surface is subjected to exposure by hydrogen carried in the solar wind. This hydrogen reduces iron oxide (FeO) in the soil to fine iron particles [16] that, should sufficient quantities exist, would be an ideal source for raw material. Iron containing ilmenite (FeTiO3) is also found in the lunar soil [17] but must be reduced, albeit by “nontraditional” methods. Some early suggestions follow terrestrial methods. Fe2O3 is a reaction product when ilmenite is subjected to molten sodium hydroxide [18], and iron can eventually be obtained through a carbochlorination process [19]. Silicon will reduce FeO to iron at 13008C. Hydrofluoric acid can be used as a leaching agent, after which iron can be recovered by electrowinning [20]. Other approaches attempt to use the unique environment of space for materials extraction. It has been suggested that the sun’s energy could be focused to reduce moon ores through vaporization [21]. More recent NASA-sponsored research [22] has focused on lunar oxygen extraction for life support and propellant where metals and silicon might be a by-product. The primary methods considered are hydrogen reduction and carbothermal reduction and molten oxide electrolysis, the latter of which has been demonstrated in the laboratory as a viable candidate for extraction of metals and silicon from lunar regolith [23, 24]. Molten oxide electrolysis of lunar regolith has the advantage of being applicable to the known regolith compositions (ore independent) and not requiring the transportation of reagents from Earth. Some disadvantages of molten oxide electrolysis are high operating temperatures (15008C and higher) and the development of suitable nonreacting electrodes. Dissolution and electrowinning of lunar, Mars, and asteroid materials can be achieved with suitable ionic liquid electrolytes at near ambient temperatures [25]. The disadvantage is that the ionic liquids may have to be transported from Earth and must be reconstituted for continuous reuse. The existence of iron ore on the surface of Mars was confirmed by results from the alpha proton x-ray spectrometer within the Pathfinder rover. The Mars soil was determined to consist of 17.5 wt % FeO, and a given rock (“Barnacle Bill”) contained 12.7 wt %. The ready presence of this ore provides a processing advantage over processes that require ore beneficiation, such as ilmenite reduction, in the moon’s stark environment. Furthermore, the carbon dioxide–rich atmosphere

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of Mars provides a very important resource. In combination with hydrogen (which, depending on the location, may or may not have to be imported), several well-known and well-characterized reactions can be implemented [26]. Now, using well-established technologies, carbon monoxide, water, and methane (CH4) can be produced and collected. Carbon monoxide (CO) will reduce (solid) FeO at a temperature below 8008C. Thus, the atmosphere of Mars not only provides a basis for life support and fuel production but could well facilitate iron production. Obviously, there will be considerable technical and financial challenges before iron, steel, titanium, and other components are produced on extraterrestrial bodies. However, this goal appears to be entirely attainable.

8.2.5 SOLIDIFICATION PROCESSING The main characteristics of the lunar and Martian environment that will affect processing techniques are lower atmospheric pressure and lower gravity. It is difficult to anticipate at this time how the price structure of the materials of interest will be altered during processing on the moon or Mars. However, it is clear that significant changes are expected in the behavior of liquid metals during processing. This in turn will affect the cost of processing. Issues that must be addressed include melting and casting techniques. 8.2.5.1

MELTING TECHNIQUES

Melt processing of metals on Earth can be done in ambient air environment, under inert gases, or in vacuum depending on the reactivity of the molten metal. On both the moon and Mars the atmosphere is extremely poor in oxygen, and thus it is anticipated that the price of melting reactive alloys will be much closer to that of nonreactive ones. This will certainly benefit titanium and magnesium and further contribute to the compression of the specific properties data in Fig. 8.5. Melt containment is another relevant issue for reactive metals because they tend to react with most ceramics used in classic processes. Recent progress in magnetic containment melting (MCM) will find a very favorable environment on the moon and Mars. The reduced gravitational acceleration will impose significantly lower requirements on the size of the coils and the energy consumption. In particular, a combination of cold-wall induction melting and MCM (Fig. 8.6) may prove to be the method of choice for melting titanium and its intermetallics. 8.2.5.2

CASTING TECHNIQUES

Most casting processes on Earth rely on gravity to help fill the mold, hence the name gravity casting. Gravity also conveniently imposes the position of shrinkage cavities in the upper part of the casting. The absence of gravity or very low gravity levels (mg) are known to create problems for scientists experimenting with

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Fig. 8.6

Cold-wall magnetic.

solidification in space. Indeed, obtaining sound samples was invariably a problem in space shuttle experiments. Some pressurization during solidification may be required to improve casting soundness in the absence of gravity. Gravity casting has a major disadvantage: metal flow is in the turbulent regime. This results in gas and solid inclusions being incorporated in the casting, which alters the quality of the cast material. To produce premium castings, countergravity casting is used. In this process the metal is fed into the mold from the bottom by applying pressure on the liquid metal (Fig. 8.7). The free vacuum on the moon and Mars will make counter-gravity casting a very competitive process.

8.2.5.3

MATERIAL PROPERTIES

Gravitational acceleration strongly influences solidification processes through Stokes flow, hydrostatic pressure, and buoyancy-driven thermal and solutal convection. Microstructural development and therefore material properties, presently being documented through ongoing research in microgravity science and applications, needs to be understood and scaled to the reduced

Fig. 8.7 Principle of counter-gravity containment melting. casting.

vacuum

pressure

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515

gravity environments. Comparison of solidification data in microgravity on orbital platforms, 1024 g on suborbital sounding rocket flights and 1022 g on parabolic aircraft trajectories with solidification data taken on Earth, have documented gravity dependence in microstructure [4]. Convection has been shown to strongly influence solute redistribution. Continual buoyancy-driven mixing of the liquid ahead of the solidification interface (for partition coefficient not equal to 1) in one gravity causes alloy macrosegregation. In low gravity a steady-state, diffusion controlled boundary layer can form, resulting in sample solute homogeneity. Eutectic alloy microstructures, for example, cast iron, are strongly dependent upon the magnitude of gravity during solidification. Spacings of eutectic fibers, flakes, and lamella; nucleation of graphite grains; and spacing of primary dendrites can be quite different from those obtained in the laboratory or foundry on Earth when solidification occurs in low gravity. Thus, handbook values for alloy mechanical and electrical properties compiled in one gravity cannot be relied upon for in situ resource processing on the moon or Mars. In summary, the moon and Mars offer rich sources of ores that can be exploited to produce metals for electrical conductors and structural materials. The new hierarchy of materials in terms of specific properties must be considered. Processing methods of choice are influenced by the low-pressure atmospheres and lower gravity present in these worlds. The influence of gravity on the microstructures created by solidifying under reduced gravity must be understood and applied before the engineering properties of these in situ produced materials can be accurately determined.

8.3 DEVELOPMENT OF AN ECONOMIC BASIS FOR MATERIALS PROCESSING IN SPACE: IN SITU PROPELLANT PRODUCTION The costs to land a pound of payload on the lunar surface are about 6–10 times higher than the cost of putting a pound of mass into low Earth orbit. This is because, for the trip from low Earth orbit to the lunar surface, six tons of propellant must be used for each ton of mass landed on the moon. Carter estimates [27] that transport from the Earth to the moon would cost $25,000 per pound in 1984 dollars. This is inflated to about $50,000 in 2010 dollars, which was about twice the cost of gold in 2011. The benefits of utilization of space resources are many. Some are beyond economic. For example, establishment of self-sustaining human habitats beyond Earth are arguably essential for human survival [2]. From the economic point of view, development of in situ products individually and in combination must be analyzed for time to profitability. Because building the first materials processing and production facilities off Earth requires launch and transport out of the Earth’s formidable gravity well, the initial costs are large. The payoff can be extraordinary (especially for space solar power satellite production from lunar or

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asteroid materials) [2]. However, because a large investment is required before income, the cost of money can be an important financial factor. Some products, for example, lunar oxygen for propellant, have the potential to substantially lower space operations costs after the production capability is established. Other products (the best example is space-collected solar power beamed for use on Earth) have the potential for extraordinary profits. These profits can be used to enrich investors and/or to dramatically accelerate the human settlement of space [2, 28]. A number of economic analysis methods have been used to study the financial feasibility of space resource utilization. To evaluate economic feasibility, often parametric analysis is performed. Simplified yearly costs of the major elements of the process are set up as parameters in an economic model. These models enable the time-integrated net present value of the program to be assessed.

8.3.1 PRODUCTION OF LUNAR OXYGEN FOR SPACECRAFT PROPELLANT: PAYING TENANT FOR A LUNAR BASE We will now discuss examples of economic analyses to determine the financial validity of space resource utilization for specific cases. The production of liquid oxygen from lunar materials has received much attention in the aerospace community for several reasons. First, oxygen, at about 40% by weight, is the most abundant element in the lunar surface regolith and rocks. Second, oxygen in its elemental state has a ready market, in that it can be used as a primary component (oxidizer) in rocket propellant. Third, the processing methods for extracting oxygen from lunar surface materials have an experiential basis in the extractive metallurgical industry on Earth. Most important from the economic standpoint, the moon is strategically located, anchoring the Earth–moon economic sphere, and has only one-sixth the gravity burden to space that transportation from Earth has. For these reasons the space resources utilization focus at NASA in recent decades has been lunar oxygen extraction for life support and propellants. Economic analysis of a single product, oxygen, for commercial viability is not strictly valid because, for example, the presence of in situ derived electrical power, food production, metals, and habitat production will change the economics substantially. To illustrate this, we will first examine an economic parametric analysis for lunar oxygen production and delivery to low Earth orbit (LEO) and then discuss the sensitivity the analysis would have to other in situ derived products. Following Simon’s parametric analysis of lunar oxygen production [29], we will consider the annual costs of producing 1000 t of oxygen on the moon and delivering as much of it as possible to LEO to supply a filling station for use by vehicles traveling from there to high Earth orbit, geocentric orbit, and beyond. A parametric economic model is designed more for flexibility than accuracy and is most useful for comparing relative costs and determining the best path to economic viability. Producing an element, oxygen, is relatively basic

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and thus allows easily modeled scenarios. The baseline scenario includes a lunar processing facility to manufacture oxygen, liquid oxygen (LO2) storage, a lunar habitat for workers, the required power system for the plant and habitat, and a transportation system and logistics to support the lunar factory and deliver its product. The costs are broken down into two categories, capital costs and operations costs. The capital costs include the development, emplacement, and installation of the lunar facility and transportation infrastructure. This cost is up front and then is amortized over 10 years. Operations costs include all the annual outlays needed to manufacture the 1000 t of LO2 and deliver as much as feasible to LEO. The operations costs include shipping the lunar oxygen to LEO. Considering the case where our only lunar product is LO2, if we assume our fuel is liquid hydrogen (LH2), then we must account for the costs of shipping 1 kg of LH2 from Earth for each 8 kg of LO2 used as propellant for the orbital transfer vehicle (OTV). Even though the OTV fuel is 80% lunar derived, Simon’s analysis found that the 20% LH2 shipped from Earth will dominate the economics because it must come “uphill” through Earth’s steep gravity well. Simon’s analysis assumes that a lunar base is in place. That is, you have some basic infrastructure on the moon onto which you can add the additional modules and power necessary to produce your product. This is a key assumption that gives oxygen production the potential for economic viability. This analysis and others [30] indicate that lunar oxygen production, in association with a lunar base, may be commercially viable if the lunar cost parameters are controlled. It speaks for the wisdom of a national or international effort to establish a lunar base as an anchor for the commercial development of Earth–moon (cislunar) space. These assumptions yield a baseline estimate (2010 dollars) of $6.2 billion capital and $1.8 billion per year in operations costs. Assuming that the OTV efficiencies allow 49% of the produced LO2 to reach LEO, a 10-year amortization of the capital, and the middle range of cost parameter estimates, the model estimates that LO2 could be delivered from the moon at about one-third the price that the space shuttle could ship it from Earth. If the worst-case cost parameters are assumed, then the costs from Earth to the moon are about the same. But the model is based on technical and programmatic assumptions. The strength of parametric analysis is that it allows assessment of the uncertainties of the cost estimates and the sensitivity of overall economics to each cost parameter. The result is that the principal capital cost driver is the power system that needs to be added to the lunar base infrastructure to operate the oxygen production facility, and the principal operational cost driver is the cost of transporting consumables (particularly LH2) from the Earth to the moon. Thus, if low-cost power and hydrogen were available, commercialization of lunar LO2 could become compelling. Let us now extend the analysis to include in situ coproduction of the most sensitive cost drivers of lunar LO2 production. Lunar polar observations since the Clementine mission have indicated relatively high concentrations of hydrogen

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in permanently shaded craters in the lunar polar regions. Methods are proposed [23] to produce solar arrays on the moon from lunar materials that could result in base power 10 times cheaper than solar arrays transported from the Earth. If these were combined with a more self-sufficient lunar base, then the costs of lunar LO2 to LEO could be well under 10 times less than supply costs from Earth. Thus, each additional space resource developed has a positive multiplier effect on the economic viability of the exploration and development of space. Let us examine the evidence for the presence of water ice on the lunar poles. In 1994 the U.S. Clementine spacecraft orbited the moon [31] and inspected the polar regions for the first time. The spacecraft beamed radio waves into the polar regions, which were then examined on Earth. The circular polarization ratio of the return radar was examined, and the results indicated the possibility of water ice in the permanently shadowed regions in lunar polar creators. The possibility of water ice existing in the lunar poles was hypothesized in the 1960s by Harrison Brown. Because the moon’s axis is stable within 1.5% relative to the sun, permanently shadowed craters exist where water from comet impacts could be stable (in a cold trap) for long periods of time. The planet Mercury has an axis that is similarly stable relative to the sun, and the presence of ice in permanently shadowed polar craters was confirmed by NASA’s Messenger mission in 2011–2012. A number of processes may be contributing to the presence of water ice in lunar polar crater cold traps. The possible sources include comets, meteorites, outgassing of the lunar interior, and water formation from the impact of the solar wind protons with the oxygenated lunar minerals. The water could form in the craters, or it could migrate by ballistic molecular trajectories across the lunar surface until it is arrested in a permanently shadowed crater near the lunar poles. The presence of water ice in polar craters was also consistent with neutron spectrometer data from the Lunar Prospector orbiter, which also indicated higher hydrogen composition at the lunar poles. However, the first definitive measurement of lunar ice was made by the Lunar Reconnaissance orbiter (LRO; 2009) and its impactor payload. The experiment included an impactor that used the LRO’s spent Centaur transfer stage, which yielded a 2300-kg kinetic payload that produced on impact 200 times the energy of the previous experiment conducted when the Lunar Prospector was crashed into the lunar polar region at the conclusion of its mission. The shepherding spacecraft first helped guide the impactor to its target and then measured the contents of the resulting cloud using imaging and spectroscopy at a distance of 5 km from the lunar surface. NASA’s LRO (2009) included the LCROSS experiment in which two lunar impacts were made in the Cebeus crater near the moon’s south pole. The significance of the presence of ice in craters at the lunar poles is twofold. First, the energy that would be required to make water on the moon is about 100 times less than that needed to make water from solar wind hydrogen. Second, the presence of nitrogen, carbon, and other light elements from comets supplements

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the deficiencies in the equatorial composition that was reported by Apollo missions. Stone describes [32, 33] a commercial venture, Shackleton Energy Co., which was formed to economically extract water from the lunar pole permanently shadowed craters to provide rocket fuel depots for cislunar operations. Water can be provided from the moon to LEO depots for 1/14 to 1/12 of the fuel needed for Earth launch. Stone estimates that about $20 billion (2009 dollars) would be necessary to set up mining operations on the moon. Financial viability assumes that: 1) inflatable structures are used for habitats and tanks, 2) multipass aero capture for the moon–Earth leg is used to minimize transport fuel percentage, and 3) the return propellant (even for the first trip) is produced on the moon. With this strategy the costs can be kept comparable with large oil development ventures on Earth. When inexpensive, lunar-derived fuel is available in depots in Earth–moon space, many ventures, such as lunar tourism and cleanup of orbital debris, become financially viable.

8.3.2 MARS IN SITU PROPELLANT PRODUCTION: MAKING A HUMAN MARS MISSION PRACTICAL To see the value of utilizing space resources for a Mars mission, one needs to only examine the exponential nature of the rocket equation [34] (also see Chapter 9), the energy available from chemical rocket propulsion to escape the Earth’s substantial gravitational field, and the resulting payload-to-fuel ratios. A rocket can have a gross weight-to-payload ratio in LEO of as much as 40 to 1. A trip to Mars’s surface could require about another 5 to 1 reduction [35]. Therefore, the gross liftoff weight-to-payload ratio to travel from Earth to Mars’s surface would be about 200 to 1. Thus, the manufacture of a needed item, for example, return propellant, in situ has a very high leverage for mission cost reduction. For example, a human Mars reference mission [36] was designed in which the return propellant was manufactured from the Martian atmosphere, which is 90% carbon dioxide (CO2) at less than 1% the atmospheric pressure on Earth. Utilizing CO2 from Mars’s atmosphere along with hydrogen feedstock brought from Earth (and in later scenarios extracted from Mars sources) was found to be technically feasible and provided a definitive cost and safety advantage. The technology required was found to be available and could be employed for the mission after a few years’ development. The concept relies on a “split mission” in which a propellant production plant is landed in a Mars landing opportunity two years before crew launch from Earth. When the plant produced the required methane propellant and oxygen for crew return to Mars orbit, the crew would be given the “go” for launch from Earth. In addition to the cost savings from in situ propellant production, the capability would enable the production of fuel for Mars surface transportation and provide a safety supplementation to oxygen life support. The Mars in situ propellant production facility can produce 30 t of LOX/CH4 propellant, which can

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provide the 5.6 km/s delta-V that is necessary to enable the crew to ascend to Mars orbit and rendezvous with the return vehicle. An additional advantage of a split mission using propellant production on Mars’s surface is that it assures a “safe haven” on Mars’s surface before the launching of a crew from Earth. This safe haven is also available for a Mars surface abort in case of mission problems in Mars’s vicinity. The alternative without Mars resource production is a six-month trip back to Earth. Of the three environments that a crew will experience (Earth orbit, deep space, and Mars surface), deep space is the most hazardous due to radiation and microgravity conditions. Orbit mechanics for a Mars mission impose a two- to three-year duration, once a crew is launched from Earth, before return to Earth is most efficient. Thus, because the transport costs of supplies to Mars is so high, the ability to produce air, water, and other life support on the Martian surface is a practical necessity. The technologies required to produce propellant from Mars CO2 are currently available and have been demonstrated in NASA laboratories. In general, the In Situ Resource Utilization (ISRU) technologies for Mars mission scenarios reduce the size and number of required Earth launches, provide dual-purpose infrastructure on Mars’s surface that support crew activities, and reduce risk. The NASA Mars Exploration Study Team lists resource utilization as an important technology for human space transportation, living in space, and planetary surfaces. These technologies include extraterrestrial mining, resource extractive processes, material preparation and handling in reduced gravity, and extraterrestrial manufacturing. The cost savings of a human Mars mission using Mars resources is about onetenth that estimated for the mission using only Earth-launched materials. This dramatic cost savings (from about $30 billion to about $3 billion per year) makes a human mission to Mars practical within contemporary NASA budgets [35]. A processing plant was designed (Fig. 8.8) that would land on Mars before the crew and make 17 times its weight in methane and liquid oxygen. Not a)

b)

Fig. 8.8 a) Artist’s rendition of propellant processing plant on Mars, b) production sequence for propellant production from Mars atmospheric carbon dioxide. (Courtesy NASA.)

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Fig. 8.9 Mars Phoenix Mission Lander arm camera of decent thruster blast area showing apparent subregolith ice. (NASA JPL.) having to carry the return propellant from Earth to Mars had the effect of allowing for smaller launch vehicles, and thus the total mission costs were reduced by a factor of 10. This savings allowed NASA to plan for human Mars missions, after the completion of the ISS, without substantial increases of the agency’s budgets. The costs of human Mars missions could be reduced further if hydrogen were recovered from Mars water ice. Thermodynamic calculations and orbital measurements from spectrometers on the Mars Reconnaissance and the Mars Express orbiters indicate that abundant water ice is present in the higher latitudes of Mars. This has been confirmed in situ by the NASA Jet Propulsion Laboratory (JPL) and University of Arizona Phoenix Lander mission, which landed 25 May 2008 at a latitude of 68.22 degrees in the northern arctic polar region of Mars. The presence of water ice was graphically illustrated by photographs (Fig. 8.9) from the lander’s arm camera of the Mars surface where the landing thrusters had blown away the surface regolith. The lander was designed to confirm the existence of subsurface water ice. The Phoenix 2.35-m robot arm with its associated soil acquisition device was used to excavate 12 trenches that uncovered an apparent water ice table under 5–18 cm of regolith.

8.3.3 OUTPOSTS VS SETTLEMENTS The degree of criticality for using space resources depends on the strategic structure of the campaign to explore and develop space. We usually term these campaigns as new living spaces, human bases, or human settlements. A space base can be defined [35] as an “outpost” or “base” where human crews stay for six months to two years and then leave or are rotated with a relief crew. The base

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is historically associated with a military occupation where a short-term expensive presence is required. A settlement, on the other hand, is historically a natural result of the proximity of a wealth-producing entity, for example, a mine, fertile land, or a commercial confluence of waterways. A settlement nominally is a place where people stay for two years to an indefinite period of time. A base relies on expensive supply from a sponsoring organization or government, but a settlement must be economically independent, relying on indigenous resources and trade. A base has by economic necessity the minimum of personnel, whereas a settlement grows more or less freely by procreation of its citizens and immigration of others to share in the wealth production. A base is ultimately unsustainable, whereas a settlement is self-sustainable. A base is often deliberately placed in an inhospitable location, whereas a settlement occurs naturally and spontaneously around places of opportunity. Thus, the economics of space (lunar, Mars, asteroid, or free space) settlements are much less burdensome to their sponsors than longterm bases would be. Sherwood and Woodcock analyze the economic input– output linear algebraic matrix for the case of a 100 t per year lunar oxygen production facility that is operated under an 18-person base mode or a more selfsufficient settlement mode. For financial viability the settled lunar habitats and infrastructure must also be constructed in situ with local resources. They conclude that the resupply needed for the oxygen production facility is a factor of 6 less when operated by a settled workforce. The small crew required to produce 100 t of oxygen would not justify the infrastructure needed for settlement; however, if other projects (e.g., the extraction of He3) were occurring and the productivity kept very high through advanced automation, the financial case for a settled lunar workforce becomes compelling. 8.3.3.1

PROCESSING SPACE MATERIALS ON A GRAND SCALE: ENERGY TO EARTH AND SPACE HABITATION

Two of the main themes that this section builds upon are that space labor and space resources are cheaper than Earth-based labor and resources to achieve large-scale monetization of space. The use of space resources is imperative to enable economically reasonable space exploration [35], settlement [37], and space industrialization and solar power development [15]. Launching advanced lightweight solar power satellites from Earth can be cheaper than using space resources when few units are launched [38]; however, Earth-launched solar power satellites, even using advanced technologies, are not economically competitive [39] unless Earth-launch costs can be greatly reduced. Space solar power satellites might be constructed autonomously from Earth but only with significant advancements in technology [40]. Using space resources, such as metals from the moon, as construction materials has a high initial cost but a much lower cost per unit produced and transported into space. This is shown graphically in Fig. 8.10, which shows the difference in cost when bringing just 50 kilotons per year to fifth Earth–moon Lagrangian point, L5, from Earth or the moon.

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Cumulative cost for sending 50 kt/year from the Earth or the moon.

The cost from the moon includes the $283 billion required for necessary lunar infrastructure. Detweiler and Curreri [41] analyzed the economics of space workers in outposts vs in settlements for the large-scale manufacturing needed to build solar power satellites to supply energy to Earth. The calculations show that for large space endeavors, space-based labor is far cheaper than Earth-based labor. Space labor in this case is the labor of people living in habitats in free space. The habitats are shielded and rotated to provide near-Earth normal radiation and gravity. The community in space creates its own products and grows its own food. The Earthbased workforce uses terrestrial supplies and bears the exorbitant costs of transporting them from Earth to space. Because of the high costs, the outpost (Earthbased) workers live in spartan conditions and must be rotated back to Earth every six months, and their salary must also be supplied from Earth. Workers living in a near-independent and moderately comfortable permanent space habitat are paid mostly using goods constructed in space. Each settlement habitat could have a large agriculture section in which food is grown. Living in permanent settlement habitats also has the added benefit of attracting individuals who are betting their future on the project. The economic consequence of this can be seen in Fig. 8.11. The cost of Earth-based workers is obtained by multiplying 614 workers by their wages of $38,420 and the cost of buying and sending 1.67 t (at $19.11 per kg) of resupply material from Earth. The cost of space-based labor is obtained by the cost of building and maintaining one space habitat for 614 workers and paying each space settler the equivalent to launching 100 kg from Earth. In the habitats for the space settlers, only a fraction of the inhabitants works in construction. The habitat in which only 44% are workers will be larger than the

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Fig. 8.11

Cost for 614 Earth-based workers vs space-based labor.

habitat in which 80% are workers. Examining Fig. 8.11, it is apparent that the costs of using Earth-based personnel (salary, transportation, and resupply costs) quickly exceed the costs of using space-based labor housed in the space habitats after less than a year of operation. Using space settlers instead of temporary workers not only begins the human settlement of space, but also provides much cheaper, more comfortable, and more dedicated workers for large-scale development of space. 8.3.3.2

ECONOMIC COMPARISON OF EARTH-LAUNCHED AND SPACE-DERIVED SOLAR POWER SATELLITES

Studies [39, 42] have attempted to achieve a viable financial model for Earth launches for the Glaser 1970 solar power satellite (SPS) design. For comparison to the O’Neill–Glaser financial model, which combines SPS and space habitat construction, a comprehensive and well-documented financial model was a NASA-led program in the 1990s commonly called the “Fresh Look Study” [38]. That study employed more advanced technologies (than Glaser’s) that included remote robotic assembly, phase array microwave pointing, state-of-the-art solar cells, and high-temperature superconductor bus lines. Figure 8.12 shows an artists’ renditions of the leading design concept Solar Disk and Sun Tower Earthlaunched SPSs compared to a Glaser-era satellite being constructed in the vicinity of an O’Neill habitat. To keep the initial costs as low as possible, the “Fresh Look Study” did not consider lunar materials, and the number of humans in space was

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Fig. 8.12 Artist’s conceptions of Solar Disc (https://commons.wikimedia.org/wiki/ File:Solardisk.jpg) (left) and Sun Tower (https://commons.wikimedia.org/wiki/ File:Suntower.jpg) (center) Earth-launched SPS and a Glaser-era space-derived SPS (https:// spaceflight.nasa.gov/gallery/images/exploration/marsexploration/html/s78_27139.html) (right) with O’Neill habitat and near-Earth asteroid miner. (Courtesy of NASA) kept as low as possible. It was found that with these advanced technologies, SPS design might be financially viable and even Earth launch costs could be aggressively reduced. The evolution of the economics of the two Earth-derived models (Sun Tower and Solar Disc) and the space-derived optimized O’Neill–Glaser model is shown in Fig. 8.13. The Earth-launch models show that after six or seven years of relatively low-cost research and development, the SPS launching begins. However, even assuming the optimistic launch costs of $400/kg, the 78 GW Sun Tower program requires 30 program years to break even, after which there is only a modest projected profit. The 30 GW Solar Disc program is even less favorable

Fig. 8.13 Economics vs time for proposed Earth-launch and space-derived SPS construction models.

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economically, with financial outlays of more than $100 billion and a projected financial break-even after 37 years. The space-derived (optimized O’Neill–Glaser) model requires an investment of $300 billion to build the infrastructure on the moon and in space to allow use of lunar materials and space-based labor. Once the infrastructure is in place, however, the SPSs are constructed more than an order of magnitude cheaper than by Earth launch. These inexpensive SPSs have a very large profit margin even when selling electricity below Earth market values. The peak investment of about $460 billion is paid back at program year 24, and by program year 27 the project is projected to realize $600 billion in profits.

8.4 SUMMARY AND CONCLUSIONS Materials science and processing research in space is a field of study that began with the sounding rocket experiments in the 1950s. Materials science studies of the lunar surface materials returned during the Apollo missions enabled the study of lunar resource utilization. The study of materials science and processing in space progressed with more than 30 years of microgravity materials processing research, which continues today in the ISS. These studies are the technical foundation that could enable lower cost human exploration through the use of in situ propellant production, the production of energy from space resources, and the eventual establishment of self-sufficient human settlement off Earth.

REFERENCES [1] Spudis, P. D., and Lavoie, A. R., “Using the Resources of the Moon to Create a Permanent, Cis Lunar Space Faring System,” AIAA Space 2011 Conference & Exposition, AIAA 2011–7185, AIAA, Reston, VA, 2011. [2] Curreri, P. A., and Detweiler, M. K., “A Contemporary Analysis of the O’Neill– Glaser Model for Space-Based Solar Power and Habitat Construction,” National Space Society Space Settlement Journal [online journal], Dec. 2011, http://www.nss. org/settlement/journal/NSSJOURNAL_AnalysisOfONeill-GlaserModel_2011.pdf [retrieved 09 Mar. 2018]. [3] Detweiler, M. K., and Curreri, P. A., “The Space Homestead and Creation of Real Estate and Industry Beyond Earth,” Space Technology and Applications International Forum 2008, edited by M. S. El-Genk, American Institute of Physics Conference Proceedings CP969, Springer New York, pp. 925–933. [4] Curreri, P. A., and Stefanescu, D. M., “Low-Gravity Effects During Solidification,” Metals Handbook: Casting, Vol. 15, 9th ed., American Society of Metals International, Metals Park, OH, 1988, pp. 147–158. [5] Charles, D., “NASA’s Busload of Science,” Science, Vol. 333, 1 July 2011, pp. 28–33. [6] Curreri, P. A., “Processing of Space Resources to Enable the Vision for Space Exploration,” Transactions of the Indian Institute of Metals, Vol. 60, No. 2, April 2007, pp. 99–102.

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[7] Heiken, G. H., Vaniman, D. T., and French, B. M. (eds), Lunar Sourcebook: A User’s Guide to the Moon, Cambridge Univ. Press, Cambridge, England, U.K., 1991. [8] Witt, A. F., Gatos, H. C., Lichtensteiger, M., Lavine, M. C., and Herman, C. J., “Steady State Growth and Segregation Under Zero Gravity: InSb,” Proceedings of the Third Space Processing Symposium Skylab Results: Vol. I & II, NASA Report No. M-74-5, p. 275, 1974. [9] Downey, J. P. (ed.), “Life and Microgravity Spacelab Final Report,” NASA CP-1998-206960, 1998. [10] Ethridge, E. C., Curreri, P. A., and McCauley, D. E., “Fourth United States Microgravity Payload: One Year Report,” NASA CP-1999-209628, 1999. [11] Curreri, P. A., Ethridge, E. C., Hudson, S. B., Miller, T. Y., Sen, S., and Sadoway, D. R., “Processing Demonstration for Lunar In-Situ Resource Utilization: Molten Oxide Electrolysis,” NASA TM-2006-214600, 2006. [12] Stefanescu, D. M., Grugel, R. N., and Curreri, P. A., “In-Situ Resource Utilization for Processing of Metal Alloys on Lunar and Mars Bases,” Space 98: The Sixth International Conference and Exposition on Engineering, Construction and Operations in Space, edited by R. G. Galloway, and S. Lokaj, American Society of Civil Engineers, Reston, VA, 1998, pp. 266–274. [13] McGannon, H. E. (ed.), The Making, Shaping and Treating of Steel, 9th ed., United States Steel Corporation, Herbick & Held, Pittsburgh, PA, 1971, p. 3. [14] Buchwald, V. F., Handbook of Iron Meteorites: Their History, Distribution, Composition and Structure, Vols. 1–3, Univ. of California Press, Oakland, CA, 1975. [15] Economou, Thanasis, “Chemical analysis of martian soil and rocks obtained by the Pathfinder Alpha Proton X-ray spectrometer,” Radiation Physics and Chemistry, Vol. 61, 2001, pp. 191–197. [16] Criswell, D. R., and Waldron, R. D., “Lunar Utilization,” Space Industrialization, Vol. II, edited by B. O’Leary, CRC Press, Boca Raton, FL, 1982, Chap. 1. [17] Rao, D. B., Bhogeswara, D., Choudary, U. V., Erstfield, T. E., Jwilling, R., and Chang, Y. A., “Space Resources and Space Settlements,” NASA SP-428, V-5, 1979, pp. 257–274. [18] Hoekje, J. J., and Kearley, A. A., Titanium Dioxide from Ilmenite by Caustic Fusion, British Patent 846,468, 31 August 1960. [19] Daubenspeck, J. M., and Schmidt, C. L., Removing Iron from Titanium Ores with Chlorine in a Fluidized Bed, U.S. Patent 2,852,362, 1959. [20] Criswell, D. R., and Waldron, R. D., “Materials Processing in Space,” Space Industrialization, Vol. I, edited by B. O’Leary, CRC Press, Boca Raton, F., 1982, Chap. 5. [21] “Summer Workshop on Near-Earth Resources,” NASA CP 2031, 1977. [22] Sanders, G., Romig, K. A., Larson, W. E., Johnson, R., Rapp, D., Johnson, K. R., Stacksteder, K., Linne, D., Curreri, P. A., Duke, M., Blair, B., Gertsch, L., Boucher, D., Rice, E., Clark, L., McCullouh, E., and Zibrin, R., “Results from the NASA Capability Roadmap Team for In-Situ Resources,” Seventh International Conference on the Exploration and Utilization of the Moon, ILC 2005, 18–23 Sept. 2005, Toronto, Canada, online http://sci.esa.int/ilewg/43268-abstracts-and-presentations/, accessed 09 Mar. 2018, European Space Agency, ESA Publications, Noordwijk, The Netherlands. [23] Freundlich, A., Ignatiev, A., Horton, C., Duke, M., Curreri, P. A., and Sibille, L., “Manufacture of Solar Cells on the Moon,” Conference Record of the Thirty-First

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IEEE Photovoltaic Specialists Conference, 3–7 Jan. 2005, Institute of Electrical and Electronics Engineers, IEEE, New York, pp. 794–797. Sibille, L., Sadoway, D. R., Sirk, A., Tripathy, P., Melendez, O., Standish, E., Doninguez, J. A., Stefanescu, D. M., Curreri, P. A., and Poizeau, S., “Recent Advances in Scale-up Development of Molten Regolith Electrolysis for Oxygen Production in Support of a Lunar Base,” Proceedings of the 47th AIAA Aerospace Sciences Meeting, 5–8 Jan. 2009, American Institute of Aeronautics and Astronautics, AIAA, Reston, VA. Paley, M. S., Karr, L. J., Marone, M., and Curreri, P.A., “Oxygen Production from Lunar Regolith Using Ionic Liquids,” Space Propulsion and Energy Sciences International Forum, NASA MSFC–2200, NASA Marshall Space Flight Center, Huntsville, AL, 2009. Sullivan, T. A., and McKay, D. S., “Utilizing Space Resources,” NASA Johnson Space Center, Houston, TX, 1991. Carter, J. L., “Lunar Regolith Fines: A Source of Hydrogen,” Lunar Bases and Space Activities of the 21st Century, edited by W. Mendell, Lunar and Planetary Institute, Houston, TX, 1985, pp. 571–582. O’Neill, G. K., “Space Colonies and Energy Supply to Earth,” Science, Vol. 10, 1975a, pp. 943–947. Simon, M. C., “A Parametric Analysis of Lunar Oxygen Production,” Lunar Bases and Space Activities of the 21st Century, edited by W. Mendell, Lunar and Planetary Institute, Houston, TX, 1985, pp. 531–541. Blair, B. R., Diaz, J., and Duke, M. B., “Space Resource Economic Analysis Toolkit: The Case for Commercial Lunar Ice Mining,” Final Report to NASA Exploration Team, 20 Dec. 2002. http://www.nss.org/settlement/moon/library/2002-Case ForCommercialLunarIceMining.pdf Spudis, P. D., “Ice on the Moon,” Space Review.com, www.thespacereview.com/ article/740/1 [retrieved 6 Nov. 2006]. Stone, W., “How the Extraction of Lunar Hydrogen or Ice Could Fuel Humanity’s Expansion into Space,” IEEE Spectrum, Inside Technology Feature, Mining the Moon, 11 June 2009, http://spectrum.ieee.org/aerospace/space-flight/mining-themoon [retrieved 09 Mar. 2018]. Sherewood, B., and Woodcock, G. R., “Cost and Benefits of Lunar Oxygen: Economics, Engineering and Operations,” Resources of Near Earth Space, edited by J. Lewis, M. S. Matthews, and M. L. Guerrieri, Boeing Defense Group, Berkeley Missouri, 1993, pp. 199–227. Noordung, H., “The Problem of Space Travel: The Rocket Motor,” edited by E. Stuhlinger, and J. D. Huntley, NASA SP-4026, NASA History Office, Washington, D.C., 1995. Rapp, D., “Refueling of Mars-Bound Vehicles in LEO with Propellants Derived from Lunar Resources,” Report by Skillstorm, Inc. in affiliation with Jet Propulsion Laboratory, 2005. Hoffman, S. J., and Kaplan, D. L., “Human Exploration of Mars: The Reference Mission of the NASA Mars Exploration Study Team,” Lyndon B. Johnson Space Center, Houston, TX, 1997. NASA SP 6107 Human Exploration of Mars: The Reference Mission of the NASA Mars Exploration Study Team, Stephen J. Hoffman,

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and David I. Kaplan, eds. JSC, July 1977. http://exploration.jsc.nass.gov/marsref/ contents.html. Billingham, J., Gilbreath, W., and O’Leary, B. (eds), Space Resources and Space Settlements, NASA SP-428, NASA, Washington, D.C., 1979. Mankins, J. C., “A Fresh Look at Space Solar Power: New Architectures, Concepts, and Technologies,” International Astronautical Federation IAF-97-R.2.03, 1997. Geuder, N., Quaschning, V., Viebahn, P., Steinsiek, F., Spies, J., and Hendriks, C., “Comparison of Solar Terrestrial and Space Power Generation for Europe,” 4th International Conference on Solar Power from Space: SPS’04, 30 June–2 July 2004, Granada, Spain, European Space Agency, ESA Publications, Noordwijk, The Netherlands, 2004. Bekey, G., Bekey, I., Criswell, D., Friedman, G., Greenwood, D., Miller, D., and Will, P., “Final Report of the NSF-NASA Workshop on Autonomous Construction and Manufacturing for Space Electrical Power Systems,” 4–7 April, 2000, Arlington, VA, http://www.nss.org/settlement/ssp/library/2000-FinalReport-Autonomous ConstructionAndManufacturing.pdf. [retrieved 09 Mar. 2018], National Aeronautics and Space Administration, NASA, 2000. Detweiler, M., and Curreri, P. A., “Habitat Size Optimization of the O’Neill–Glaser Model for Space Solar Satellite Production,” Space Manufacturing 14: Critical Technologies for Space Settlement, NASA Marshall Space Flight Center, Huntsville, AL, 2010. Rouge, J. D., “Space-Based Solar Power as an Opportunity for Strategic Security,” Report to the Director, National Security Space Office, 10 Oct. 2007, http://www.nss. org/settlement/ssp/library/final-sbsp-interim-assessment-release-01.pdf [retrieved 9 Mar. 2018].

CHAPTER 9

Thermal Protection Systems and Hot Structures for Hypersonic Vehicles David E. Glass† NASA Langley Research Center, Hampton, Virginia

Thermal protection systems (TPS) and hot structures are required for a range of hypersonic vehicles ranging from ballistic reentry to hypersonic cruise vehicles, within both Earth’s atmosphere and non-Earth atmospheres. The focus of this chapter is on air-breathing hypersonic vehicles in the Earth’s atmosphere. This includes single-stage-to-orbit (SSTO) and two-stage-to-orbit (TSTO) accelerators, access to space vehicles, and hypersonic cruise vehicles. This chapter will start out with a brief discussion of aerodynamic heating and thermal management techniques to address the high heating, followed by an overview of TPS for rocket-launched and air-breathing vehicles. The argument is presented that as we move from rocket-based vehicles to air-breathing vehicles, we need to move away from the “insulated airplane” approach used on the Space Shuttle Orbiter to a wide range of TPS and hot structure approaches. The primary portion of the chapter will discuss issues and design options for ceramic matrix composite (CMC) TPS and hot structure components, including leading edges, acreage TPS, and control surfaces. The current state of the art will be briefly discussed for some of the components. The two primary technical challenges impacting the use of CMC TPS and hot structures for hypersonic vehicles are environmental durability and fabrication, and they will be discussed briefly.

9.1 INTRODUCTION The Hypersonics Project within Fundamental Aeronautics Program at NASA has as one of its reference systems TSTO access to space vehicles. The concept of operations for this particular class of vehicles is illustrated in Fig. 9.1 as a function of altitude vs Mach number.  This

chapter was originally presented at NASA Headquarters in March 2007 as part of the NASA Aeronautics Research Mission Directorate Technical Seminar series. It is intended as a high-level tutorial on CMC TPS and hot structures. † Structural Mechanics and Concepts Branch, MS 190, AIAA Associate Fellow.

This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Fig. 9.1

TSTO concept of operations as a function of altitude and Mach number.

The TSTO will accelerate to a staging Mach number (M ), indicated by the shaded area, in the range 6  M  13. At that point, the first stage and second stage separate. The first stage continues on to orbit powered by the rocket booster, and the first-stage air-breather returns to the launch site. The propulsion system is a turbine-based combined cycle (TBCC) engine with both a turbine, operating up to around M ¼ 3.5, and a scram-jet engine operating up to around M ¼ 12. The maximum Mach number for the turbine and the minimum and maximum Mach numbers for the scram-jet are dependent on their demonstrated capability. These air-breathing vehicles are usually optimized to accelerate along a constant dynamic pressure. Dynamic pressure is a function of the density of the air times the velocity squared, and the aerodynamic heating is a function of velocity cubed. Lines of constant dynamic pressure shown in Fig. 9.1 correspond to 100 psf, 500 psf, and 2000 psf. When flying low in the atmosphere at high velocities, high dynamic pressure and high aerodynamic heating are obtained. One of the challenges for airbreathing vehicles is to balance the desire to fly at a high dynamic pressure, resulting in both increased performance due to increased airflow into the engine with the desire to limit aerodynamic heating to the leading edges and airframe. Figure 9.2 shows the aerodynamic heating on a reference 1-ft-diameter sphere for three different trajectories: a trajectory flown by a crew exploration vehicle (CEV) capsule entering from low Earth orbit, an SSTO air-breather during ascent, and the Space Shuttle Orbiter during descent. The heating shown in the figure is not the actual heating observed on the vehicle but is representative of a 1-ft-diameter sphere. For example, consider a basketball (1 ft diameter) and “fly” it through each of the three different trajectories. For the CEV, a very large spike in the heat flux is evident, but for a relatively short time. For the

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.2

533

Heating on a reference 1-ft-diameter sphere for three different trajectories.

Space Shuttle Orbiter, the heating occurs over a much longer time. The SSTO vehicle has a long duration trajectory and a high heat flux. The rocket boost is shown by the dashed line and is accelerating the second stage into orbit. For airbreathing vehicles, the radius of the leading edges will be much smaller than 1 ft. As discussed later in this chapter, the heat fluxes will be much higher than shown in Fig. 9.2 because the heat flux is a strong function of the radius. Heat load is another important parameter. If the heat flux is integrated over time, the heat load, which is the total heat into the system, is obtained. The heat load is shown in Fig. 9.3 for the three different trajectories for the CEV capsule, the SSTO air-breather, and the Space Shuttle Orbiter on reentry.

Fig. 9.3

Heat load on a reference 1-ft-diameter sphere for three different trajectories.

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9.2 BACKGROUND There are multiple options for dealing with the severe thermal environments encountered during hypersonic flight. Passive, semipassive, and actively cooled approaches can be used. Although the differences between rocket and airbreathing propulsion systems are well known, the differences between rockets and air-breathers can have a significant impact on the TPS and hot structures. As we move toward air-breathing hypersonic vehicles, the severe thermal structural challenges require a new approach to thermal management, one that includes both TPS and hot structures.

9.2.1 THERMAL MANAGEMENT There are three types of thermal management that can be used to cool hypersonic vehicles: passive, semipassive, and active. Passive and semipassive thermal management may include a phase change. Actively cooled structures would include a pumped coolant. 9.2.1.1

PASSIVE

An insulated structure is used for moderate heat fluxes for relatively short periods of time. The figure on the left in Fig. 9.4 shows surface heating, with thermal radiation as the mechanism to remove the heat. The objective of the insulation is to minimize heat reaching the structure, which remains cool. The high emittance coating on the surface enables most of the heat to be radiated away, with only a small amount of heat conducted through the insulation to the structure. As an example of an insulated structure, a photograph of the Space Shuttle Orbiter elevons, taken on the STS-116, is shown in Fig. 9.4. The heat sink structure, shown in Fig. 9.5, is used with a moderate heat flux for a transient situation. When surface heating occurs, some heat is radiated

Fig. 9.4 Schematic and photograph (Space Shuttle Orbiter elevons) of an insulated structure.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.5

535

Schematic and photograph (X-15) of a heat sink structure.

away and some heat is absorbed by the structure. If a heat sink structure is heated for a long period of time, enough heat could be absorbed to overheat the structure. The example shown is the X-15, where the leading edges used a heat sink structure. Figure 9.6 illustrates a hot structure used for moderate heat flux for long periods of time. In contrast to a heat sink structure, hot structures can be used for a higher heat load for long periods of time, allowing the structure to reach steady-state conditions. Again, the heat is both radiated away and conducted inward. The entire structure will increase to elevated temperatures. 9.2.1.2

SEMIPASSIVE

If high heat fluxes persist over a small region, such as a leading edge, for long periods of time, a semipassive approach may need to be used, such as a heat pipe. Heat is transferred by a working fluid to another region of the heat pipe where the heat is radiated away. As with hot structures, the structure operates hot. An example, shown in Fig. 9.7, is a heat-pipe-cooled wing leading edge. Ablation is another semipassive approach to thermal management. The purpose of the ablator is to keep the structure cool. Ablators are for very high

Fig. 9.6

Schematic and image of a hot structure.

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Fig. 9.7

Schematic and photograph illustrating a heat-pipe-cooled leading edge.

heat fluxes but for relatively short periods of time and for single use. As an example, the Orion capsule (shown in Fig. 9.8) includes an ablative heat shield. This is the same approach that was used on the Apollo capsule, where the heat is blocked by ablation and the ablator is consumed. Heat is also absorbed by the ablation process. 9.2.1.3

ACTIVE

For still higher heat fluxes and for long periods of time, active cooling is required. Convective cooling is often used. An example shown in Fig. 9.9 is the space shuttle main engine (SSME). Here, convective cooling is used in the propulsion system, where the heat is transferred into the coolant. The coolant heats up and carries the heat away. The structure operates hot but is maintained within its temperature use limits by the active cooling. Film cooling, used inside a propulsion system, is another approach used for high heat fluxes for long periods of time. For film cooling, the coolant is injected into the flow, usually at an upstream location and at a single, discrete location

Fig. 9.8

Schematic and illustration of an ablative heat shield.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.9

537

Schematic and photograph of actively cooled structure (SSME).

(Fig. 9.10). It operates as a thin, cool, insulating blanket. The structure will reach high temperatures and operate hot. Transpiration cooling is the final active cooling approach. This cooling method is used for high heat fluxes and long periods of time. An example in Fig. 9.11 shows a cooled carbon/carbon (C/C) combustion chamber test article. For transpiration cooling, the coolant is continuously injected into the hot gas flow through a porous structure over large areas as opposed to a discrete location with film cooling. The coolant also decreases the heat flux to the structure, and again the structure operates hot. Different thermal management techniques are applied to different flight vehicles. Figure 9.12 shows the temperature vs exposure time for multiple vehicles. Shown in the top left of Fig. 9.12 are reentry capsules Apollo/CEV (Orion) and Mercury, which all use(d) ablators due to very high temperatures but relatively short times. The X-15 used a heat sink approach, due to short times and relatively moderate temperatures. The SR-71 experienced longer flight times resulting in a hot structure design, and, again, the temperatures were moderate. The Space Shuttle Orbiter is an insulated aluminum airframe with moderate temperatures

Fig. 9.10

Schematic of film cooling and drawing of a hypersonic vehicle.

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Fig. 9.11 Schematic of transpiration cooling and a C/C cooled combustion chamber test article [1]. and exposed for longer times. Moving forward with air-breathing hypersonic vehicles, as illustrated by an SSTO vehicle, a new approach will be needed to thermally protect the vehicles. The new approach will inevitably encompass a wide range of different approaches, including hot structures, insulation, and active cooling.

9.2.2 THERMAL PROTECTION SYSTEMS FOR ROCKET-LAUNCH VEHICLES The Space Shuttle Orbiter, shown in Fig. 9.13, uses a conventional, skin-stringer aluminum aircraft structure. The structural temperatures are required to stay 5000

Apollo/ CEV

4000

Hot Structure Insulation Active Cooling

Ablators

Mercury 3000 T, ¡F

SSTO

Insulation

2000 Heat sink

Shuttle

1000 Hot structure

X-15 0

SR-71 1 Exposure Time, hr

2

Fig. 9.12 Thermal management approach for several vehicles as a function of temperature and exposure time [2].

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.13

539

Nominal maximum temperatures on the surface of the Space Shuttle Orbiter [2].

below 3508F for reuse purposes. To keep the temperatures down to 3508F, reusable surface insulation tiles are used primarily on the windward surface, and reusable blankets are used primarily on the leeward surface. For the leading edges and the nose cap, where temperatures are greater than 23008F, reinforced carbon/ carbon (RCC) is used. The X-33 was a suborbital experimental vehicle that NASA funded (but canceled before flight) that was intended to be the predecessor for an SSTO rocket vehicle. The TPS on the X-33 was similar to the Space Shuttle Orbiter TPS, except that it had a metallic TPS on the windward surface. As shown in Fig. 9.14, blankets were used on the leeward surface, and a metallic TPS was used on the windward surface of the X-33. Both vehicles use(d) C/C leading edges, nose cap, chin, and skirt.

Fig. 9.14

Thermal protection system used on the X-33.

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9.2.3 TPS AND HOT STRUCTURES FOR AIR-BREATHING HYPERSONIC VEHICLES For an air-breathing vehicle, the aerothermodynamics, propulsion system, and airframe, including much of the TPS and hot structures, are highly integrated, as shown in Fig. 9.15 where the entire underside of the vehicle is part of the propulsion system. This is very different from most rocket-based systems. The differences between rocket-based and air-breathing vehicles have significant impacts on the airframe structures. The differences in rockets and air-breathers (Fig. 9.16) impacts how thermal management and TPS are handled. Rockets accelerate but do not cruise while in atmospheric flight and are usually launched vertically. They leave the atmosphere quickly (“up and out”) and generally fly at a low dynamic pressure. Conversely, air-breathers accelerate and cruise in the atmosphere. They are often launched (or take off) horizontally and fly at a high dynamic pressure because they fly low in the atmosphere at high velocities to capture the air for the engine. For rocket-based vehicles (such as the Space Shuttle Orbiter), high drag is not a problem on ascent and is desired on descent for deceleration. Air-breathers, on the other hand, are optimized for low drag and thus have thin, slender bodies with low thickness-to-chord ratios. Weight and volume are another difference between rockets and air-breathers. Rockets are extremely weight sensitive, and the structural mass fraction is on the order of 10% of the gross takeoff weight (GTOW). This means that if 1 lb of structure is added to a rocket-based vehicle, then 10 lbs of weight will be added to the total system. Air-breathers tend to be more volume sensitive, which impacts the drag. The structural mass fraction is in the range of 30% of the GTOW. Weight is also extremely important for air-breathers, but the volume is what drives the drag.

Fig. 9.15

Image of a hypersonic vehicle highlighting multiple disciplines.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.16

541

Typical rocket and air-breathing vehicles.

The difference in trajectories impacts the TPS. Rockets are driven by high descent heating. Ascent heat loads are generally low due to the short ascent time and the “up and out” trajectory. Air-breathers are driven by ascent, descent, and cruise and are exposed to high heat loads due to their long ascent times at high dynamic pressure. On a rocket-based vehicle (Shuttle), leading edges tend to be blunt due to a desire for a high descent drag and low heat flux. Air-breathers have sharp leading edges due to the desire for low drag and low thickness-to-chord ratio. The heat fluxes are high because of the sharp radii (Fig. 9.17) and high dynamic pressure.

9.2.4 THERMAL–STRUCTURAL CHALLENGES Thermal–structural challenges can be quite severe on air-breathing vehicles. One of the primary thermal–structural challenges results from large thermal gradients. For example, in a cryogenic tank containing liquid hydrogen as a fuel, the liquid hydrogen will be –4238F, and the outer surface of the TPS may be between 20008F

Fig. 9.17

Rocket-based blunt leading edges vs sharp leading edges on X-43.

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and 30008F. With different materials operating at a wide range of temperatures, attaching the various components (tank, insulation, structure, TPS, etc.) that are growing and shrinking at different magnitudes is challenging. Control surfaces are often hot and are often connected to an actuator inside the vehicle that is much cooler. On some structures, there are thin cross sections (due to the need for low drag) at high mechanical loads. These high mechanical loads are often imposed at elevated temperatures. The stability of the outer mold line is important because it can impact performance. For example, sharp nose leading edges generate shocks, which are necessary to maximize airflow into the engines. As a result, leading edges should not ablate, both to generate the desired shocks and to enable reuse. Steps and gaps can be a problem. Gaps may potentially allow sneak flow, where hot plasma leaks into the structure. Forward-facing steps may result in local hot spots, thus locally increasing the surface temperature. Thermal expansion of the propulsion system creates other issues. On air-breathers, the propulsion system includes much of the undersurface of the vehicle, is very long, and can grow several inches. It must be attached to the airframe and allow for the differential growth between the propulsion system and the airframe. Other vehicle and structural challenges include affordability. Production costs, life-cycle costs, and inspection and maintenance costs are all important. Other issues to consider are damage tolerance, low-speed impact such as tool drop, foreign object damage on a runway, hypervelocity impact, weather, and reuse potential. All of these requirements lead to a new approach to thermal protection and move us away from the Space Shuttle Orbiter type of insulated airframe. History shows the use of new material systems enables vehicles. Figure 9.18 illustrates the SR-71 where titanium was used. The manufacturing processes and

Fig. 9.18

Examples of where new material systems have helped enable new vehicles.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

543

Fig. 9.19 Material specific strength as a function of temperature for several material classes. databases had to be extended to higher temperatures to enable its use. The same is true for the X-15 with Inconel, where manufacturing processes and databases had to be extended to higher temperatures to enable its use. The Space Shuttle Orbiter is an aluminum “airplane” with ceramic tiles and blankets, which were developed to enable this type of vehicle to be flown. The required material attributes for hypersonic air-breathing vehicles are high temperature capability (20008–40008F), high strength at those elevated temperatures, high toughness, light weight, and environmental durability. Figure 9.19 shows material specific strength, which is strength divided by density, vs temperature. High specific strength at elevated temperature is the goal, which is obtained with high strength and low density and is shown in the shaded region in the top right-hand corner of the figure. Metallic options include metallic matrix composites (MMCs), superalloys, and titanium. These all have good specific strength, but it drops off by the 20008F range. Grouped together as CMCs, the C/silicon carbide (SiC) material, advanced carbon/ carbon (ACC), and SiC/SiC provide high strength at elevated temperatures, which is key for air-breathing vehicles. A CMC is illustrated in Fig. 9.20, where fibers inside of a matrix are evident. The fibers carry the load, and the matrix transfers the external load to the fibers. Depending on the processing and fabrication approaches, the fiber/matrix interface could have an interface coating that, if present, is the key to the toughness of the structure. It provides a weak mechanical interface for increased toughness and graceful failure. One of the key components of the CMC is the environmental barrier coating, which protects the materials from oxidation at high temperatures.

9.3 TPS AND HOT STRUCTURE COMPONENTS As discussed previously, much of the outer surface of hypersonic vehicles is subjected to severe aerodynamic heating. As a result, the airframe must be either protected from the heat or designed to operate even when exposed to extreme

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Fig. 9.20

Schematic of a ceramic matrix composite with fibers inside of a matrix.

heating. The airframe components discussed here include leading edges, acreage TPS and aeroshells, and control surfaces.

9.3.1 LEADING EDGES Figure 9.21 shows typical ascent leading-edge heat flux for an SSTO vehicle. The wing leading edge heat flux is 500 Btu/ft2-sec and the nose 5000 Btu/ ft2-sec. The cowl leading edge maximum heat flux is 50,000 Btu/ft2-sec, due to a Type IV shock–shock interaction. In comparison, the Space Shuttle Orbiter wing leading-edge maximum heat flux is 70 Btu/ft2-sec, and the

Fig. 9.21

Typical ascent leading-edge heat flux for an SSTO vehicle.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.22

545

Effect of radius on heat flux.

CEV heat shield maximum heat flux on lunar direct return is expected to be 700 Btu/ft2-sec. The leading-edge radius has a significant impact on the heat flux. Heat flux is inversely proportional to the square root of the radius. Figure 9.22 shows the heat flux as a function of radius, with a 1-in. radius having a heat flux of 500 Btu/ft2-sec. As the radius increases, the heat flux drops. As the radius becomes smaller, which is required for most air-breathing vehicles, the heat flux increases significantly. Sharp leading edges impact the chordwise heat flux distribution, and the radius impacts the maximum value. Area “A” in Fig. 9.23 shows the distribution around a sharp leading edge of a hypersonic vehicle. The stagnation line, indicated by the arrow “D”, would be basically at zero angle-of-attack. For a blunt leading edge, such as on the Space Shuttle Orbiter and shown in area “B”, the heating distribution is more spread out around the leading edge. Also, for the Space Shuttle Orbiter distribution, the maximum heat flux (arrow “C”) is on the lower surface due to a high angle-of-attack. The net effect of the sharp leading edge is thus an intense, localized heating. For sharp leading edges, passive, semipassive, and active thermal management options may be used to manage the intense localized heating. In Fig. 9.24, for passive leading edges, reuse limits are indicated by the horizontal line, below which the materials can be used for multiple missions. Above that temperature, the material becomes an ablator and thus single-use. The sequence of preferred options is passive, then heat-pipe or semipassive, and then active. Active cooling results in increased costs, complexity, and weight but is required for high heat flux.

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Fig. 9.23

9.3.1.1

Effect of chordwise position on heat flux for a sharp and blunt leading edge.

PASSIVE LEADING EDGES

Figure 9.25 illustrates the net energy flow on passive leading edges. There is a net heat input into the leading edge in the stagnation region. The heat is conducted aft of the stagnation region, where it is then radiated to space. The amount of heat transferred from the stagnation region aft is a function of the thermal conductivity of the leading-edge material. The goal is to attempt both 1) to have hightemperature materials to increase the capability of the leading edges and 2) to manipulate the thermal properties to reduce the temperatures, resulting in temperatures below the reuse temperature of the leading edge.

Fig. 9.24

Leading-edge thermal management options.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

547

Energy Radiated To Space

Heat Flux

Energy Transfer

Wing / Fuselage Structure

High-temperature coatings The SiC-based coatings such as those used on the Space Shuttle Fig. 9.25 Net energy flow on passive leading edges. Orbiter leading edges are useable to 30008F. On air-breathing vehicles, there may be situations where the temperatures will be above 30008F, in which case a different class of materials will be required (e.g., carbides, oxides, and diborides of Hafnium (Hf ) and Zirconium (Zr), etc.). These materials, such as Hf and Zr, could be used as part of a composite matrix. Some materials, such as Iridium (Ir), are more appropriate as a coating, as shown in Fig. 9.26. Thermal conductivity The thermal properties of the material can also have a significant impact on the surface temperatures. One such thermal property is the thermal conductivity of the fiber and the weave architecture. There are multiple grades of carbon fiber with different thermal conductivity. The fibers can also be woven with different architectures. High thermal conductivity fibers in an unbalanced weave help spread heat, avoid hot spots, and survive high heat flux environments. Figure 9.27 illustrates 3:1, 4:1, and 5:1 plain weaves. In the 3:1 (75/25 unbalanced) plain weave, there are three times as many fibers in the 3 direction as in the 1 direction.

Fig. 9.26

Photomicrograph of a cross section of a C/C leading edge with an Ir coating.

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3:1 plain weave

k3 = 190 Btu/hr-ft-¡F k1 = 70 Btu/hr-ft-¡F

Fig. 9.27

4:1 plain weave

k4 = 200 Btu/hr-ft-¡F k1 = 60 Btu/hr -ft-¡F

5:1 plain weave

k5 = 210 Btu/hr-ft/¡F k1 = 50 Btu/hr-ft/¡F

Illustration of plain weaves and photographs of arc-jet tested leading edges.

For the 3:1 plain weave, the thermal conductivity is 190 Btu/hr-ft-8F in the 3 direction and 70 Btu/hr-ft-8F in the 1 direction. The 4:1 and 5:1 plain weaves show the thermal conductivity increasing in the strong (or primary) direction and decreasing in the weak direction. These leading edges with the 3:1, 4:1, and 5:1 architecture were all arc-jet tested. The 5:1 weave architecture was able to remove more heat from the leading edge than the 3:1 architecture, thus demonstrating a better thermal performance and survivability with the highly unbalanced architectures. Although the thermal performance is enhanced in the strong direction, the strength in the weak direction is reduced. Emissivity Emissivity is another important parameter that affects the surface temperatures. The radiation equilibrium temperature, shown in Fig. 9.28, is a function of the surface heat flux and the emissivity. The figure illustrates a radiation equilibrium condition, where the convective heat flux equals the heat that is radiated away from the surface. As an example, if the incoming convective heat flux and the radiated heat flux are 50 Btu/ft2-sec, with an emissivity of 0.8, the surface temperature will be 30008F. If the emissivity is lower, for example 0.3, the surface temperature will increase to near 40008F to radiate away the same 50 Btu/ft2-sec. Recombination (catalytic) efficiency The recombination (catalytic) efficiency is another important parameter. There are disassociated species in the hypersonic flow. For example, the air will be disassociated, and there will be recombination of that flow. The surface will serve as a catalyst to increase the recombination. This recombination can be exothermic, meaning the reaction gives off heat. Heat will be released at the surface, thus providing heat to the surface. In addition to recombination of the gas flow,

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.28

549

Effect of emissivity on radiation equilibrium temperature.

recombination of surface atoms can result in increased heat flux. Silicon gas released from the surface coating reacting with an oxygen atom to form SiO is an example of a reaction that is exothermic and releases 605 Btu/mol. SiðgÞ þ O $ SiO A plot of heat flux vs recombination efficiency is shown in Fig. 9.29 [3]. The data represent an arc-jet test at Mach 4 with an emissivity of 0.84 and show how the recombination efficiency impacts the surface heat flux. When the recombination efficiency is close to zero, the surface is referred to as noncatalytic, and the heat flux is 10 Btu/ft2-sec. For a fully catalytic surface with a recombination efficiency of 1, the heat flux is 28.3 Btu/ft2-sec. The catalycity of the surface resulted in nearly a factor of three increase in the surface heat flux. Thermal expansion Another important property is thermal expansion. As shown in the top row of Fig. 9.30, a rod that is initially at a uniform temperature To and is heated up to a higher temperature T1 will increase in length DL. It is constrained at one end and free to grow on the other end. The strain in the rod is the growth DL divided Fig. 9.29 Plot of heat flux vs recombination (catalytic) efficiency from an arc-jet test [3].

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by the initial length, while the stress in the rod would be zero because it is allowed to grow. If the rod is constrained on both ends (assume one dimension) and heated to a higher temperature, the strain is zero because it is not growing. However, the stress in the rod is now nonzero. The stress is equal to the modulus of elasticity E times the thermal expansion a times the temperature difference DT. The third row in the figure illustrates one of the main challenges in thermal structures, which results from a temperature gradient. In the third row of the figure (illustrating a two-dimensional example vs one dimension in the first two rows), the growth is constrained in one direction, but there is also a temperature gradient across the rod in the other direction. The strain is nonzero because it is growing, and the stress is nonzero. The thermal stress is generated due to a materials thermal expansion, a temperature differential, and a structural or mechanical restraint of the thermal growth. Thermal stresses can be difficult to address because they differ from mechanical stresses in that with mechanical stresses, “beefing” up the structure can lower the stresses, whereas for thermal stresses, a thicker structure can actually increase the stresses. Oxidation Oxidation is a key factor in high-temperature structures due to the desired long lifetimes. To maximize the life of the structure, the goal is to operate in the passive oxidation regime. The first equation in Fig. 9.31 shows silicon carbide as a solid, which is the coating material on many refractory composite structures. If there is high oxygen pressure, then there is plenty of oxygen. The silicon carbide reacts with two oxygen molecules to form silicon dioxide (silica), SiO2, which is a protective barrier for the structure. The silica is a solid and forms a protective scale on the surface. There is also emission of CO2 gas as a by-product of the reaction.

Fig. 9.30

Impact of thermal expansion on stress and strain.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.31

551

Comparison of active and passive oxidation.

If operating in a low-pressure environment, active oxidation can occur at high temperatures. In a low-pressure environment, there is not enough oxygen available, so that for each silicon atom, only 1.5 oxygen atoms are available (vs 2 in a high-pressure environment). The result is the formation of SiO instead of the protective SiO2. The result is active oxidation. Because SiO is a volatile gas, it does not form a protective scale (as does the SiO2). The silicon carbide coating is attacked as SiO gas is formed and leaves the surface. The photographs in Fig. 9.31 illustrate a pretest specimen with a silicon carbide coating and the specimen after testing in an active oxidation environment. The transition between the passive and active oxidation regime depends on the temperature and the oxygen partial pressure. The top graph in Fig. 9.32 shows the temperature vs time and position vs time for a specimen in an arc-jet environment [4]. As the specimen is gradually moved closer to the nozzle (x-position decreasing), the temperature gradually increases. Once the temperature and pressure cross over into the active oxidation regime, there is a significant jump in the surface temperature. This effect has been noticed at multiple temperature/pressure combinations. The lower plot shows that the flow field energy remained relatively constant during the test. X-43 leading edge As an example of a passive leading-edge design approach, consider the X-43 Mach 10 vehicle nose leading edge, which was designed to reach nearly 40008F during the short 130-sec flight [5]. As described previously, high conductivity fibers can be woven into an unbalanced architecture to help conduct heat away from a leading edge, thus reducing the maximum temperatures. This approach was used on the X-43 Mach 10 vehicle, where a 3:1 (75/25 unbalanced) plain weave was used to obtain a high thermal conductivity in the chordwise direction, as shown in Fig. 9.33.

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Fig. 9.32 Temperature vs time (top) and flow enthalpy vs time (bottom) illustrating a large temperature rise at the passive to active oxidation interface [4]. For the X-43, the other approach to dealing with the high heat flux was to use high-temperature material systems. Thirteen different high-temperature materials were evaluated to find one that would survive the flight [5]. The flight conditions (Mach 10 at 105,000 ft) were simulated in the H2 arc-jet facility at Arnold

Fig. 9.33 Photograph of X-43 leading edge with superimposed schematic drawing showing the 3:1 weave architecture of C/C leading edge.

THERMAL PROTECTION SYSTEMS AND HOT STRUCTURES FOR HYPERSONIC VEHICLES

Fig. 9.34

553

Arc-jet coating evaluation of X-43 leading edges.

Engineering Development Center, Tullahoma, Tennessee. The nose radius for the arc-jet test was 0.03 in. The heat flux for the test was 1300 Btu/ft2-sec, for about 130 sec. The photograph on the bottom left in Fig. 9.34 shows a leading-edge test article during arc-jet testing. The figure also shows a successfully tested specimen and one that failed, where the coating came off and resulted in severe oxidization of the C/C substrate. High thermal gradients can result in high thermal stresses, as was the situation on the leading edges of the Hyper X Mach 10 vehicle. The white area in Fig. 9.35 illustrates the original geometry, and the black area illustrates a deformed geometry due to the thermal expansion resulting from the high temperatures. Thermal– structural modeling predicted the leading edges would fail due to the high compressive thermal stresses resulting from the constraint of the thermal expansion. The predicted high thermal stresses raised several questions that needed to be answered before committing to flight. These questions included the following: .

Would the substrate, with coating, fail due to these high thermal stresses?

.

Would the coating buckle and spall?

.

Would the coating fall off, resulting in oxidation of the substrate?

Fig. 9.35 Illustration of spanwise compressive thermal stresses on the X-43 leading edge.

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Would slotting be required to keep the leading edge from buckling? If so, what kind of spacing would be needed, and would the slotting prevent spalling of the coating?

A series of tests at different scales confirmed that the leading edges would survive [5]. Shuttle Orbiter wing-leading-edge repair After the Space Shuttle Columbia accident, the Columbia Accident Investigation Board recommended that a wing-leading-edge repair capability be developed. Both a crack filler material and plug repair capability were developed and flew on the return to flight mission. The plug repair consists of several different 7-in.-diam C/SiC cover plates attached to the RCC leading edges via a TZM refractory metal attachment mechanism (Fig. 9.36). A high-temperature oxidation barrier was developed to protect the C/SiC during the reentry. The C/SiC cover plates were designed to maintain flexibility to reduce the number of plates taken to orbit (Fig. 9.36). In addition, the protuberance was minimized to prevent excessive aerodynamic heating. The plug repair was validated via arc-jet testing and was stowed on board the International Space Station (ISS) [6]. 9.3.1.2

HEAT-PIPE-COOLED LEADING EDGES

When heat fluxes result in temperatures above the multi-use temperature of structures, other approaches need to be evaluated. Heat-pipe-cooled leading edges,

a)

b)

Fig. 9.36 Photograph of Space Shuttle Orbiter wing-leading-edge plug repair, a) C/SiC cover plates and TZM attachment mechanism, and b) C/SiC cover plate formed around simulated leading edge.

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Container

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Working fluid

Wick

Vapor flow Liquid

Heat input Evaporator

Fig. 9.37

Heat output Condenser

Schematic diagram illustrating the operation of a heat pipe.

which are semipassive, use heat pipes to reduce leading-edge temperatures below the reuse limits of the materials. A heat pipe is a self-contained, two-phase heat transfer device composed of a container, a wick, and a working fluid. A schematic diagram illustrating the operation of a heat pipe is shown in Fig. 9.37. Heat input locally to one section of the heat pipe, the evaporator region, is conducted through the container and into the wick/working-fluid matrix, where it is absorbed by the evaporation of the working fluid. The heated vapor flows to a slightly cooler section of the heat pipe where the working fluid condenses and gives up its stored heat. The heat is then conducted through the wick/working-fluid matrix and container and is rejected. The location of the heat pipe where heat is rejected is called the condenser region. The cycle is completed with the return flow of liquid condensing back to the heated region (evaporator) by the capillary pumping action of the wick. During normal operation, heat pipes operate as devices of very high effective thermal conductance and maintain a nearly uniform temperature over the entire heat-pipe length. Heat pipes provide cooling of stagnation regions by transferring heat nearly isothermally to locations aft of the stagnation region, thus raising the temperature aft of the stagnation region above the local radiation equilibrium temperature. When applied to leading-edge cooling, heat pipes operate by accepting heat at a high rate over a small area near the stagnation region and radiating it at a lower rate over a larger surface area, as shown in Fig. 9.38. The arrows in Fig. 9.38 indicate the net heat flux. The use of heat pipes results in a nearly isothermal leading-edge surface, thus reducing the temperatures in the stagnation region and raising the temperatures of both the upper and lower aft surfaces. Heat-pipe-cooled wing leading edges were studied during the National Aerospace Plane (NASP) program [7]. The leading edges had a C/C structure with Molybdenum-Rhenium (Mo-Re) heat pipes as the container, shown in

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Fig. 9.38

Leading-edge heat-pipe operation.

Fig. 9.39. Three straight Mo-Re heat pipes, 30-in. long, were fabricated and embedded in C/C [8]. The heat pipes had a D-shaped cross section and used Li as the working fluid. A single leading-edge-shaped heat pipe was later fabricated and tested. Figure 9.40 shows a heat pipe in the shape of a leading edge that was fabricated for an advanced space transportation system [9]. Figure 9.41 shows how this

Fig. 9.39

Photograph of Mo-Re tube embedded in C/C.

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Fig. 9.40 Leading-edge-shaped heat pipe with uniform color illustrating uniform temperature. heat pipe would fit and be used on a leading edge. This heat pipe used a Hastelloy-X container and sodium, a liquid metal, as the working fluid. The radio frequency heating coils are evident at the nose region in Fig. 9.40. The entire heat pipe is a uniform red color, indicating it is isothermal. Material compatibility and thermal stresses are two of the major challenges with heat pipes embedded in composite materials such as C/C. Two different materials (carbon and Mo-Re for the container) at elevated temperatures for long durations are problematic. If the Mo-Re is in contact with the C/C, molycarbides will start to form. The formation of the molycarbides is a function of both time and temperature. Carbon may also diffuse through the Mo-Re into the heat pipe and degrade the heat pipe’s operation. This diffusion is also a strong function of both time and temperature. Thermal stresses present another problem. The Mo-Re has a larger coefficient of thermal expansion than C/C, and will thus be constrained by the C/C. Because of the constraint in growth, there is the possibility that the flat section of the heat pipe could buckle during operation. The difficulty here is that good thermal contact (through-the-thickness) is required to transfer the heat from the outer surface into the heat pipe, but a loose fit is desired to minimize thermal stresses due to the mismatch in coefficients of thermal expansion. 9.3.1.3

ACTIVELY COOLED LEADING EDGES

For the highest heat fluxes, active cooling is required. Metallic actively cooled leading edges have been fabricated, and the challenges include hydrogen interaction and multicycle life. A photograph of the internal portion of an actively cooled leading edge is shown in Fig. 9.42. Fig. 9.41 Schematic drawing showing placement of heat pipe on leading edge.

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Photograph of the internal portion of an actively cooled leading edge.

With a goal of light weight and high-temperature capability, actively cooled composites can provide significant benefits. As a first step, metallic tubes have been embedded in CMCs to hermetically seal the naturally porous CMCs. Materials that have been used to contain the coolant include superalloys and refractory metals. The concepts evaluated, shown in Fig. 9.43, include a composite with metallic tubes inserted to contain the coolant, co-processed metallic tubes in a composite, and braised tubes to obtain intimate contact with a composite. An all-composite actively cooled structure would provide the lightest weight and highest temperature capability. However, numerous challenges exist, including optimum through-the-thickness conductivity, cooling containment, manifolding, oxidation protection, long life, and material compatibility. A photograph of an all-composite actively cooled test article is shown in Fig. 9.44.

9.3.2 ACREAGE TPS AND AEROSHELLS There are several approaches to acreage TPS on hypersonic vehicles. One is an insulated structure, as used on the Space Shuttle Orbiter. A standoff TPS approach as used on the X-33 is another option. The standoff approach uses internal insulation below the outer surface. Finally, a load-bearing aeroshell could be used. Inserted metal tubes in composite

Co-processed metallic tubes Intimate contact metallic tubes

Fig. 9.43

Approaches for actively cooled composites.

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Fig. 9.44 Photograph of an all-composite actively cooled test article.

9.3.2.1

INSULATED STRUCTURE

An insulated structure has successfully been used on the Space Shuttle Orbiter for years. The insulators, consisting of tiles and blankets, are attached directly to the cold structure to form the outer mold line. The insulation transfers the aero loads to the structure. A strain isolation pad is required between the tiles and aluminum structure due to the different thermal expansion of the two materials. The insulation is segmented (individual tiles), and each tile is on the order of around 6 in.  6 in. Figure 9.45 shows a photograph of STS-114 with an excellent view of the tile insulation system. (The lighter color wing leading edges were replaced before the flight and had not yet experienced a reentry.) NASA advanced tile NASA has developed an improved tile called TUFROC (toughened uni-piece reinforced oxidation resistance composite). It has a higher temperature capability

Fig. 9.45

Photograph of STS-114 and the tile insulation system.

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(30868F) than the highest temperature tiles currently available (26908F). The tile consists of a carbonaceous cap composed of ROCCI (refractory oxidationresistant ceramic carbon insulation) that provides dimensional stability to the outer surface and a fibrous inner layer that provides optimal insulation of the substructure. The higher temperature capability of the tile may permit its use on some leading edges or nose caps. Arc-jet testing of the TUFROC has been performed and the surface is relatively noncatalytic [10]. EADS Astrium flexible external insulation Blanket insulation is used on much of the upper surface of the Space Shuttle Orbiter. There has been some effort to improve the state of the art of blanket insulations. Work has been performed by EADS Astrium (Europe) to improve their flexible external insulation (FEI) with both new high-emittance coatings for improved stability and a less toxic waterproofing for both initial and re-waterproofing [11]. A photograph of the FEI insulation is shown in Fig. 9.46. Boeing conformal reusable insulation Boeing has developed a pair of conformal reusable insulations for potential use on both windward and leeward surfaces. The Type I insulation has a maximum use temperature of 18008F for multi-use and 20008F for single use. The Type II insulation has a maximum use temperature of 22008F for multi-use and 24008F for single use. 9.3.2.2

STANDOFF TPS

The standoff TPS approach can be used on a cold or warm structure. One advantage of the standoff TPS is that the outer aeroshell may have a different contour

Fig. 9.46

Photograph of repaired EADS Astrium FEI insulation [11].

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Fig. 9.47 Schematic drawing of the structural layout of the X-33 RLV illustrating the support structure used for stand-off TPS.

than the support structure. The substructure on the X-33 supporting the metallic TPS that formed the aeroshell is shown in Fig. 9.47. The TPS panels, shown in Fig. 9.48, are attached to standoff brackets to create the desired aeroshell. The critical challenge with this approach is to transfer the aero loads, but not the thermal loads, to the structure. As with insulating tiles, the panels are segmented but are much larger, around 18 in.  18 in. The metallic TPS is limited to regions where the use temperatures are in the range of 18008F, depending on the metallic alloys used. Numerous papers have been written concerning metallic TPS both in the United States and Europe [12–16]. For hypersonic vehicles, there is often the

Fastener access covers

Panel-to-panel seal Outer facesheet

Attachment standoff brackets

Encapsulated insulation External stiffeners

Fig. 9.48 Schematic drawing of X-33 metallic TPS illustrating stand-off TPS attachment to substructure.

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need for higher use temperature acreage TPS, thus giving rise to the desire for a CMC TPS. In the United States, the metallic TPS is at a higher technology readiness level (TRL) than the CMC TPS, whereas in Europe, the CMC TPS is at a higher TRL than the metallic TPS. A CMC standoff TPS is illustrated in Fig. 9.49 [17]. An external CMC panel with oxidation protection forms the aeroshell, and the standoff TPS includes an internal insulation. The insulation could be an efficient high-temperature fibrous insulation. High-temperature seals and an attachment are used to attach this external CMC panel to the internal structure. The primary challenge with any standoff TPS is attaching the external panel to the substructure and transferring aero and pressure loads but not thermal loads. In addition, the entire system must withstand acoustic and vibration loads. Snecma Propulsion Solide (SPS) (Bordeaux, France) has developed a CMC shingle TPS, separating the mechanical function and the thermal function [17–19]. The mechanical function components, the CMC shell, the fasteners, and the standoffs have good mechanical properties but are not optimized to minimize heat transfer. The thermal function components, the internal insulation, the seals, and the insulating washers are not intended to be load bearing but are very efficient thermal insulators. The TPS has a total system mass of 3.7 lb/ft2. The attachment approach is shown in Fig. 9.50. The attachment enables expansion mismatches between the CMC shingle (panel) and the structure by adapted flexibility of the standoff, prevents large CMC deformation through sufficient stiffness, and transfers loads from the CMC panel to the structure. The attachment is designed for easy replacement, with a portion of the attachment integrated with the CMC panel and a portion integrated with the structure. SPS has fabricated a CMC shingle TPS array and has exposed it to mechanical, dynamic (modal, vibration, and acoustic), thermal, thermal–mechanical, micro-meteoroid orbital

Fig. 9.49

Schematic drawing of Snecma Propulsion Solide (SPS) CMC shingle TPS [17].

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Fig. 9.50 Schematic drawing of SPS shingle TPS attachment mechanism to the substructure [19]. debris, and arc-jet testing (see Fig. 9.51). The array was 31.5 in.  17 in., with the CMC 0.039-in. thick, and was fabricated with one main panel and one secondary panel, allowing a linear seal to be tested. The total arc-jet testing time was equivalent to 11 reentries. SPS also fabricated and tested two test articles to address the interface between a leading edge and the shingle TPS. These test articles were subjected to arc-jet testing. The maximum temperature measured on the flat panel behind the leading edge was 21928F, whereas the temperature was maintained below 2308F at the structure side (i.e., the interface between the structure and the standoffs), demonstrating the thermal efficiency of the concept. The flat array in Fig. 9.52

Fig. 9.51 SPS CMC TPS array for mechanical, dynamic, thermal, and thermal– mechanical testing.

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Fig. 9.52 Snecma Propulsion Solide CMC TPS array during arc-jet testing. was 13.4 in.  13.4 in., with the CMC 0.039-in. thick, and was tested at the EADS Bordeaux SIMOUN arc-jet facility. Two linear seals were tested between panels, with a “triple point” at the panel intersection. A photograph of the array during arc-jet testing is shown in Fig. 9.52. The objective of the testing was to investigate the effects of steps and gaps at the seal. Eleven tests were conducted with different step and gap configurations. A curved test article is shown in Fig. 9.53 during and after arc-jet testing at the Scirocco facility at CIRA in Italy. It was 23.6 in.  16 in., with the CMC again 0.039-in. thick and a radius of 9.8 in. Again, two liner seals were tested with a triple point at the panel intersection. The purpose of this testing was to evaluate the seal at the leading-edge/panel interface and to orient the seal in the flow direction. Assembly and disassembly with access only to the outer surface was also demonstrated. 9.3.2.3

INTERNAL INSULATION

Internal insulation is required for the standoff TPS to reduce the heat load to the substructure. The internal insulation should be lightweight and flexible and have a a)

Fig. 9.53

b)

Snecma Propulsion Solide CMC TPS array a) during and b) after arc-jet testing.

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high heat capacity and low thermal conductivity. It must also be capable of longduration flight at elevated temperatures, implying that it is either nonoxidizing or protected from any oxidizing environment. High-temperature fibrous insulation is often used as the internal insulation. MT Aerospace internal multiscreen insulation (IMI) A more efficient insulation is a high-temperature multilayer insulation (MLI), with reflective layers separated by layers of fibrous insulation. MT Aerospace (Augsburg, Germany) has developed two different high-temperature MLIs, referred to as internal multiscreen insulation (IMI), one for use at 18328F and one for use at up to 29128F [20]. The medium-temperature MLI (18328F) uses Nextel 312 fabric for containment of the MLI, whereas the high-temperature version uses Nextel 440 fabric. A schematic diagram of the medium-temperature IMI using nanoporous material (NPM) as the cold-side insulation is shown in Fig. 9.54. Steve Miller and Associates Research Foundation opacified fibrous insulation An alternative to the MLI is an opacified fibrous insulation (OFI), developed by Steve Miller and Associates Research Foundation (SMARF) (Flagstaff, Arizona). The OFI is a flexible insulation system consisting of opacified fibrous insulation felts quilted between layers of high-temperature fabric. Combinations of opacified silica/alumina/zirconia felts can be used up to 30008F. The standard OFI consists of opacified alumina and silica felts, with a continuous use temperature of 27008F, encased in Nextel fabric cloth. The nominal combined density of a 1-in.-thick blanket in this configuration is 18 lb/ft3 including fabric and thread, with an areal density of 1.5 lb/ft2, but density can be varied from 6 to 24 lb/ft3.

Fig. 9.54

Schematic drawing of the MT Aerospace medium-temperature IMI [20].

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Fig. 9.55 Thermal conductivity of the SMARF OFI as a function of temperature illustrating the impact of atmospheric pressure.

The thermal conductivity of the fibrous insulation is a function of temperature and pressure. A higher air pressure increases the effective thermal conductivity, as shown in Fig. 9.55. 9.3.2.4

LOAD-BEARING AEROSHELL

A load-bearing aeroshell carries the aerodynamic and vehicle axial loads. Insulation may be either incorporated in the structure or separate, below the aeroshell. This type of aeroshell has the potential for reduced weight. The image in Fig. 9.56 shows the DARPA/Air Force Falcon HTV-2 vehicle, which uses a C/C aeroshell [21, 22]. Another example of a load-bearing aeroshell is the C/SiC aeroshell for the UK Ministry of Defense–sponsored Sustained Hypersonic Flight Experiment (SHyFE), shown in Fig. 9.57 [23–25]. The objective of this flight experiment is to design and fly a prototype ramjet capable of sustained hypersonic flight. The

Fig. 9.56

DARPA/Air Force Falcon HTV-2 vehicle, which uses a C/C aeroshell.

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Fig. 9.57 C/SiC load-bearing aeroshell used on the SHyFE [Credit: Fig. 1, Sustained Hypersonic Flight Experiment (SHyFE), AIAA-2006-7926, Ref. 24]. vehicle weighs approximately 30 kg and is 1.5 m long, with a diameter of 7 in. The aeroshell is being fabricated by MT Aerospace (Augsburg, Germany) utilizing chemical vapor infiltration (CVI). Fuel flows in an annulus between two C/SiC tubes. The fin roots are bonded to the aeroshell. Many of the attachments used on previous designs have been eliminated in the approach used here. 9.3.2.5

STRUCTURALLY INTEGRATED TPS

The final TPS approach discussed is a structurally integrated TPS, as illustrated in Fig. 9.58. Here, the outer and inner walls carry the airframe loads, with the outer wall operating hot and the inner wall insulated. It is thermally integrated, has a higher structural efficiency, and is potentially lower maintenance. The outer surface is a robust, high-temperature material. The wall thickness helps provide stiffness, and the large integrated structure eliminates/reduces surface gaps and steps and also provides low part count. Although there may be some advantages to this approach, it should be considered the lowest TRL of the approaches discussed here.

9.3.3 CONTROL SURFACES Control surfaces are required on vehicles to control the vehicle during flight. Figure 9.59 illustrates a hypersonic vehicle with elevons and rudders. For hypersonic vehicles, drag is very important, thus driving toward thin control surfaces. There are several different approaches for designing and fabricating control surfaces, including insulated, hot structure, and hybrid. Fig. 9.58 Schematic drawing of structurally integrated TPS.

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Fig. 9.59 Image of hypersonic vehicle showing rudders and elevons, two types of control surfaces. An excellent paper on control surface requirements, design, and validation has been written by Steinacher et al. [26], some of which is summarized here. The requirements for control surfaces are a function of the vehicle design and the entire mission, from pre-takeoff ground operations to postlanding cool down. For some insulated structures, the highest temperatures are reached after landing/touchdown. Reusability requirements include integrity of the oxidation coating and lifetime of the various components. It is important to ensure that the life of the individual components exceeds the life of the control surface or that it is designed for easy replacement. The geometry of the control surface is impacted directly by flight control (aerodynamics) and vehicle interfaces. Mass, stiffness, and structural integrity have an indirect impact on the geometry requirements. The mechanical requirements are generated by the pressure differential between the windward and leeward surfaces, whereas the thermal requirements include the maximum temperature, thermal stresses, oxidation (active and/or passive), interfaces with cold structures such as actuators, and thermal radiation from opposing surfaces. Functional requirements, such as number of cycles and flap actuation, also impact the requirements. As an example, the control surfaces on the X-38 reentry vehicle had a deflection range of +208 and a rate of 0.1–0.7 Hz and were designed for 874 cycles. The approach used to attach control surfaces to a vehicle also impacts design. For example, control surfaces could be attached behind the vehicle, underneath the vehicle, or at the trailing edge. Static and dynamic seals are also required around hinge lines, around the actuation rod, and at supports to prevent sneak flow into the vehicle. Some of the requirements can be validated by analysis and some by review and inspection, whereas some must be validated by testing [26]. 9.3.3.1

INSULATED CONTROL SURFACE

Insulated control surfaces have been successfully used on the Space Shuttle Orbiter for many years. Some of the advantages of an insulated control surface include its suitability for very large structures and its minimal thermal expansion issues. Disadvantages include weight, low thermal margin because the structure is a low-temperature structure, and thick cross sections [27]. In Fig. 9.60, the image on the left shows a metal or polymer matrix composite (PMC) control surface with tile on it as insulation. The tiles can also be seen on the picture of the orbiter elevons [27].

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CERAMIC MATRIX COMPOSITE HOT STRUCTURE CONTROL SURFACE

The CMC hot structure control surface approach provides the lowest weight and thinnest cross section, minimal thermal expansion mismatch problems, and a good thermal margin. This approach is also advantageous because of its sufficient strength and stiffness and because no external insulation is required. Disadvantages to this approach include the high manufacturing and tooling costs for the box structure and scale-up. This approach may not be recommended for very large structures. Its repair capability is limited, and the manufacturing risk is high in cases of production failure or damage. Access for coating, inspection, and maintenance of internal areas is also a disadvantage [27]. A mechanically assembled CMC control surface can be fabricated and assembled using multiple smaller parts. Advantages to this approach are that the tooling is relatively simple and damaged components can be replaced without a complete scrap of the entire control surface. A disadvantage to this approach is that tolerance buildup can be problematic in assembly of numerous separate parts. High part count is another disadvantage of this approach. General Electric and Materials Research & Design C/SiC body flap A C/SiC body flap was fabricated by General Electric Energy Ceramic Composite Products (GECCP), LLC (Newark, Delaware) and designed by Materials Research & Design (MR&D) (Wayne, Pennsylvania). This control surface used C/SiCfastened joints and a thin ply torque tube and box structure along with gusset members for load transfer. Figure 9.61 shows the individual parts and the fasteners used for assembly, and Fig. 9.62 shows a photograph of the assembled body flap. MT Aerospace C/SiC body flap MT Aerospace developed C/SiC body flaps for the X-38 reentry vehicle, shown in Fig. 9.63 [28, 29]. The body flap was fabricated entirely of CMCs, including

Fig. 9.60 Schematic [27] and photograph (Space Shuttle Orbiter) of insulated control surface.

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a)

b)

Fig. 9.61 Photographs of a) individual parts used to assemble control surface and b) fasteners used for the assembly.

Fig. 9.62

Fig. 9.63

Photograph of a mechanically assembled body flap (GECCP and MR&D).

Photograph of the X-38 C/SiC body flaps fabricated by MT Aerospace [30].

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fasteners and bearings. Several key technologies were developed and qualified as part of the body-flap development. These include manufacturing of large C/SiC structures, manufacturing of thick C/SiC structures, oxidation protection, hightemperature bearings, highly loaded joints with ceramic bolts and nuts, interface of cold metallic to hot ceramic parts, and high temperature dynamic seals. The bearing assembly was qualified via both hot and cold testing. The fullscale body flaps were qualified for flight under ascent and descent vibration, thermal, and static pressure loads. MT Aerospace has more recently updated the body flap design for the Pre-X flight demonstrator [30]. 9.3.3.3

INTEGRATED FABRICATION

An integrated fabrication is an alternative fabrication approach. Advantages to this approach include fewer joints and better mechanical performance. Disadvantages include the complex tooling, the associated fabrication expense, and the risk of damage during fabrication. An integrated fabrication approach has been demonstrated by MT Aerospace (Augsburg, Germany) on a C/SiC body flap segment, shown in Fig. 9.64. The part was fabricated utilizing a 2-D prepreg of carbon fabric that was cured and pyrolyzed and then further densified via a CVI process. The integrated fabrication permitted the part to be fabricated without any fasteners that would increase mass [30]. 9.3.3.4

HYBRID CONTROL SURFACE

Hybrid control surfaces combine a CMC with a lower temperature material such as titanium or a superalloy. An advantage of a hybrid control surface approach is that manufacturing may be more affordable for large structures. Also, it may not require a TPS on the upper surface, and the CMC leading and trailing edges can be replaced if they are damaged. Disadvantages are the thermal mismatch between the cool structure (metal/PMC) and the CMC. The weight increase over an all-CMC control surface can amount to 30%–40% [27]. In addition, the box structure’s insulation can result in increased thickness, leading to either a reduced structure thickness and smaller moment of inertia or a thicker cross section. X-43 hybrid control surface The X-43 used the hybrid control surface approach, as shown in Fig. 9.65. A Haynes 230 superalloy was used for Fig. 9.64 Photograph of the side and top view of an integrated body flap segment demonstration [30].

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Fig. 9.65

Schematic diagram of the X-43 hybrid control surfaces.

the structure, and C/C was used for the leading edges. There was no insulation on the control surface. A spindle connected the control surface to an actuator in the fuselage. This was a relatively small control surface structure, on the order of 2 ft. Dutch Space X-38 hybrid control surface Dutch Space (Leiden, The Netherlands) has designed and fabricated hot metallic control surfaces (rudder) for the X-38 flight vehicle, shown in Fig. 9.66. The initial rudder was fabricated of PM-1000, with a use temperature up to 21908F. When the requirements for the rudder changed, Dutch Space designed and qualified a hybrid Ti/ceramic tile

Fig. 9.66

Photograph of Dutch Space hot rudder designed for the X-38 [31].

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rudder in one year. For the crew return vehicle, a hybrid metallic/CMC rudder was planned [31].

9.4 KEY TECHNICAL CHALLENGES CMCs are the materials that will enable many of the components on air-breathing hypersonic vehicles. CMCs will be used for leading edges, acreage, and hot structures and in the propulsion system. For most CMC structures, there are two primary materials and structures technical challenges: fabrication and environmental durability.

9.4.1 FABRICATION There are several processes for fabricating C/C, C/SiC, and SiC/SiC, each with their own challenges. Small coupons can be fabricated with relative consistency. However, a coupon built today may or may not be similar to one that is built six months from now. The key point is that a state-of-the-art material is not the same as a state-of-the-art structure. The fact that the material can be fabricated does not guarantee that a structure can successfully be fabricated and that it would survive the required mission life. This is illustrated in Fig. 9.62, which shows the simplicity of a CMC test coupon vs the fabrication complexity of a CMC body flap. The following fabrication challenges, again process dependent, arise in going from a small material coupon to a large flight structure: thickness (density uniformity throughout the structure), complex curvature, large scale, low interlaminar properties, delaminations, critical flaw size, nondestructive inspection, tooling, assembly methods and tolerances, reproducibility, fabrication modeling, manufacturability of structures design, and affordable (cost and schedule) fabrication techniques. Meeting the flight requirements is another challenge that impacts the structural maturity of a component. Operation has a significant impact on the ability to use these materials as structures on flight vehicles. Stated another way, the TRL is a function of the requirements. Challenges to consider for flight requirements include thermal loads, thermal gradients, mechanical loads, acoustic and vibration loads, pressure (oxidation), combined loads, number of cycles, and total mission life. During flight, many of the loads are combined, thus generating potentially severe loading conditions.

9.4.2 ENVIRONMENTAL DURABILITY The primary environment durability challenge is oxidation resistance, which has a major impact on mission life. Oxidation of CMCs was discussed earlier in this chapter. The number of cycles required under combined loads, inspection and

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repair, and the ability to predict mission life all present challenges. Figure 9.34 shows a failed test article after severe oxidation during an arc-jet test.

9.5 CONCLUDING REMARKS Hypersonic air-breathing vehicles provide some unique thermal and thermal– structural challenges, such as sharp leading edges and structures with thin cross sections. Solving these challenges will require moving beyond an insulated aluminum “airplane,” such as the Space Shuttle Orbiter, to a vehicle with multiple TPS and hot structure approaches. The ability to build and fly these hypersonic airbreathing vehicles successfully will depend on the ability to use multiple types of CMC structures (leading edges, control surfaces, actively and passively cooled propulsion structures, and acreage TPS) but will occur only after solving the environmental durability and fabrication challenges. Finally, the success in addressing these challenges, as indicated by the TRL, is a function of the specific vehicle requirements.

ACKNOWLEDGMENTS Numerous individuals provided input to the presentation given at NASA Headquarters titled “TPS and Hot Structures for Hypersonic Vehicles,” from which much of this paper was derived. The help of Kim Bey and Kevin Rivers at NASA Langley Research Center was instrumental to the development of that presentation, and thus this chapter. Susanne Waltz, NASA Langley, provided exceptional graphics support. Valuable input was also received from Keith Belvin, Keith Bird, Max Blosser, Walt Bruce, Craig Ohlhorst, Jeff Robinson, Steve Scotti, Wally Vaughn, and Kathryn Wurster from NASA Langley Research Center; Craig Stephens and Larry Hudson from NASA Dryden Flight Research Center; Brian Sullivan from Materials Research & Design; Brian Zuchowski from Lockheed Martin Aeronautics; George Cunnington from Cunnington and Associates; Ray Dirling from Science Applications International Corporation; John Koenig from Southern Research Institute; Bob Klacka from General Electric Energy Ceramic Composite Products; and Steve Miller from Steve Miller and Associates Research Foundation. The assistance provided by each one is greatly appreciated.

REFERENCES [1] Hald, H., Ortelt, M., Fischer, I., Greuel, D., and Haidn, O. J., “Effusion Cooled CMC Rocket Combustion Chamber,” AIAA/CIRA 13th International Space Planes and Hypersonics Systems and Technologies, AIAA-2005-3229, AIAA, Reston, VA, May 2005.

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[2] Thornton, E. A., Thermal Structures for Aerospace Applications, AIAA Education Series, AIAA, Reston, VA, 1996. [3] Clark, R. K., Cunnington, G. R., Jr., and Wiedemann, K. E., “Determination of the Recombination Efficiency of Thermal Control Coatings for Hypersonic Vehicles,” Journal of Spacecraft and Rockets, Vol. 32, No. 1, 1995, pp. 89–96. [4] Hilfer, G., “Ceramic Thermal Protection Materials—How Far Can We Go? New Aspects on the Oxidation Behavior During Re-entry Flight,” Second International Symposium on Reentry Vehicles and Systems, Arcachon, France, March 26–29, 2001. [5] Ohlhorst, C. W., Glass, D. E., Bruce, W. E., Lindell, M. C., Vaughn, W. L., and Smith, R.W., “Development of X-43A Mach 10 Leading Edges,” 56th International Astronautical Congress, Fukuoka, Japan, Oct. 2005, IAC-05-D2.5.06. [6] Rivers, H. K., and Glass, D. E., “Advances in Hot Structures Development,” Fifth European Workshop on Thermal Protection Systems and Hot Structures, Noordwijk, The Netherlands, May 17–19, 2006. [7] Glass, D. E., and Camarda, C. J., “Preliminary Thermal/Structural Analysis of a Carbon/Carbon Refractory-Metal Heat-Pipe-Cooled Wing Leading Edge,” Thermal Structures and Materials for High-Speed Flight, edited by E. A. Thornton, Vol. 140, Progress in Astronautics and Aeronautics, AIAA, New York, 1992, pp. 301–322. [8] Glass, D. E., Merrigan, M. A., and Sena, J. T., “Fabrication and Testing of Mo-Re Heat Pipes Embedded in Carbon/Carbon,” NASA CR-1998-207642, March 1998. [9] Glass, D. E., Merrigan, M. A., Sena, J. T., and Reid, R. S., “Fabrication and Testing of a Leading-Edge-Shaped Heat Pipe,” NASA CR-1998-208720, Oct. 1998. [10] Stewart, D. A., and Leiser, D. B., “Lightweight TUFROC TPS for Hypersonic Vehicles,” 14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference, AIAA-2006-7945, AIAA, Reston, VA, Nov. 2006. [11] Antonenko, J., Rieck, U., and Ritter, H., “Improvement of Flexible External Insulation Concerning Re-Usability,” Fifth European Workshop on Thermal Protection Systems and Hot Structures, Noordwijk, The Netherlands, May 17–19, 2006. [12] Blosser, M. L., “Fundamental Modeling and Thermal Performance Issues for Metallic Thermal Protection System Concept,” Journal of Spacecraft and Rockets, Vol. 41, No. 2, 2004, pp. 195–206. [13] Dorsey, J. T., Poteet, C. C., Wurster, K. E., and Chen, R. R., “Metallic Thermal Protection System Requirements, Environments, and Integrated Concepts,” Journal of Spacecraft and Rockets, Vol. 41, No. 2, 2004, pp. 162–172. [14] Poteet, C. C., Abu-Khajeel, H., and Hsu, S.-Y., “Preliminary Thermal-Mechanical Sizing of a Metallic Thermal Protection System, Journal of Spacecraft and Rockets, Vol. 41, No. 2, 2004, pp. 173–182. [15] Fischer, W. P. P., “ULTIMATE: Metallic Thermal Protection System for Future RLV’s – Design and Performance Verification Approach,” Society of Automotive Engineers TP Series, 2004-01-2566, SAE International, 34th International Conference on Environmental Systems (ICES), Colorado Springs, CO, July 19–22, 2004. [16] Fatemi, J., Birjmohan, S., Rijkeboer, M., and Bakker, M. C. M., “Analysis and Testing of a Metallic Thermal Protection System of RLV Applications,” Fourth

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International Symposium Atmospheric Reentry Vehicles & Systems, Arachon, France, March 2005. Pichon, T., Soyris, P., Barreteau, R., Foucault, A., Parenteau, J. M., and Lacoste, M., “Generic CMC Thermal Protection System Pre-Development for Reentry Demonstrator Vehicle,” Second European Conference for Aerospace Sciences (EUCASS), Brussels, Belgium, July 2007. Coperet, H., Soyris, P., Lacoste, M., Garnett, J., and Tidwell, D., “MMOD Testing of C-SiC Based Rigid External Insulation of the X-38/CRV Thermal Protection System,” IAC-02-I.3.06, 53rd International Astronautical Congress, Houston, TX, Oct. 2002. Pichon, T., Soyris, P., Faucault, A., Parenteau, J. M., Prel, Y., and Guedron, S., “C/SiC Based Rigid External Thermal Protection System for Future Reusable Launch Vehicles: Generic Shingle, Pre-X/FLPP Anticipated Development Test Studies,” Fifth European Workshop on Thermal Protection Systems and Hot Structures, Noordwijk, The Netherlands, May 17–19, 2006. Keller, K., Pfeiffer, E., Handrick, K., Weiland, S., Ullmann, T., and Ritter, H., “Advanced High Temperature Insulations,” Fifth European Workshop on Thermal Protection Systems and Hot Structures, Noordwijk, The Netherlands, May 17–19, 2006. Walker, S. H., and Rodgers, F., “Falcon Hypersonic Technology Overview,” AIAA/ CIRA 13th International Space Planes and Hypersonics Systems and Technologies, AIAA-2005-3253, AIAA, Reston, VA, May 2005. Glass, D. E., Dirling, R., Croop, H., Fry, T. J., and Frank, G. J., “Materials Development for Hypersonic Flight Vehicles,” 14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies, AIAA-2006-8122, AIAA, Reston, VA, Nov. 2006. Cain, T., and Walton, C., “The Sustained Hypersonic Flight Experiment,” AIAA-2003-7030, AIAA, Reston, VA, Dec. 2003. Dadd, G., Owen, R., Hodges, J., and Atkinson, K., “Sustained Hypersonic Flight Experiment (SHyFE),” 14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference, AIAA-2006-7926, AIAA, Reston, VA, Nov. 2006. Goodman, J., and Ireland, P., “Thermal Modelling for the Sustained Hypersonic Flight Experiment,” 14th AIAA/AHI Space Planes and Hypersonic Systems and Technologies Conference, AIAA-2006-8071, AIAA, Reston, VA, Nov. 2006. Steinacher, A., Lange, H., Weiland, S., and Hudrisier, S., “Development of CMC Body Flap for Future Re-Entry Vehicles,” 58th International Astronautical Congress, Hyderabad, India, IAC-07-C2.4.01, Sept 24–28, 2007. Trabandt, U., Schmid, T., and Werth, E., “CMC and Metallic Hot Structure Hybrid Components for RLV,” 54th International Astronautical Congress, Breman, Germany, 2003. Pfeiffer, H., and Peetz, K., “All-Ceramic Body Flap Qualified for Space Flight on the X-38,” 53rd International Astronautical Congress: The World Space Congress—2002,” IAF-02-I.6.b.01, Houston, TX, Oct. 2002. Pfeiffer, H., and Dogigli, M., “All-Ceramic Body Flap for the X-38: A Revolutionary Step Forward,” Third European Conference Structures and Technologies Challenges for Future Launchers, Strasbourg, Germany, Dec. 2001.

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[30] Lange, H., Steinacher, A., Handrick, K., Weiland, S., Sygulla, D., Guedron, S., and Salmon, T., “Status of Flap Development for Future Re-Entry Vehicles (Pre-X),” Fifth European Workshop on Thermal Protection Systems and Hot Structures, Noordwijk, The Netherlands, May 17–19, 2006. [31] Mooij, E., Vo, Q. T., and Offerman, J. W. P. I., “Technology Status and Developments of Metallic Hot Structures and Thermal Protection Systems,” 5th European Workshop on Thermal Protection Systems and Hot Structures, Noordwijk, The Netherlands, May 17–19, 2006.

CHAPTER 10

Aero Engine Materials James C. Williams University of North Texas, Denton, Texas

10.1 INTRODUCTION Materials have been a major enabling technology since the inception of the use of the gas turbine engine for aircraft propulsion. Although the focus of this chapter is on gas turbines it should be recognized that there are many smaller turbines that are fitted with a gear box and used as a highly efficient source of torque to power turboprop aircraft and rotary wing aircraft. The basic principles of operation for these smaller turbines and for the larger ones used to propel commercial aircraft are similar. In some cases, the smaller scale of these turboshaft engines can pose manufacturing challenges to create features such as air cooling passages in turbine blades. Considerable progress has been made in turbine engine design, performance, reliability, and durability since the Boeing 707, equipped with four Pratt and Whitney JT-3C engines and entered into service on 26 October 1958. Several metrics can be used to quantitatively track this progress. They are helpful in assessing the evolution of the efficiency and technical sophistication of the gas turbine engine. The first of these is simply the maximum thrust the engine produces divided by the weight of the engine. This metric is called thrust to weight (T/W). Figure 10.1 shows T/W as a time series for military engines. This figure shows that T/W has increased by nearly threefold since the inception of the J-57 engine used for the B-52 military bomber and the JT-3C for the first model of the Boeing B-707. Here, military engines have been selected to illustrate the change in T/W as a function of time because they have a much more common architecture, and the comparisons from model to model are more consistent. The second metric is the set of operating temperatures at two critical locations in the engine. These temperatures are the best indicators of the temperature capability of the materials being used. There are two critical temperatures in a turbine engine that indicate the sophistication of the materials of construction. The first of these is the compressor exit temperature, known as T3, and the second is the turbine inlet temperature, known as T41. A schematic of a turbofan engine is shown in

Copyright # 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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Fig. 10.1

Thrust to weight of military engines as a time series.

Fig. 10.2a, and a cutaway of a modern, large turbofan engine is shown in Fig. 10.2b. A time series for T3 and T41 values spanning the period 1959–1994 are shown in Table 10.1, along with two other key parameters: the overall pressure ratio (OPR), which is analogous to the compression ratio of an internal combustion engine, and the bypass ratio (BPR), which is defined as the ratio of the mass of air that passes around the engine through the fan duct to the mass of air that passes through the engine core and participates in the combustion process. These parameters will be described in more detail later. The original turbine engines used for aircraft propulsion were known as turbojets because all of the air that enters the engine passes through the engine core a)

b)

Fig. 10.2 (a) Schematic of a turbofan engine showing the major sections. (b) Cutaway of a turbofan engine.

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TABLE 10.1

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TIME SERIES OF KEY OPERATING PARAMETERS FOR SEVERAL AIRCRAFT ENGINES

Entry Into Service

Engine Model

T3 (88 C)

T41 (88 C)

OPR

BPR

1958

JT-3C

550 (est.)

870

11.5

0

1970

CF-6

590

1345

15

6

1985

CFM56-3

1994

GE90-92B

500

1400

13

4.5

695

1425

38

9

and participates in the combustion process to create the hot gas stream that produces the propulsive force (thrust). Soon after the turbojet was fielded, it was realized that the propulsive efficiency would be enhanced if the engine also contained a front fan that essentially acted like a ducted propeller. Engines equipped with a fan became known as turbofan engines. One characteristic parameter of a turbofan engine is the bypass ratio, which was defined earlier. For example, in a turbojet all of the airflow passes through the course because there is no fan. Thus, a turbojet has a BPR of zero. To a significant degree, turbofan engines were enabled by the availability of higher quality materials that allowed the manufacture of larger rotating parts such as the fan disk without increasing the risk that the rotor contained flaws. The material of choice for these rotating parts of the fan, except the fan shaft, is a titanium (Ti) alloy, of which there are several to choose from. Ti alloys are preferred because they have 50% lower density with comparable strength when compared to steel. This makes them a more structurally efficient choice. A Ti alloy fan disk for a high-BPR turbofan engine is shown in Fig. 10.3. This disk has been machined from a forging. Both the forging and the billet used to make the forging have been ultrasonically inspected to exacting standards to minimize the possibility of the presence of any melt-related defects or other flaws. Since the introduction of the turbofan engine, the BPR of each generation of turbofan engines has continued to increase. This increased BPR has had a major effect on reducing the fuel consumption normalized for engine size. Today, engines used on large, twin-engine aircraft, such as the one shown in Fig. 10.2b, have BPR values approaching 10:1. Here, it is equally important to note that the high-BPR turbofans have the additional benefit of being considerably quieter. This noise reduction is achieved because the fan air has a lower velocity than the core or jet air, which is essentially “encapsulated” by the fan air as it exits the engine exhaust nozzle, making the engine much quieter. For passenger aircraft engines, in addition to operating temperatures and the fan BPR, other important parameters relate to the sophistication, maturity, and reliability of a turbofan engine. These include three important parameters that are directly related to engine reliability: in-flight shutdowns, unscheduled engine removals, and engine-caused delayed dispatch or departure. For any

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Fig. 10.3

Titanium alloy fan disk for a large turbofan engine.

particular airplane–engine combination, these three parameters are central considerations when the Federal Aviation Administration (FAA) issues permission for the airplane to fly long distances overwater. This rating is called extended twin-engine operations (ETOPS). The ETOPS rating is stated in terms of minutes. The number of minutes is the time an aircraft is allowed to fly from one safe landing zone to the midpoint of the next safe landing zone. For example, an airplane–engine combination with a 180-min ETOPS rating can fly 6 h between safe landing zones. Because of the significant increase in reliability of aircraft engines, it now is possible to fly twin-engine aircraft nonstop from Los Angeles to Sydney, Australia. Twenty years ago, this possibility would have been unthinkable. The improvement in the reliability of engines is illustrated in Figs. 10.4 and 10.5. The data in these figures clearly show the tremendous improvement in engine reliability that has occurred since the inception of turbinepowered flight. From the standpoint of an airline operator, another critical parameter is the specific fuel consumption (SFC), defined as the amount of fuel consumed per pound of delivered thrust. SFC is mainly an engine-related parameter, although aircraft weight and drag are also factors in the overall fuel consumption of aircraft and engine combination. All of these parameters affect the efficient and economical operation of turbine-powered aircraft. It is important to note that the aircraft manufacturers typically set the minimum values for each of these engine-related

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Fig. 10.4

583

Percent on-time departures without engine-caused delays as a function of time.

parameters. The technical approaches and design decisions required to produce an engine that meets these expectations lies squarely in the province of the engine manufacturer. Here again, the substantial improvements in the SFC over time are shown in Fig. 10.6. The curve is derived from discrete data points, An alternate way to evaluate SFC improvements is to track this parameter as a function of the aircraft entry into service date as is shown in Fig. 10.7.

Fig. 10.5

In-flight engine shutdown (IFSD) rate as a function of time.

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Fig. 10.6 Specific fuel consumption for engines developed in the given time frames. In this chapter, the progress in turbine engine materials over approximately the past 30 years will be described and discussed. Throughout and where feasible, the aforementioned parameters will be related to the metrics described previously

Fig. 10.7 Specific fuel consumption by engine model. Plot includes demonstrator and concept engines.

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to track progress. Initially, in the late 1940s and early 1950s, the progress was driven by the need for more powerful and reliable military engines, particularly for fighter aircraft. During this period, the U.S. government invested heavily both in engine and materials research and development (R&D) and in manufacturing technology to create the supply chain that produces both the high-quality raw materials and the cast and forged components that go into modern aircraft engines. The supply chain also is critical to the success of aircraft manufacturers such as Boeing, although their focus has traditionally been on aluminum alloys. However, over the past two decades, there has been a steady growth in the use of polymer matrix carbon fiber composites (PMCs) as a lighter weight, more durable replacement for Al alloys. Compared with that of Al alloys, the supplier base for PMCs is still in its relative infancy but is now growing rapidly. The accelerated growth of the PMC supplier base has largely been prompted by the new Boeing B-787, the structure of which is composed of more than 50% PMCs, mostly replacing conventional Al alloys. One long-term consequence of this significant and systematic government investment in the supplier base was the creation of a U.S. material and component supply system for the aircraft and aircraft engine industry that is still unsurpassed anywhere else in the world. More recently, the U.S. aircraft and aircraft engine manufacturers, particularly those making commercial products, have become global businesses that consistently supply more than half the aircraft and aircraft engines sold in the world every year. Consequently, commercial aircraft and aircraft engines are now a major factor in the net positive export balance for this industrial sector of the United States. This example shows that proper investments in R&D and technology can be good for the country in terms of economic security as well as national security. Today, Boeing commercial airplanes account for about 30% of the total export sales of the United States. The majority of these Boeing airplanes are powered by engines produced either by General Electric or Pratt and Whitney. More recently, government policies related to technology control and export have become a greater consideration in the global aircraft engine market, both commercial and military. Although protection of the most critical enabling technologies is appropriate from a national security standpoint, there is, at times, a fundamental conflict between national security and economic security. Further, in the global economy, partnering with foreign aircraft engine producers has become an essential aspect of gaining market access in other countries. There are well-documented times when such partnerships have been impeded by complex and sometimes arcane export control rules imposed by the federal government. Nevertheless, the three Western turbine engine producers, Pratt and Whitney, Rolls-Royce, and General Electric, clearly lead the world in the design and production of turbine engines used to power both commercial and military aircraft. Each of these three companies has developed into a global enterprise. For example, the General Electric aircraft engine business (GE Aviation) has partnerships with companies located in France, Germany, Italy, Sweden, Japan, and

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China. The activities in these partnerships range from the coproduction of military engines to the design and production of engine modules that are incorporated into commercial engines bearing the GE emblem. It also is important to note that GE Aviation also sells engines to airlines in each of these countries. Clearly, these partnerships have enabled General Electric to have a very successful and profitable aircraft engine business. The other two makers of large commercial aircraft engines, Pratt and Whitney and Rolls-Royce, have similar partnerships, some of which involve the same global partners as General Electric. This can lead to some interesting arrangements to ensure that each company’s proprietary information is protected and not inadvertently shared with its direct competitors.

10.2 BRIEF HISTORICAL PERSPECTIVE ON AIRCRAFT ENGINE EVOLUTION Although there were a few turbine-powered military aircraft at the end of World War II, it was not until the Korean conflict that turbine-powered fighter aircraft became a critical aspect of air superiority during aerial combat. Two of the early and highly successful fighter aircraft were the American F-86 Sabre jet and the Russian MIG-15. The military engine requirements justified and drove the aforementioned R&D and infrastructure investment to create the supply chain and the engine design and production technology. Initially, there were several turbine engine manufacturers in the United States, including Westinghouse. However, the military jet engine market was too small to support a large number of producers, and, over time, Pratt and Whitney and General Electric prevailed in the United States, as did Rolls-Royce in the United Kingdom and several engine producers in Russia, now the former Soviet Union. The effectiveness of these early turbine-engine-powered military fighter aircraft served as a proof of concept that promptly led to the development of jet-powered commercial aircraft. The speed and comfort of these jet-powered aircraft soon became apparent. The first jetpowered commercial aircraft entered service in the late 1950s. Two of these were in Europe, and two were in the United States. These aircraft were the U.S.-made Boeing 707 and Douglas DC-8, the UK-made de Havilland Comet, and the Frenchmade Sud Aviation Caravelle. The comet used a Rolls-Royce Avon engine, whereas the other three aircraft used variants of the Pratt and Whitney JT-3C. Since the early stages of jet engine development, it was known from thermodynamics that higher operating temperatures meant better fuel consumption characteristics and higher power density (T/W). Over the years, the ability to operate engines at higher temperatures has largely been limited by materials capability. This limitation has been incrementally mitigated, and each new turbine engine design has incorporated the next generation of materials and design features. The primary improvements that have enabled this progress are higher temperature capability materials, designs that incorporate high-pressure turbine air cooling to create increasingly larger gas metal temperature differentials, and the manufacturing capability for making turbine airfoils with internal cooling passages. Also,

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experience gained over the past 40 years or so has proven that material quality is the limiting feature for engine reliability. This chapter traces the evolution of improved materials synthesis and processing during this same period. Concurrently, the improved reliability and durability of each generation of engine has been a critical part of safe and affordable commercial air travel. In this regard, the Pratt and Whitney JT-3C powered Boeing 707 can be used as a representative baseline for early jet-powered commercial flight. The initial goal for the Boeing 707 was to have an engine capable of staying on wing without removal for overhaul for 500 flight hours. Today, a modern 747 class engine made by any of the three Western aircraft engine producers mentioned earlier is capable of staying on wing for as long as 25,000 hours, depending to some degree on the routes flown by the particular aircraft. Needless to say, this 50-fold increase not only has enormous economic benefits for the airline operators but also is reflected in lower airline fares to the benefit of the traveling public. This tremendous improvement is the result of more robust engine designs, improved engine component capability, and more capable materials. The focus of this chapter will be on the latter of these three factors, although there is a degree of interdependence among the three. Particular attention will be devoted to the evolution of jet engine materials capability and quality over the past 35–40 years.

10.3 EVOLUTION OF JET ENGINE MATERIALS It is perhaps simplest to discuss the materials used in a modern aircraft engine by relating them to the section of the engine where these materials are predominantly used. Accordingly, this chapter will consider materials used in the principal sections of the modern turbofan engine as identified in Fig. 10.2a: the fan, the compressor, the combustor, and the turbine. Although there is some overlap in material types used in these different sections, this is still probably the most straightforward way to describe jet engine material usage. Improvements in materials capability for jet engines are manifested as improved alloys, improved processing methods, and reduction of the presence of materials defects. In reality, these factors are intimately connected, and it is unusual when one or the other is solely responsible for significant improvement. For example, new alloy formulations often pose significant challenges to produce as productionscale ingots and therefore require major improvements in alloy-melting technology. As will be seen, in the limiting case for nickel (Ni)-base alloys, the transition from ingot metallurgy production to powder metallurgy production of billet for rotor forgings is required.

10.3.1

FAN

As described in this chapter’s introduction, early jet engines such as the JT-3C turbojet did not have a fan, but that quickly changed in the interest of improved

588

J. C. WILLIAMS

fuel efficiency. Later, the fan BPR increased in the interest of thrust growth, noise reduction, and continued improvement in fuel efficiency. The fan contains three critical rotating components: the fan shaft, the fan disk, and the fan blades or airfoils. The fan shaft of a modern high-BPR engine is made of high-strength alloy steel because it is required to transmit enormous amounts of torque at maximum power settings. Typically, fan shafts are made from maraging or secondary hardening steels, both of which are used in conditions having yield stress values in excess of 1700 MPa. It is well known that steels at such strength levels are prone to reduced fracture resistance if the inclusion size and content are not carefully controlled. To meet the level of fracture resistance necessary to ensure reliability of the fan shafts made from these high-strength steels, the producers of these materials have developed “clean melt” practices. The result is a highquality, very high-strength material with acceptable fracture resistance. The fan disk is the largest single rotating stage in a turbofan engine and typically is made of a Ti alloy because of the need for high strength, light weight, and good damage tolerance. As mentioned in the introduction, one of the early barriers to producing higher-BPR turbofan engines was concern about the quality of larger Ti ingots required to make forgings as large as fan disks. Over the past 30 years, there has been a concerted effort among the major Ti producers to improve the quality of rotor-grade Ti ingots. Major reductions in defect rates have been achieved by controlling the input materials used to make the first melt electrode used in multiple vacuum arc remelting (VAR) and by carefully monitoring and controlling the melt practice during each of the multiple VAR operations. Even so, rotor grade Ti alloys made by multiple VAR still contain a greatly reduced but finite number of melt-related defects. The most harmful of these defects and the most difficult to completely eliminate are the defects that result from local nitrogen enrichment. These defects are known as hard alpha or Type I defects and consist of a core of hard TiN surrounded by interstitial stabilized alpha phase. The TiN core is extremely brittle and fractures at very low plastic strain levels. If one of these defects is located at a position in the fan disk that experiences stresses in excess of the yield stress, the TiN fractures and the resulting crack propagates into the surrounding material under subsequent cyclic loading during service. This leads to a significantly shortened component life that can culminate in failure during service. Because the fan disk is such a large rotating component, it is impossible to design an aircraft engine that is capable of containing the pieces of a fan disk that fails in service. A disk failure during service can have disastrous consequences. Therefore, the materials producers, forging companies, and engine makers take extreme measures to minimize the probability of having a Type I defect in a rotating component. A second type of materials defect is traceable to the ingot stage of material production. This defect is called a high-density inclusion (HDI) and is a region that is rich in heavy metal, typically tungsten (W). There are several potential origins of HDIs, but all are related to the input material used in the first melt electrode. After a long period of policing the input material and identifying numerous

AERO ENGINE MATERIALS

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sources of W-rich inclusions, the recurring sources are W electrode tips used in gas tungsten arc welding and tungsten carbide cutting tool fragments. Both enter the input material via the revert that is used to make up as much as 34% of the first melt electrode. Today, all of the revert used to make rotor-grade material is x-rayed to identify and remove the HDIs before the material is included in the first melt electrode. This has led to a major reduction in the frequency of HDIs in the final mill product. Unlike Type I defects, HDIs are tightly bound to the Ti matrix and therefore are not readily detected during ultrasonic inspection. More details of Ti metallurgy and processing are given in Chapter 2. Beginning in the early 1990s, a new melting technology for rotor-grade Ti alloys was introduced. This melting method is called cold hearth melting and is a completely different technology from VAR. Essentially, the metal is melted by independent heat sources, either electron beams or plasma torches. The melting is done in an evacuated or argon (Ar) filled chamber by containing the molten Ti alloy in a large, water-cooled copper hearth (the cold hearth), which always maintains a thin layer of solid Ti alloy that is in direct contact with the molten metal. The molten Ti flows down the hearth and into an ingot mold, which can be circular or rectangular in cross section. One of the benefits of cold hearth melting is that it permits control of the time that the Ti alloy is in a molten state, which is not possible with VAR. Rotor-grade Ti alloys made by cold hearth melting have demonstrated reductions in Type I defects by as much as two orders of magnitude. Another important difference between VAR melting and hearth melting is that in a VAR ingot, all of the material that is part of the VAR electrode becomes a part of the ingot, as there is no other place for the material to go. In hearth melting, the molten metal flows along the hearth, and any input pieces that are significantly dense sink and are trapped in the mushy skull. They therefore are eliminated from the ingot. As a result, the current rate of HDI “finds” in hearth-melted materials is essentially zero. Consequently, cold hearth melting has become a widely used, even preferred, method of producing rotor-grade Ti alloys for the engine industry today. Until recently, all of the fan blades for high-BPR turbofan engines were made from solid Ti alloy forgings. Even for the second-generation turbofan engines with BPRs on the order of 6, such as the GE CF-6, the Rolls-Royce RB-211, and the Pratt and Whitney JT-9, the fan blades were solid Ti forgings. Examples of solidforged Ti fan blades are shown in Fig. 10.8. More recently, as the overall size of the engines used on large, double aisle, twin-engine aircraft such as the Boeing 777 has increased, the BPR and the fan blade chord have also increased. As a result, these engines require very large, wide fan blades and must have a lighter-weight construction than can be achieved with solid-forged Ti blades. In response to this new requirement, two completely different fan blade designs have emerged. One uses hollow Ti fan blades, whereas the other uses PMC fan blades. Both technologies, although relatively new, are now quite mature and have been used successfully in commercial service for more than 15 years. Examples of large hollow Ti and PMC fan blades are shown in Fig. 10.9.

590

Fig. 10.8

J. C. WILLIAMS

Solid-forged Ti fan blades: 747 class (left) and 737 class (right).

The fan frame, or front frame, is the remaining important component in the fan section. It is a static component that contains the front main shaft bearing and often the front engine mount, making it comparable in criticality to the three rotating components already described. Over the years, there has been a shift from a fabricated fan frame to a one-piece investment cast Ti alloy hub frame for 737-class engines and cast central hubs with separately cast struts for larger engines. The struts are attached to the hub by welding. This change was enabled by a major advance in investment casting

Fig. 10.9 Large 777-class fan blades: hollow Ti (left) and PMC with thin Ti erosion guard on leading edge (right).

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591

technology for Ti alloys. The benefit of these cast frames is that they replaced fabricated frames consisting of a number of individual pieces that were mechanically fastened or welded together. The fabricated frame was a heavier and very laborintensive component that was expensive to make and was less structurally efficient. Further, the holes for mechanical fasteners are intrinsic stress concentrators that frequently can serve as fatigue crack initiation sites. These features require periodic inspection, which increases the maintenance cost for this component. Examples of investment cast Ti frames are shown in Fig. 10.10. Pratt and Whitney recently introduced a geared turbo fan (GTF) engine in which the fan shaft is connected to the low-pressure turbine (LPT) through a gearbox that reduces the fan rotational speed relative to the LPT. This permits the LPT to operate at more optimum (higher) speeds, which has associated efficiency gains. Several new regional aircraft have elected the GTF engine because of the improved fuel efficiency (claimed to be 10%). A GTF engine also is being offered as one of two options for the improved version of the Airbus A-320, called the A-320neo (new engine option). This concept places 100% confidence on the gearbox reliability, but Pratt and Whitney has done a lot of testing so that this is not a major concern. An open question is how large the GTF engine can become because of the need for an oil cooler for the gearbox oil. A very large turbofan generates .100,000 shaft horsepower (SHP) in the fan shaft at takeoff, and so the heat generated by the gearbox, even when it is extremely efficient, is significant. Only time will tell how this plays out and how far up the thrust

Fig. 10.10 Investment cast Ti frames. Frames as shown are cast plus lightly pickled. Feature definition is done by casting.

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scale the GTF engine concept can be efficiently used. Clearly, a gearbox transmitting 100,000 SHP will add weight to the propulsion system that will need to be “bought back” by efficiency improvements and weight reductions in the LPT.

10.3.2

COMPRESSOR

Among the three large engine producers, Pratt and Whitney, Rolls-Royce, and General Electric, there is a divergence in design philosophy regarding the overall turbofan engine architecture. Pratt and Whitney and General Electric use a so-called two-spool architecture in which the fan and low-pressure compressor are driven by a shaft connected to the low-pressure turbine (the low-pressure spool). The high-pressure compressor is driven by a separate, concentric shaft that is connected to the high-pressure turbine (the high-pressure spool). On the other hand, Rolls-Royce uses a so-called three-spool architecture that includes a lowpressure spool in which the low-pressure turbine drives the fan and low-pressure compressor, an intermediate-pressure spool in which the intermediate-pressure turbine drives an intermediate-pressure compressor, and a high-pressure compressor spool that is driven by the high-pressure turbine. Both of these design philosophies have proven to be effective. Nevertheless, there are arguments for and against both architectures, largely promulgated by the advocates of the two- or three-spool design philosophies. In the end, the gas temperature gradient from the fan inlet to the exhaust nozzle exit is similar for both engine architectures and depends mainly on the overall pressure ratio of the engine design. As will be described later in this chapter, the materials challenge in the compressor is dominated by the gas exit temperature (T3). Consequently, all of the subsections of the compressor will be generically referred to as the compressor in this chapter. The compressor airfoils have a very thin cross section because their function is to “pump” the air, thereby compressing it. Consequently, compressor airfoils are produced by a wrought process, either forging or another deformation process. The need to manufacture the airfoils by wrought processing imposes a limit on the alloys that can be used, in contrast to the situation in the turbine, which will be discussed later. Also, the slender leading edges of the compressor airfoils make them susceptible to foreign object damage, which is another reason more ductile alloys are attractive for this application. In any circumstance, the temperatures in the front of the compressor are low enough that Ti is used extensively, both for the rotor and the airfoils, because of its excellent strength-to-weight characteristics at temperatures up to and including 6008C. Beyond 6008C, Ni-base superalloys are used because of their superior temperature capability and because the embrittlement of bare (uncoated) Ti alloys above 6008C that is caused by oxidation is a major concern. Until fairly recently, the Ni-base alloys used in the compressor were wrought alloys produced by ingot metallurgy methods. More recently, as the compressor exit temperature (T3) and turbine inlet temperature (T41) continue to increase in the interest of improved fuel consumption characteristics, the temperature capability of

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wrought Ni alloys, such as Waspaloy, In-718, and U-720, has become marginal. In place of these wrought alloys are the newer, more highly alloyed Ni-base superalloys such as MERL-76, R-88DT and a powder metallurgy (PM) version of U-720. Figure 10.11 compares the temperature capability of several Ni-base alloys used for rotors in the back of the compressor and all of the turbine. For a variety of reasons, each engine company has a preferred alloy; the aforementioned alloys in each class are mainly used by Pratt and Whitney, General Electric, and Rolls-Royce, respectively. From a materials standpoint, the last compressor stage presents the greatest challenge because there is no provision to cool the compressor rotor and airfoils as is done in the high-pressure turbine. (Cooling of the turbine components will be described in the turbine section of this chapter.) Consequently, the gas temperature and the metal temperatures are the same in the compressor. Further, the T3 values of the latest engine models challenge the temperature capability of even the PM Ni-base alloys. (Refer to the Chapter 2 section on superalloys for more details.) Processing of these alloys can have a profound effect on properties, which creates the opportunity to tailor the materials for a particular design situation. For example, grain size can be controlled by a combination of forging practice and heat treatment. The four critical properties for a rotor material are ultimate tensile strength (UTS), creep strength, low cycle fatigue (LCF) life, and fatigue crack growth rate (da/dN). Figure 10.12 shows the effect of grain size on these four properties. This figure shows that two of the properties (creep strength and crack growth rate) have a positive dependence on grain size (i.e., the property improves with increasing grain size), whereas the other two properties (UTS and LCF life) show a negative dependence. Historically, the typical processing routes for Ni-base rotor alloys has aimed to create a mean grain size of 10–20 mm. However, if better creep strength is needed, so-called coarse grain processing can be used to provide material with a 40–60 mm grain size. Such processing is

Fig. 10.11 Ni-base disk alloy temperature capability for ingot metallurgy (IM) and powder metallurgy (PM) processing.

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Effects of grain size as controlled by processing on four critical properties for

helpful but comes at the expense of two of the other properties (UTS and LCF life). Coarse grain processing also requires a more uniform plastic strain in the forging than standard processing to ensure a uniform recrystallization response. This can pose limitations on the forging geometry that can increase the buy-to-fly ratio of the part. These newer generation, higher temperature capability alloys all contain higher concentrations of strengthening solutes, which imparts better hightemperature strength to them. All of these newer, solute-rich alloys are made by PM methods because of the difficulty encountered in making large-diameter ingots that are free from solute segregation after solidification. The introduction of PM Ni alloys for rotating parts is a major undertaking because it adds significant cost but also requires a number of other changes in design practice and service life calculations. For example, the gas atomized powder used for these forgings typically contains indigenous oxide inclusions formed by oxidation of the reactive alloying elements (e.g., Al) during the atomization process. Also, some inclusions are incorporated into the powder during handling. To manage the inclusion size distribution, the powder is screened after atomization to remove inclusions larger than the smallest screen opening. One typical screen opening is 53 mm (–270 mesh). Powder screened to this maximum particle size is then assumed to have only oxide inclusions smaller than this, and so the remaining question is the concentration or number density of these particles and their size distribution under the maximum 53 mm size imposed by screening. There are a variety of methods to determine these data, and all users of PM Ni-base alloys employ at least one method. The reason these data are so important is that fatigue cracks preferentially nucleate at these inclusions, and there is good

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but qualitative evidence that larger particles are more potent crack initiation sites. Thus, there is a stochastic element in estimating fatigue life in PM rotating parts. This moves the life methods out of the deterministic domain into the use of probabilistic methods. This is a big change, but one that has been successfully accomplished by analysts in the engine companies.

10.3.3

COMBUSTOR

As shown in Fig. 10.2a, the combustor is located between the high-pressure compressor and the high-pressure turbine. In the combustor, the hot, highly compressed air discharged from the compressor is mixed with fuel, and the mixture is ignited to create the hot gas stream from which the turbine sections extract the necessary work to drive the compressor and fan. The operating temperatures in the combustor are the highest of any location in the engine, but fortunately the operating stresses are relatively low because the combustor is essentially a static component and therefore has none of the stresses associated with a rotating part. Nevertheless, the temperature and stress combinations are high enough to exceed the capability of the metallic materials used to make the combustor. As a consequence, a liner and cooling air are used to control the temperature of the combustor case and liner at levels that are within the capability of the respective materials. The combustor liner is a relatively thin structure that shields the combustor case from the direct flame in the combustor. The combustor case is essentially a static pressure vessel that operates at a high temperature, but one that is lower than that of the combustor liner. Until now, all-metal (except for ceramic coatings) combustors have served the turbine engine industry well. Over time, significant progress has been made in the combustion efficiency of the combustor, including reducing the emissions created by the burning of the fuel. In the early days of jet-powered aircraft, for example, a trail of dark smoke, caused by incomplete burning of the fuel, was commonly observed emanating from the engine exhaust stream, especially during maximum power settings such as takeoff and climb. Today, there is essentially no smoke observable from the modern aircraft engine exhaust. This is due in part to more efficient burning of the fuel and in part to a more effective mixing of the combustion products as they exit the combustor. Notwithstanding the progress just mentioned, the management of the fuel–air mixture in a modern combustor is still a challenge, in part related to the use of cooling air mentioned earlier. The emissions of concern are carbon monoxide (CO) and oxides of nitrogen (NOx). As the sensitivity toward the emissions from aircraft engines continues to grow, there is a clear need for further innovation with regard to combustor design. The range of these design possibilities will be largely dictated by the availability of materials that are capable of operating at higher temperatures, thereby eliminating or dramatically reducing the need for cooling air for the combustor liner. In this regard, ceramic matrix composites (CMCs) are a promising class of materials for combustor liners in next-generation, low-emission combustors. CMC systems

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Fig. 10.13

CMC combustor liner after running in a factory demonstrator engine.

capable of withstanding 14508C will be required in order for the combustor to operate without cooling air on the liner. Today there are ceramic fiber and matrix combinations capable of retaining their strength at such operating temperatures, but it is unclear whether these material systems will retain their fracture toughness after extended exposure periods (e.g., 10,000 h at temperatures as high as 14508C). A considerable research effort is underway in government agency laboratories such as NASA. Aero engine companies are also working on developing CMC systems, with the initial application most likely being combustor liners. An example of a CMC combustor liner for a small engine that has been run in a factory test is shown in Fig. 10.13. This liner is made from SiC fibers with a SiC matrix that has been deposited by vapor deposition. There are other fibers, such as aluminum oxide (Al2O3), but these have different thermal expansion values and its poor thermal conductivity makes Al2O3 a poor choice for a matrix. Once a suitable ceramic composite system (fiber, matrix, and fiber coating) is identified, the supplier base for ceramic fibers and for CMC components also will require substantial further development. The relatively small market for CMCs will limit the willingness of private investment to create such a supplier base. A design-related issue associated with the combustor that impinges on materials and on engine efficiency is the uniformity of the temperature at the exit of the combustor or inlet to the turbine. An uneven temperature distribution is referred to as hot streaks, and these become the effective T41 value that limits the engine operating temperatures.

10.3.4

HIGH-PRESSURE TURBINE

The high-pressure turbine (HPT) is the most technologically sophisticated section of the turbofan engine, largely due to the operating temperatures and stresses

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experienced by components in this section. The newest and most fuel-efficient turbofan engines have turbine inlet temperatures (T41) that exceed 14508C. As shown in Table 10.1, this is an increase of about 5508C since the introduction of the Pratt and Whitney JT-3C engine for the Boeing 707. When considering T41 values, it is important to remember that although the more recent T41 temperatures are the gas temperature, these are well above the temperature capability of even the most advanced cast monocrystal turbine airfoils made from the newest Ni-base alloys. The origin of this temperature differential will be discussed later in this section. Early in the evolution of the aircraft gas turbine, it became clear that the benefits of reduced fuel consumption that accrue to higher T41 operating temperatures were a critical factor for the success of jet-powered commercial aircraft. Consequently, engine companies diligently pursued various means of achieving higher T41 values. The resulting innovations can be divided into materials and process solutions and design solutions. The materials and process solutions are twofold: 1) development of higher temperature capability turbine blade alloys and 2) the development of investment casting technology for turbine airfoils that permits control of the grain structure of the casting. The design solution, which complements the materials and process solutions, is to air cool the turbine airfoils to create a gas–metal temperature differential. Although these two classes of solutions are related, they will be discussed separately here for clarity. The evolution of HPT blade materials and investment casting technology over roughly the past 40 years is an interesting case study. In the 1960s, cast HPT airfoils replaced the earlier airfoils that were forged. The motivation for introducing cast turbine airfoils was twofold: 1) the newer, higher temperature capability alloys were richer in solute and were extremely difficult, if not impossible, to forge; and 2) the introduction of the intricate internal passages for air cooling the airfoils could not be created in the solid-forged article by machining. However, like the forgings they replaced, these cast airfoils also had an equiaxed (EQ) grain structure (as is the case with most castings). The creep strength of these first EQ cast airfoils was considerably better than that of the forgings they replaced. This was mainly because of the larger as-cast grain size and the higher solute content of the new alloys, but the creep strength of these EQ castings was still limited by grain boundary sliding. As turbine blade investment casting technology evolved and was improved, the capability of casting turbine airfoils with columnar grain structures was demonstrated and then reduced to practice. This advance is largely attributable to the development of casting furnaces capable of creating a very steep temperature gradient at the liquid–solid interface as the investment casting shell was withdrawn from the furnace hot zone. This columnar structure only has grain boundaries oriented parallel to the turbine blade axis (i.e., there are no grain boundaries oriented to have a high shear stress that drives grain boundary sliding). This columnar structure is often called directionally solidified (DS), and DS airfoils exhibited considerably better creep strength than the EQ structure of the first investment cast turbine airfoils. However, Ni-base superalloys have a

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tendency to exhibit grain boundary cracking when subjected to transverse stresses (i.e., normal to the DS grain boundaries) or to shear stresses. Such stresses are attributable to anisotropic thermal expansion or to mechanical loads induced when the rotating airfoils interfere with the static components, causing a rub. A rub causes bending about an axis perpendicular to the DS grain boundaries. This grain boundary cracking can be mitigated by adding the so-called grain boundary strengthener elements such as boron, hafnium, or zirconium to the alloy in low concentrations. These grain boundary–strengthening additions are effective in preventing grain boundary cracking, but each of them lowers the melting temperature of the turbine airfoil alloys. As a consequence, the effective homologous operating temperature (expressed as a fraction of the alloy melting temperature) of the columnar grain structure turbine airfoils was higher because of these alloying additions. These additions are harmful to the creep strength at maximum temperatures because of the attendant increase in diffusion rates. It was then realized that these additions could be eliminated if the highangle grain boundaries were completely eliminated from the cast structure. This led to the further development of casting technology capable of casting Ni-base turbine airfoils that were free of high-angle grain boundaries. Such castings are known as monocrystals (MX). Once the capability to cast MX turbine airfoils was demonstrated, the grain boundary strengtheners were removed and the temperature capability of the turbine airfoils was substantially improved. Figure 10.14 shows examples of cast turbine airfoils with EQ, DS, and MX structures. In the aero engine industry, turbine designers are perpetually interested in higher operating temperatures because of the beneficial effects higher temperatures have on engine performance, particularly reduced fuel consumption. Accordingly, several generations of MX turbine blade alloys have been developed,

Fig. 10.14 Investment cast and macro-etched turbine airfoils with three different grain structures: equiaxed, directionally solidified, and monocrystal.

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Fig. 10.15 Temperature capability of eight turbine airfoil alloys with different composition and processing as noted. each of which has a higher temperature capability than the previous one. The improvement in temperature capability of cast turbine airfoil alloys over time is shown in Fig. 10.15. The data shown in Fig. 10.15 pertain to EQ, DS, and MX castings and to the effects of alloy composition. This figure allows the effects of both alloy composition and grain structure to be compared. Improved MX alloy temperature capability is obtained through the addition of higher concentrations of refractory metal alloying elements, in particular Rhenium (Re). This is mainly because refractory metals diffuse more slowly than the other alloy components (e.g., Ni, Al, and Cr). In MX turbine airfoils, where grain boundary sliding is eliminated, diffusion is still the rate-controlling parameter for high-temperature creep because it controls dislocation motion, including climb. However, an associated issue with these advanced, refractory element–rich turbine blade alloys is that the refractory elements displace other alloying elements that improve the oxidation resistance of the Ni-base alloys. In particular, the alloys become more dilute in chromium. These refractory element–rich alloys also have been shown to be prone to microstructural instability during longtime elevated temperature service. This instability can be manifested either as the formation of secondary recrystallization zones or the formation of topologically close packed phases such as sigma (s) and pi (p). As in most metallurgical problems, this becomes a trade-off in which oxidation resistance and alloy stability is traded against creep resistance. In the end, Cr- and Al-rich coatings can restore much of the oxidation resistance, whereas the improved creep properties are intrinsic to the alloy composition. However, it is important to remember that the higher creep strength turbine blade alloys have lower intrinsic oxidation resistance. Thus, if the coating

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is breached in service, environmental attack of the exposed base metal occurs more rapidly. This is a particular problem in the case of the turbine blade tips where a rub between the rotating turbine blade and the static structure will breach the coating. This results in more rapid oxidation of the tip than the rest of the airfoil and causes the clearance between the turbine blade and the static structure to increase, leading to a time-dependent efficiency loss of the engine. As mentioned earlier, starting several generations ago, with each new engine model the turbine designers opt for such high T41 values that there are no Ni-base superalloys available that are capable of withstanding these high temperatures for the long-time intervals between overhauls. Consequently, a design solution that creates a gas–metal temperature differential was introduced. The response to this requirement was to introduce air cooling passages into the turbine airfoils to maintain a temperature differential between the hot gas and the metallic airfoil. Air cooling of turbine airfoils enables engine operation at T41 values well in excess of the turbine blade alloy temperature capability, as can be seen from Table 10.1 and Fig. 10.15. These cooling schemes use air, which is bled from the high-pressure compressor, routed to the high-pressure turbine, and forced through the internal cooling passages that are cast into the turbine airfoils. Small holes are created that connect these passages and the airfoil surface. Some of the cooling air is transmitted through these holes to form a film of cool air on the airfoil surface that shields the underlying metal from the hot gas. This technique of protecting the airfoil is called film cooling. Examples of these holes in the surface of an MX airfoil are shown in Fig. 10.16a. Further, examples of these internal passages are shown in Fig. 10.16b, which is an MX airfoil with the exterior surface removed to expose the internal air passages. These passages are created during casting by using ceramic cores shaped like the desired cooling passages. These cores are placed where they are desired inside the investment casting shell. The shell is filled with molten metal and withdrawn from the casting furnace with the cores embedded inside the metal casting. After casting, the cores are removed chemically using chemical solutions, such as concentrated KOH, that attack and dissolve the ceramic core but not the metal casting. The desire for higher T41 values is essentially insatiable; consequently, more and more elaborate cooling schemes have been devised. Given that all high-pressure turbine airfoils in modern turbofan engines are cast as MX, the advances in casting technology required to produce articles that remain MX but still incorporate these intricate internal passages are quite remarkable. Engines with as much as 1508C gas– metal temperature differentials are in service today. Notwithstanding the benefits realized from improvements in turbine airfoil alloys, including the successive introduction of several generations of higher temperature capability alloys, and in casting technology starting with EQ and ending with MX and the implementation of cooled turbine airfoils, new innovations are always welcomed by the turbine designers. Therefore, a coating technology has also evolved for air-cooled HPT airfoils. This technology involves applying a thin (125 mm) layer of ceramic coating to the surfaces of the

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Fig. 10.16 Investment cast turbine airfoils showing (a) film cooling holes and (b) internal cooling channels cast into the airfoil.

turbine airfoils that are exposed to the hot gas stream. This coating is called a thermal barrier coating (TBC) because it limits the heat transfer from the hot gas stream to the underlying metallic turbine airfoil. The use of a TBC enables the engine to operate with a larger hot gas–metal temperature differential. Figure 10.17, shows a cross section of a turbine airfoil with a TBC applied on the surface that faces the hot gas stream. From this it can be seen that the TBC is a multilayer coating system consisting of a ceramic top coat and a bond coat to enhance adherence of the ceramic top coat to the Ni-base airfoil. TBCs are now standard in all modern HPTs. However, as with ceramic coatings on metals in general, there is a thermal expansion compatibility problem, and spalling of the ceramic coating during service is an ongoing concern. The use of a metallic bond coat improves the adherence of the ceramic TBC and has been helpful in this regard. Nonetheless, spalling of the TBCs still occurs over time. Therefore, TBCs today are used essentially as a life extension feature rather than as a prime reliable coating, which would allow the maximum gas–metal temperature differential to be realized. In the end, TBCs are still very helpful in improving longtime turbine durability at high operating temperatures. They also enable a reduction in the amount of cooling air used to cool the turbine airfoils. It should be mentioned that the cooling air that is bled from the compressor is air that the compressor pumps but that does not participate in or contribute to the thrust produced by

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the engine. Therefore, minimization of cooling air consumption is an important contributor to improving the fuel consumption characteristics of an engine. In this regard, TBCs are very helpful.

10.3.5

LOW-PRESSURE TURBINE

The LPT drives the fan and low-pressure compressor, but because it is downstream from the HPT, it also operates at considerably lower temperatures. This is because much of the energy from the hot gas stream has already been extracted by the HPT. Except in some military engines, the LPT airfoils are solid, EQ investment Ni-base alloy castings (i.e., they are not air cooled). Because the hot gas stream expands as it moves toward the engine exhaust nozzle, the LPT has a larger diameter than the HPT. Consequently, the LPT airfoils are longer and therefore more prone to casting defects if the casting process is not tightly controlled. The lower operating temperatures of the LPT permit the use of Ni-base alloys that have better castability, which tends to offset the tendency for casting defects. Because the LPT airfoils are longer and uncooled, they are also considerably heavier than HPT airfoils and therefore present an opportunity for weight reduction. In this regard, there is a relatively new Ti-based intermetallic compound (TiAl) that has been shown to have comparable temperature capability to the most common Ni alloys used for conventionally cast LPT blades. The importance of this new material is that it has half the density of the comparable Ni-base alloys with nearly equivalent creep strength. These lightweight TiAl LPT airfoils permit elimination of as much as 200 lbs per LPT stage in a 60,000-lb thrust class aero engine. Much of this weight reduction is due to the reduced mass of the Ni-base disk because the reaction forces caused by the lighter-weight TiAl airfoils are so

Fig. 10.17 Cross section of cast turbine airfoil with thermal barrier coating applied, showing (top to bottom): ceramic topcoat, metallic bond coat, diffusion zone, and turbine airfoil alloy.

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Fig. 10.18 Cutaway of GEnx LPT showing last two stages of cast TiAl airfoils. much smaller. The first commercial engine to enter revenue service using TiAl LPT airfoils is the GEnx, which is used on the new Boeing 787 and the stretched version of the Boeing 747 (the B747–8). The particular intermetallic alloy General Electric is using for this application has a composition Ti-48Al-2Cr-2Nb (stated in atomic percent). A picture of the GEnx LPT stage with TiAl airfoils installed is shown in Fig. 10.18. Because it is an intermetallic compound, TiAl and variants thereof have lower ductility than conventional Ni or Ti alloys. Consequently, the LPT airfoils in the GEnx engine are cast because it would be very challenging to forge them without cracking. Not only does the use of cast airfoils aid in manufacturability, but the coarser structure of the casting also imparts improved creep strength.

10.4 MATERIALS PROCESSING The preceding paragraphs have described generically the types of alloys used in the different sections of the modern aero engine. In some cases, the processing method used to produce them has also been briefly described. For a reader unfamiliar with the general subject of aircraft engine materials, this account may sound quite straightforward. In reality, the technology necessary to produce highquality materials with reproducible properties has had a long evolutionary path. The capability for reproducible processing of Ti- and Ni-base alloys of the quality used in jet engines, particularly for rotating parts, is one of the true success stories of metallurgical engineering. The critical nature of the rotating parts, such as the disks used in the fan, compressor, and turbine sections, makes the performance and quality requirements for these materials and components among the most demanding for any manufactured product in the world. In service, a large rotating part such as a fan disk has a tremendous amount of stored energy because it is essentially a flywheel turning at very high rotational speeds. It is impossible to design an aircraft engine with sufficiently robust casings to contain the pieces of a disk should it fail during operation of the engine. In the rare instances when a disk does fail, these events are known as uncontained failures. Fortunately, these events are very rare and are

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becoming even less common as years go by. A major contributing factor to the reduction in uncontained failures has been the evolution and improvement of melting and forging technology for both Ti- and Ni-base alloys. Although the details of the melting methods used for each of these two material classes are quite different, the common goal of eliminating melt-related defects is pervasive. Today, the class of materials used for rotors made of Ni alloys (known as rotor grade) is typically triple melted using three different melting techniques sequentially. The first melt is accomplished by vacuum induction melting (VIM), the second melt is accomplished using electro-slag remelting (ESR), and the third melt is accomplished by VAR. In combination, this triple-melting method produces a minimal concentration of inclusions in ingots of good homogeneity. In the case of Ti alloys, VAR has been the standard practice for many years and is still practiced today but has been supplanted by cold hearth melting. For many years the standard practice for melting rotor-grade Ti alloys was 3X VAR. However, cold hearth melting has become the preferred melting practice for much of the rotor grade melted today. Cold hearth melting reduces the number of melt-related defects that are unique to Ti alloys, principally interstitial stabilized inclusions (also known as Type I) and HDIs. All major Ti suppliers that produce rotor-grade Ti alloys for the aero engine manufacturers now have the capability for cold hearth melting of Ti alloys. Acceptance of cold hearth melting is not universal, but the process is in wide use today and its acceptance is growing. The production of rotor-grade Ni and Ti alloy forgings used in jet engines is also a highly specialized business that requires meticulous attention to detail. This detail begins with high-quality ingots produced by the melting techniques outlined previously. Ingots are broken down at high temperatures in a large forging press with considerable attention given both to the amount of plastic strain and the temperature at which the strain is being introduced. Once ingot is reduced to a large mill product form, it is heat treated to recrystallize the material with the intent of creating a uniform microstructure that is finer than the as-cast structure. The recrystallized material is then further reduced in size to produce a cylindrical product known as forging billet. The diameter of the forging billet is typically in the 10–25 cm range. This is a critical stage in the production of the material because the macro- and microstructure are largely determined by the practice used to produce the forging billet. Different materials producers have varying types of equipment in their plants, and, consequently, there is no single “recipe” for the production of the rotor-grade forging billet. Irrespective of the details of the equipment being used, several guiding principles must be followed when producing rotor-grade forging billet. One is the careful attention to the workpiece temperature and the amount of strain being introduced during working operations. The temperature, strain, and strain rate combination used to qualify the billet production process is known as the process window. If processing is not done in such a way that the material is maintained in the process window, the resulting billet is not acceptable. For example, if the strains are too large or the

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working temperature becomes too low, the billet can contain small voids known as strain induced porosity. These small voids are typically smaller than the detection limit for ultrasonic inspection, and therefore the final forging can contain the small voids. If this happens, the voids can serve as early fatigue crack initiation sites depending on their location, leading to a shortened life of the rotor.

10.5 TRENDS AND OUTLOOK The foregoing sections make it clear that higher operating temperatures are the key to continued improvements in fuel efficiency. This need is mitigated by the concurrent necessity of containing ownership costs of commercial aircraft, which include first cost and engine maintenance cost. Therefore, increases in operating temperatures that cause more rapid deterioration of the turbine will be faced with reluctance from the operators (airlines and leasing companies) and will be difficult to implement. The best place to test this may be in new, large military aircraft engines where range is more important with each passing year. Because the number of places where the United States can rely on having foreign military bases continues to decrease, the global reach of these military aircraft is becoming increasingly important from a strategic standpoint. In such cases, cost becomes a secondary consideration. It is clear that metallic turbine components are approaching their operating temperature limits, even with improved cooling schemes. Further, as the OPR of each generation of engine increases, the “cooling air” temperature increases along with T3. Much study has been devoted to the use of “cooled cooling air,” which would be achieved by extracting the cooling air from the compressor, routing it through a heat exchanger, repressurizing it, and injecting it into the turbine as cooling air. From a reliability standpoint, this increase in complexity is not attractive. Even today, a majority of reliability issues are related to malfunction of components outside the engine flow path. Adding complexity without sacrificing reliability is always a challenge. Looking forward to prospects for new materials and process technology in turbine engines, the use of ceramic composites, as mentioned earlier, is the most attractive way to defeat the metals temperature limits. The challenge will be in finding material systems (fiber, matrix, and fiber coating combinations) that have long-term stability at temperatures above those where metals can operate. Also, the ability to manufacture airfoils with cooling capability that is equivalent to that being realized in metal turbine airfoils will pose a significant challenge in a fiber-reinforced material. Nevertheless, CMCs are the class of material most likely to be used in the hot section of new turbine engines. Recent generations of engines have made increasing use of PMCs for the fan and now the fan case. Additional PMC applications will no doubt be found, but the long-term temperature stability of these materials will limit how far back in the engine they can be used.

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The other trend that must be mentioned is the geared turbofan that Pratt and Whitney is offering in the market. This concept offers a significant improvement in fuel efficiency, which will almost certainly induce a response from General Electric and Rolls-Royce. This kind of competition makes everyone better, and the ultimate beneficiaries are the airline operators and the flying public.

10.6 SUMMARY This chapter has attempted to trace history of the aircraft gas turbine engine from its inception in the late 1950s through the present. The emphasis is on the contributions of materials and processing to the substantial progress that has been made, but the combined results of design and improvements in materials and materials processing methods are the real story. The increases in T/W and the improvements in SFC nicely summarize the performance improvements. For commercial aircraft, the improvements in reliability and reduction in cost of ownership are equally important. These improvements also are discussed and illustrated. The ability to operate twin-engine aircraft long distances overwater provides a compelling testimony to the importance of these improvements. The quest for better materials and designs will continue, but the practical range of new opportunities narrows with each new generation of material. The aircraft gas turbine is a monumental success story of engineering teamwork. It is a story that should be better told to the general public when the nature of engineering is being explained.

BIBLIOGRAPHY Because of the general nature of this chapter, specific references to each point are not practical. Instead, the following resources cover everything described in the chapter (and more) in considerable detail. NI-BASE ALLOYS Huron, E. S., Reed, R. C., Hardy, M. C., Mills, M. J., Montero, R. E., Portella, P. D., and Telesman, J. (eds), Superalloys 2012, Wiley, New York, 2012. Reed, R. C., The Superalloys: Fundamentals and Applications, Cambridge Univ. Press, New York, 2006. Reed, R. C., Green, K. A., Caron, P., Gabb, T. P., Fahrmann, M. G., Huron, E. S., and Woodard, S. A. (eds), Superalloys 2008, Wiley, New York, 2008. Sims, C. T., Stoloff, N. S., and Hagel, W. C., Superalloys II, Wiley, New York, 1987.

TI ALLOYS Donachie, M. J., Titanium: A Technical Guide, ASM International, Materials Park, OH, 2000.

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Leyens, C., and Peters, M. (eds), Titanium and Titanium Alloys, Wiley, New York, 2003. Lu¨tjering, G., and Williams, J. C., Titanium, 2nd ed., Springer-Verlag, New York, 2007. Niinomi, M., Akiyama, S., Hagiwara, M., Ikeda, M., and Maruyama, K. (eds), Ti-2007 Science and Technology, Japan Institute of Metals, 2007. Zhou, L., Chang, H., Lu, Y., and Xu, D. (eds), Ti-2011 Science and Technology, Chinese Institute of Metals, 2012.

CHAPTER 11

Materials for Solid Rocket Engines Diep V. Trinh NASA Marshall Space Flight Center, Huntsville, Alabama

11.1 INTRODUCTION Solid rocket engines or solid fuel rocket motors are appropriately named, as they are powered by solid fuel propellants. Such propulsion systems are more generally referred to as solid rocket motors (SRMs). When SRM engines are used in the booster stage of a launch system, they are referred to as solid rocket boosters (SRBs). Because some of the surviving structures on the space shuttle’s SRB engines were recovered, refurbished, and reused for other missions, these booster motors were often referred to as reusable solid rocket motors (RSRMs). Shuttle RSRM systems were some of the largest SRM engines ever built for flight, until the even larger five-segment reusable solid rocket motor (RSRMV) was developed for NASA’s Space Launch System (SLS). A substantial amount of technical information has been extensively documented for the RSRM configuration, and this provides an excellent resource of SRM/SRB knowledge. Thus, this discussion makes use of the insight available regarding large-scale, state-of-the-art SRM technology relative to the RSRM platform. However, it should be noted that SRMs come in a variety of sizes and configurations, as they are also used in various military hardware, missile systems, launch vehicles, and unique small motor applications throughout the space, aerospace, and defense industries. A simplified overview of the primary components in modern SRMs is shown in Fig. 11.1. The components include 1) a casing or casing segments, which are often composed of steel (other designs have used filament-wound composite casings); 2) an insulation layer, a rubbery material applied to the inner casing walls to provide thermal insulation during the rocket burn cycle; 3) a liner layer, which is applied to the insulation to facilitate binding between the propellant and insulation; 4) propellant grain, which provides combustion/thermal energy for the rocket; 5) an igniter motor, which is incorporated at the front or forward end of the casing to ignite the propellant grain at the beginning of the burn cycle; and 6) the nozzle section at the back or aft end of the motor, which is responsible for converting thermal energy from the combustion process into kinetic energy, providing thrust for the rocket. This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Fig. 11.1 Basic illustration of the space shuttle’s reusable solid rocket motor showing the major components. (Courtesy of NASA/ATK Thiokol.) Monolithic, single-case configurations have often been the rocket designer’s dream. Not only do they improve grain installation and packing uniformity, but they also improve manufacturability in many respects as well as safety factors by the elimination of joints and joining mechanisms that are necessary when multiple segments are used. As history has shown, issues with joints, seals, and compartmentalized stage sections have sometimes produced undesirable consequences. Ironically, the manufacturing requirements for modern, large-scale Earth-to-orbit SRM platforms are often governed by constrictions that are imposed or highly influenced by logistics factors and/or domestic producibility limitations. Such constraints have generally forced the industry to manufacture large SRM structures in multiple segments and segment assemblies that can be manufactured, packed, and shipped to the launch site for mating and final assembly operations. Shuttle SRB structures have always been fabricated using multiple sections and segments. Consequently, the specific technologies and mechanisms required to effectively join and seal these segments together during assembly has always been quite critical. Modern large-scale SRM platforms must make use of the very best in state-of-the-art joining and sealing methodologies because such SRM casings are subjected to extreme pressures during the burn cycle. Although the specifics are not yet well defined, next-generation SRM technology for space launch vehicles will include many improvements over the RSRM approach, as evidenced during design phases of the RSRMV. Some advanced ideas and concepts may eventually find applications in upgrades for SLS and future vehicles.

11.2 SOLID ROCKET COMPONENTS 11.2.1

PROPELLANT GRAIN

For solid fuels, the final mixture is in the solid (or semisolid) state, and most of the constituents are composed of solid particles. Although there may be a number of

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components in any given solid propellant grain formulation, modern solid fuel mixtures are designed around three functional constituents: 1. Fuel sources: One or more of the powders act as solid fuel to the burning process. 2. An oxidizer or oxidizing agent: This crystalline component oxidizes the fuel components via combustion, which then causes the flame or explosion. 3. Binders or binding agents: These hold the fuel-oxidizer particle mixture together. Binders are often derived from resins or polymers that are cured to a semisolid or elastomeric form. During grain formulation, compounding, and curing, the fuel and oxidizer particles become suspended in the rubbery binder network (the matrix), which maintains uniformity and localization of the reactive constituents throughout the propellant body. An igniter motor is installed at the forward end or the dome region of the case assembly to initiate the combustion process at the beginning of the burn cycle. Archaic solid fuel mixtures have been around since the ninth century and were originally based on “black powder,” which is a mixture containing charcoal (carbon), sulfur, and potassium nitrate (KNO3), a very strong oxidizing agent. During the burn process of devices using these mixtures, KNO3 oxidizes the charcoal and the sulfur, both of which act as fuel. (The presence of sulfur also reduces the ignition temperature.) This is the basic fuel formulation that has always propelled classical consumer fireworks, such as sky rockets and bottle rockets. In these low-tier devices, the dry powder is often packed into the paper housing or “casing” without a binder. Basic gunpowder, developed in the 19th century, is also based on KNO3 in solid charcoal. Nitrocellulose (cellulose nitrate), also known as guncotton, can perform all three functions required for rudimentary solid-state propulsion. Because of its tacky nature, nitrated cellulose has been used exclusively as the binder in many pyrotechnic compositions. By itself, the organic cellulosic portion of this molecule makes a good fuel material, whereas the protruding nitro groups abundant along the cellulose backbone (beta-bonded glucose) act as strong oxidizing agents. Model toy rockets have often employed solid KNO3 mixtures in which sugar (such as sucrose, ordinary table sugar) is the principal fuel component. In these formulations, the sucrose melts and binds or fuses the mixture together to form the solid fuel pack. However, none of these low-tech or early history examples are good representatives of modern SRM compositions. Since the 1960s, large-scale SRM formulations, including the space shuttle’s RSRM boosters, have typically used specially compounded particle/powder-based composite propellant mixtures. These SRM formulations consist of an oxidizer along with combustible fuel components, which are blended and bound together in a rubbery/elastomeric-type polymer binder/plasticizer, such as a modified polybutadiene. For RSRM, the solid propellant formulation includes particles of

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ammonium perchlorate (NH4ClO4) and one or more particle size classes of aluminum metal combined within a binder/matrix of polybutadiene acrylonitrile (PBAN) copolymer (alternately referred to as polybutadiene–acrylic acid– acrylonitrile terpolymer). All perchlorates are strong oxidizing agents, but NH4ClO4 is an especially potent oxidizer. In contrast, elemental aluminum is a strong reducing agent, particularly in fine particle form when much of the particle surfaces are exposed and made available to the oxidizer. Thus, the NH4ClO4 acts as the oxidizer while the aluminum and PBAN binder become the fuel. The solid fuel burning process is a very powerful, high-temperature oxidation/reduction chemical reaction that is essentially a controlled explosion designed to last for several minutes. As the burn cycle in an SRM engine progresses, the propellant grains undergo deflagration and recession; in other words, the grains progressively burn down as their volumes recede. Thus, their initial sizes, shapes, and surface areas become important factors in the burning process and must be appropriately considered during the SRM design phases. In general, the amount of grain surface area exposed to the flame is proportional to the thrust of the SRM but is inversely proportional to the burn cycle time. Therefore, obtaining a very high thrust for long periods of time can become a challenge. Grain geometry, formulation, mixing, and casting are critical factors in the design and formulation of solid propellant SRM engines. Generally, once a solid composite fuel mixture is ignited, it cannot be regulated or turned off and on “at will.” This contrasts with liquid propellant engines (see Chapter 12), which can be designed to provide real-time throttling, shutdown, and reignition. However, newer solid motor designs are beginning to emerge that can reduce some of these limitations. In hybrid rocket platforms, the propellant reactants are in different physical states, such as liquid oxygen with solid polymer/aluminum mixtures. These engines can often be throttled, extinguished, and reignited by use of the appropriate control mechanisms. Also, more powerful solid propellants are now available, many derived from cyclic nitroamine oxidizers that provide substantially higher oxidizer-to-fuel ratios and energy densities. These compounds are of great interest to the military but have not yet been seriously pursued for use in space launch systems.

11.2.2

FUEL

Metals are used as fuel in high-energy propellant to enhance density and to increase the total heat release during combustion. Many propellants need nominal amounts of metal to suppress unstable combustion. Today, aluminum is the most commonly used fuel due to its relatively high combustion efficiency, high heat of combustion, low cost, and ready commercial availability. The amount of aluminum in formulation ranges from 2% to 21%. Beryllium is sometimes used in propellants because it has high heat of combustion and low molecular weight combustion products; however, the beryllium oxide in the exhaust

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product is toxic. Berylliosis, a condition caused by repeated inhalation of beryllium or beryllium compounds, is associated with reduced lung capacity similar to the effect of long-term exposure to asbestos. Thus, beryllium fuels can be used only for space-based applications such as space intercept weapons. Magnesium is being used in some modern “clean” formulations requiring reduced exhaust emissions. However, low-smoke formulations used in tactical propulsion systems use very little (if any) metallic powder. Table 11.1 summarizes typical fuels used in composite propellant.

11.2.3

OXIDIZERS

High oxygen content is a preferable aspect for an oxidizer and is also one measure of its overall performance. Certain physical properties of common oxidizers used in composite propellants are described in Table 11.2. Ammonium perchlorate (AP) is the most popular oxidizer in use today due to its moderate performance, fairly low cost, and good process ability. One drawback of this compound is that the chlorine in the oxidizer reacts with hydrogen to form hydrochloric acid (HCl) in the exhaust product. Water vapor can combine with the HCl cloud from the large solid motor to form acid rain in the area near the launch pad and can cause ozone depletion in the upper atmosphere. There is ongoing effort within the industry to develop new propellant formulas with less

TABLE 11.1 METAL FUELS USED IN COMPOSITE PROPELLANTS Fuel

Chemical Symbol

Molecular Mass, kg/kmol

Density, kg/m3

Remarks

Aluminum

Al

26.98

2700

Low cost, common fuel

Aluminum hydride

AlH3

30

1420

Difficult to make

Beryllium

Be

9.01

2300

Toxic exhaust products

Beryllium hydride

BeH2

11.03

650

Toxic exhaust products

Boron

B

10.81

2400

Inefficient combustion

Magnesium

Mg

24.32

1750

Clean propellant applications

Titanium

Ti

47.9

4500

Low performance but high r

Zirconium

Zr

91.22

6400

Low performance but high r

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TABLE 11.2 Compound

OXIDIZERS USED IN COMPOSITE PROPELLANTS

Chemical Formula

Density, kg/m3

Oxygen Content Mass %

Design Considerations

Ammonium nitrate (AN)

NH4NO3

1730

60

Low cost, moderate performance

Ammonium perchlorate (AP)

NH4ClO4

1950

59.5

Low cost, moderate performance

Nitronium perchlorate

NO2ClO4

2200

66

Very reactive, unstable

Potassium nitrate

KNO3

2110

47.5

Low cost and performance

Potassium perchlorate

KClO4

2520

46.2

Low regression rate, moderate performance

Sodium nitrate

NaNO3

2170

56.4

High cost, moderate performance

than 1% HCl in exhaust. Next to AP, ammonium nitrate (AN) has the most widespread use among the remaining oxidizers in the table. Even though AN is not as energetic as AP, it provides a clean exhaust (at least in terms of HCl) and is less expensive. The issues with AN include its characteristic low burn rate as well as susceptibility to a phase change near 308C. In today’s motors, the rest of the oxidizers in Table 11.2 are not currently used. The overall propellant properties depend greatly on the size of oxidizer particle. Processing cost is driven by the complexity of the specified particle size distribution in the propellant formulation. Some of the high-solid-loaded formulations of composite fuel binder propellant used a trimodal AP system. The term modal refers to the number of peaks (or modes) in a plot of the particle size distribution. In addition, the oxidizer shape characteristic can in turn influence the propellant burning rate and propellant process ability. Hence, the burning rate can be controlled in two ways: by varying the ratio of coarse to fine oxidizer particles or by varying the oxidizer particle size.

11.2.4

BINDER

The primary function of a binder is to hold the entire formulation in a structurally stable grain that can endure temperature variation as well as acceleration and pressure forces during flight. Binders that have both high combustion energy and low density are preferable. Moreover, the best binder materials provide

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satisfactory structural integrity while at the same time occupying a minimum volume. This property is referred to as solid loading, which is the ratio of the total mass of fuel plus oxidizer as a percentage of the total propellant mass. Solid loading of composite propellants ranges from 84%–90%, and only 10%– 16% of propellant mass corresponds to binder and other minor ingredients, such as curatives and modifiers (see Sec. 11.2.5). Table 11.3 summaries typical binder materials used in composite propellants for SRMs. Binders are normally long-chain polymers that can hold the crystal oxidizer and propellant powder in place by forming a continuous matrix through cross-linking and polymerization. Cross-linking takes place by incorporating a small amount of curative into the propellant just before casting into the motor. To further promote cross-linking reaction, the casting grain is cured at a high temperature for a prescribed amount of time. The required cure temperature and cure time depend on the selection of binder material. Referring to Table 11.3, the most common binder used today is hydroxyterminated polybutadiene (HTPB). CTPB, PBAN, and especially HTPB are used in many modern technologies, whereas PS, PEPU, and PBAA are used for older applications. PBAN commonly is used for service conditions from 08F to 1208F and fairly high elongations approximately on the order of 30% true elongation. CTPB is used for desired performance at temperatures from 658F to 1508F. The new binder GAP has not yet been used in space application even though designers have proposed it for tactical rockets.

11.2.5

MINOR INGREDIENTS

For the polymerization process to take place, a small amount of curative is required in the propellant formulation. A cure catalyst is regularly included TABLE 11.3

TYPICAL BINDERS USED IN COMPOSITE PROPELLANT

Binder Designation

SRM Applications

Hydroxy-terminated polybutadiene (HTPB)

IUS, Peacekeeper, Star 48

Polybutadiene acrylonitrile (PBAN)

Titan and shuttle SRMs

Polybutadiene acrylic acid (PBAA)

Older rockets

Nitrocellulose (plasticized) (PNC)

Minuteman and Polaris

Nitrate ester polyether (NEPE)

Peacekeeper, SICBM

Carboxy-terminated polybutadiene (CTPB)

Minuteman Stg. 2,3

Polyether polyurethane (PEPU)

Polaris Stg. 1

Polysulfide (PS)

Older rockets

Glycidal azide polymer (GAP)

New binder

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with the curative to increase the rate of chemical reaction, which is initiated by adding this constituent. A cure reaction must promote extensive cross-linking in the final binder to give satisfactory mechanical properties and to avoid high temperature flow. The selection of the cure system depends on the chemically active groups on the polymer molecules. The carboxyl group (ZCOOH), hydroxyl group (ZOH), and mercaptyl group (ZSH) are the common chemical groups in polymers. In the PBAN case, a cross-linking agent such as epoxy is preferred, whereas imines such as tris[1-(methyl) aziridinyl] phosphine oxide or MAPO are the cure agent for CTPB. Plasticizers are also used in certain propellants to improve physical properties at low temperatures. A darkening agent (such as carbon black) is required in many double-base formulations to make the translucent propellant darker and avoid excessive thermal radiation through the propellant itself. A small number of modifiers or a burn rate catalyst is sometimes needed in the formulation to achieve the desired burn rate in the propellant. Ferric oxide is broadly used as the burning rate catalyst in composite propellants. Copper chromite (actually a mixture of cupric oxide and chromic oxide) is also used extensively even though it is incompatible with some polybutadiene binders and with double-base binders. These additives can be very important because variations among different lots of oxidizer, fuel, and binder materials can lead to significant changes in the burn rate. The amount of catalyst required to achieve the desired burning rate in the motor is normally determined by doing subscale tests on the new batch of propellant. To avoid binder oxidation reactions in propellant over long storage periods, antioxidant is often required in the formulation. Finally, a bonding agent is used in a few propellants to improve the bond between the binder and oxidizer. To select the right propellant type for a specific mission, the designer needs to consider the required performance, internal ballistic, flame temperature, burning rate, specific impulse, mass flow, thrust, and mechanical properties, as well as the necessary storage stability. The response to hazards and any other specialized properties (e.g., smokeless, etc.) also must be considered. Engineers regularly use test data to predict a motor’s expected performance properties rather than modeling only. Propellant is processed in three stages: pre-mixing, final mixing, and casting. During propellant processing, personal safety is a main concern, therefore equipment and facilities amenable to remote operation are mainly employed. Mixers used for manufacturing solid propellant are normally either vertical planetary mixers or Sigma-blade mixers, commonly used to make large batches of stiff bread dough. Figure 11.2 shows a vertical mixer used for 600-gallon batches. After mixing is complete, the propellant is then ready to cast into segment of motor. The cast pit for a reusable solid rocket motor forward segment is show in Fig. 11.3. Most composite propellant motor are cured at elevated temperature. It is preferable to keep the temperature as nearly constant as possible throughout mixing, casting and curing.

MATERIALS FOR SOLID ROCKET ENGINES

Fig. 11.2

Mixing propellant. (Courtesy of NASA/ATK Thiokol.)

Fig. 11.3 Casting pit for a reusable solid rocket motor forward segment. (Courtesy of NASA/ATK Thiokol.)

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11.3 MOTOR CASES Motor cases for space and missile applications have been fabricated from both metals and nonmetals. Steel and composite casings have become common. The RSRM motor case assembly is composed of D-6AC steel alloy that is a half inch thick and uses a total of 11 individual case sections. Several of the smaller sections are joined into larger segments at the factory before installation of the insulation material, liner, and propellant grain into each of the segment assemblies. The dome is joined to the forward segment at the factory. Then, at the launch site (in the field), the segment assemblies are joined together, the igniter is installed to the forward segment, and the nozzle is installed to the aft segment to make the final SRM structure. For the space shuttle’s RSRM casing, there are a total of seven factory joints and three field joints. Thus, the final RSRM case assembly consists of the forward segment, the forward-center segment, the aftcenter segment, and the aft segment. Additional motor joints include the igniter-to-case, the nozzle-to-case, and the nozzle’s “Joint 1” (forward exit cone to aft exit cone). The RSRM configuration is referred to as a four-segment motor. Historically, this configuration has effectively accommodated the manufacturing and shipping requirements associated with the production, logistics, and assembly of the RSRM program. In comparison, the RSRMV configuration has been designed as a five-segment motor utilizing four case field joints and consisting of the forward segment (with dome), the forward-center segment, the center-center segment, the aft-center segment, the aft segment, and the aft exit one. The use of composite materials, typically carbon fiber-reinforced epoxy, has entered many areas that were traditionally metallic. This has expanded to sporting goods and aerospace components including large commercial passenger jets. Composite materials have the freedom to customize performance based on the mechanical properties, ply orientations, and number of layers. The application of carbon fibers to motor cases is very effective due to the superior tensile strength of carbon and the fact that a motor case is functioning essentially as a pressure vessel. Pressure vessel performance is dominated by the tensile strength of the carbon fibers. Filament-wound composite motor cases are used worldwide for aerospace and defense applications. NASA had a large-scale composite motor program for the space shuttle in the mid-1980s. This carbon fiber-based motor built on the developments from the Kevlarw-based Peacekeeper missile. The composite RSRM for NASA was being developed by Hercules and the Thiokol Corporation; each had independent development efforts. The composite motor case was going to save more than 60,000 lbs mass compared with the metallic case, and this would have allowed approximately an additional 8,000 lbs payload capability for polar orbits. The composite RSRM was developed from the materials available at the time. The carbon fiber was Hercules AS-4, which is considered mediocre compared with materials currently available. Today there are several highperformance carbon fibers that have a higher tensile strength and higher

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modulus, which will result in even more mass savings. There have also been advances in resin chemistry that will allow the resin matrix to have a high glass transition temperature (Tg). The higher Tg resins allow the carbon fiber to translate properties better and reduce the amount of insulation in the case, thus saving additional mass. Composite motor cases have now become an industry standard for commercial launch vehicles and tactical missiles. The proper selection of fibers and resins allows the user to customize the design and fine-tune performance. The fiber properties and their ply angles allow the designer to adapt the case to specific needs, including changes in the operational pressure of the motor, alterations in the reactivity of the propellant, adaptations of the operational environments, and so on.

11.4 INSULATION MATERIALS The insulation materials (see Figs. 11.4–11.6) are specially formulated rubber materials often based on elastomeric polymers derived from nitrile butadiene rubber (NBR). Legacy RSRM insulation rubbers have used white asbestos fibers embedded in NBR matrix/binders that are cured in place inside the casing. These materials have been abbreviated as ASNBR. Binder rubbers derived from ethylene-propylene dimer monomer (EPDM) have also been used in low-cost rocket/missile insulators formulated with carbon or silica fibers. To address concerns associated with the use of asbestos, alternative formulations have incorporated the use of organic fibers spun from liquid crystal polymers such as Kevlarw (polyaramid) and polybenzimidazole (PBI). Specially formulated insulators have been extensively evaluated for use on the RSRMV motor. During installation and curing of NBR-based insulators, cross-linking reactions take place within the rubber, which vulcanizes and bonds the material onto the inside of the steel case segments. During the burn cycle, this layer thermally insulates the casing from the heat of propellant combustion, which burns at about 60008F (note that steel melts at about 25008F). The rubber insulator is applied at varying thickness throughout the inner SRB structure to ensure that minimum thermal safety factors are met (Fig. 11.5). Insulator thickness can vary from fractions of an inch in cooler regions up to several inches in the dome region.

Fig. 11.4

Schematic layout of RSRM components.

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a)

Solid rocket motor

b)

Field joint area

Fig. 11.5 Solid rocket motor configuration showing field joint area. (Courtesy of NASA/ATK Thiokol.) Two primary rubber matrices are used for motor insulation: EPDM and NBR. Both of these rubber materials are blended with a list of constituent materials to form a composite material that is flexible, soft, and tacky in its green state. This allows the material to be formed and molded to specific and complex geometric configurations needed (Fig. 11.6) for proper design performance and optimization of mechanical performance of the motor. A primary choice of constituent material in these composite rubber materials is that of fiber reinforcement. The fiber reinforcement is critical to the mechanical performance of the insulator and affects the thermal insulation and erosion performance. Optimizing the fiber choice for performance (discussed in the performance section) requires finding the right balance between tensile strength, elasticity, and erosion resistance. Five primary fiber choices have been extensively tested and used on full-scale SRMs. They are 1) asbestos fiber, 2) silica fiber, 3) carbon fiber, 4) Kevlarw fiber, and 5) PBI (polybenzimidazole) fiber. The insulation composite materials are identified with acronyms that identify the base rubber material and its fiber constituent. The insulations that have been tested by full-scale SRMs are identified as follows: 1) ASNBR (asbestos fiber NBR), 2) SFEPDM (silica-filled EPDM), 3) CFEPDM (carbon fiber EPDM), 4) KFEPDM (Kevlarw fiber EPDM), and 5) PBINBR (PBI fiber-filled NBR). SFEPDM is primarily used as thermal protection material around the booster and has not been used in large-scale SRM insulation applications due to its

Fig. 11.6 Insulation at the reusable solid rocket motor joint. (Courtesy of NASA/ATK Thiokol.)

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poor erosion performance capability. ASNBR was first tested and produced in the 1960s and 1970s for the initial SRM programs. As a result of discoveries related to the asbestos fiber effects on human health, the Occupational Safety and Health Administration has mandated stringent processing safety requirements that are costly to processing. In addition, there are concerns over the environmental effects of burning SRMs with asbestos fibers. To find a replacement for ASNBR, the Kevlarw and PBI fiber-based insulation development began in the late 1990s. The performance differences of these materials will also be discussed in the performance section of this chapter. Asbestos is a natural inert and inorganic crystalline material that has a high melting point. Asbestos in fiber form increases strength and absorbs heat energy as its temperature is increased and approaches its melting point. Kevlarw and PBI fibers are synthetic aromatic fibers that can be fabricated to precise dimensional requirements. They are both organic and have been successfully tested in rubber matrices for thermal insulation. However, the organic nature of the two fibers has produced processing differences with ASNBR.

11.4.1

PROCESSING

The insulation material is compounded at the supplier and provided to the motor manufacturer for installation and processing. There are essentially four stages to processing: 1) installation, 2) devolatizing and debulking, 3) cure, and 4) testing. 11.4.1.1

INSTALLATION

Depending on the substrate of the motor case, the materials used for installation and bonding to the case can be different. However, the basic process remains the same. The case is either composite or metal. In the case of space shuttle RSRMs, the case is steel. The process for installation of the initial rubber sheet to the case wall is to prepare the case with a coat of primer. There are industry standard primers for bonding NBR rubber materials to steel substrates. To ensure a clean surface, the NBR material is wiped with an approved solvent. The solvent may also provide a tackier surface for the green rubber during installation of the sheets to hold in place until cure. 11.4.1.2

DEVOLATIZING AND DEBULKING

The green rubber sheets remain in place due to their own tack and tackifying agents until cure. Before cure at raised temperature, the segment and material are subjected to slightly raised temperatures and vacuum to remove volatiles (devolatize) that may be left behind from solvent application and to remove air pockets (debulk). The need for this process and its pressure, time, and temperature are wholly dependent on the size of the motor and the materials chosen for case, bonding, liner, and propellant.

622

11.4.1.3

D. V. TRINH

CURE

The vulcanized rubber requires heat cure to cause the sulfur within the composite to react with the polymer chain and create a cured solidified material. Each rubber composite has an optimum cure profile for temperature and time to achieve the material flow (reduces internal voids) and then convert to vulcanized rubber. This cure also activates the primer for the insulation bonding to the case as well as rubber to rubber sheet bonding. 11.4.1.4

TESTING

Before processing and installation, all lots of green rubber material undergo receiving inspection and minimal chemistry screening testing to verify the material is consistent with past material delivered. Any significant deviation is identified and addressed before use. The insulation has several postcure test samples that are run in parallel with the segment to verify the integrity of the specific lot/batch used on the motor and to verify that the processes have produced an insulated segment meeting the design requirements. The testing is composed of 1) small tensile specimens to verify mechanical properties are within processing limits and design minimum/maximum requirements, 2) shore hardness specimens, and 3) small panel testing to verify cure and bonding of the insulation to itself and the substrate.

11.4.2

PERFORMANCE PARAMETERS

There two primary performance parameters for the insulation. First, it must have the mechanical properties required to prevent structural failure during fabrication, processing, and transportation. Second, it must perform thermally to protect the case hardware and appropriate sealing zones around case joints. 11.4.2.1

MECHANICAL PERFORMANCE

The simplest property of performance is mechanical. Numerous mechanical tests are performed at various environmental conditions to determine the tensile and strain capability. These are simple tests and performed on numerous batches and lots to validate performance. The analysis of the data provides a statistical population limit, called the A basis, which is used in the structural analysis of the motor design. The definition of A-basis values resides in MIL-HBK-17 and NASA-STD-6016. The A-basis requirement is levied upon all materials for qualification in aerospace applications. For new materials in development such as PBINBR, generation of A-basis mechanical properties is typically required. (At least 99% of the population of values is expected to equal or exceed the A-basis mechanical property allowable, with 95% confidence.) The A-basis values are then used in finite element analysis (FEA) models that produce structural safety

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Fig. 11.7 Finite element analysis model for reusable solid rocket motor forward segment. (Courtesy of NASA/ATK Thiokol.) factors for all loading conditions specific to the geometry of the design. Figure 11.7 shows an example of an FEA model identifying load values and points throughout the motor: propellant, liner, and insulation. This type of analysis is required to certify the motor design for use in its intended environments. 11.4.2.2

THERMAL PERFORMANCE

Thermal performance analysis requires more complex and sophisticated tools. During motor firing, the insulation strength is no longer as critical due to the compression pressures within the motor. However, the case bond line temperature cannot exceed 2008F at any point during the motor firing. This is to ensure that the insulation remains in place until the motor burn is complete to protect the structural integrity of the case for reuse. Because the insulation becomes exposed to high temperatures throughout the motor at different times during the burn, and because the insulation erodes away as it is exposed to high-heat loading environments, the model for predicting the performance and verifying the temperature limits are not exceeded is complex. Figure 11.8 depicts the model and the variables used for predicting thermal performance. Basic principles for modeling the material decomposition include convective and radiative heating that causes material decomposition and charring. The resulting char layer provides protective thermal resistance. Figure 11.9 is a graphical representation of surface erosion. The char surface recedes due to both chemical erosion (reactions between the combustion gases, pyrolysis gases, and char) and mechanical erosion (char failure due to mechanical loads, such as flow shear stress). Calculation of the erosion rate is critical for predicting the material decomposition rate. Calculation of the material decomposition rate is required to verify that the bondline and motor case do not exceed the temperature limit at any point during motor burn. Analytical model inputs are required to predict ablation and insulation through the thickness isotherms during motor burn. These include 1) thermal boundary conditions (radiation and convection heat transfer), 2) chemical boundary conditions (elemental composition at boundary layer edge and diffusion through the boundary layer), 3) char recession (both chemical and mechanical), 4) energy balance (both surface and through the thickness), and 5) material decomposition performance in environments.

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Fig. 11.8

Model and variables used for predicting thermal performance.

The required material properties for use in thermal performance modeling can be grouped into the four categories listed in Table 11.4. Erosion performance is a critical parameter in insulation design and materials selection due to the extreme operational environments of the motor. Research continues in refining predictions and determining material variations due to fiber choice (organic vs inorganic) and environment within the motor (case wall vs aft dome). The char formation and ablation as well as chemical erosion

Fig. 11.9

Graphical representation of surface erosion.

MATERIALS FOR SOLID ROCKET ENGINES

TABLE 11.4

625

MATERIAL PROPERTIES USED IN MODELING THERMAL PERFORMANCE OF INSULATION

Combustion Gas Properties

Material Thermal/Chemical Propertiesa

Density

Enthalpy vs temperature

Viscosity

Specific heats vs temperature

Specific heat

Thermal conductivity vs temperature

Thermal conductivity

Species mass fraction

Species concentration Recovery enthalpy Emissivity Material Decomposition Properties

Pyrolysis Gas Properties

Virgin density

Enthalpy vs temperature

Char density

Species concentrations

Activation energy Arrhenius preexponential factor Reaction order a

These material properties should be quantified and characterized in both the virgin material and the char material at both ambient and operational conditions.

due to reaction with propellant products during motor burn as a function of fiber choice and fiber material thermal properties is not fully understood. In general, all insulators perform well thermally along the case wall, but choosing an insulator that thermally performs in the aft dome environment is more challenging. This is due to the propellant and slag impingement directed off the nozzle into the aft dome (shown in Fig. 11.10 to the right) and recirculation of the products around the aft dome circumference and axially. Currently, the best practices are limited to using empirical data collected from full-scale motor testing for validation of the thermal model predictions and assigning statistical uncertainty at the system performance level. Several subscale tests are in development at NASA Marshall Space Flight Center to better segregate material variability from environment/design variability. In addition, there is a need to develop the criteria for development of thermal performance properties for full rigorous qualification (A-basis values) similar to mechanical properties. This development would necessarily require definition of thermal performance properties and test the definition to generate the property values for population into NASA’s online Materials and Processes Technical Information System (MAPTIS).

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Fig. 11.10

Schematic of nozzle and aft case. (Courtesy of NASA/ATK Thiokol.)

11.5 LINER AND INHIBITOR The liner layer is installed on top of the cured insulation layer (see Fig. 11.4), and its purpose is to facilitate binding interactions between the propellant grain and the primary insulation layer. After wet liner installation, the liquid rubber material is cured and then the propellant grain is cast and cured on top on of the liner. Liner compositions may bear substantial similarities to underlying insulator formulations, as they are also often derived from polybutadiene-type rubber binders analogous to acrylonitrile/butadiene copolymers (NBR). Both RSRM and RSRMV rubber liners are based on carboxyl-terminated polybutadienes (CTPB) in a mixture formulated with amine (or imine) curatives/crosslinkers that interact with the carboxyl groups to form a soft, rubbery matrix. Typical liners are applied via spraying or roll application; thus, they are relatively thin. Finally, propellant slurry is applied to the cured liner surfaces followed by additional curing of the layered assembly at temperatures similar to those used to cure the liner. To a certain extent, the insulator, liner, and propellant are co-cured together to form the packed segment. The liner material adheres very well to both the rubber insulator and the propellant phase, as all three phases contain compatible binders. Without the liner layer, adhesion of propellant to the rubber insulator would be less than desirable.

11.5.1

MATERIALS AND PROCESSES

The primary ingredient (70%–90% by weight) in SRM liners is the base organic polymer. Two curatives are added to fuse the liner between the propellant and the insulation. The second largest ingredient is the fiber filler. This performs the same function as in the insulation in that it changes the mechanical properties

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for strength and elasticity. The type of fiber also affects the mechanical properties. Just as in the insulation, the liner for space shuttle solid motors used asbestos fibers. Development of liners with organic fibers is ongoing. Catalysts may be added to control the cure time for processing efficiencies. Finally, a thickening agent may also be added based on the application methods, design geometry, and facility configuration. This thickening agent enables the liner to stay in place during processing until final cure with the propellant. In general, the liner also undergoes an application process, curing process, and testing process. The liner is applied either by hand or through a sling lining operation that closely resembles a spray method. The method of application is based on the geometry of the design area. The more complex the geometry, the more likely the application will be by hand. The cure process is in two phases, the precure and the co-cure. The precure enables the liner to be moved to the casting facility for the propellant. The liner is not fully cured because the process is intended to cure the propellant to the insulation. During precure the temperature is raised to between 1008 and 1508F just to begin the polymerization process. The segments are then sent to the propellant facility to have propellant cast, and the full co-cure of the liner and propellant is completed. The testing of the cured liner material is in two parts: liner and bonding systems. The liner testing consists of simple tensile specimens to verify mechanical properties and a shore A hardness specimen that simply verifies that the material is fully cured. The system bonding specimens ensure that the liner material and process produced a proper cure to the insulation, the propellant, or the case substrate (in the configurations where little or no insulation is present in the design).

11.5.2

PERFORMANCE PARAMETERS

The performance requirements for liners are simple and straightforward. The material has to be able to bond the propellant to the insulation or case, must be sufficiently strong to hold the weight of the propellant, and must not be so stiff as to cause a fracture/failure within the liner that could go undetected. The primary performance requirements are the ability to process the material to properly cure (verified in the process control testing described in the previous section) and adequate mechanical properties such as tensile strength and elongation. These are generally driven by statistical methods of A-basis derivation for analysis as described in the insulation section of this chapter.

11.6 NOZZLE The nozzle section or aft exit cone converts thermal energy generated during the combustion of the propellant into kinetic energy by expansion of the combustion gases. The ratio of throat area to exit plane area is critical in the design of rockets because it must be optimized to achieve overall system performance.

628

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For reference, the RSRM nozzle is composed of seven major subassemblies with four dual O-ring joints, and the ability to carry pressure loading is provided by steel and aluminum housings to which ablative liners are bonded. These liner structures provide thermal protection for the metal housings and are fabricated from autoclave cured, tape-wrapped, carbon cloth phenolic (CCP) matrix composites. The principal flame surface articles are composed of CCP, including the throat and the inner aft exit cone skirt. Supporting nozzle structures include lower temperature composites derived from silica cloth phenolic and glass cloth phenolic configurations. An RSRM nozzle weighs about 24,000 lbs and can vector up to 88 in any direction, providing directional thrust capability. RSRM nozzle throat and nozzle joints are is shown in Fig. 11.11.

11.6.1

FUNCTION

The nozzle of an SRM serves to regulate the internal pressure of the motor and direct the flow of combustion products out of the motor. Both functions are critical to achieving the desired motor thrust and payload capability. The nozzle will typically have a converging-diverging aerodynamic contour on the flame surface and may be fixed in position/orientation or movable to allow thrust vector control. Additionally, the nozzle may be submerged, with a significant portion of the nozzle structure inside the motor, to reduce the overall length and inert mass of the rocket. A submerged nozzle will reduce the mass of propellant that can be held in the motor. As combustion gases from the burning motor propellant flow into a nonsubmerged nozzle, they first traverse the inlet region of the nozzle contour. The inlet is a converging geometry that accelerates the subsonic gas flow. The transition from the converging portion of the nozzle aerodynamic flame surface to the diverging portion is the geometric throat. Near to the geometric throat, depending on thermodynamic and compressibility factors, the velocity of gas flowing through the nozzle reaches the speed of sound. The diverging portion of the nozzle is the exit cone. For an optimized design condition, the aerodynamic contour of the exit cone is neither overexpanded nor underexpanded, and the

Fig. 11.11

RSRM nozzle throat and nozzle joints. (Courtesy of NASA/ATK Thiokol.)

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combustion gases accelerate through the exit cone to leave the aft end, or exit plane, with supersonic velocities. An optimized design condition will feature a specific curved (often parabolic) exit cone contour derived to prevent flow separation or other disturbances. An arbitrary parabolic contour will not provide an optimized design condition. If an optimized design condition is not necessary, an overexpanded or underexpanded straight cone (frustum) may suffice. A straight cone may typically be manufactured with less effort, time, and cost. In general terms, if the nozzle is submerged, a plane tangent to the forwardmost portion of the nozzle flame surface, called the nose tip, will divide the incoming flow of combustion gas into two streams. Flow outboard of the nose tip will pass along the outer surfaces of the submerged portion of the nozzle and remain inside the motor. Flow inboard of the nose tip will pass along similar surfaces to a nonsubmerged nozzle. Because the flow conditions vary with location from the nose tip through the throat and onward to the exit plane, the temperatures and pressures on the nozzle surface also vary with location. Combustion gas temperatures (typically 5,0008F) and pressures (typically 1,000 psig or more) are sufficient to generate ionized plasma. The combustion products are chemically reactive with the nozzle surface and may include significant amounts of molten slag particles from the propellant, which may impact the nozzle surface. The internal environment of the SRM nozzle is so severe that it may degrade or actually remove material from the nozzle surface at a significant rate. In highperformance designs with sufficiently long burn durations, this material removal, called erosion, may result in considerable macroscopic changes in the nozzle flame surface contour. Because the cross-sectional area of the throat acts to regulate the internal pressure of the motor and higher pressures provide more thrust and payload capability, the functional performance of the rocket motor can significantly decline during motor operation if there is significant erosion of the throat.

11.6.2

MATERIALS AND PROCESSES

Because of the severe internal environment, one common design solution is to form the nozzle interior flame surface from an insulating liner, with a higher strength structure or housing on the exterior to carry the loads and pressures. Typically, the liner will be fabricated from a single nonmetallic material or layers of multiple nonmetallic materials. The housing is most often metallic, although nonmetallic materials (such as fiber-reinforced polymer matrix composites) may be an option for some designs. An ablative nozzle liner absorbs some of the combustion gas thermal energy via convection and radiation, dissipates that energy in the pyrolysis (thermochemical degradation) of a thin layer of the nozzle liner material, and injects cooler gaseous pyrolysis products into the boundary layer above the nozzle liner surface. This boundary layer cooling improves the insulating function of

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the nozzle liner. The best ablative nozzle liner materials will naturally form an open-cell porous structure in the degrading, or heat-affected, layer of material nearest to the flame surface. The porosity reduces the bulk conductivity of the material, and an open cell structure provides transport pathways for the coolest pyrolysis gases generated deepest in the material, thus improving the boundary layer cooling efficiency. Phenolic resin is a common choice for polymer matrix composite nozzle liner applications due to its excellent char generation performance. Unlike many other thermoset polymers, phenolic typically achieves its final material properties long before reaching a uniform degree of polymerization. Thus, the polymer molecules in a service part have a broad random distribution of molecular weights. Consider the matrix material at an arbitrary depth near the flame surface of the nozzle liner as a simplified but illustrative example. As the heat flux on the nozzle surface increases the internal temperature at this location, the phenolic first undergoes an endothermic postcure thermochemical condensation reaction, which dissipates some of the incident heat. With continuing heat flux, the material begins to pyrolyze. Side groups begin to break off from the polymer chains and new secondary reactions can occur, dissipating more of the incident thermal energy. The material density decreases as liquid and gaseous reaction products are generated, and, with continuing heat flux and rising temperature, porosity increases and the material softens. The presence of porosity inhibits the heat flux through the material, thereby insulating the cooler material below. When the temperature of the material at this depth rises far enough, noncarbon species are eventually eliminated from the long, twisted carbon chains that were once the back-bone of phenolic polymer molecules. Sufficient thermal energy exists for those carbon chains to begin reorganizing; the matrix material becomes intrinsically more dense and rigid. CCP composite ablative insulators employ carbon fibers in a woven form to improve retention of the charred phenolic matrix material near the nozzle liner flame surface. The most historically prevalent carbon fiber used in CCP for nozzle liners is carbonized from rayon. For many applications, the carbonization process may produce fibers with adequate mechanical properties while retaining a primarily amorphous internal structure, which enhances the insulating performance of the nozzle liner. In other applications, it may be desirable to perform a more extensive, higher-temperature carbonization process to obtain fibers with more ordered/crystalline internal structure, accepting some increase in thermal conductivity for the sake of higher strength. Tape wrapping is a typical manufacturing process used to produce nozzle liner components from CCP. Glass cloth phenolic (GCP) composite ablative insulators employ fiberglass in a woven form to improve retention of the charred phenolic matrix material near the nozzle liner flame surface. GCP has a significantly lower temperature capability than CCP and may be a practical design selection only for short-duration motors with low performance requirements. However, GCP may be an excellent choice for a co-cured backup insulator behind a CCP nozzle liner.

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Silica cloth phenolic (SCP) composite ablative insulators employ more expensive, higher-purity glass fibers in a woven form to improve retention of the charred phenolic matrix material near the nozzle liner flame surface. SCP has higher temperature capability than GCP and may be a better design solution for nozzle liner applications. However, it has a much lower temperature capability than CCP and exhibits an erosive sensitivity to high-speed impingement of alumina/slag particles within the motor plume. For small motors such as the booster separation motor (BSM) in NASA’s Space Shuttle program, graphite has been used as the throat insert material. The BSM, originally manufactured by Pratt and Whitney’s Chemical Systems Division (CSD), is a fairly simple SRM that aids in separation of the two SRBs from the shuttle. There are clusters of four motors at the top and bottom of each booster motor. The BSMs fire for approximately 0.8 s to provide the necessary impulse to separate the boosters safely. The basic throat design involved a one-piece ATJ Graphite throat with shallow conical interface on its outside diameter retained by aluminum housing. The nozzle has a fixed angle to the motor centerline that is provided by asymmetric closure geometry. There is no vectoring. The forward motors are significant in that they fire across the front of the orbiter with a potential for creation of debris in a critical area. This factor led to a major redesign after the BSM contract was awarded to Alliant Techsystems (ATK) in Promontory, Utah. Between the CSD and ATK manufactured BSMs, more than one thousand successful firings were performed during the shuttle program. The graphite formulation used for the BSM program was originally manufactured by UCAR (Union Carbide) as a hot-press-molded material manufactured at the company’s Niagara Falls facility. In the late 1980s the transition was made to an isostatic molding process at the UCAR (now GrafTech) Clarksburg, West Virginia, facility. Once completed, it became Isostatic Molded ATJ Graphite, or more commonly, ATJ. The primary constituent materials are fine-grain petroleum coke and coal tar pitch. The process includes a molding of the coke particles into bulk geometry and a multiple-cycle heat treat with coal tar pitch impregnation, with the final process being a graphitization cycle. The ATJ cycle is a specific set of proprietary raw materials and processes unique to this material. GrafTech also makes other formulations with differing designations as well as formulations based on different coke particle size and/or different impregnation cycles to produce properties for other industrial applications. In 2010, Graf Tech elected to discontinue the ATJ grade of graphite and is currently sampling a replacement grade to the aerospace community.

11.6.3

HISTORICAL ATJ USAGE IN AEROSPACE INDUSTRY

The aerospace usage of ATJ represented a very small percentage of the market for ATJ and as such presented specification issues. In general, it has been a consistent product that has performed very well over many years and in thousands of SRMs.

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These have included the NASA space shuttle BSMs as well as other separation motors, numerous tactical motor applications, and a few larger motors such as the Minuteman III third stage. There are other variations of bulk graphite motors within the industry, some of which have also been qualified for SRM use. Most of the SRM applications have used ring shapes well under a foot in diameter with moderate cross sections, not too thin and not too thick, based on experience developed over years of test and nozzle design. Most of the applications are throats for relatively short-duration motors, from under 1 s to a 10 s burn time. A significant exception to this is the Minuteman III third-stage nozzle which uses larger diameter parts at longer firing durations for backup to its Tungsten throat. The isostatically molded product is available as manufactured in cylindrical logs as well as square and rectangular cross sections with side dimensions up to 25 in. and lengths up to 84 in. This compares to more typical limitations of geometry to a little over 1 ft in the hot-press-molded ATJ. The early isostatic material was based on a material and process formulation originally referred to as TS-1792; this material was manufactured until raw materials issues forced a change in the early 2000s. A revised materials and process formulation referred to as TS-5245 was used after the discontinuation of TS-1792. All of these were formulated to have similar properties and have been referred to, at least in the bulk graphite industry, as ATJ or isostatic-molded ATJ. By its nature, the graphite grains have a highly oriented planar structure with properties in the across-grain orientation (“c” direction) being dramatically different (softer, higher thermal expansion, weaker, less thermal conductivity, etc.) than those in the with-grain (“a–b”) orientation. As the billets are fabricated, these grains are oriented preferentially in a planer isotropic structure and aligned in the with-grain plane. The development of the isostatic material process included an effort to bring properties in the two orientations closer to an isotropic state, but differences still exist. Larger billets, particularly those with square and rectangular sections, can present additional challenges in maintaining consistent properties without significant gradients both within plane and axially.

11.6.4

DATA SETS

Significant sets of data exist for several populations of ATJ, but each has limitations to its usefulness and applicability. Property sets include work done by Southern Research Institute (SRI) in Birmingham, Alabama, on the ATJ-S material for reentry vehicle applications. T1792 and TS-5245 formulations were evaluated and characterized by SRI for the BSM program. Also, these two materials were evaluated for the U.S. Army Tactical Missile System (ATacMS). Characterization and acceptance testing of the TS-5245 by SRI for the BSM program represented the first BSM procured and sampled from three subsequent production lots. There were two sets of data for the same TS-5245 population: one from SRI for BSM and one from GrafTech that was completed in 2007. All data sets provide insight into the property trends of the material. Some data sets track

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633

better through time, and some have larger numbers of data points but are limited to a single time frame. The most extensive data sets have tended to concentrate on inter- and intra-billet variability rather than variability with time. The majority of the BSM work is based on across-grain (c) tension at room temperature. Historically, this tends to be the most critical parameter with respect to the original BSM design as well as the most discriminating property with respect to both strength and positional gradients. Additional work includes with-grain (a–b) tension; torsional shear; ring tension; compression, including limited elevated temperature work; thermal expansion; thermal conductivity; and fracture studies.

11.7 IGNITERS The igniter is a small rocket motor that ignites the propellant grain. The RSRM igniter motor (shown in Fig. 11.12) is only about 48 in. long and 17 in. wide. However, its construction is similar to that of the main motor. The igniter is composed of a casing, an insulation layer, a liner, and propellant grain. The insulation and liner layers may be similar in composition to those used in the primary SRB motor. The igniter also includes a safe and arm (S&A) device mounted to the forward end. This device ensures that the motor fires only when appropriate command is given and provides the first ignition pulse via a pyrotechnic charge. The igniter and S&A device also contain components composed of silica/phenolic and carbon phenolic composites to form the nozzle and throat inserts.

Fig. 11.12

Components of a pyrogen igniter. (Courtesy of NASA/ATK Thiokol.)

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The function of the igniter is to induce the controlled combustion of solid propellant fuel at high pressures and temperatures in a controlled and predictable manner and at a stipulated rate. Many types and variations of solid rocket igniter have been used successfully to ignite solid propellant rocket motors. Virtually all igniters used in SRMs are of the pyrogen variety. The primary components of most solid rocket igniters are an initiation system; an energy release system; and the containment hardware, which also provides for attachment to and/or mounting within the rocket motor. In pyrogen igniters, the igniter itself operates as small SRM to provide hot gases and high pressures for igniting the main propellant grain. The propellant grain used in a pyrogen igniter is similar to propellant grain in SRMs. Smaller SRMs, such as BSMs used in RSRMs, simply use BKNO3 pellets to provide adequate energy to start combustion. An electrical signal ignites the pellets pyrotechnic and then propagates the ignition from the initiator to the pyrogen propellant grain. Reaction products from the pyrogen grain are expelled through the nozzle and impinge on the surface of the motor propellant to begin combustion. In the design of a pyrogen igniter, the main objectives are to obtain the necessary energy output while keeping the igniter as small and lightweight as possible. The igniter housing and nozzle must be able to withstand the internal pressure and intense heat while the igniter is in operation. Furthermore, if retainment of the hardware is desired, the igniter hardware must also be well insulated to endure the chamber environment during the entire SRM burning. Normally, igniter materials, such as insulation, casing, propellant, and liner, are similar to the material used in large motors.

11.8 HYBRID ROCKET PROPULSION SYSTEMS Hybrid rockets combine two different forms of propellants: solid and liquid. The oxidizer is a liquid and the fuel is a solid for the common hybrid rocket. The combustion within a hybrid rocket differs from that of either liquid or solid rockets due to the separation of fuel and oxidizer into two different physical states. The oxidizer-to-fuel ratio (O/F) varies down the length of the fuel port; therefore, the hybrid burns as a macroscopic diffusion flame. Hybrid rocket motors generally have the ability to be throttled, shut down, or restarted. Other advantages include safe operation, propellant versatility, environmental cleanliness, grain robustness, low sensitivity to temperature, and low cost. Like any other propulsion system, hybrids have some drawbacks, such as limited combustion efficiency, low regression rate, and low bulk density. The major considerations for selection of hybrid propellants are thermodynamic properties, regression rate behavior, and O/F shift during a burn. Carbon-based polymers in the form of plastic or rubber are typical fuels for hybrid rockets, including polymethyl-methacrylate (PMM), polyethylene (PE), and polybutadiene (PB). The terminator of each polymer chain of these polymers

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is either a carboxyl or hydroxyl ion. The molecular structures of these polymers are shown in Fig. 11.13. Today, hydroxyl-terminated polybutadiene (HTPB) is the most popular solid propulsion polymer because it is extremely safe to handle and highly energetic. PMM is used most often for combustion research because of its low cost, easy accessibility, and transparency. Normally, the fuel density of simple polymer fuel is much less than that of solid rocket fuel. Powdered metal, such as aluminum, is sometimes added to the polymers to increase propellant density, and this leads to decreased hybrid rocket motor volumes. Oxidizers can be gaseous or liquid. Liquid oxygen and hydrogen peroxide (H2O2) as well as nitrogen tetroxide (N2O2) and fluorine are suitable oxidizers. Table 11.5 lists a few oxidizers and their properties. “Reverse hybrids” also exist, where the oxidizer is solid and the fuel is liquid. In this case, ammonium perchlorate is commonly used as the solid oxidizer, whereas kerosene, hydrazine, or liquefied hydrogen is used as the liquid fuel. Table 11.6 lists a few selected fuels for reverse hybrid propulsion.

Fig. 11.13

Chemical structure of typical fuels.

636

TABLE 11.5 Oxidizer

Symbol

Fluorine

Fl

Hydrogen peroxide

OXIDIZERS FOR HYBRID PROPULSION SYSTEMS

TBP, K

TFP, K

Pvap, Pa

Density, kg/m3

Storage

Handling

Stability

Material Compatibility

85

53

5.57 MPa @144 K

1509

Cryogenic

Very toxic, flammable

Good

Al, SS, Ni, brass

H2O2

419

267.4

345 @289 K

1414

Decomposes at 1%/ year

Burns skin, flammable

Unstable at T . 414 K

Al, SS, Ni, Kel-F

Nitrogen tetroxide

N 2 O2

294

261

0.765 MPa @344 K

1440

Good when dry

Burns skin, toxic

Temperature dependent

Al, SS, Ni, Teflon

Oxygen

O2

90

54

5.07 MPa @154 K

1142

Cryogenic

Good

Good

Al, SS, Teflon, Kel-F, Ni, Cu

D. V. TRINH

Fuel

Symbol

TBP, K

TFP, K

FUELS FOR REVERSE HYBRID PROPULSION Pvap, Pa

Density, kg/m3

Storage

Handling

Stability

Material Compatibility

Hydrogen

H2

20.3

13.8

1.294 MPa @32.8 K

71

Cryogenic

Flammable

Flammable

Al, SS, Ni, Kel-F

Hydrazine

N2H4

386

274

19300 @344 K

1010

Good

Toxic, flammable

Toxic, flammable

Al, SS, Teflon, Kel-F, polyethylene

Monomethyl hydrazine

CH3NH-NH2

359

220

60657 @344 K

878

Good

Toxic

Toxic

Al, SS, Teflon, Kel-F, polyethylene

MATERIALS FOR SOLID ROCKET ENGINES

TABLE 11.6

637

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11.9 SUMMARY, TRENDS, AND OUTLOOK The SRM designer’s disciplines are the interrelationships among molecular structures of constituent materials; formulation; processing; and mechanical properties of propellant, liner, and insulation. Rocket motor storage and operation place structural loads on the grain, the propellant-liner-insulation (PLI) system interface adhesions, and the insulation-to-case bond. As a result, the designer sometimes must weigh ballistic performance against structural load design considerations. In the future, there will be higher demand to go beyond low Earth orbit, requiring more energetic propellant with higher heat flux; therefore, HTPB and HTPE may be better choices than the currently used PBAN. Replacements for asbestos in rubber insulation systems may introduce new challenges for processing. Propellant voids and PLI unbonds or separations still occur and remain a major concern in SRM manufacturing. Composite insulator rubber research continues to be an active area of interest as manufacturers strive to produce insulation systems able to endure higher heat flux for future space launch systems. Monolithic filament-wound composite SRM cases are already in use in smallto moderate-sized SRM systems. More powerful lightweight motors of the future may benefit greatly from the availability of segmented composite cases, wherein significant advancements in joint technology may be necessary. Advancements in nondestructive evaluation techniques capable of rapidly screening composite cases used in flight hardware for hidden damage from accidental impacts or rough handling will also be essential. CCP has been a frequently used ablative insulating nozzle material for SRMs, protecting primary structural components from the extreme thermal environment during motor operation. Production of rayon, the typical precursor for carbon fibers, results in environmentally unfriendly by-products; therefore, most major domestic manufacturers have ceased operations. High-quality rayon fibers are still produced in Europe by one manufacturer, and some lower grades are produced in Asia. Much effort has been placed on investigating rayon replacement technology. New fibers such as Lyocell appear to be promising. Staple fiber versions, which are more resistant to particle impact, can now be fabricated and tested at laboratory scale. Usable versions may soon be in commercial production.

BIBLIOGRAPHY Advisory Group for Aerospace Research and Development, “Combustion of Solid Propellants,” AGARDLS-180, Advisory Group for Aerospace Research and Development, Neuilly Sur Seine, France, 1991. ATK Thiokol, “RSRM Design and Manufacturing Baseline Short Course Presentation,” ATK Thiokol, Promontory, Utah, 2008.

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“Bulk Modulus Testing of Thiokol Polyisoprene DL1514A and Thiokol Natural Rubber TR3005A,” Report U-91-4475, 14 Feb. 1991. “Characterization Effort of NARC Material Evolution Series, Volume IV: NARC HRPF,” Southern Research Institute Report, SRI-MME-93-182-7033.4, Feb. 1993. “Characterization of Glass and Silica Phenolic Two-Dimensional Composites,” Volume I: Glass Phenolics, Southern Research Institute Report, SRI-MME-91-327-6272 “Characterization of Glass and Silica Phenolic Two-Dimensional Composites,” Volume II: Silica and Low Resin Glass Phenolics, Southern Research Institute Report, SRI-MME91-1115-6272 Humble, R. W., Henry, G. N., and Larson, W. J., Space Propulsion Analysis and Design, Space Technology Series, McGraw-Hill, New York, 1995. McDonald, A. J., Bennett, R. R., Hinshaw, J. C., and Barne, M. W., “Chemical Rocket and the Environment,” Aerospace America, Vol. 25, No. 5, 1991, p. 33. Moore, G. E., and Berman, K., “A Solid Liquid Rocket Propellant System,” Jet Propulsion, Vol. 26, 1956, pp. 965–968. “MX6001 Glass Cloth Phenolic Material Characterization,” Southern Research Institute Report, SRI-ENG-99-55-9806.00. NASA, “Solid Propellant Grain Design and Internal Ballistics,” NASA SP-8076, 1972. NASA, “Solid Propellant Selection and Characterization,” NASA SP-8064, 1971. NASA, “Solid Rocket Igniters,” NASA SP-8051, 1971. NASA, “Solid Rocket Motor Metal Cases,” NASA SP-8025, 1970. NASA, “Solid Rocket Motor Nozzles,” NASA SP-8115, 1975. NASA, “Solid Rocket Motor Performance Analysis and Prediction,” NASA SP-8039, 1971. Siegel, B., and Schieler, L., Energetics of Propellant Chemistry, Wiley, New York, 1964. Stokes, E. H., “The Physical, Mechanical and Thermal Properties of the FM 5055 404 Aft Inlet Ring, Volume IV: Thermal Tests, Gas Permeability and Thermal Expansion,” Southern Research Institute Report, SRI-EAS-88-201-6032-1, March 1988.

CHAPTER 12

Materials for Liquid Propulsion Systems John A. Halchak Los Angeles, California

James L. Cannon NASA Marshall Space Flight Center, Huntsville, Alabama

Corey Brown Aerojet-Rocketdyne, West Palm Beach, Florida

12.1 INTRODUCTION Earth to orbit launch vehicles are propelled by rocket engines and motors, both liquid and solid. This chapter will discuss liquid engines. The heart of a launch vehicle is its engine. The remainder of the vehicle (with the notable exceptions of the payload and guidance system) is an aero structure to support the propellant tanks that provide the fuel and oxidizer to feed the engine or engines. The basic principle behind a rocket engine is straightforward. The engine is a means to convert potential thermochemical energy of one or more propellants into exhaust jet kinetic energy. Fuel and oxidizer are burned in a combustion chamber where they create hot gases under high pressure. These hot gases are allowed to expand through a nozzle. The molecules of hot gas are first constricted by the throat of the nozzle (de Laval nozzle), which forces them to accelerate; then as the nozzle flares outward, they expand and further accelerate. The mass of the combustion gases times their velocity, reacting against the walls of the combustion chamber and nozzle, produces thrust according to Newton’s third law: for every action there is an equal and opposite reaction [1]. Solid rocket motors are cheaper to manufacture and offer good value for their cost. Liquid propellant engines offer higher performance, that is, they deliver greater thrust per unit weight of propellant burned. They also have a considerably higher thrust-to-weight ratio. Because liquid rocket engines can be tested several times before flight, they have the capability to be more reliable, and their ability to shut down once started provides an extra margin of safety. Liquid propellant engines also can be designed with restart capability to provide an option for orbital maneuvering. In some instances, liquid engines also can be designed to be reusable. On the solid side, hybrid solid motors also have been developed  Consultant.

Copyright # 2017 by the American Institute of Aeronautics and Astronautics, Inc. The U.S. Government has a royalty-free license to exercise all rights under the copyright claimed herein for Governmental purposes. All other rights are reserved by the copyright owner.

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with the capability to stop and restart. Solid motors are covered in detail in Chapter 11. Liquid rocket engine operational factors can be described in terms of extremes: temperatures ranging from that of liquid hydrogen (–4238F) to 60008F hot gases, enormous thermal shock (70008F/s), large temperature differentials between contiguous components, reactive propellants, extreme acoustic environments, high rotational speeds for turbomachinery, and extreme power densities. These factors place great demands on materials selection, and each must be dealt with while maintaining an engine of the lightest possible weight. This chapter will describe the design considerations for the materials used in the various components of liquid rocket engines and provide examples of usage and experiences with each.

12.2 LIQUID ROCKET ENGINES Liquid rocket engines are either monopropellant or bipropellant. Monopropellant engines either use a straight gaseous system or employ a catalyst to decompose the propellant in an exothermic reaction. An example was the reaction control system on the Mercury capsule in which each small thruster used hydrogen peroxide decomposed by a silver catalyst to provide attitude control for the vehicle [2]. Monopropellant thrusters are usually used only for low-thrust systems such as satellite propulsion systems. Bipropellant systems use either hypergolic propellants or propellants that require a source of ignition. Hypergolic propellants usually are storable, that is, they are sufficiently stable to remain in their tanks for a considerable period of time under normal Earth or space conditions. The most commonly used hypergolic propellants are nitric acid, nitrogen tetra oxide (NTO), hydrogen peroxide, monomethyl hydrazine (MMH), and unsymmetrical dimethyl hydrazine (UDMH). Nonhypergolic propellants that have been most commonly used are liquid oxygen, alcohol, kerosene, and liquid hydrogen. Liquid rocket engines differ greatly depending on mission, materials of construction, and launch vehicle design. The four broad categories into which liquid engines may be placed are 1) first-stage or core (booster) engines, 2) upper-stage engines, 3) satellite propulsion engines, and 4) attitude control systems. Each of these categories has its own unique requirements that significantly influence the engine design. In a liquid engine, the fuel and oxidizer propellants must be delivered to the combustion chamber at a pressure significantly greater than that in the combustion chamber. For relatively low-thrust engines (e.g., satellite propulsion or attitude control), this may be done by pressurizing the propellant storage tanks. For higher thrust, high-performance engines, pumps must be used. To obtain the necessary power to drive high-performance pumps, a turbine drive is needed. To power the turbine, hot gases or high-pressure fluids are needed. There are many methods by which to provide the gases or fluids to drive the turbines. One of the most common methods is to provide hot gases by tapping off and

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combusting a small portion of propellants in a separate chamber called a gas generator. Another common method, the expander cycle, uses the energy of expanding fuel, heated by the combustion chamber and nozzle, to drive the turbines. Expander cycles require a low vapor pressure propellant, which limits them to propellants such as hydrogen or methane. Another cycle, the staged-combustion cycle, partially combusts almost all of the fuel with a portion of the oxidizer before entering the main combustion chamber where it is mixed with the remaining oxidizer to complete the combustion process. An alternative staged-combustion cycle reverses the process by utilizing oxidizer-rich precombustion with combustion completed by injecting the remaining fuel into the main combustion chamber. This version of staged combustion is favored by Russian designers [2]. Staged combustion is used when large power inputs are required to produce very high combustion chamber pressures (e.g., 3000þ psi). No matter what the cycle, ducting is needed for the transfer of propellants to various engine components. To control the flow of propellants, one needs valves that are quick acting and able to operate against very high-pressure fluids. To actuate the valves, hydraulic, pneumatic, or electrical actuators are necessary. A method is needed to uniformly inject and mix the propellants in the combustion chamber to ensure uniform, efficient, and stable combustion; this is called the “injector” of the combustion chamber. Unless hypergolic propellants are used, a method to initiate combustion is required. Common methods to do this include the use of a pyrotechnic device, the injection of a hypergolic compound, or the use of spark ignition. A nozzle and sometimes a nozzle extension are needed to capture as much reactive thrust as practical. The combustion chamber and nozzle must be cooled to prevent them from melting from the heat of combustion. Cooling methods include regenerative, dump, film, transpiration, and ablative, or some combination of these. Commonly used for cooling is a double-walled design through which fuel (and sometimes oxidizer) is circulated to extract heat. For directional control of the vehicle, gimbaling of the engine or vanes in the exhaust stream may be used; either of these requires actuators. Sensors are required at key locations to monitor engine performance and make appropriate propellant flow changes or provide for rapid shutdown if an anomaly is detected. To manage this, a control system is needed, which usually takes the form of an electronic computer. All of these elements must come together in an efficient, lightweight package that operates smoothly, efficiently, and reliably. The materials of construction are a key element to achieving the goals of performance, reliability, and relatively light weight.

12.2.1

PROPELLANT SELECTION

The choice of propellant combination for a liquid rocket stage is of fundamental importance because it helps to determine the size, weight, performance characteristics, and cost of the resulting vehicle. Liquid rocket propellants can be divided into two broad categories: storable propellants, which are in the liquid state at room temperature and pressure, and cryogens, which must be kept cold to

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prevent them from boiling. Storable propellants may be kept in the vehicle for extended periods before launch, and many storable propellant combinations are also hypergolic, which means they ignite on contact and do not require an ignition source. Cryogens, on the other hand, offer higher specific impulse (Isp) than storable propellants and are often used for applications where performance is critical [3]. The selection of propellant combination is influenced by two important factors: the engine’s operating environment (booster or upper stage) and its thrust size. Upper stages have often used liquid hydrogen as a fuel because the performance benefits of this cryogenic fuel frequently outweigh the difficulty and cost of maintaining it in the liquid state. The performance of upper-stage engines is critical because any loss in efficiency will require more propellants to lift a given payload, impacting the design of both the upper stage and the booster stage that must lift it. For small stages, however, performance has less of an impact on the overall vehicle, and many small engines have been designed to use storable propellants. Booster stages, for which performance is somewhat less critical, often have used RP-1, a refined form of kerosene, or storable fuels such as hydrazine. The denser fuel reduces the physical size of the stage, which is beneficial because the booster stage is the largest stage in the vehicle and dictates the size of the handling and launch facilities. Stages using either kerosene or hydrogen fuel generally use liquid oxygen, which can be obtained and stored fairly easily, as the oxidizer, whereas those stages using storable fuels such as hydrazine typically use nitrogen tetra oxide (N2O4) oxidizer, creating a propellant combination that is both storable and hypergolic.

12.2.2

HISTORICAL PERSPECTIVES

All of the early rocket pioneers, including Konstantin Tsiolkovsky, Robert H. Goddard, and Hermann Oberth, concluded that to achieve their performance aims, liquid propellant engines would be needed. These three men also independently conceived the concept of staging for the practical delivery of a payload into Earth orbit. Of these three, only .Goddard put theory into practice. Goddard was the first person to fly a liquid-propelled rocket, in March 1926, using a crude rocket that burned liquid oxygen and gasoline. Interestingly, all early rocket scientists recognized that hydrogen would provide the highest performance as a fuel, some 50 years before the first successful hydrogen-fueled rocket engine was developed: the Pratt and Whitney RL-10. Throughout the 1930s, Goddard designed, built, and tested an increasingly sophisticated series of test rockets. He received more than 200 patents for various features of his rockets. Dr. Goddard was the quintessential 19th-century lone-wolf inventor: he attempted to do everything himself, and he patented everything that he invented. However, he was living in the 20th century, and rocketry was a technology that, to successfully advance, required interdisciplinary teams of engineers and scientists working together. Still, Goddard was a prolific inventor,

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and the rockets that he designed and built pioneered most of the common features of today’s launch vehicles. In the 1930s, while Goddard was laboring in the New Mexico desert in virtual secrecy with a small amount of funding from the Guggenheim Foundation [4], teams of engineers in Germany and the Soviet Union began work on rockets with substantial government funding. It is interesting to note that in the mid-1930s, the most advanced rocket engines in the world were under development in the Soviet Union. However, that ended abruptly in 1937 with Stalin’s great purge of the Soviet military and intelligentsia. Because of their association with key military leaders whom Stalin considered potential rivals, the rocket design bureaus were caught in the purge. Many of their top scientists were sent to the gulags while the military officers who supported them were executed. Work on advanced rocketry in the Soviet Union did not begin again until after the end of World War II. In contrast, a well-funded rocket development effort was underway in Germany under the leadership of Col. (later General) Walter Dornberger and the brilliant Dr. Wernher von Braun. The young von Braun had recognized that rocketry was “big science” that required funding only a government could provide. When the German army offered him a position in its recently formed rocket development program, he readily accepted. The final result of the German effort was the remarkable A4 ballistic missile, better known as the V-2. The V-2 was the world’s first man-made object to leave the Earth’s atmosphere and enter outer space. Although not a decisive weapon, it ravaged London and Antwerp and affected Allied war strategy. At the end of the war, both the Americans and the Soviets captured components for about 100 V-2 missiles and brought them to their countries for study and development. Each country hired more than 100 German engineers to aid in missile development. Led by von Braun, almost all the top German engineers signed on with the Americans. The Soviets obtained the second-tier engineers, but they too were very skilled and significantly aided Soviet rocket development. The Soviets immediately gave rocket development a high priority, as it was perceived as the way to counter the commanding lead of the United States in manned bombers. By way of contrast, in the United States, rocket development progressed at a leisurely pace. In the late 1940s, there were four major American rocket engine companies: Reaction Motors, which was formed by members of the American Rocket Society around 1940; Aerojet, which was founded around 1942 by Theodore von Karman and some of his graduate students at California Institute of Technology; General Electric, which had the contract from the U.S. Army to oversee the launching of V-2 rockets from White Sands proving ground in New Mexico; and the relative newcomer to the field, North American Aviation (NAA), a leading military aircraft company located in Los Angeles. In the 1950s, two additional companies worthy of note entered into liquid rocket engine development and production: Bell Aircraft (Bell Aerospace Div.) and the Marquardt Company, which originally was founded to produce ramjets but entered the realm of small rocket engines in 1959.

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NAA entered the rocket engine business because of a contract from the U.S. Air Force to produce an intercontinental cruise missile (the X-10/SM-64 Navajo) that needed a rocket boost to activate its ramjet propulsion. The other rocket engine companies were competitors, and so NAA had no choice but to do it itself. The corporation made the logical decision to build upon V-2 technology. The key to the success of NAA’s rocket engine business, later to be named Rocketdyne, was that the corporation was both willing and able to invest large amounts of company funds in the business, up front, with the vision that contracts would eventually follow. The other companies either had difficulty obtaining financing or had corporate management reluctant to invest company funds into development efforts without obvious contracts on the horizon. In 1946, NAA began by testing small rocket motors and then built three duplicate versions of the V-2 engines, followed by the development of an improvement on the V-2 engine that had higher thrust yet lower weight [5]. In the late 1940s and early 1950s, Aerojet built engines for the Aerobee sounding rocket and the Corporal ballistic missile. Reaction Motors built the engine for the Viking sounding rocket. This engine had the distinction of being the first rocket engine to use gimbaling for directional control. Reaction Motors also made the small rocket motors for a series of piloted research planes, including the Bell X-1 and X-2 and the Douglas Skyrocket. Some years later, it provided the rocket motor for the X-15, the first piloted aircraft to enter outer space. Edward Nu, an engineer at Reaction Motors, developed and patented the concept of the tubular-wall combustion chamber; however, he never applied it beyond small demonstration models. Reaction Motors became a subsidiary of Thiokol in 1958 and went out of business in 1972 for lack of contracts. In 1950, NAA developed the first high-thrust American engine (75,000 lbs thrust), the XLR43-NA-1, which evolved into the A6 and A7, better known as the Redstone engine [5]. This was followed by the first production tubular-wall combustion chambers for the Navajo missile. All of the early rocket engines used either alcohol or aniline as a fuel. NAA’s 1953 investment in corporate funds to gain basic knowledge in kerosene-fueled engines paid off in 1954 when the Air Force unexpectedly released a request for quotes for engines for Convair’s Atlas intercontinental ballistic missile (ICBM). NAA’s groundwork research on rocket engines fueled with kerosene enabled it to easily win the competition. This was followed within a year by contracts for engines for both the Thor and the Jupiter intermediate range ballistic missiles (IRBMs). The increased business justified the formation of a separate division, which was given the name Rocketdyne in 1955. For its part, Aerojet obtained the engine contract for the Titan ICBM a year after the Atlas engine contract (1955). Aerojet then made a significant improvement in the Titan engines by converting propellants from liquid oxygen (LOX) and kerosene to the hypergolic propellants NTO and MMH. General Electric (GE) made some singular contributions in injector design and provided the first-stage engine for the Vanguard launch vehicle, which was slated to launch a satellite in 1958. Because of a corporate decision to focus on

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turbojet engines, GE left the rocket engine business around 1966. Marquardt became a leader in providing small thrusters for satellites and manned space vehicles; however, the company became a victim of the end of the Cold War downturn, and its existence ended in the early 1990s. Bell Aerospace produced a number of relatively small monopropellant and bipropellant thrusters and, most notably, the Agena family of upper-stage engines. Bell Aerospace was purchased by Atlantic Research Corporation in 1983 [2]. In the Soviet Union, the engine design bureau headed by Valentin Glushko also built on V-2 technology and developed many different engines, including the RD107/108 family of engines in 1957. This four-combustion chamber, LOX–kerosene engine, developing 220,000 lbs thrust, powered Sergei Korolev’s R7 rocket to launch the world’s first artificial satellite. When the Soviet Union beat the United States in launching the first Earth satellite, the unexpected propaganda coup spawned the space race. To partially counter the Soviet accomplishment, von Braun used the U.S. Army’s Redstone missile, with its reliable A7 engine, to launch America’s first satellite. The same Redstone then was used to place the first American astronauts into space with the Mercury suborbital flights. The Atlas and Titan ICBMs were turned into launch vehicles for both human flight and satellite launching. For their part, the Soviets used the R-7 with its reliable RD107/108 engines to launch the first man into space, followed by many “firsts” in manned spaceflight. Interestingly, the R-7 with its RD107/108 engines remains to this day the only man-rated Russian launch vehicle. In 1958, another company, Pratt and Whitney, entered the rocket engine business with the world’s first hydrogen-fueled rocket engine, the 15,000-lb thrust RL-10. In 1959, Rocketdyne began work on the F-1, which at 1.5 million lbs thrust, would eventually power the first stage of the Saturn V moon rocket and hold the title of the world’s largest rocket engine for almost 20 years. In 1962, Rocketdyne also began development of the J-2, a 230,000-lb thrust hydrogen-fueled engine. For the Apollo moon mission, Rocketdyne provided booster engines for the Saturn V vehicle: first stage (five F-1 engines), second stage (five J-2 engines), and third stage (one J-2 engine). Aerojet provided the all-important engine for the Command and Service Module. TRW provided the Lunar Excursion Module (LEM) descent engine, whereas a Bell–Rocketdyne team furnished the LEM ascent engine. Marquardt was the supplier of the reaction control system (RCS) engines on the LEM, whereas Rocketdyne provided the RCS engines for the Apollo capsule. Following the successful Apollo moon mission, design of the main liquid propulsion engines for the space shuttle began, with Rocketdyne receiving the development and production contract. Later, Pratt and Whitney received a contract to provide improved high-pressure fuel and oxidizer pumps for the shuttle engines. When designed in the early 1970s, the space shuttle main engine (SSME) was the most advanced rocket engine in the world. At the time of development, the SSME had the highest Isp, could be throttled between 60% and 109% of rated thrust, and was also reusable. This was the first engine in the West to use

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the staged-combustion principle, wherein the majority of the propellants were first partially combusted in a “preburner,” the gases from which were used to drive high-pressure turbopumps before entering the main combustion chamber to be mixed with the remaining propellants and complete the combustion process. The staged-combustion process enabled very high combustion chamber pressures, which greatly improved the efficiency of the engine. This staged-combustion, high-combustion chamber pressure concept had first been developed in the United States by Pratt and Whitney. In the Soviet Union, the staged-combustion concept was pioneered around 1958, and this concept was adopted by the leading rocket engine designer, Valentin Glushko, who also decided to shift from LOX–kerosene to hypergolic propellants in the early 1960s. Glushko had had considerable difficulty with combustion instability in large, LOX–kerosene engines. Mitigating instability was the reason for dividing the RD107/108 into four separate combustion chambers, thereby separating combustion into smaller chambers wherein instability was more manageable. The multiple-chamber design remains a feature on most Russian LOX–hydrocarbon engines. Glushko was attracted by the higher performance and inherent stability of hypergolic propellants. This led to a major conflict with the Soviet chief vehicle designer Sergei Korolev, who favored kerosene-fueled vehicles. Korolev brought the Kuznetsov design bureau into the picture to produce the LOX–kerosene engines for his N-1 moon rocket. This design bureau adopted the staged-combustion cycle for LOX–kerosene engines. Before this, most large engines had powered their turbopump turbines with gas generators. By utilizing all of the fuel and oxidizer in a two-stage combustion process, the staged-combustion cycle provided the highest performance attainable in LOX–kerosene engines. The Soviets also had elected to use oxygenrich preburner combustion to control turbine temperatures, as opposed to the West, where fuel-rich gas generator and preburner combustion was used. The use of oxygen-rich combustion, although riskier than fuel-rich, provided an extra margin of performance due to the higher mass flow. Although the Soviets belatedly developed hydrogen-fueled engines, their main focus was, and remains, kerosene–oxygen and hypergolic propellant engines. It is interesting to note that, in the 1960s race to the moon, engineers in the United States elected to go with the relatively simple, gas-generator cycle for their F-1 and J-2 engines for the Saturn V–Apollo launch vehicle. This required tackling the issues of combustion instability and learning how to deal with liquid hydrogen but offered the advantages of reliability and high performance (in the case of hydrogen) in highly producible engine designs. In contrast, the Soviets expended considerable effort on the design complexities of staged combustion, which in the final analysis may have adversely affected the development schedule and reliability of their N-1 moon rocket. In the 1970s, other countries began the development of their own indigenous rocket engine technology. These included Japan, European countries (France, Germany, Sweden, etc.), and China. These developments borrowed heavily

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from American and Soviet technology, but each country also contributed its own unique design features to its engines. In recent years, an emphasis has been placed on cost as opposed to strictly focusing on performance. In this, Russia has a distinct advantage due to the exchange rate of the ruble. In the West, Japan, the United States, and Europe have made significant efforts to produce lower cost engines. One example is the RS-68 engine, which may be the first rocket engine in which, as a design factor, cost was given equal consideration as performance and the overall weight of the engine. Today we see the emergence of the commercial space launch companies, such as SpaceX, Blue Origin, Orbital Sciences, and Virgin Orbit. These companies are creating a new, low-cost paradigm, using new operational philosophies and advanced materials in their designs. These companies are set to begin an entirely new chapter in the space launch business.

12.3 GENERAL DESIGN CONSIDERATIONS FOR MATERIALS SELECTION The most basic division among liquid rocket engines is between pressure-fed and pump-fed engines. In the first type, the propellant tanks are pressurized to provide the desired combustion pressure, whereas in the second type, pumps are used to raise the pressure of the fuel and oxidizer after they leave the tanks. Pressure-fed engines are inherently simpler and have found wide application for small upper-stage engines. To avoid making the propellant tanks too heavy, the combustion pressure for pressure-fed engines is generally limited to 100– 200 psia. Table 12.1 lists some existing pressure-fed engines and the thrust size and combustion pressure for each. The higher propellant tank weights inherent in the pressure-fed approach, along with the low combustion pressures, limit the applicability of this type of engine. Large stages tend to accept the higher complexity of pump-fed systems to obtain lower stage weights. In addition, low combustion pressures limit the nozzle expansion ratio for booster engines because the exhaust flow will separate from the nozzle wall if its pressure drops too far below the ambient air pressure. For this reason, booster engines typically use the pump-fed approach. Once a pump-fed engine has been selected, it is necessary to determine how the power to drive the pumps will be obtained. In most pump-fed engines, a gas turbine is used to drive the pump; the source of the gas for this turbine determines the “cycle,” the general architecture of the engine. The rocket cycles in most common usage today can be grouped into three broad categories: 1) expander, 2) gas generator, and 3) staged combustion. In the expander cycle, one of the propellants—typically the fuel—flows through passages in the thrust chamber, picking up heat to keep the chamber cool. The propellant warms up, gasifies, and is used to drive the turbine. The turbine temperatures are fairly low, which is good for hardware life, but

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TABLE 12.1 Engine

HISTORICAL PRESSURE-FED ROCKET ENGINES

Vehicle

Propellants

Thrust, lbf

Chamber Pressure, psia

AJ10-118 K

Delta II

N2O4/ Aerozine-50

9800

125

RS-18

Apollo lunar module (ascent stage)

N2O4/ Aerozine-50

3500

122

LMDE

Apollo lunar module (descent stage)

N2O4/ Aerozine-50

9850

103

AJ10-137

Apollo service module

N2O4/ Aerozine-50

20,500

97

Aestus

Ariane 5

N2O4/MMH

6744

160

because the amount of heat that can be picked up by the coolant is limited, the expander cycle usually operates at moderate to low combustion pressures. Expander cycles may be either open or closed; in the open type, the turbine discharge flow is routed overboard or into the divergent part of the nozzle, whereas in the closed type, the turbine flow is routed to the combustion chamber to be burned. The closed expander cycle is shown schematically in Fig. 12.1. In Figure 12.1, the fuel first passes through a two-stage pump, which is driven by a gas turbine. The fuel then cools the thrust chamber, which increases the fuel’s temperature and converts it from a liquid to a gas. Next the fuel passes through the turbine, where it provides the power to turn the fuel and oxidizer pumps. The fuel then flows through the main injector, enters the combustion chamber, and is burned. The oxygen first passes through a single-stage pump, which is driven by the single turbine via a gear assembly. The oxygen then passes through a control valve, which is used to vary the relative amounts of fuel and oxidizer flowing through the engine. Then the oxygen flows through the main injector and into the combustion chamber. In the open expander cycle, a relatively small fraction of the fuel cools the thrust chamber, then drives the turbines and is discharged overboard, or into the divergent section of the nozzle. In this way, higher turbine pressure ratios can be achieved, allowing a more complete release of the energy in the coolant flow and enabling higher chamber pressure operation. The additional chamber pressure allows a higher nozzle expansion ratio within a given envelope, helping to offset the performance loss incurred by routing the turbine flow around the combustion chamber. Figure 12.2 shows a schematic of an open expander.

MATERIALS FOR LIQUID PROPULSION SYSTEMS

Fig. 12.1 Closed expander engine schematic.

Fig. 12.2

Open expander engine schematic.

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Expander engines offer a simple engine architecture but have a limited range of applicability. Only fuels that gasify readily and cleanly, such as hydrogen or methane, are appropriate for expander cycles; kerosene fuel, one of the most common rocket propellants, does not satisfy this requirement. Also, because the thrust chamber surface area increases more slowly than its internal volume as the engine’s thrust size goes up, the relative amount of heat pickup in the coolant circuit decreases and reduces the achievable chamber pressure. As a result, expander engines have typically been used for smaller upper-stage engines, although some large booster applications are being studied. To get more energy into the flow driving the turbine, some rocket engines have an auxiliary combustor, called either a “gas generator” or a “preburner,” which burns some of the propellants and increases the temperature of the gas flowing into the turbine. The difference between the gas generator and preburner of staged-combustion cycles lies in what happens to the turbine exhaust. In gas generator cycles, the turbine exhaust is routed overboard, either through an external pipe or into the divergent section of the nozzle. This method is fairly simple and allows a high-pressure ratio across the turbines, but it is wasteful because the turbine gases are not fully expanded in the nozzle. In the staged-combustion engine, the turbine exhaust is routed to the main injector, where it is burned with the rest of the propellants. This method is more efficient, but because it requires that the turbine exhaust pressure be higher than the main combustion chamber pressure, system pressures tend to be quite high and the engine system is complex. A schematic of a gas generator engine is shown in Fig. 12.3. In this engine, a small part of the fuel exiting the pump is routed to the gas generator, where it is burned with a small amount of oxygen to provide the hot

Fig. 12.3

Gas generator engine schematic.

MATERIALS FOR LIQUID PROPULSION SYSTEMS

Fig. 12.4

653

Staged-combustion engine schematic.

gas to drive the turbines. The rest of the fuel goes to cool the thrust chamber, then to the injector where it is burned in the combustion chamber along with the oxygen not used in the gas generator. The hot gas from the gas generator drives the fuel and oxygen turbines in series. The turbine exhaust is then routed into the divergent section of the nozzle, so that it may expand and contribute some thrust before leaving the engine. Figure 12.4 shows a schematic for a staged-combustion engine. In this cycle, most of the fuel goes to cool the nozzle and then moves on to the preburner. The remainder of the fuel cools the combustion chamber and then goes to the main injector. On the oxidizer side, a small portion of the flow goes to the preburner, with the rest passing through the oxidizer control valve and then to the main injector. The hot gas produced by the preburner powers the turbines for the fuel and oxidizer pumps. Once it has left the turbines, the hot gas moves to the main injector and into the combustion chamber, where it joins the rest of the fuel and oxidizer. To keep the hot gas temperatures within acceptable limits for the turbine hardware, staged-combustion engines run the preburner with large amounts of excess fuel or oxygen; the excess flow does not burn and reduces the combustion temperature to acceptable levels. The choice of whether to run the preburner fuel rich or LOX rich is made to obtain the maximum possible energy extraction in the turbines and depends on the particular fuel being used. Hydrogen-fueled engines like the SSME use fuel-rich preburners, whereas kerosene-fueled engines like the RD-180 run the preburner LOX rich. Gas generator engines, on the other hand, tend to run the gas generator (GG) fuel rich, whether the fuel is hydrogen or kerosene, because this minimizes the fuel routed overboard through the turbine circuit, thereby maximizing overall engine performance.

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Fig. 12.5 Chamber pressure range for various pump-fed cycles. There are many variations on the schematics shown in Figures 12.1 through 12.4. The fuel and oxidizer pumps may be separate or on a single shaft. If separate, each may be driven by its own turbine or by a single turbine through a gear train. Staged-combustion engines may have one preburner or two. The exact configuration is chosen to optimize the final engine design, taking into account the relative impacts of cost, complexity, weight, and performance, as well as the technological limitations placed on the engine design, such as material limits for operating stress and temperature. The choice of cycle depends largely on how the engine will be used. Booster engines, which must operate at high pressure to avoid flow separation in the nozzle, usually use the GG or staged-combustion cycles. Upper-stage engines, which operate in a near-vacuum and are not susceptible to flow separation, can operate at the lower pressure of the expander cycle. Figure 12.5 shows the range of chamber pressure for various types of cycles, including existing rocket engines and some that did not progress beyond the conceptual stage. It can be seen that a broad range of chamber pressures exists for each cycle, and that there is a fair amount of overlap between the different types. The requirements of each application will drive not only the choice of cycle but the choice of chamber pressure as well.

12.4 MATERIALS Materials were very important from the very beginning of rocketry. Sir William Congreve’s solid rockets (“the rockets’ red glare” of the American national anthem) were an improvement over the Indian war rockets because they used iron as the casing material. Goddard used an eclectic combination of

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materials: aluminum tubing, ceramic-lined aluminum combustion chambers, ceramic-coated tool steel nozzles, asbestos thermal protection, pump housings made of aluminum, brass and steel impellers, aluminum alloy turbines, and so on. Materials usage progressed to increasingly higher strength and higher temperature alloys, as pressures and temperatures increased to achieve higher performance. More recently, nonmetallic materials such as ceramics, ceramic matrix composites, and even polymeric composites have been considered for specific applications. At one time or another, just about every engineering material known has been employed in rocket engine construction. Environmental factors are a major concern in the selection of materials for liquid rocket engines. Most of the propellants are reactive fluids. Oxygen, hydrogen, nitrogen tetraoxide, hydrazine, and red fuming nitric acid all have potentially detrimental effects on materials. Their effects vary considerably from material to material, but at one time or another, each has caused problems. The specific environmental effect of a propellant must be taken into account in the materials selection and design process [6]. One of the challenges is to develop suitable screening tests to evaluate the effect of a propellant, under expected operational conditions, on the materials of choice. Efforts also have been made to develop materials that are tailored for the operational environment. Considerable work has been expended in developing materials such as hydrogen-embrittlement-resistant and oxygen-ignition-resistant alloys. Development of a new material for rocket engine application requires a lengthy development process. This involves advancing the material technology readiness from understanding the fundamental material properties to developing mechanical properties over the anticipated operating range, developing manufacturing processes, and developing a design database. Considerable financial investment also is required to develop, certify, and incorporate a new material into components for a flight system. The materials selection for a liquid rocket engine is determined by five general factors: the size of the engine, the engine duty cycle (expendable or reusable), the propellants, the turbine drive cycle, and the stage in which the engine will be used (booster or upper stage). Large engines today almost always require extensive use of metals. Small engines (a few pounds to a few thousand pounds thrust) can be made from ablatives, high-temperature alloys, or high-emissivity alloys. In general, the higher the combustion chamber pressure in a rocket engine, the greater the material issues, for a couple of reasons. First, higher pressures mean that material strength, particularly specific strength (strength-to-weight ratio) becomes very important. At high operating pressures, material thickness and component weight can become unwieldy unless materials of suitable specific strength can be employed. Second, high-strength materials tend to be more complex and difficult to fabricate. With some notable exceptions, high-strength materials also tend to have toughness and ductility issues. They also tend to be more susceptible to environmental effects. Hence, careful screening and thorough material characterization are necessary.

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For materials selection, the engineer must first know and understand the duty cycle of the particular component. The operating temperature range must be considered; both cryogenic and elevated temperatures pose unique issues. The propellants and combustion products must be taken into account for their potential effects on material properties. The principal limiting structural loads must be taken into account, whether they be high or low cycle fatigue, fracture toughness, specific strength, stress rupture, or some other property. Depending on the design and operational factors, external environmental effects such as galvanic corrosion or stress-corrosion cracking susceptibility of the materials being considered for use must be taken into account. In LOX systems, oxygen compatibility must be known. In hydrogen systems, depending on the operating temperature, environmental hydrogen embrittlement, internal hydrogen embrittlement, or hydrogen reaction (hydride formation) must be considered. Other propellants, such as nitrogen tetraoxide, hydrogen peroxide, nitrous oxide, methane, red fuming nitric acid, and hydrazine, may also present their own compatibility problems. For rotating components in pumps, high specific strength and ductility are important. For pump housings, castings often are the most cost-effective method of production; therefore, alloys with adequate specific strength and good castability are preferred. For ducting and lines, good specific strength, weldability, and formability are important. For valves handling liquid oxygen, poppets and seats with high resistance to ignition in LOX are mandatory. Combustion chambers require liners with good thermal conductivity and thermal fatigue resistance. Many materials options are available for nozzles, and the particular design and cost constraints will determine whether regenerative cooling, film cooling, ablation, or radiation cooling will be used. Each of these choices dictates a specific type of material: conventional metal alloy, refractory metal, ceramic composite, silica-phenolic ablative, and so on. Materials selection for major components of liquid rocket engines is discussed in the next sections [6].

12.5 THRUST CHAMBER MATERIALS The basic elements of a thrust chamber include the propellant inlet and distribution manifolds, injector, ignition device (for nonhypergolic propellants), combustion chamber, de Laval-type converging–diverging nozzle, and expansion nozzle (Figs 12.6 and 12.7). For almost all liquid rocket engines, the combustion chamber and nozzle are integral. For many engines, the expansion nozzle, or part thereof, is also integral with the combustion chamber and nozzle. The nozzle extension, by definition, is a separate structure designed to capture the exhaust energy that might otherwise be lost at high altitudes or in space. Because all are interrelated, they will be discussed together in this section. The combustion chamber is where the propellants are mixed, ignited, and thereby form the pressurized hot gases that then are accelerated and ejected through the nozzle, reacting against the injector face and expansion nozzle

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Fig. 12.6 F-1 rocket engine showing a) the expansion portion of the nozzle, and b) the bolted-on nozzle extension. (Courtesy of NASA image archives.) (and, if present, the nozzle extension) to propel the rocket. The combustion chamber includes the injector as well as the chamber itself. The function of the nozzle is to convert the chemical-thermal energy generated in the combustion chamber into kinetic energy. The de Laval contraction–expansion nozzle converts relatively slow-moving, high-pressure, high-temperature gas in the combustion chamber into high-velocity gas of lower pressure and temperature. The expansion portion of the nozzle, as well as the nozzle extension, are there to capture as much of the exhaust energy as is practical. The expansion nozzle of a rocket engine is not an inconsequential component, for it can account for as much as 40% or more of the total thrust generated by the engine [3].

Fig. 12.7 Rocket engine thrust chamber. (Courtesy of NASA image archives.)

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INJECTORS

The purpose of the injector is to introduce the propellants into the combustion chamber and atomize and mix them as thoroughly as possible to provide efficient and stable combustion. Injectors are of a wide variety of designs, depending on the type and nature of propellants to be mixed, the required operational characteristics of the engine, and the individual judgment of the design engineer. A detailed discussion of the theory behind injector designs is beyond the scope of this chapter. However, some general design characteristics and the materials used will be discussed here. The first injectors were simple spray nozzles that incompletely mixed propellants so that long combustion chambers were necessary to ensure complete combustion. The V-2 introduced a somewhat more sophisticated injector design with its 18 “burner cups” depicted in Figs 12.8, 12.9, and 12.10. In the V-2, the alcohol fuel and liquid oxygen oxidizer were injected through brass nozzles, and combustion initiated in 18 burner cups welded to the top of the combustion chamber. This design was an expedient that was adopted when the German engineers at Peenemunde were unable to obtain stable combustion with flat-faced injector designs (Private communication with Konrad Dannenberg, retired NASA engineer and former Peenemunde design engineer, 2002).

Fig. 12.8 Cutaway of the double-wall V-2 combustion chamber. (Original German drawing, courtesy of the Huntsville Space Museum archives.)

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Fig. 12.9 Cutaway of a V-2 combustion chamber head, showing the triple-wall configuration and “burner cup” injector arrangement. (Photograph taken at the Kansas Cosmosphere Museum.) The burner cups, like the rest of the thrust chamber, were a Cr-V alloy steel . The injector nozzles themselves were made of brass and threaded into the walls of the burner cup. The burner cup design was difficult and inefficient from a production standpoint, as each had to be welded separately onto the dome of the combustion chamber. The design also resulted in the “plumber’s nightmare” of 18 separate aluminum oxygen lines feeding the burner cups. Each burner cup had its own set of nozzles and was, in effect, a separate combustion chamber where combustion was initiated and then completed in the main chamber. In the 1940s, combustion instability was not understood, but the burner cup arrangement effectively negated instability for the V-2.

Fig. 12.10 Cutaway of a V-2 burner cup and a brass LOX injector head. (Photographs taken at the Kansas Cosmosphere.)

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In the cut-away drawing of Fig. 12.8, the alcohol fuel enters at the lower manifold and makes one upward pass to the injector nozzles at the top of the combustion chamber. Fig. 12.10 shows a detailed cross section of a burner cup, along with a brass injector head for liquid oxygen. The single liquid oxygen injector nozzle can be seen at the top of the burner cup. Also visible are the two rows of alcohol nozzles arranged along the sides of the burner cup. It is noted that the German engineers had attempted to develop a flat-faced injector but encountered combustion instability, which they simply described as “high-frequency vibrations” (Private communication with Konrad Dannenberg, former Peenemunde engineer, 2002). This burner cup type of design, in modified forms, continued to be used on a variety of engines designed in the late 1940s and throughout the 1950s. A significant improvement in injector design was the development of the flatfaced, concentric-ring injector of the A6/A7 “Redstone” engine, depicted in Fig. 12.11. The rings alternated between fuel and oxidizer and could be manifolded such than no joints were needed to separate fuel from oxidizer. This design provided better propellant mixing and considerably simplified production [2, 9]. Initially, the rings were 4130 Cr-Mo alloy steel, nickel plated, and brazed with pure copper into a backup body also of nickel-plated 4130. A significant achievement of the Redstone injector was its freedom from combustion instability. As engines grew in size, baffles were added to enhance combustion stability, and materials were changed to oxygen-free high thermal conductivity (OFHC) copper rings brazed into an austenitic stainless steel backing body (Figs 12.12 and 12.13) [5].

Fig. 12.11 Concentric-ring injector. Rings alternate: fuel, oxidizer, fuel, oxidizer, etc. (Photograph taken at the Huntsville space museum.)

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Fig. 12.12 archives.)

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Baffled, concentric-ring injector for an Atlas launch vehicle. (Courtesy NASA

To keep the injector face from melting, injector elements are designed to maintain the fire of combustion a short distance from the injector face. However, heat extraction is still necessary; therefore, it is desirable to have materials with good thermal conductivity or some form of active cooling. As mentioned previously, OFHC copper, with its high thermal conductivity, was often used in injector faces. For the Rl-10, which was the first hydrogen-fueled engine, transpiration cooling was devised for the injector face, using a unique porous metal named Rigimesh. Rigimesh is a commercial product mainly used for filters that was adapted for the face of rocket engine injectors. It consists of multiple layers of fine, austenitic stainless steel screens that are compressed and diffusion bonded to produce a quite strong yet porous metal product. This material is used for the face of many hydrogen-fueled rocket engines. For injectors that must mix a gas with a liquid, tube within a tube (coaxial) injectors are used. Examples of this are the J-2 and Vulcain engines (Fig. 12.14) and the various Russian kerosene-fueled, staged-combustion engines [2]. J-2 engine used austenitic stainless steel for the coaxial injector. A wide variety of metal alloys has been used for injectors, including alloy steel, copper alloys, stainless steels, aluminum alloys, titanium alloys, cobalt-base alloys,

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Fig. 12.13

Aluminum alloy injector with baffles. (Courtesy NASA archives.)

and nickel-base alloys. Any of these materials may be acceptable, depending on the propellant and the duty cycle of the engine. Recent advances in manufacturing technology, such as additive manufacturing, are being explored to help reduce the cost and time to develop propulsion hardware such as injectors. Using conventional materials, such as the Inconels, injectors that previously required hundreds of pieces can now be fabricated as a single piece or several pieces as shown in Fig. 12.15. This 3-D printed injector was successfully hot-fire tested in LOX/LH2 with no noticeable loss of performance over a conventionally fabricated injector [8]. 12.5.2

COMBUSTION CHAMBERS, NOZZLES, AND EXPANSION NOZZLES

Combustion chambers, nozzles, and expansion nozzles are of many designs, including tubular, channel wall, sandwich wall, solid one-piece (sometimes ceramic coated), or ablative. Materials selections that have been used include aluminum alloys, low-alloy steel, stainless steel, pure nickel, nickel-base alloys, cobalt-base alloys, titanium alloys, copper alloys, niobium alloys, carbon– carbon, ceramic matrix composites, glass-phenolic, beryllium, and refractory metals. These also can be regenerative cooled, dump cooled, film cooled, radiation cooled, or ablative cooled. Nozzle extensions also use these cooling techniques, but they rarely are regenerative cooled. Fabrication techniques include machining, tube forming, welding, brazing, diffusion bonding, composite layups with autoclaving, or woven fabric infiltrated by chemical-vapor deposition.

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Fig. 12.14 Coaxial injector face of the J-2 engine, consisting of 614 tube-within-a-tube elements. (Courtesy NASA image archives.) Originally, American, European, and Japanese designs tended to favor tubular construction for the combustion chambers and nozzle of larger engines, although welded sheet metal with ceramic coatings was sometimes used. Tubular designs have used nickel 200, stainless steels 316 and 347, Inconel 600, X750, and A286. Newer engines operating at higher pressures have gone to channel-wall construction for combustion chambers, while sometimes retaining tubular

Fig. 12.15 Left: 3-D printed rocket injector from the selected laser melting printer. Right: Injector after finishing and polishing. (Source: Courtesy NASA/MSFC.)

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expansion nozzles. Carbon–carbon composites have been used for nozzle extensions on upper-stage engines, such as the RL-10 and Vinci engines. Russian designs traditionally have used channel-wall construction for combustion chambers, using a combination of copper-chromium alloy inner liner and lowalloy steel, stainless steel, or nickel-base outer shells [9]. The MC-I (previously known as Fastrac) program provided a low-cost, 60,000-lb (60 K) thrust rocket engine to the aerospace community. Part of this low-cost design is an ablative chamber/nozzle assembly that is actively film cooled with kerosene The chamber/nozzle is designed for one-time use only and will be replaced after every flight. The baseline chamber/nozzle consists of a tape-wrapped silica phenolic liner with a filament-wound carbon epoxy overwrap added for extra strength. A filament-wound glass phenolic overwrap was also developed [10]. Dr. Robert Goddard’s first rocket used a combustion chamber made of aluminum with a layered liner consisting of asbestos and alundum/alumina (Al2O3) sleeves. The nozzle, which was threaded into the combustion chamber, was made of alloy steel with asbestos and alundum (Al2O3) sleeves (Fig. 12.16). Goddard used tool steel for his combustion chambers and nozzles, then later nickel or Monel. For thermal protection, the interior surfaces of his combustion chambers were usually coated with alumina. Although Goddard understood regenerative cooling, he considered it too complex and used only film cooling. Because of this he was constantly plagued by burn-through of the chamber walls. In some of Goddard’s designs, the copper tubing was wrapped around and soldered to the exterior of the combustion chamber. However, this was not for cooling purposes. Instead, it was to heat gases, such as nitrogen, to power gyroscopes or pump turbines [11]. The V-2 rocket engine had a very short expansion nozzle. This was a result of limitations on the overall missile length. It was a double-wall configuration that was integral with the Fig. 12.16 Replica of Goddard’s first combustion chamber and nozzle. (Photograph taken at the Huntsville Space Museum.)

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Fig. 12.17 V-2 combustion chamber and nozzle. (Photograph taken at the Huntsville Space Museum restoration shop.) combustion chamber. It was made of welded sheet metal, low-alloy steel (6000 series, 0.9Cr-0.15V-0.3C-1.0Mn). All things considered, 6000 series steel probably was the best material available in the 1940s for a sheet metal rocket engine. The steel producers were held to tight specification requirements in spite of the exigencies of World War II in Germany (Private communication with Rudy Schlidt, former Peenemunde quality engineer, 2001). At the burner cup end, a triple-wall configuration was used (Fig. 12.9). This area was dump filled with alcohol before engine start to keep the injectors from melting before fuel regenerative cooling took effect. The alcohol fuel and liquid oxygen oxidizer were injected through brass nozzles, and combustion initiated in 18 burner cups welded to the top of the combustion chamber. For cooling, the fuel (75% alcohol–25% water) was circulated through the double-wall sides of the combustion chamber. As this was not entirely adequate, film cooling also was used in “hot spots.” Aluminum tubing was used to transfer LOX from the turbopump to the combustion chamber. As 18 of these tubes were needed to feed each of the 18 burner cups, it resulted in the complex assembly so apparent in V-2 engines—see Fig. 12.17. In this figure the curved tubes around the nozzle are fuel lines made of steel. The smaller tubes at the top of the combustion chamber are liquid oxygen lines made of aluminum alloy. The lower exit nozzle portion of the thrust chamber, below the fuel inlet manifold, was only film cooled (Fig. 12.8). In place of fuel circulation, fiberglass insulation filled the space between the double walls of the lower chamber. Steel tubing carried the alcohol fuel to a lower manifold, which then circulated upward through the double walls and into the injectors. Both oxidizer and fuel tubes contained bends to allow for thermal expansion (Fig. 12.17). Developed in the late 1940s, the NAA A6/A7 engine (Redstone engine) was a significantly improved version of the V-2 engine. Like the V-2, it followed the welded, double-wall, sheet metal configuration, with a relatively short expansion

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Fig. 12.18 NAA (Rocketdyne) A6 rocket engine with double-wall, welded sheet metal. (Courtesy NASA photo Archives.) nozzle (Fig. 12.18). For the thrust chamber, 4130 low-alloy steel (Cr-Mo) sheet metal was used with an aluminum alloy LOX dome bolted onto the injector end. The most important improvement of this engine was the development of a stable, flatfaced injector. The chief injector designer, George Sutton, created an injector consisting of 19 concentric rings made of nickelplated 4130 rings brazed into a 4130 injector body, using pure copper as the brazing medium. The key to stability was the orifice pattern, which provided intimate mixing and rapid combustion of the fuel and oxidizer [2]. A humorous (although it was not thought so at the time) anecdote involving materials selection for the Redstone engine occurred during its first test. In front of U.S. Army generals and corporate executives, the first engine test resulted in an explosion that destroyed the engine. It seems the designer had selected alloy steel for the LOX dome, which, at an operating temperature of –2908F, was well below the ductile-to-brittle transition of the steel. Startup shock fractured the dome and resulted in the explosion. From then on, LOX domes were made from aluminum alloys. Because of high-strength materials and more efficient design, the A6/A7 engine weighed 34% less than the V-2 while producing 78,000 lbs thrust as opposed to the 56,000 lbs thrust of the V-2 (both sea-level thrust values). The Reaction Motors XLR-10 rocket engine for the U.S. Navy’s Viking sounding rocket also had a V-2-like design, employing a welded double-walled configuration. This engine used nickel 200 sheet metal instead of steel. When engines went to higher thrust levels, the sheet metal configuration reached a limit: walls thin enough to maintain heat transfer would buckle, whereas walls thick enough to resist buckling would have insufficient heat transfer. The answer in the United States was to go to a tubular configuration (Fig. 12.19). Tubes were brazed to each other and to a metal shell or hat-bands

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for stiffening. The result was a lightweight yet flexible structure that resisted buckling and had good heat transfer characteristics. At first, tubes were made of nickel 200 and the assembly was hand brazed. Later, tubes were made of stainless steel or nickel-base alloys and furnace brazed (Fig. 12.24). Joining the tubular assembly by welding also was attempted but was not successful until many years later, when Volvo developed a successful welding procedure for nozzles [9]. In the late 1950s and throughout the Apollo program in the 1960s, the combustion chamber and nozzle were made of the same construction; for example, both were of tubular construction, with the tube wall of the combustion chamber simply extending to form the nozzle. When the SSME began development, it was apparent that the higher combustion chamber temperature and pressure required much stronger construction, as well as an extremely high heat transfer capability. This dictated a channel-wall construction along with a copper alloy hot wall for high thermal conductivity. A new copper alloy, Narloy-Z, containing 3% silver and 0.5% zirconium, was developed to meet these requirements. Narloy-Z provided better strength and thermal fatigue at the operating temperature than existing copper alloys, while retaining a thermal conductivity 80% that of high-purity copper. It was the key to obtaining the combustion efficiency of the SSME [6, 12]. In the Soviet Union, channel-wall or sandwich-wall combustion chambernozzle configurations were used. Usually, the inner liner was a copper–chromium (Cu  3%Cr) alloy into which slots or channels were milled, and then this was brazed to a cover/close-out sheet of low alloy steel, stainless steel, or

Fig. 12.19 Comparison of welded sheet metal nozzle to tubular nozzle. (Author’s illustration.)

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Fig. 12.20 Russian combustion chamber and nozzle, showing channel-wall and corrugation-sandwich configurations. (Author’s illustration.)

nickel-base alloy. Figure 12.20 illustrates both the channel-wall and sandwich-wall configurations. Usually, the channel-wall design is used for the combustion chamber, although on higher performance engines it also may be used for the expansion nozzle. For medium-performance engines, such as the RD-107, the expansion nozzle is a sandwich structure in which the inner liner is a Cu-Cr alloy; a corrugated sheet metal is used as the divider; and the outer shell can be alloy steel, stainless steel, or nickel-base alloy. The entire assembly is brazed and the corrugations provide flow passages for coolant circulation. The channelwall design provides both strength and superior heat transfer. Channel-wall and sandwich-wall designs are also reputed to be cheaper to fabricate than tubular configurations. The drawback to channel-wall or sandwich nozzles is that they are heavier than tubular nozzles. Originally, in the 1930s, the liner and outer shell were only bolted together and the resultant leakage between channels was simply tolerated. Following World War II, a pressure brazing method was developed to securely bond the milled channel-wall liner to the outer shell. The development of this brazing method, termed solder-welding in Russia, was a technological improvement that made possible effective, efficient, and reliable rocket engines (Fig. 12.21). Initially, almost all nozzles were straight cone shapes. The conical nozzle provided ease of manufacturing but at the expense of thrust loss due to the divergent flow from the axis of thrust. As engines grew in size, the increasing nozzle length and thrust losses of conical nozzles became unacceptable. Dr. Rau, at what was then NAA’s aerophysics laboratory and would later become Rocketdyne, developed the “Rau equations” that generated the optimum contour shape for a

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Fig. 12.21 Photograph of a Russian RD107 engine showing the copper alloy inner liner of the nozzles. (Photograph taken at the Kansas Cosmosphere museum.) bell-shaped nozzle. The contoured or bell-shaped nozzle provided quick turning of exhaust gases for optimum thrust recovery and significantly shortened the overall length of the nozzle (Fig. 12.22). The disadvantage of the contoured

Fig. 12.22

Conical vs contoured nozzle. (Author’s illustration.)

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nozzle was the requirement for tubing with complex, tapered shapes. A tubing manufacturer located in Los Angeles, which specialized in golf club shafts, developed the technology for producing tubing with the required shape and taper for rocket nozzles. An interesting combination of tubular wall and sandwich wall nozzle construction is found in the F-1 engine, wherein the combustion chamber and initial part of the nozzle are brazed tubes and the nozzle extension is a type of sandwich wall construction (Fig.12.23). Tubes are of alloy X750. Sandwich construction is Hastelloy C. The nozzle of the SSME was a brazed tubular assembly, consisting of 1080 tubes of A286 stainless steel that were brazed to each other and to a 718 alloy backup shell. Brazing was performed in a retort under a cover gas of hydrogen (Fig. 12.24). This photograph shows the outer retort still glowing red-hot after completion of the brazing cycle. Brazing was a considerable challenge, as the process had to accommodate the difference in thermal expansion coefficient between the A286 tubes and the 718 shell. In Europe, the expansion nozzles for the Vulcain engine are made by Volvo to a unique tubular design. The Inconel 600 tubes have a square cross section; the tubes are wrapped in a spiral pattern to obviate the need for tapering. Instead of brazing, tubes are joined by gas-tungsten arc (GTA) fillet welds (Figs 12.25

Fig. 12.23 F-1 rocket engine. Combination of a tubular and a sandwich-construction. (Courtesy NASA photo archives.)

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Fig. 12.24

Fig. 12.25

Brazing the F-1 nozzle. (Courtesy NASA photo archives.)

Assembly of a Volvo Aero tubular nozzle. (Courtesy of Volvo Aero.)

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Fig. 12.26

Completed Volvo Aero tubular nozzle. (Courtesy of Volvo Aero.)

and 12.26). A newer Volvo design for a nozzle extension to the uprated Vulcain engine involves sandwich construction in which the inner liner is a milled channel configuration that is closed out by a cover sheet stitched in place by burn-through laser welds (Figs 12.27 and 12.28). The material of construction is reported to be stainless steel [9, 13, 14]. Solid metallic nozzles and nozzle extensions also are used. Some examples are the European Aestus upper-stage nozzle extension made of Haynes 25 sheet metal

Fig. 12.27 Schematic of the Volvo welded square tube nozzle design. (Courtesy of Volvo Aero.)

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Fig. 12.28

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Volvo laser-welded sandwich expansion nozzle. (Courtesy of Volvo Aero.)

and the Apollo Service Propulsion engine nozzle is made of a welded assembly combination of niobium alloy (C103) expansion nozzle and titanium alloy (Ti-6Al-4V) nozzle extension. (Figs 12.29 and 12.30). F-1 engine used a nozzle extension of Hastelloy-C sheet in a type of welded, double-wall configuration in which cooling was obtained by ducting warm turbine exhaust gas in a dumpcooling arrangement. Small thrusters have been made of beryllium because of its unique heat capacity and thermal radiation characteristics (Fig. 12.31). Carbon–carbon nozzle extensions are used on engines such as the RL-10B2 (Fig. 12.32) and the Vinci upper-stage engines. Ablative nozzles are exemplified

Fig. 12.29 Apollo service propulsion module engine. (Photograph taken at the Huntsville Space Museum.)

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Fig. 12.30 AJ-10 Apollo service propulsion module engine nozzle and nozzle extension. (Photograph taken at the Huntsville Space Museum.) by the glass-phenolic nozzle extension of the RS-68 booster engine and reaction control system (RCS) and orbital maneuvering system (OMS) engines such as the SE-6 for the Gemini program and SE-8 for the Apollo Command Module (Fig. 12.33). Typical ablative materials are carbon cloth phenolic, glass cloth phenolic, and silica cloth phenolic (Fig. 12.33 ). Fabrication of ablatives is similar; for example, the carbon cloth is impregnated with a phenolic resin and carbon filler and then oven cured [15]. In the late 1990s, an ablative chamber/ nozzle was developed for the Fastrac engine. The chamber/nozzle was built as one piece with an ablative liner and a composite overwrap. The ablative liner is tape wrapped using silica phenolic tape and

Fig. 12.31 Small thruster with beryllium combustion chamber and high-temperature alloy nozzle. (Courtesy NASA photo archives.)

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Fig. 12.32 RL-10 engine with extendable carbon–carbon nozzle extension. (Courtesy NASA photo archives.)

then cured on the tape-wrap mandrel. After completing the liner, it is machined, and the injector attachment flange then is bonded to the exterior of the liner. A carbon-epoxy overwrap is then filament wound around the entire structure and cured. The nozzle is then machined to the correct length. Interface hardware fabricated from 304 stainless steel is bonded to the composite nozzle with mechanical locks in high-temperature and high-stress locations (Fig. 12.34) [10, 16].

Fig. 12.33 Cross section of a small ablative reaction-control (RCS) engine. (Courtesy NASA photo archives.)

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Fig. 12.34 Fastrac engine hot-fire test at NASA Marshall Space Flight Center. (Source: NASA/MSFC.)

12.6 TURBOPUMP MATERIALS 12.6.1

INTRODUCTION

Turbine-driven pumps became necessary when the rocket engine requirements (thrust level, mission duration, ability to be throttled to various power levels, etc.) exceeded the capability of pressurized propellant feed tank systems. The word turbopump was coined to describe these pump/turbine components. The common theme of all turbopumps is that the power to pump the propellant is provided by passing a fluid, usually (but not always) an expanding gas and usually (but not always) at elevated temperatures, through a series of turbine blades to impart rotational torque to power either a centrifugal or axial flow pump. The goal of all turbopumps is to produce the required pumping power and flow with as lightweight and compact a package as possible while maintaining adequate structural margin over the anticipated mission duty cycle [17, 18]. The methods used to develop the turbine drive fluid depend on the type of system used, which has been described previously. However, the propellants (fuel and oxidizer) being “pumped” are necessarily liquid so that the pumping elements can increase the pressure and flow rates of the propellant to levels required by the engine system. The starting pressures are low, usually dependent on the vehicle’s propellant storage tank volume, but must be raised to sufficiently support the required combustion levels after being forced through various ducting arrangements throughout the engine system. The propellants end up at the main

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injector before being directed into the main combustion chamber to be ignited, producing the smoke and fire rockets are famous for [17]. The materials used in the turbine end of the turbopumps are dependent on the chemical composition of the drive gas used, its pressure, temperature, and flow rate. All will influence how the media will interact with the metallic and nonmetallic mechanical components along the way. Also, by the very nature of a rotating pump, there will be fatigue life implications that must be considered in addition to expected (and unexpected) material strength requirements. Likewise, the fluids being pumped, the oxidizer and fuel, present their own set of challenges dependent on the materials used, which are likewise dependent on the pressure, temperature, flow rates, chemical stability or lack thereof, and so on [6, 12]. Turbine-driven pumps may be directly coupled to the pumping elements using a single shaft or may be indirectly coupled with the turbine-generated power directed to the pumping elements using a gear speed reduction transmission arrangement as was used on the RL10 and the Atlas MA-2, MA-3, and MA-5 engine series. Gear arrangements add complexity to a system, and lubrication issues are common. Finally, the type of rocket system being constructed will impact the materials selection decisions greatly. Rocket engine systems are usually separated into two general groups: reusable and expendable. A reusable rocket engine system, for instance, is intended to be used over and over again, thus greatly lengthening the required life of the hardware. Being reused again and again subjects the engine system to a potentially large number of “hot fires” or starts that tend to load the hardware to the highest stress levels. Also, the increased hot-fire duration of multiple missions will subject the hardware to longer fatigue-related damage over the engine’s useful life. Although individual high-wear components may be changed out after each flight, the major components will be reused with little more than a cleanup and inspection. The SSME is an example of such an engine system. An expendable rocket system is intended to be used for a single mission with a minimum of individual number of starts before actual flight. And after the mission is completed, the spent first-stage or “booster” engines are usually jettisoned into the ocean or a remote land, never to be used again. These engines are usually designed to take advantage of the minimum mission life expectancies by operating higher on the material’s strength and fatigue life capability and may allow higher levels of propellant interaction than what would be considered prudent for a reusable system. This type of system would be appropriate for ballistic missiles as well as satellite or cargo delivery rockets, such as the current Atlas, Delta, H-II, Long March, and Ariane families of rocket systems. In considering materials selection, there are two distinct portions of the turbopump: the pump end and the turbine end. Usually, the materials selection considerations for these are quite different. The pump operating temperatures vary from room temperature to cryogenic temperatures. Environmental effects of the fluids being pumped must be considered, but often the effects are limited to

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temperature alone, as the cryogenic temperatures often ameliorate environmental effects by reducing chemical reactivity. On the other hand, high-pressure oxygen presents the unique danger of an oxygen fire if a source of ignition is present. For this reason, the oxygen compatibility of the materials employed must be considered [6]. The turbine end of the turbopump operates at temperatures that can vary from cryogenic (for turbines driven by propellants in their liquid state) to significantly elevated temperatures. The reactivity of propellant gases often can be exacerbated by elevated operating temperatures. This summary is intended to give an overview of some of the different materials that are used for a variety of turbopump engine systems with various types of drive systems pumping various types of propellants [6, 17, 18].

12.6.2

SAMPLE TURBOPUMP CONFIGURATIONS

To get a better feel as to how turbopump designs can vary, several representative pumps from different eras are shown with a short description of features, starting with Fig. 12.36. 12.6.2.1

ROBERT GODDARD’S TURBOPUMPS: LATE 1930S

Goddard was the first to use turbopumps to deliver propellants (gasoline and LOX) to a rocket engine combustion chamber. Goddard used impellers made of brass and steel. Aluminum alloy 2017 (4Cu-0.5Mn-0.5Mg) was used for turbines and pump housings. Turbines were machined from 2017 and driven by gaseous oxygen. Dowmetal (Magnesium-Al-Mn) was also used for pump housings. The risk of using a magnesium alloy in an oxygen pump was unknown in Goddard’s day and may have impacted the reliability of his systems. 12.6.2.2

A4 (V2) TURBOPUMP: 1940S DESIGN

As it was the world’s first ballistic missile, the turbopumps for the V-2 were initially an unknown quantity. The engine required pumps that could quickly start and rapidly develop a high head. To the surprise of the German engineers, fire pumps fit their requirements, and the pump design progressed without much difficulty, although the incorporation of a turbine drive proved to be a challenge. The turbopump for the V-2 engine was a rather unique design, having a direct-drive, two stage turbine situated between the fuel and oxidizer pump with pump inlets entering from the center of the pump instead of from the ends (Fig. 12.35). The V-2 turbopump developed 665 hp at 3800 rpm. The turbine drive gas was produced in a separate system and consisted of steam created by the catalytic decomposition of hydrogen peroxide, using potassium permanganate as the catalyst. Pump impellers were of the shrouded centrifugalflow type. The pump housings and impellers were castings of an aluminum 13% silicon–0.3 Mn alloy, generally called Silumin, which is a family of aluminum

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Turbopump for the A-4 (V-2) missile. (Huntsville Space Museum Archives.)

casting alloys that were commonly used in 1930–1940 (but still are in use today) [2, 8, 17]. The shaft and steam manifold were low-alloy steel, whereas the turbine disk was an aluminum alloy having the composition Al-2Mg-1.4Mn. The turbine blades were aluminum permanent mold castings of 13X alloy (Al-13%Si). In those days, permanent mold casting was performed in air, and so each turbine blade developed a thin layer of oxide on its surface. This thin oxide layer had the effect of providing a sort of “ceramic” coating that protected each blade from the hot steam (400þ 8F) for the duration of the powered portion of the flight (56 s). This same turbine blade feature was copied for the Rocketdyne A6/A7 Redstone rocket engine turbine. 12.6.2.3

REDSTONE ENGINE A6/A7 TURBOPUMP: EARLY 1950S DESIGN

The turbopump for the NAA A6 engine, which became the Redstone engine, can be looked upon as an improved version of the German V-2 turbopump. The turbine remained located between the fuel and oxidizer pumps, and the turbine drive gas remained steam from decomposed hydrogen peroxide, except a silvercoated screen was used as the catalyst. Also, the permanent mold 13X cast aluminum turbine blades on the V-2, with their oxide layer coating, were retained. The turbine disk alloy was changed to 7075 aluminum. The pump inlets were reversed

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Cutaway of a XLR43-NA-1 engine turbopump. (Courtesy NASA image archives.)

to a more logical configuration at the ends of each pump, and inducers were added to increase performance. Fig. 12.36 is a cutaway of a XLR43-NA-1 engine turbopump, which is the same design as used on the A6 and A7 Redstone engines. Impellers and housings were changed to 356-T6 cast aluminum alloy, a significantly stronger material than the Silumin used in the V-2 pumps. The design changes, combined with higher strength materials, provided a lighter yet higher performing turbopump. The A7 turbopump produced 836 hp at 4840 rpm vs the 665 hp of the V-2 [5]. 12.6.2.4

ATLAS TURBOPUMP: LATE 1950s DESIGN

With the advent of high-performance rocket engines, the demands on materials increased dramatically. Gas generators now provided hot gases for turbine drives, requiring materials with good high-temperature properties. The Mark 3 was the first high-performance turbopump in the Western world (Fig. 12.37). Originally designed for the Atlas ICBM, it has been used, in various modified forms, for the engines of the Atlas, Thor, Jupiter, Delta I, Delta II, and Delta III rockets. The Mark 3 featured fuel and oxidizer pumps on a common shaft, driven through four gears by a two-stage turbine [3, 17]. It produced 2550 hp

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at 6700 rpm. Turbine speed was 30,000 rpm. For the first time, high-purity kerosene was used for the fuel along with LOX oxidizer. The high-strength, cast aluminum alloy Tens-50 (the precursor to A357) was used for the impellers and pump housings. Inducers were 7075 aluminum forgings. The common pump shaft was 4340 high-strength, low-alloy steel. Turbine disks were forged 16-25-6, an early austenitic high-temperature alloy (Fe-16Cr-25Ni-6Mo). Turbine blades were investment castings of Stellite 21 (Co-27Cr-5Mo-2.5Ni) and were welded to the turbine disks. Gears were 9310 alloy steel, carburized. 12.6.2.5

F1 TURBOPUMP (MARK 10): LATE 1950S DESIGN

The 1.5 million-lb thrust F-1 engine for the Saturn V launch vehicle required new thinking for turbopump design (Fig. 12.38). The extremely high power requirements of the Mark 10 pump would place an impossible load on any gear train [17]. Gearing was abandoned in favor of direct drive, with the turbine coupled by a single shaft to both the fuel (RP-1) and oxidizer (LOX) pumps. Mark-10 Turbopump for the F-1 Engine generated 53,000 hp at 5500 rpm. Flow rates were 25,000 gpm for LOX and 15,600 gpm for RP-1. The Mark 10 was the first turbopump where many parts were too large for a workman to lift, necessitating the use of cranes. The Tens-50 aluminum pump housings and impellers were cast in specially designed, highly chilled molds

Fig. 12.37 Mark 3 turbopump used on the Atlas, Thor, Jupiter, and Delta launch vehicles. (Courtesy NASA image archives.)

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Fig. 12.38

Mark-10 Turbopump for the F-1 Engine. (Courtesy NASA image archives.)

that provided very rapid solidification rates to obtain never-before-seen strength and ductility in aluminum castings of that size. The turbine disks and manifold used the newly developed high-temperature alloy, Rene 41 (Ni-19Cr-12Co10Mo-3Ti-1.5Al-.12C). This alloy produced challenges in welding, which eventually were overcome by postweld vacuum heat treatment. 12.6.2.6

HYDROGEN TURBOPUMPS (RL-10 AND MARK 15): LATE 1950s TO EARLY 1960s DESIGNS

The RL-10 (Fig. 12.39) was the world’s first hydrogen-fueled rocket engine. It was an expander cycle engine with a moderate thrust of 15,000 lbs. Because it was an expander cycle, driven by the expansion of gasified hydrogen, turbine temperatures were near room temperature and pressures were relatively low. Therefore, aluminum alloys could be used for the turbine as well as the pumps. The RL-10 is remarkable in that it remains in production to this day, a time span of more than 55 years. In contrast, the world’s second hydrogen-fueled engine, the J-2, was a highthrust (230,000 lbs.) rocket engine using a gas generator cycle. This necessitated unique design and the materials challenge of having to deal with both a –2538C/–4238F fluid and hot turbine gases. The Mark 15 fuel turbopump for the J-2 engine (Fig. 12.40) was an axial flow pump driven by a two-stage turbine [17]. It generated 7900 hp at 27,000 rpm and pumped 3000 gpm liquid hydrogen at 1238 psi discharge pressure. The choice of an axial flow pump was dictated as much by space limitations as by performance requirements. When the J-2 was designed, there were very few data on the mechanical properties of materials at liquid hydrogen temperatures (–2538C/–4238F). The selection of

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Monel K-500 was made because it had sufficient strength and, as an alloy of only copper and nickel, the metallurgical staff was certain that it would retain sufficient ductility at these cryogenic temperatures. The J-2 engine also was notable because it incorporated one of the first applications of alloy 718 in the form of forgings. Alloy 718 was selected for the fuel pump turbine disks as well as the injector back plate. Alloy 718 originally had been developed by INCO as a sheet alloy, and it was air melted. When in the form of sheet, inclusions and a certain degree of segregation were tolerable because the short transverse (thickness) direction is not relevant. However, in forgings, short transverse properties are very important, and it soon became evident that mechanical properties were substandard in this direction. The problem was solved by instituting processing changes which included double-vacuum melting and high-temperature homogenization.

Fig. 12.39 RL10A turbopump assembly. Source: NASA SP 8107 Turbopump Systems for Liquid Rocket Engines. (Courtesy NASA image archives.)

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Mark 15 fuel turbopump for the J-2 engine. (Courtesy NASA image archives.) SSME TURBOPUMP: 1970S DESIGN

The turbopumps for the Space Shuttle Main Engine (SSME) presented significant challenges from design and materials standpoints. The turbopumps were to be reused many times over, pressures were very high (6000–8000 psi), and turbines had to endure hydrogen-rich combustion products at temperatures on the order of 8408C (15508F). Figure 12.41 shows an early version of the high pressure hydrogen turbopump. It had a two stage turbine that generated 70,000 hp at 38,000 rpm. The three stage pump had a discharge pressure of 6800 psi. Hydrogen embrittlement became an issue for the turbines, and the high pressures in the oxygen pump presented a serious potential for oxygen fires if an ignition source, such as a severe rub, was present [6, 12]. 12.6.2.8

SSME ALTERNATE TURBOPUMP: 1990S DESIGN

The late 1980s saw the start of a new SSME high-pressure oxidizer turbopump (HPOTP) design. This new turbopump was specifically designed to improve flight life and eliminate critical vulnerabilities in the original HPOTP. A heavier allowable weight permitted a much stronger, stiffer rotor system, as well as more robust pump and turbine housings. Material advances in casting technologies

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permitted the incorporation of complex, fine-grained castings that allowed the elimination of 293 welds within the turbopump, including all 250 welds for which there was no root side access for inspection at fabrication. Silicon nitride rolling elements were used in the pump-end bearing, virtually eliminating bearing wear and fatigue concerns. Single-crystal alloy blades with thin-walled, hollow airfoils were incorporated to address blade cracking. This new HPOTP design first flew in 1995 [21]. In the mid-1990s, a redesigned high-pressure fuel turbopump (HPFTP) was restarted as part of the SSME Block II program. It incorporated improvements similar to those on the alternative (Block I) HPOTP. The new HPFTP design had no welds and used silicon nitride rolling elements in both of its bearings. Similar to the Block I HPOTP, single-crystal alloy turbine blades with thin airfoils essentially eliminated blade cracking in the Block II HPFTP. This new HPFTP design first flew in 2001 [21]. 12.6.2.9

TECHNOLOGY DEVELOPMENT TURBOPUMPS: LATE 1990S TO EARLY 2000S

Other turbopump technology programs in the mid to late 1990s and early 2000s focused on turbomachinery technology to reduce part count and improve overall design and manufacturing cycle time. The NASA Low Cost Boost Technology (LCBT) Project objective was to significantly lower the cost of access to space for small payloads. Two projects were funded to completion, producing hot-fire test results to address these challenges: the Fastrac engine turbopump and the

Fig. 12.41 archives.)

The original high-pressure fuel turbopump for the SSME. (Courtesy NASA image

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Bantam turbopump. These projects focused on adapting common manufacturing techniques and existing commercial, off-the-shelf hardware to the aerospace applications [20]. The MC-1 engine, a NASA-led in-house design (formerly known as the Fastrac 60 K), is a pump-fed liquid rocket engine with fixed thrust and gimballing capability. The engine was initially designed for the LCBT Project and small space vehicles. The engine burns a mixture of RP-1 hydrocarbon fuel and LOX propellants in a GG power cycle. The MC-1 Fastrac turbopump design stepped away from the traditional gearbox design and, like the F-1 turbopump, incorporated the oxidizer impeller, the fuel impeller, and the turbine on a single shaft [21]. The Bantam turbopump also was designed to meet the MC-1 engine requirements but took a different approach by using twin rotating shafts within a single housing. NASA contracted with Rocketdyne, which teamed with Barber Nichols Inc. to design, fabricate, and test the Bantam turbopump [22]. The U.S. Air Force, NASA, and leading aerospace industry contractors joined forces to develop the integrated powerhead demonstrator (IPD), a risk-reduction effort to develop engine technologies for a full-flow, hydrogenfueled, staged-combustion rocket engine in the 250,000-lb thrust class. The IPD engine employed dual preburners that provide both oxygen-rich and hydrogenrich gases for the staged-combustion engine. The IPD project was intended to address two major turbomachinery technological challenges, turbine life and bearing wear, which are traditional life limiters for rocket engine turbomachinery. The high-performance turbomachinery developed for the IPD demonstrator included hydrostatic bearings that fully support the rotor of both the fuel and oxidizer pump, instead of ball bearings or roller bearings that are typically used. The use of oxygen-rich steam to power the IPD oxygen turbopump is intended to dramatically increase the safety of engine system operation, limiting seal failure between the pump and the turbine that could leak extremely hot gases into the turbine and cause them to burn prematurely. The IPD program introduced materials to address oxygen-rich environments and hydrogen-rich environments and bearings to address hardware life issues experienced in other high power density rocket engine turbomachinery [23]. During the NASA Reusable Launch Vehicle (RLV) program in the late 1990s and early 2000s, materials technology projects were established to explore the use of aluminum (Al) and copper (Cu) metal matrix composites (MMCs) and carbon fiber-reinforced silicon carbide (C/SiC) ceramic matrix composites (CMCs). These technology efforts were started to address the program goals for reducing engine component weight and reducing hardware and operational costs. The Al MMCs were considered for turbopump housings applications. Housings for highperformance rocket engine turbopumps make up a very high percentage of a turbopump’s total weight. MMCs offered the potential benefits of high specific strength, tailoring properties, propellant compatibility, reducing cost, and producibility [24, 25].

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Fig. 12.42 Ceramic matrix composite blisk. (Courtesy NASA/ MSFC Turbopumps—2010.)

The C/SiC CMCs for turbopump applications were also investigated during the RLV program. The use of ceramic composites in rocket engine turbines was to address turbine life issues on reusable turbopump systems, to increase safety and reduce costs. One intriguing characteristic of the C/SiC CMC is its ability to withstand damage. In extensive testing, the ceramic blisk continued to perform normally despite a crack in one of the blades. The ceramic blisk also can withstand temperatures of 20008F (10938C), considerably higher than the 12008F (6498C) temperature limit for nickel-alloy turbines. A 7.6-in.-diam C/SiC CMC blisk (Fig. 12.42) was tested in a rocket engine turbopump repeatedly, accumulating data that will be used to predict CMC blisk life [26]. Recent advances in additive manufacturing have made it possible to make complex parts for rocket engine turbomachinery hardware. The idea is to make complex parts quicker at lower cost and be able to test hardware in a representative environment to drive down risk. In 2015, an LH2 turbopump designed using 3-D printed parts (Fig. 12.43) was hot-fire tested as a component and then as part of a breadboard engine. The LH2 turbopump is for a 35K lbf expander cycle engine. The 3-D printed turbopump has 45% fewer parts than similar pumps made with traditional welding and assembly techniques. The pump design speed is 90,000 rpm [27].

12.6.3

TURBINE DRIVE METHODS AND COMPONENTS

To drive the pump(s), high-pressure gas must be introduced to the turbine portion of the turbopump. How the gas is generated depends on the type of engine system being used. The three major methods are gas generator cycle, staged-combustion cycle, and expander cycle [9, 18]. Basically, gas is ducted to the turbine inlet portion of the pump. The gas is introduced to a pressure vessel housing that directs the flow, usually through a nozzle and into the turbine blades. Depending on the propellants used, various component materials may be used. Except for expander cycles, all must have a reasonably high strength at elevated operating temperatures (which could be as high as 15008F) and have a sufficient resistance to high and low cycle fatigue. And should hydrogen be the fuel and the hot gas fuel rich, a review of the selected materials’ resistance to hydrogen embrittlement is necessary. Also, some designs

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Fig. 12.43

3-D printed liquid hydrogen turbopump. (Source: NASA/MSFC.)

rely on an oxidizer-rich hot gas, which would be sensitive to catastrophic ignition depending on the temperature, pressure, and possibility of entrained particles (potential impact ignition sources). Examples of previously selected materials include housings composed of nickel-based superalloys, such as Inconel 625 and 718; precipitation hardening iron-based stainless steels, such as A286 and Incoloy 903; and a variety of cobaltbased alloys, such as Haynes 188. Depending on the need for light weight, the housing may be fabricated assemblies composed of many different parts usually welded together. If weight is not a major design driver, suitable castings (usually investment) may be employed. The penalty of added weight is offset by a reduction in the number of weldments, eliminating the quality issues and difficult inspection requirements that usually accompany a welded assembly. The turbine blades themselves may be separately cast as individual blades, and some may be welded onto the appropriate hub or even integral to the forged hub known as a blisk (short for bladed disc). Some blisks have been cast, but the preference is for forged components in high-stressed rotating applications. The RS27 engines used Stellite 21 cast blades welded on to a 16-25-6 stainless hub. The original SSME turbines used cast individual MAR-M-246 blades with fir tree bases secured onto Waspaloy hubs. The original MAR-M-246 directionally solidified

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turbine blades of the SSME proved to be susceptible to hydrogen embrittlement. These were replaced in the alternative turbopump by single-crystal PW1480 blades that proved much more resistant to hydrogen effects. Some systems have such high power loads that the individual blades may extract up to 600 hp each from the hot turbine gas. Also, designs may require multiple rows of blades to extract the necessary power from the available gas. Such systems employ interstage stators (stationary structures with aerodynamic vanes) to help direct the hot gas flow through the blades. For hydrogenfueled engines, such as the SSME, to protect the components from hydrogen environmental embrittlement damage and internal hydrogen embrittlement, coatings impermeable to hydrogen (gold or copper) or overlays (iron based) are used [6, 12].

12.6.4

PROPELLANT PUMPING ELEMENTS

The power provided by the turbine must be linked to the propellant pump portions of the turbopump. As mentioned, these may be directly coupled to the pumps using a single shaft connecting the turbine to the pump. The F-1 engine used 4340 steel alloy forged shafts for this purpose. The F-1 application was a little less challenging because the fuel is RP-1, an ambient temperature kerosene type of fuel, hence the 4340 steel shaft is fully suitable. Shafts of any nature require rotational and axial support to maintain proper alignment during operation. Bearings of various configurations such as roller bearings, ball bearings, and hydrostatic bearings are all employed. Materials for the respective races and rolling elements are usually extremely hard: 440C stainless steels or an earlier 52100 1C-1.5Cr bearing steel have been used; more recently, Cronidur 30 stainless steel (15Cr-.5Ni-.4N2-.3C) has also been used for bearings in SSME and RS68 pumps. The rolling elements are often made of 440C high-hardness stainless steel. Some pumps have begun to employ bearings using ceramic (silicon nitride) rolling elements. If the connection is an indirect one whereby a gear set is used, the gears may be low-carbon steel (such as 9310) with a carburized layer (case) to produce highhardness, wear-resistant gear teeth surfaces while maintaining a tough core. During operation, lubrication is an issue because many systems use the propellants themselves to provide the necessary lubrication, which usually is not very effective. The pumped propellant’s density determines the pump’s ability to do work on the fluid to increase its pressure and flow rate; a less dense fluid, with hydrogen being the least dense, requires additional effort to get the needed pump discharge levels. Fuels include such materials as gasoline, alcohol, hydrazine, RP-1, liquid hydrogen, and others [3, 6]. Gasoline and its relative kerosene are somewhat benign fuels from a materials selection perspective because they are relatively warm (ambient temperatures), whereas liquid hydrogen is cryogenic (–4238F),

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which may eliminate some materials’ ductility and promote embrittlement as well, leading to completely unexpected failures. Oxidizers, the other half of the combustion equation, also vary widely, with such materials as hydrogen peroxide, nitrogen tetroxide, nitric oxide, and LOX providing free oxygen to support the combustion process [3]. As with the fuels, oxidizers vary in their reactivity, and LOX is cryogenic (–2908F), which requires the consideration of low temperature ductility. The high reactivity of LOX makes unexpected ignition a concern if an energy source of ignition is present, such as a severe rub of an impeller or an impact from foreign object debris. Propellants are delivered to the pump inlet housing from the tanks at relatively low pressure and increase the pressure using either axial or centrifugal (radial) pumping elements, the latter being far more common. In most turbopumps, the propellant is flowed through an inducer for a moderate pressure increase before flowing into the impeller for a bigger boost in pressure. Inducers guarantee a solid head of propellant to ensure that cavitation does not occur in the impellers. If the propellant is dense enough, a single-stage impeller may be all that is required. If not, such as for liquid hydrogen, multiple stages of impellers will be required with diffusers (also known as crossovers) in between, before exiting the pump at the required pressure and flow rates. High specific strength is a key property for impellers, as it determines the maximum rotational tip speed, and therefore the pumping efficiency, that can be attained. The materials involved for the various components have to be compatible with the respective propellants while providing the necessary function. Because weight is a key issue with any rocket engine, high-specific strength alloys of aluminum and titanium are frequently used, which would not be suitable for high-temperature turbine components. For impellers, high specific strength is a key design criterion, as it governs the maximum tip speed that can be attained. Housings have been produced from lightweight aluminum alloys, including Tens-50 investment castings for the F-1 engine, and various nickel-based superalloys, including centrifugally cast alloy 625 for the RS68 fuel turbopump. Other materials, such as investment cased alloy 718, have also been used. Inducers have been produced using nickel-copper alloys such as K-500 Monel forgings for LOX pumps (to prevent ignition) or titanium alloys Ti-5Al–2.5Sn forgings for the fuel pump. Impellers can be composed of both forged and cast aluminum alloys such as A357, which was mentioned previously [8]. The SSME fuel turbopump uses three titanium alloy (Ti-5Al-2.5SnELI) forged impellers. Between each impeller is a (usually cast) diffuser or crossover, which takes the high-pressure impeller discharge at the impeller outside diameter and routes (and slows) the hydrogen back near the shaft so that it will engage properly with the next stage for the next pressure boost. SSME used A357 cast aluminum for this purpose, whereas the latest pumps use the latest, beryllium-free version of this alloy, F-357.

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12.7 VALVES Propellant valves are a key to controlling the flow of propellants and gases to various parts of a rocket engine. They may control the flow of propellants to main thrust chambers and gas generators or preburners. Other functions include control of bleed flow, spark igniter flow, and various thermal conditionings. Although most valves are usually of the two-position (open–closed), normally closed design, they may include an intermediate opening position to meet specific sequencing requirements. For thrust- or mixture-control purposes, a continuously variable opening position may be required on some propellant system valves. Prime design considerations for propellant valves include the following objectives: propellant compatibility, structural integrity, no leakage of propellant through the valve when closed, proper actuating time during opening and closing in accordance with the requirements of the control system, minimum pressure drop, and fulfilment of fail-safe and/or fail-operational system requirements [1]. Among the great variety of propellant valve types available, each has certain characteristics that make it suitable for a specific application. Frequently used propellant valves can be classified according to their design configurations: butterfly, ball, poppet, Venturi, or gate. Valves are actuated by mechanical actuators, which in turn are driven hydraulically, pneumatically, electrically, or even explosively. A discussion of the detailed design, operation, and advantages of each valve type is beyond the scope of this chapter. However, certain common considerations from a materials standpoint are shared by all valves. One consideration is the ability to reliably seal when closed. Sealing is usually achieved by a hard surface going into a softer one, such as a metal on a polymeric or elastomeric material (e.g., Teflon or Kel-F) or a high-strength metal (sometimes with a hard coating, such as tungsten carbide) into a soft metal, such as a copper alloy. If the valve is operating in oxygen service, extreme care must be taken to ensure that both mating surfaces are oxygen compatible. Examples of ball, butterfly, and poppet valves are given in Figs 12.44, 12.45, and 12.46, respectively. Recent work with 3-D printing has resulted in valves being produced with fewer parts and reduced cost than using conventional materials. Hot-fire testing of LOX and liquid hydrogen has shown positive results [27]. Over the years, a variety of materials have been used in valves. A favorite material for valve housing has been aluminum alloy, either cast or wrought. Cast valve housings have typically been of either A356 or A357 high-strength aluminum. For wrought housings, forged 7075 or 2024 are common. When using aluminum alloys, care must be taken to use only heat treatments that are not susceptible to stress-corrosion cracking. For higher strength needs, nickelbase alloys such as 625 and 718 or titanium alloys such as Ti-6Al-4V, or Ti-5Al-2.5Sn ELI for liquid hydrogen, have all been used. In the SSME main

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Cross section of the SSME main fuel valve. (Courtesy NASA image archives.)

fuel valve (Fig. 12.44) housing is titanium-5Al-2.5Sn alloy, ball/shaft is 718, and main ball seal is Kel-F. A perennial favorite for valve internals such as poppets has been precipitation hardening stainless steels, 17-4Ph and 15-5Ph. However, their susceptibility to hydrogen environment embrittlement and poor toughness at cryogenic temperatures limits the use of Ph stainless under these conditions.

Fig. 12.45

Typical butterfly valve. (Courtesy NASA photo archives.)

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Fig. 12.46

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Typical poppet valve. (Courtesy NASA image archives.)

The bottom line for materials selection for valves is to use whatever material makes sense from a strength, compatibility, and cost standpoint.

12.8 LINES AND DUCTS There is a large variety of lines and ducts on a liquid rocket engine: high- and lowpressure propellant ducts, hydraulic lines, lines to carry purge gases, sensor lines, and drain lines. Like valve bodies, a number of materials has been used for lines

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and ducts, including stainless steels, aluminum alloys, nickel-base alloys, iron-base superalloys, and even titanium alloys. Probably the favorite material for lines and ducts has been austenitic stainless steels, due to their weldability, ready formability, and reasonably good corrosion resistance. Staged-combustion engines, with their high-combustion chamber pressures fed by high-pressure pumps, require propellant ducting with a high specific strength in order to limit engine weight. The high-strength nickel-base 718 alloy is often chosen for these applications. A notable example of how a combination of a design change and material change can result in a significant cost reduction occurred during the F-1 engine program. This was the successful replacement of the articulated stainless steel propellant ducts with rigid aluminum alloy ducts. The four original stainless steel ducts, which cost in excess of $100,000, were replaced with four 6061 aluminum ducts costing only $5000. The actions to accomplish this included a redesign of the ducts, innovative materials processes, and creative manufacturing fit-up procedures. As evident in the J-2X photograph (Fig. 12.47), lines and ducts take on compact complex shapes to package the engine design. These complex shapes challenge traditional manufacturing techniques such as casting, forging, roll

Fig. 12.47

J-2X engine. (NASA/SSC photo.)

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Fig. 12.48

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J-2X DMLS gas generator discharge duct. (NASA/MSFC photo.)

forming, spinning, and welding. With an increased emphasis on cost reduction and improved manufacturing time, rapid manufacturing techniques such as direct metal laser sintering (DMLS) are being adapted and evaluated for fabrication of complex parts such as the J-2X GG exhaust duct (Fig. 12.48). The GG exhaust duct was made from alloy 625 using DMLS at a reduced cost and schedule over one made by conventional means. This duct was hot-fire tested and nondestructively inspected after hot-fire testing with positive results [28].

12.9 SUMMARY The importance of materials in liquid rocket engines is fairly obvious and will continue to be so in the future. With the present emphasis on commercial space, cost has become an important parameter and is now often considered as a variable of equal importance with weight and performance. Therefore, materials and fabrication costs have become prime considerations in any materials selection decision. This has resulted in considerable interest in additive manufacturing (AM) as a means to reduce both fabrication costs and the amount of material used to produce a component. The projected cost and schedule savings resulting from AM are extremely attractive. However, as with any new, potentially revolutionary process, many issues need to be addressed, such as materials characterization, development of reliable process parameters, an understanding of the metallurgy of the deposits, and quality control procedures. Metals and metal alloys remain the primary materials of construction. Ceramics, ceramic composites, and polymeric composites have found limited usage

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and hold promise for the future, but, although offering much promise, these nonmetallic materials have not yet reached their potential for widespread application in rocket engines.

REFERENCES [1] North American Aviation, An Introduction to Rocket Missile Propulsion, North American Aviation, 1958. Canoga Park, CA [2] Sutton, G. P., History of Liquid Propellant Rocket Engines, AIAA, Reston, VA, 2006. [3] Hutzel, D. K., and Huang, D. H., Modern Engineering for Design of Liquid Propellant Rocket Engines, Progress in Astronautics and Aeronautics, Vol. 147, AIAA, Washington, DC, 1992. [4] Durant, F. C., “Robert Goddard: Accomplishments of the Roswell Years, 1930– 1941,” AAS 89-268, American Astronautical Society History Series, Chap. 19, 1986. [5] Ezell, W. F., and Mitchell, J. K., “Engine One,” Threshold, Summer 1991, pp. 53–63. [6] Shoemaker, M., and Somerville, J., “A Material Difference,” Threshold, No. 4, Spring 1989, pp. 29–33. [7] Sutton, G. P., and Biblarz, O., Rocket Propulsion Elements, 8th ed., Wiley, New York, 2010. [8] Boen, B. (ed.), “Hot-Fire Tests Show 3-D Printed Rockets Parts Rival Traditionally Manufactured Parts,” NASA, http://www.nasa.gov/exploration/systems/sls/ 3dprinting.html [retrieved 24 July 2013]. [9] Stenholm, T., “The Development of the Vulcain Nozzle Extension,” 34th Joint Propulsion Conference, AIAA-98-4013, AIAA, Reston, VA, 1998. [10] Peters, W., Rogers, P., Lawrence, T., Davis, D., D’agostino, M., and Brown, A., “FASTRAC Nozzle Design, Performance and Development,” AIAA-2000-3397, AIAA, Reston, VA, 2000. [11] Ball, I. M., “Construction of a Replica of Robert H. Goddard’s First Successful Liquid- Propellant Rocket,” AAS 92-339, American Astronautical Society History Series, Chap. 14, 1993. [12] Lewis, J., “Materials and Processes for Space Shuttle’s Engines,” Metal Progress, Vol. 107, March 1975, p. 4. [13] Haggander, J., “Current Status of Laser Welded Sandwich Nozzle,” 38th Joint Propulsion Conference, AIAA-2002-4146, AIAA, Reston, VA, 2002. [14] Ryden, R., “The Volvo Aero Laser Welded Sandwich Nozzle for Booster Engines,” 55th International Astronautical Conference, IAC-04-IAC-S.3.07, Canadian Aeronautics and Space Institute, Kanata, ON, Canada, 2004. [15] NASA, “Application of Ablative Composites to Nozzles for Reusable Solid Rocket Motors,” NASA Practice No. PD-ED-1218, NASA Marshall Space Flight Center, Huntsville, AL. [16] NASA, “Turbopump Systems for Liquid Rocket Engines,” NASA Space Vehicle Design Criteria (Chemical Propulsion), NASA SP-8107, Aug. 1974. [17] Stangeland, M., “Turbopumps for Liquid Rocket Engines,” Threshold, No. 3, Summer 1988, pp. 35–43.

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[18] Sutton, G. P., “Turbopumps, a Historical Perspective,” 42nd AIAA/ASME/SAE/ ASEE Joint Propulsion Conference, AIAA 2006-5033, AIAA, Reston, VA, pp. 3–4. [19] Van Hooser, K. P., and Bradley, D. P., “Space Shuttle Main Engine: The Relentless Pursuit of Improvement,” AIAA SPACE 2011 Conference, AIAA, Reston, VA, 2011. [20] NASA, “Lower Cost Is Goal of Bantam Technology,” Marshall Star, Vol. 37, No. 27, 1997, pp. 1–3. [21] Ballard, R. O., and Olive, T., “Development Status of the NASA MC-1 (Fastrac) Engine,” AIAA 2000-3898, AIAA, Reston, VA, 2000. [22] Barber-Nichols, “Rocket Engine Turbopumps” [online brochure], http://www. barber-nichols.com/products/rocket-engine-turbopumps [retrieved March 31, 2018]. [23] NASA Facts, “Next Generation Propulsion Technology: Integrated Powerhead Demonstrator, FS-2005-01-05-MSFC, Jan. 2005. [24] Lee, J., and Elam, S., “Development of Metal Matrix Composites for NASA’s Advanced Propulsion Systems,” Proceedings of the 4th Conference on Aerospace Materials, Processes, and Environmental Technology, NASA/CP-2001-210427, edited by D. E. Griffin, and D. Cross Stanley, NASA Marshall Space Flight Center, Huntsville, AL, 2001. [25] Effinger, M., Clinton, R. G., Dennis, J., Elam, S., Genge, G., Kiser, J. D., Eckel, A., Jaskowiak, M., and Land, J., “Fabrication and Testing of Ceramic Matrix Composite Propulsion Components,” Proceedings of the 4th Conference on Aerospace Materials, Processes, and Environmental Technology, NASA/CP-2001-210427, edited by D. E. Griffin, and D. Cross Stanley, NASA Marshall Space Flight Center, Huntsville, AL, 2001. [26] NASA, “Ceramic Matrix Composite Turbine Disks,” Advanced Space Transportation Technology Summary, Pub 8-1279, FS-2001-04-76-MSFC, http:// www.nasa.gov/centers/marshall/pdf/100395main_ceramic_matrix.pdf [retrieved March 31, 2018]. [27] Mohon, L. (ed.), “Piece by Piece: NASA Team Moves Closer to Building a 3-D Printed Rocket Engine,” NASA News Release, 17 Dec. 2015. [28] Betts, E. M., Eddleman, D. E., Reynolds, D. C., and Hardin, N. A., “Using Innovative Technologies for Manufacturing Rocket Engine Hardware,” JANNAF 6th Liquid Propulsion Conference, Distribution Statement A, 7 Dec. 2011.

CHAPTER 13

Advanced Materials for In-Space Propulsion Les Johnson and Tiffany Lockett NASA Marshall Space Flight Center, Huntsville, Alabama

13.1 INTRODUCTION In-space propulsion begins after separation from the launch vehicle and continues with performing the functions of primary propulsion, reaction control, station keeping, precision pointing, and orbital maneuvering. The main engines used in space provide the primary propulsive force for orbit transfer, planetary trajectories, and extraplanetary landing and ascent. The reaction control and orbital maneuvering systems provide the propulsive force for orbit maintenance, position control, station keeping, and spacecraft attitude control. Advanced in-space propulsion technologies enable much more effective exploration of our solar system and permit mission designers to plan missions with greater reliability and safety. With a wide range of possible missions and candidate propulsion technologies, the selection of a few technologies to meet the requirements of all future missions is a difficult one. A better approach is to provide a portfolio of propulsion technologies that will provide optimum solutions for a diverse set of missions and destinations, and each will have its own, likely unique, set of advanced materials requirements. A large fraction of the rocket engines in use today are chemical rockets; that is, they obtain the energy needed to generate thrust by chemical reactions to create a hot gas that is expanded to produce thrust. A significant limitation of chemical propulsion is that it has a low specific impulse (Isp, thrust per mass flow rate of propellant) relative to more advanced technologies. Historically, these propellants have not been applied beyond upper stages due to issues with storage, transfer, and control in space. Concepts for advanced propulsion technologies, such as electric propulsion, are commonly used for station keeping on commercial communications satellites and for prime in-space propulsion on some scientific missions because they have significantly higher specific impulse. The higher specific impulse can greatly lighten the mission propellant mass requirement; however, electric propulsion systems generally have very small values of thrust and therefore must be operated for long durations to provide the total impulse required This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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by a mission. In this chapter, several advanced propulsion technologies are described that offer significantly better performance than that achievable with chemical propulsion. In-space propulsion represents technologies that can significantly improve a number of critical metrics. Space exploration requires reaching a destination safely and quickly in an affordable manner with enough mass for consumables and scientific discovery. The simple act of “getting there” requires the employment of an in-space propulsion system, and the other metrics are modifiers to this fundamental action. In almost all cases, new materials are required.

13.2 CHEMICAL PROPULSION (LIQUID STORABLE AND LIQUID CRYOGENIC) Harold Gerrish NASA Marshall Space Flight Center, Huntsville, Alabama

13.2.1

INTRODUCTION

Chemical propulsion systems have been used for many years for in-space propulsion. Chemical propulsion systems are used for orbital velocity change (DV), ascent/descent, guidance, navigation, and control (otherwise known as reaction control system). Most chemical in-space propulsion systems use liquid propellant. Common operating requirements include multiple starts, starting in a cold vacuum of space, engine throttling, and high area ratio nozzles for higher specific impulse. In addition, the thrust chamber will be exposed to space environment effects (e.g., vacuum, galactic cosmic radiation, thermal heating or cooling, micrometeoroid orbit debris). Large engines usually have regenerative cooled thrust chambers, whereas small engines are usually radiative cooled. All chemical in-space propulsion systems provide thrust by expelling the propellant through a de Laval nozzle. Figure 13.1 shows the pressure balance. The integrated pressures around the thrust chamber control volume result in a total thrust in the direction of motion. Depending on applications, thrust requirements can

Fig. 13.1

Pressure balance on the chamber and nozzle wall [1].

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Fig. 13.2 Typical pressures (P), temperatures (T), flow velocities (V), and Mach number (M) in a de Laval Nozzle [2]. vary from less than a pound to thousands of pounds. Figure 13.2 shows how the static temperatures and pressures decrease throughout the nozzle while the propellant velocities and Mach numbers (.1) increase. High exit velocity means high specific impulse and is dominated by the stagnation temperature and propellant average molecular weight. High area ratio nozzles (.150:1) allow more surface area for thrust and a velocity increase that results in a higher specific impulse. Hypergolic bipropellant systems commonly use nitrogen tetroxide (NTO) as an oxidizer and monomethylhydrazine (MMH) as a fuel. Hypergolic means the burn occurs when the propellants come in contact with each other. These propellants are space storable for long-duration missions, but they are both toxic and have material compatibility challenges. They can be used for the main DV and/or the reaction control system (RCS) burns and are commonly pressure-fed with helium. Systems like this have been used on Apollo’s service module and the Space Shuttle’s Orbital Maneuvering System, and are planned for use on Orion’s service module. The most common monopropellant systems use hydrazine as the storable propellant. Flowing hydrazine over a catalyst initiates the decomposition reaction for energy release. Thus, an oxidizer is not required. For small spacecraft, this system can be used for both DV and RCS burns. They are pressure fed with helium using a tank diaphragm. The Orion crew module is currently slated to use monopropellant thrusters for reentry and landing. Commercial and NASA science missions also use a combination of the two aforementioned systems, referred to as a dual-mode system where the DV is provided with hypergolic propellants (e.g., hydrazine and NTO) and RCS maneuvers are provided with the monopropellant fuel only. Advantages include higher Isp for the DV mission phase and higher reliability for the RCS function. Cryogenic bipropellant systems are nontoxic, provide specific energy, and can use liquid hydrogen or liquid methane as the fuel and liquid oxygen (LOX) as the oxidizer. Methane has been considered as a possible in situ resource for Mars exploration missions because it can be produced from the carbon dioxide atmosphere with the addition of hydrogen. Accordingly, much ground testing has been done with LOX/methane engines and systems.

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Cryogenic-based systems store propellant at very low temperatures and require special insulation and dedicated cryogenic fluid management (or a cryostat) to minimize propellant boil-off, which is usually vented and lost. These systems are commonly pump fed and are used for DV burns with higher performance requirements than pressure-fed systems. An example of a current cryogenicbased in-space engine is the J2-X engine, with its high area ratio nozzle, which has planned use for future orbit transfer stages.

13.2.2

MATERIALS REQUIREMENT

The various bipropellants in current use for in-space propulsion are shown in Table 13.1. The specific impulse and combustion temperatures are ideal limits. Thermal soak back temperatures after burning must also be considered. The thrust chamber material must be compatible with the propellants and their combustion products. In addition, materials including any coatings used for oxidation resistance must be compatible with the planned operating temperature extreme, operating durations with margin, and induced thermal cycles. Ablative-cooled thrust chambers with pressure-fed systems are simple and often used for shortduration space missions. Silicon carbide with throat inserts is compatible with oxygen-based oxidizers. Other ceramic fibers can also be used. Radiative-cooled chambers are often made of columbium and must radiate to open space. Engines with high chamber pressure and long life requirements can have regenerative cooling used in their extremely hot regions. Hypergolic chemical propulsion systems have material compatibility and toxicity issues. They are commonly used with NTO and MMH. NTO can react with any iron and water impurities and create ferric nitrate, which can clog line filters and affect system performance. The titanium alloy (Ti-6Al-4V) has been shown acceptable for both NTO and MMH. However, the only nonmetal (used for seals) that seems compatible with NTO is Teflon, which has been known to swell during long exposures. This also makes it difficult for other measurement devices, such as mass gauges, to be placed inside the propellant tank. The thrust chamber is usually radiative cooled with columbium material with aluminide or disilicide coating (good up to 25008F). Platinum liners have also been used. Large engines often have regeneratively cooled stainless steel thrust chambers with columbium nozzles. Large area nozzles have employed all columbium, columbium with titanium extension, or carbon. Higher performance hypergolic engines that use NTO and hydrazine and operate at temperatures above 25008F could have a thrust chamber materials combination of platinum/ iridium, platinum/rhenium, or iridium-coated rhenium. Hydrazine is a common monopropellant with properties very similar to MMH: temperatures less than 19008F due to hydrazine decomposition, inlet pressures up to 600 psi, and thrust chamber pressure range of 150–300 psi. A common catalyst is “S-405” composed of ceramic beads (e.g., alumina) coated with iridium. Common material for the monopropellant thrust chamber is

a

LIQUID PROPELLANT COMBINATIONS FOR IN-SPACE PROPULSION Mixture Ratio (O/F)

Combustion Temperature (8F)

Remarks

Specific Impulse (s)

Density Impulse (s-gm/cc)

Freezing and Boiling Points (8F), Oxidizer/Fuel (FP)(BP)/ (FP)(BP)

F2/H2

474

241

(2364)(2307)/(2435)(2423)

9.3

6440

Hypergolic

b

O2/H2

456

143

(2362)(2298)/(435)(2423)

4.7

5110

Nonhypergolic

OF2/B2H6

430

424

(2371)(2299)/(2265)(2135)

3.5

7010

Hypergolic

OF2/B2H6

420

357

(2371 )(2299)/(2265)(2135)

2.15

5910

Hypergolic

F2/N2H4

419

551

(2364)(2307)/(35)(236)

2.4

7285

Hypergolic

OF2/CH4

417

451

(2371)(2299)/(2300)(2260)

5.6

6700

Nonhypergolic

O2/B2H6

408

303

(2362)(2298)/(2265)(2135)

2.0

5960

Nonhypergolic

N2H4/B2H6

402

254

(35)(236)/(2265)(2135)

1.2

4085

Nonhypergolic

N2O4/B2H6

375

340

(11)(70)/(2265)(2135)

2.9

5710

Nonhypergolic

b

N2O4/N2H4

341

409

(11)(70)/(35)(236)

1.23

5513

Hypergolic

b

MON/EMHF

341

407

(223)(29)/(276)(144)

2.2

5330

Hypergolic

b

MON/MMH

338

401

(223)(29)/(263)(189)

2.4

5370

Hypergolic

b

N2O4/50–50

339

408

(11)(70)/(18)(170)

2.1

5175

Hypergolic

b

N2O4/MMH

339

407

(11)(70)/(263)(189)

2.3

5290

Hypergolic

Propellant Combination

703

Notes: MON ¼ Mixed oxides of nitrogen, 85% N2O4 2 15% NO. MMH = Monomethylhydrazine, CH3 . N2H3, 50 – 50 ¼ 50% UDMH, (H3C)2N2H2 2 50% N2H4. EMHF ¼ Eutectic mixture of hydrazine fuels ¼ 87.6% MMH 2 12.4% N2H4. a Based upon theoretical shifting equilibrium at 150-psia nozzle stagnation chamber pressure and 40:1 nozzle expansion area ratio in the vacuum. b Propellant technology and application are well established.

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

TABLE 13.1

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Inconel and Haynes 230. Other advanced nontoxic monopropellants are being investigated (e.g., AF-M315E and LMP-103S). Cryogenic bipropellant engine systems require augmented spark igniters due to the nonhypergolic nature of the propellants. Pump-fed designs can have chamber pressures 1500 psia. The chamber material can be NARloy-Z (copper-silver-zirconium alloy) or stainless steel. Iridium-coated rhenium can resist oxidation up to 33008F. Other investigated high-temperature materials include the following: hafnium/iridium-lined rhenium; tungsten rhenium substrate with possible hafnium carbide for high-temperature strenght; iridium/ rhodium or platinum/rhodium for improved oxidation resistance; and hafnium oxide (stable high-temperature oxide), zirconium oxide (low conductivity, high temperature resistance), and hafnium carbide/silicon carbide (proven hightemperature stability) for a termal barrier protective layer.

13.3 GRIDDED ION THRUSTERS Dan Goebel Jet Propulsion Laboratory, California Institute of Technology, Pasadena, California

13.3.1

INTRODUCTION

Ion thrusters represent a class of electric propulsion engines that produce thrust by ionizing a large fraction of the propellant in a discharge or ionization chamber, and then electrostatically accelerating these ions into a beam using a series of closely spaced high-voltage grids [3]. Several versions of these thrusters use different techniques to ionize the neutral gas propellant and deliver the ions to the accelerator grids while containing the energetic electrons in the plasma discharge to enhance the thruster efficiency. Other versions of ion thrusters use radio frequency discharges, microwave discharges, and surface ionization to ionize the propellant [4]. Electrons in the discharge are confined by magnetic fields produced by rings of permanent magnets to increase their path length in the discharge chamber and improve the probability of an ionization collision occurring before loss to the anode. Ions generated in the discharge chamber flow freely to the grids and are inhibited from hitting the walls due to ambipolar effects from the electron motion in the magnetic fields. The ions that enter the high transparency grid system (.70%) are accelerated to high velocity by the voltage applied to the grids. The final component of the ion thruster is the electron neutralizer, which injects electrons into the ion beam to prevent a negative charge from building up on the spacecraft due to the expulsion of high-energy ions by the engine. Ion thrusters operate at accelerator voltages of about 1 kV–10 kV and feature a higher efficiency (60% to .80%) and higher specific impulse (2000 to more than 10,000 s) than other thruster types. Ion thrusters are characterized by power

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density levels of up to 12 W/cm2 of grid area and thrust-to-power levels of up to 40 mN/kW and have been built with grid diameters from about 3 cm to nearly 60 cm. Flight thrusters presently operate at power levels from 100 W to 4.5 kW. Ion thrusters used for flight missions produce thrust as low as several mN at an Isp of 2500 s to as high as 170 mN at an Isp of 3600 s. Ion thrusters developed in industry are being used for station keeping in commercial communications satellites [5] and for drag compensation in low Earth orbit science missions [6]. The next generation of thrusters in development for flight is represented by the GRC NEXT thruster [7] that operates at up to 7.2 kW at an Isp of more than 4000 s and an efficiency of 70%. Ion thrusters have been operated at very high power, with recent versions achieving nearly 30 kW of power at an Isp of 6000 to 9000 s [8]. However, the inherent low thrust of these devices requires that long operating times are needed to produce the total impulse required for most missions and applications. The NSTAR ion thruster demonstrated more than 30,000 h in a life test, and the NEXT ion thruster has achieved more than 500 kg propellant throughput to date.

13.3.2

MATERIALS REQUIREMENTS

The performance and life of ion thrusters is determined primarily by the material properties of the accelerator grids in all ion thruster types and also the cathode electrode life in electron discharge ion thrusters. Material properties used in the discharge chamber also largely determine the thruster mass and cost. 13.3.2.1

ACCELERATION GRIDS

The electrostatic acceleration grids must extract the ions from the discharge plasma and focus them through the downstream grid(s) with minimal direct interception to form the thrust beam. The grids must also have a high transparency to minimize the electrical cost of producing the ions and a low sputtering yield for long life. Finally, the grids must hold off high voltage and withstand occasional arcing events and vibrations associated with launch loads. To achieve high transparency and low interception, the grids must be relatively thin and configured with many small, aligned apertures to electrostatically focus the ions through the grids. The grids must also maintain a small, uniform grid gap of the order of 0.5 to 1 mm after experiencing 10 Grms random launch vibration loads and thermal heating during use. The grids are also typically dished so that grid flexing and contact is minimized during launch vibrations, and aperture alignment is maintained during differential thermal expansion of the grids. The high voltage is applied between the plasma generator and the middle grid, called the acceleration grid, which is the most negative electrode in the system and therefore will have the most severe ion bombardment. Acceleration grids typically have lower transparency (,30%) to trap neutral gas in the discharge chamber and have thicknesses from 0.5 to 1 mm to provide the desired

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life in the presence of some ion bombardment erosion. The first grid, the screen grid, shadows the accelerator from ions flowing from the plasma generator and must be very thin (,0.5 mm) and transparent (.70%) to allow the electric field from the acceleration grid to penetrate through the screen apertures to extract a large fraction of the incident ions. In some systems a third grid, called the deceleration grid, is placed downstream from the acceleration grid to shadow it from ions backflowing from the ion beam. The deceleration grid must also have a high transparency and be well aligned with the other grids. A realization early in the development of ion thrusters was that refractory metal grids can provide the strength and low erosion yields desired for good grid performance. Since the 1970s, grids have been fabricated of arc-cast molybdenum that has good voltage hold-off characteristics [9], low sputter yields under xenon ion bombardment, and can be readily dished by hydroforming or hotpressing. Molybdenum grids are standard in all NASA-developed thrusters [5, 8] and the XIPS thrusters [6]. The life of ion thruster grids is primarily determined by sputtering of the grids by ion bombardment. The first grid receives low-energy ion bombardment due to interception of some of the ions produced in the plasma generator, and the second and third grids receive ion bombardment from slow ions produced by charge-exchange (CEX). Ions produced by CEX in the grid gap cause barrel erosion of the aperture inner diameters, and ions produced by CEX in the near thruster plume cause “pit-and-groves” erosion of the downstream grid surface. The NSTAR grids demonstrated more than 30,000 h of operation, and the characteristic grid erosion patterns due to these two processes are observed in the grid photos taken at the end of the test and shown in Fig. 13.3. To reduce the erosion and extend the life of the grids, alternative materials have been pursued that feature lower sputtering yields. Low atomic mass materials with a large mass difference from xenon tend to have lower sputtering yields [11], which leads to graphite/carbon, beryllium, and titanium being considered as replacements to molybdenum. Although beryllium has a very low sputtering yield and good material properties for grids, it is considered too dangerous to machine and handle for this application. Titanium grids were fabricated and tested on an NSTAR ion thruster [12] in 2000. Fig. 13.3 New grid photograph (top) and eroded grid photograph (bottom) taken after 30,000 h from the NSTAR extended life test [10].

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Reasonable beam-generation performance was measured, but titanium’s larger thermal expansion and smaller thermal conductivity caused the grids to expand during use, which increases the dish radius and beam divergence and tended to make the grids contact during startup due to differential heating [11]. Carbon has a lower sputtering yield from xenon ion bombardment than from titanium [13] and is therefore expected to have a longer life, but it has less mechanical strength. Graphite grids in early tests could not survive launch vibrations, but this problem was solved by vibration isolation techniques, and graphite is used successfully for the acceleration grid in the small-diameter (10-cm) Qinetic T5 thruster [7] and in the larger, 22-cm T6 ion engine [12]. Higher strength carbon–carbon composite grids were developed and successfully tested with the 30-cm NSTAR engine [14] and the 57-cm NEXIS engine [15]. The 57-cm NEXIS grids also passed flight-level vibration tests and a 2000-h wear testing. Pyrolytic graphite grids were developed for 8-cm and 30-cm ion thrusters [16] and survived 9.2 Grms random vibration tests in spite of significant grid clashing during the tests. Testing of the voltage-hold-off capability of carbon-based grid materials in the presence of arcing [13] has shown reduced breakdown voltages compared with refractor metal grids. Additional development of carbon-based grids is required to improve the manufacturing methods, minimize the grid thicknesses to improve the transparency, and increase the voltage standoff and arcsurvival capability for optimum grid performance. New grid materials and/or long-life grid coatings compatible with ion bombardment, thermal expansion, and launch vibrations are needed. 13.3.2.2

CATHODES

Hollow cathodes are used in the plasma generator of electron-discharge ion thrusters to provide the discharge current for ion production. Because of the higher current required of these devices (5 to .30 A), they are typically configured as thermionic hollow cathodes [3]. Hollow cathodes are also used as the neutralizer cathode in every type of ion thruster to inject electrons into the beam without the need for immersing the emitter in the beam. Neutralizer cathodes use thermionic electron emitters in electron-discharge ion thrusters and microwave hollow cathodes in microwave ion thrusters. The life of the thruster is often determined by the evaporation characteristics of the thermionic electron emitter at the temperatures required to produce the desired currents. Barium-oxide dispenser cathodes inserts have achieved more than 30,000 h in testing [17] in the NSTAR thruster, and alternative insert materials such as lanthanum hexaboride are projected to provide higher current capabilities and longer life [18]. The life of both the discharge and neutralizer cathodes is also determined by the sputtering properties of the materials used as the electrodes in the hollow cathode construction [19, 20]. Figure 13.4 shows photographs of the NSTAR discharge hollow cathode at the beginning and end of the 30,000-h extended life test [18]. The cathode structure features an internal cathode tube and orifice plate that

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Fig. 13.4 Photograph of the NSTAR DCA assembly after 30,352 h indicating complete keeper plate erosion [21]. encloses the thermionic electron emitter, and an exterior tube and orifice plate called the keeper electrode that protects the cathode tube from ion bombardment and aids in starting the cathode discharge. Figure 13.4 shows that the molybdenum keeper electrode was fully eroded away by the end of the test, and the cathode orifice plate wrapped around the tube has been seriously damaged. This erosion has been attributed to energetic ions generated in the near cathode plume [22]. Modifications of the keeper material to materials with lower sputtering yields such as tantalum and graphite are projected to reduce the erosion rate; modifications to the cathode structural design to minimize the production of energetic ions are considered critically important to achieve longer life. 13.3.2.3

DISCHARGE CHAMBER

The discharge chamber in electron discharge and microwave ion thrusters is typically manufactured of stainless steel or magnetic steel as required for structural integrity and the electron confinement technique used. Flight thrusters have also used titanium to reduce the mass of the discharge chamber and made use of spin forming and e-beam or laser welding as fabrication techniques to reduce both mass and cost. Alternative materials such as aluminum and graphite have been proposed for flight thruster discharge chambers, but development is needed. Microwave and radio-frequency (RF) ion thrusters have discharge chambers that are compatible with the coupling of electromagnetic radiation to the plasma. Whereas laboratory RF ion thrusters typically use quartz tubes as the discharge chamber wall, flight RF ion thrusters use alumina ceramic discharge chambers with RF coils wrapped around the outside to inductively couple the RF energy into the discharge [23]. Alternative materials that reduce the mass of the RF discharge chamber while surviving launch loads and maintaining RF coupling to the plasma are desirable. Microwave thrusters use an antenna structure inserted into the discharge chamber edge to couple the microwave energy to

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

709

the plasma and use ceramic or quartz windows to provide a vacuum interface between the discharge chamber and the waveguide and microwave system. Alternative materials that reduce the mass of this system are also desirable in these thrusters.

13.4 HALL PROPULSION SYSTEMS Hani Kamhawi NASA John H. Glenn Research Center at Lewis Field, Cleveland, Ohio

13.4.1

INTRODUCTION

Hall-effect thrusters (HETs) belong to the family of electric thrusters. Electric thrusters use electric energy to heat, ionize, and accelerate injected propellant [23]. HETs are categorized as electrostatic thrusters because they use an applied electric field to accelerate the ionized propellant. HET experiments in the United States and former Soviet Union began in the early 1960s [24, 25]. The first flight tests of a HET were those of the SPT-60 Soviet thruster in 1971 [26]. Two distinct HET configurations have been extensively researched and demonstrated: single-stage magnetic layer thrusters (MLTs), sometimes called stationary plasma thrusters (SPTs), and two-stage-anode layer thrusters (TALs) [27, 28]. In this section, the focus will be on the MLT configuration because it is more widely researched, developed, and flown. In the TAL configuration, an intermediate electrode is used to decouple the ionization and acceleration processes and stages. Most of the materials requirements for the MLT thruster configuration apply for the TAL; therefore, the focus in the next section will be on materials requirements for the MLT thruster configuration. HETs can operate at discharge voltages between 200 V and 1000 V demonstrating total thrust efficiencies up to 65% and specific impulse up to 4000 s. Hall thrusters are characterized by power density levels of up to 80 W/cm2 and thrust-to-power levels of up to 90 mN/kW and have been built with outer diameters of up to 46 cm. U.S. manufacturers of flight Hall propulsion systems include Aerojet, Loral Space Systems, and the Busek Company. The highest power flight HET propulsion system is Aerojet’s Hall Thruster Propulsion System (HTPS). Aerojet’s HTPS employs a 4.5-kW flight HET, and the Hall propulsion system is used aboard the U.S. Air Force’s Advanced Extremely High Frequency (AEHF) communication satellite launched in August of 2010. Busek BHT-200 is a 200 W HET and the first U.S.-manufactured HET flying in space on the Air Force’s TacSat 2 satellite launched in 2006 [29]. Although current state-of-the-art (SOA) HETs fulfill and meet current mission requirements, future missions will require advancements and improvements in materials and manufacturing techniques employed in Hall propulsion

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systems. The next section will highlight major materials requirements for the various subsystem elements and will briefly discuss areas of research that can enhance our ability to build lower mass, longer life, lower cost, higher power, and higher performing Hall propulsion systems.

13.4.2

MATERIALS REQUIREMENTS

Hall propulsion system performance, life, and cost can be improved by improving on SOA materials and manufacturing processes and/or developing new materials. The materials requirements discussed below are relevant to low-, mid-, and highpower Hall propulsion systems. 13.4.2.1

HALL THRUSTERS

In an MLT HET device, performance and life are determined by hollow cathode performance and life, discharge channel design, insulator materials selection, magnetic field topography, anode/propellant distributor design, and the propellant used. 13.4.2.2

HOLLOW CATHODE

In Hall thrusters, the hollow cathode serves the critical function of ionizing the injected propellant and neutralizing the exiting ion beam. As with ion thrusters, the hollow cathode life and performance are mainly governed by the heater life, thermionic emitter life, and erosion of critical hollow cathode components due to energetic ion bombardment. Sec. 13.3.2.1 addressed the major materials issues associated with hollow cathodes. However, higher power Hall thrusters pose the additional challenge of requiring emission currents up to 250 A. These high emission currents, when compared to typical SOA of 30 A, will require new emitter configurations that will maintain the emitter peak temperatures within operational limits that are consistent with long life and will require selection of materials with high sputter erosion resistance due to the energetic nature and higher density of the generated ions in high current cathodes. Discharge channel walls The ceramic discharge chamber walls serve the function of containing the injected propellant, providing voltage isolation between the anode and the magnetic circuit elements, and protecting the magnetic circuit inner and outer front poles and electromagnets against ion bombardment that ultimately causes HET end of life. Discharge channel wall materials greatly influence the performance and lifetime of a HET. It is desired that the insulating discharge channel walls have high mechanical strength, high dielectric strength, and good thermal conductivity; be resistant to thermal and mechanical shock; and possess low sputter yield magnitudes [30]. In addition, the secondary electron emission properties of the

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711

discharge channel material at various temperatures greatly influence HET performance and discharge stability [31, 32]. Erosion of the discharge channel walls due to impinging ions exposes the magnetic circuit elements in a HET, eventually resulting in end of life. In situ discharge channel replacement and magnetic field shaping have been demonstrated as potential approaches for eliminating discharge channel erosion as a life-limiting mechanism in HETs, but they result in potentially more complicated thruster designs [33, 34]. Boron nitride is the material of choice for the various SOA flight HETs. Various studies have been performed to assess different insulator discharge channel wall materials, including various grades of boron nitride, silicon carbide, alumina, graphite, and polycrystalline diamond [35, 36]. These studies have demonstrated that potential performance and life gains can be attained by use of alternate materials, but additional research and development are needed to confirm the preliminary results. Additionally, most of the testing has been performed with low-power Hall thrusters (,2 kW), and more research should be directed to testing with higher power HETs. 13.4.2.3

MAGNETIC CIRCUIT

The magnetic circuit elements generate an applied radial magnetic field with the desired field topography (field strength and shape). The radial applied magnetic field interacts with the axially applied electric field (between anode and cathode), forcing the electrons to execute an azimuthal drift (Hall current) under the influence of the EB force. Collisions between the azimuthally drifting electrons and injected propellant create unmagnetized ions that are accelerated by the applied electric field. The topography of the applied magnetic field greatly impacts the performance, discharge stability, life, and mass of the HET. The magnetic field strength in a particular thruster design is dictated by the range of thruster operating conditions. The shape of the magnetic field streamlines influences thruster operation and greatly impacts the discharge channel erosion profile. The major elements of the magnetic circuit include the inner and outer electromagnets, inner core and front pole, back pole, outer ring and outer front pole, and inner and outer magnetic screens. An optimum magnetic circuit design entails attaining the desired magnetic topography with the lowest possible electromagnet number of windings and applied current (amp-turns) and lowest potential magnetic circuit elements mass. The magnetic circuit elements (excluding electromagnets) are usually manufactured from a soft ferromagnetic material [37]. Desired properties include high permeability, high saturation flux, small hysteresis loss, and a high Curie temperature. In addition, materials selection has to take into account that the magnetic circuit elements, which usually serve as the thruster housing structure, must be able to withstand the structural loads during launch. Improving the magnetic permeability of SOA alloy ferromagnetic materials will reduce the mass of HETs and

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will permit attaining the desired magnetic field strength at a lower amp-turn magnitude. Ferromagnetic alloys with high Curie temperature will permit operation at higher power densities than SOA. In addition, higher power HETs will require advancements in manufacturing and heat treatment processes of large ferromagnetic thruster pieces and parts to retain the desired magnetic properties. In HET designs, minimizing the amount of space occupied by the electromagnets will result in reduced thruster mass. This can be achieved by using smaller gauge-sized wires. However, this results in higher internal electromagnet temperatures due to the higher resistivity that results from the smaller wire cross section, which increases the probability of wire insulator damage or an electromagnet short circuit. This will distort the magnetic field topography and will alter and disrupt the HET operation and affect its performance, eventually causing end of life. As such, developing lower resistivity conductor materials will greatly improve the ability to manufacture space-efficient electromagnets. In addition, developing higher temperature wire insulation materials (typical SOA is 4508C) will further enable operation at higher electromagnet currents, thus resulting in more efficient and compact magnetic circuit designs. 13.4.2.4

ANODE/PROPELLANT DISTRIBUTOR

The anode in a HET constitutes the positive electrode and also serves the function of uniformly distributing the injected propellant in the annular discharge chamber. Uniform propellant flow is critical for optimized thruster performance and eliminates any nonuniform discharge current concentrations. Typical anode designs include a manifold base, an orifice plate, and a discharge plate. Limited information is found in the open literature regarding anode materials and designs mainly because anode designs are typically considered by the company or institution. Higher power HETs will require larger anodes and must be fabricated from materials that can withstand higher heat loads and temperatures. 13.4.2.5

PROPELLANTS

HETs have been operated with a variety of propellants, including xenon, krypton, bismuth, and magnesium. The desired properties in HET propellant are a low ionization potential, a large ionization cross section, a low boiling point, and a relatively high storage density. Table 13.2 lists properties of xenon, krypton, bismuth, and magnesium propellants. Xenon has been the propellant of choice for SOA HETs. It has a moderate first ionization potential, is not toxic, and is easy to handle and store. HETs operating on xenon have demonstrated thrust efficiencies up to 70%. Extensive testing has been performed with krypton because its lower atomic mass results in higher HET specific impulse. However, krypton’s higher first ionization potential when compared with xenon results in lower thruster efficiencies [38, 39]. Bismuth is another propellant that has been widely researched and tested with HETs [40]. Bismuth’s large atomic mass, low first ionization

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

TABLE 13.2

713

PROPERTIES OF SELECTED HET PROPELLANTS

Propellant

Xenon

Krypton

Bismuth

Magnesium

Iodine

Atomic mass, g/mol

131.3

83.8

209

24.3

126.9

First ionization potential, eV

12.1

14

7.3

7.6

10.5

Boiling point, 8C

2111.7 (@10 Pa)

2153.2 (@10 Pa)

768

500

9 (@10 Pa)

Density, gm/cm3

1.6 (@14 MPa, 508C)

0.5 (@14 MPa, 508C)

9.8

1.7

4.9

potential, and high storage density make it very attractive for HET missions that require high thrust-to-power performance. However, its high boiling point complicates the propellant feed system and adds additional complexity to the HET propulsion system. Preliminary HET testing with magnesium has been performed, but conclusive findings on the benefits of implementing magnesium are awaiting additional testing results [41, 42].

13.5 SOLAR THERMAL PROPULSION Harold Gerrish NASA Marshall Space Flight Center, Huntsville, Alabama

13.5.1

INTRODUCTION

Solar thermal propulsion (STP) is an in-space propulsion concept that uses solar energy to heat the propellant for thermal expansion through a nozzle. A variety of propellants have been considered (e.g., hydrogen, ammonia, methane), but hydrogen has the lowest molecular weight and highest specific impulse (.700 s). The specific impulse is about twice that of conventional chemical thrusters. The major materials challenges include maximizing the operating temperatures and material compatibility with the propellants. Figure 13.5 shows how the concept works. Solar energy is collected by a concentrator (reflector or lens) and focused inside an absorber cavity. The absorber cavity heats up to very high temperatures (.2700 K) and transfers heat to the propellant. Figure 13.6 shows what the STP concept could look like. STP was primarily considered as an upper stage to take payloads from low Earth orbit (LEO) to geosynchronous Earth orbit (GEO). Having a specific impulse about twice that of conventional chemical in-space engines allows for more payload weight on the launch vehicle. However, the volume of liquid hydrogen took up more volume in the payload shroud and the thrust level was 2–4 lbs,

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L. JOHNSON AND T. LOCKETT

Fig. 13.5

Solar thermal propulsion operating principle [43].

which yields a time for a LEO to GEO transfer mission of 30 days. Other considered STP applications include using the heat from the system in bimodal for electric power generation and as a transfer stage to the moon and other solar system destinations. The primary technical challenge facing STP projects is choosing the engine materials that meet mission performance design requirements (many hours of exposure to propellant at temperatures greater than 2700 K).

13.5.2

MATERIALS REQUIREMENTS

For STP to be considered for future missions, it must outperform conventional chemical propulsion systems. Current capabilities limit the capture of solar

Fig. 13.6

Solar thermal propulsion upper-stage concept [44].

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

Fig. 13.7

715

Melting point of various materials of interest to solar thermal propulsion [45].

energy to the kilowatt range and limit the heating to a small flow rate of propellant to extremely high temperatures, as is necessary for high propulsion performance. Without high operating pressure or thrust, one must operate for hundreds of hours to build up orbit transfer velocity, depending on the mission. The challenging STP materials requirements for the engine are as follows: .

material to withstand temperatures greater than 2700 K;

.

endurance for a few hundred hours at low operating pressure (25–100 psi);

.

material compatibility with hydrogen, ammonia, or methane;

.

minimum permeability; and

.

high thermal conductivity.

High-temperature operation is the most important materials requirement. Figure 13.7 shows the melting point of various materials. Past STP work focused on refractory metals (e.g., tungsten, rhenium, tungsten/rhenium alloys), graphite with rhenium coatings, and carbides (e.g., tantalum carbide).

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L. JOHNSON AND T. LOCKETT

It is also noted that various material properties and complex design geometries are driven by the material fabrication process (e.g., sintering, chemical vapor deposition, hot isostatic press, vacuum plasma spray, electron beam melting, etc). Figure 13.8 shows the components of a direct gain STP engine with the long shell components made of 100% tungsten. Direct gain engines are those that thrust during solar exposure. The complex geometry was achieved using vacuum plasma spray. Tungsten is low cost, has the highest melting point of all refractory metals, and is compatible with hydrogen. However, tungsten is very dense and heavy, recrystallizes at low temperatures, oxidizes easily, has a very low coefficient of thermal expansion at high temperatures (which makes joining with other materials difficult), and is not ductile enough for machining at room temperatures. Figure 13.9 shows two direct gain STP engines with major components made of 75% tungsten/25% rhenium. The major components of this engine were also vacuum plasma sprayed. The alloy was selected because rhenium adds much better ductility to the machining. 13.5.2.1

THERMAL STORAGE ENGINES

Thermal storage engines build up heat over time before actually thrusting. The heat exchanger was made of open cell rhenium foam. Rhenium is compatible with hydrogen and has good ductility but is very costly. Other STP thermal storage engines are made of graphite with rhenium coatings. The graphite holds heat well, and the rhenium helps protect the surface from hot hydrogen. Permeability of hot hydrogen through the rhenium quickly erodes the carbon from behind the rhenium and affects the heat transfer. The challenge is having a coating process that can apply a thin, dense, uniform coating with a coefficient of thermal expansion similar to the base material. The ultimate performance that can be achieved with STP is with carbides (e.g., tantalum carbide) or material with high-temperature carbide coatings (e.g., niobium carbide). At temperatures greater than 3200 K, hydrogen propellant dissociates to small quantities of monatomic hydrogen and has a lower average molecular weight, which increases its specific impulse. Future work in this field

Fig. 13.8

100% tungsten solar thermal propulsion engine components [46].

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

Fig. 13.9

717

75% tngsten and 25% rhenium direct gain solar thermal propulsion engines [47].

is needed to obtain more material property data associated with the carbide fabrication process used and effects of using more hot monatomic hydrogen. Preliminary results show various carbides can hold up to the hot hydrogen exposure. 13.5.2.2

CONCENTRATORS

The large solar energy concentrator, with a capture area of 4 m  6 m (Fig. 13.10), poses additional materials challenges. Two concentrator concepts that are being considered are reflectors and Fresnel lens. Rigid concentrators can also be made, but they are heavy and take up valuable payload volume. Past work looked more at inflatable concentrators, which are lightweight and can be folded into a small container. The best material was found to be a clear polyimide Fig. 13.10 4 m 3 6 m inflatable off-axis concentrator [48].

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L. JOHNSON AND T. LOCKETT

CP. A thin coating of aluminum is used for it to be a reflector. Many lightweight materials (e.g., Mylar) have been considered for STP concentrators, but polyimide CP can withstand the space environment effect [e.g., exposure to charged particles and ultraviolet (UV) radiation] best and maintains specular reflectivity. More work is needed with materials that can inflate and then become rigidized with UV exposure. This avoids concerns with puncturing an inflatable in orbit from micrometeoroid orbit debris.

13.6 NUCLEAR THERMAL PROPULSION Mike Houts NASA Marshall Space Flight Center, Huntsville, Alabama

13.6.1

INTRODUCTION

Nuclear thermal propulsion (NTP) represents the next “evolutionary step” in highperformance liquid rocket engine development, yet it uses many of the same technologies found in today’s chemical engines and stages [e.g., liquid hydrogen (LH2) turbopumps, nozzles, propellant tanks]. Chemical engines, however, produce their energy through the combustion of an oxidizer-fuel mixture, whereas NTP derives its energy from the fissioning of uranium-235 atoms contained within fuel elements that comprise the engine’s reactor core. A representative “expander cycle” NTP engine is shown in Fig. 13.11.

Fig. 13.11

Nuclear Thermal Rocket Element Environmental Simulator (NTREES) [49].

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

719

High-pressure LH2 flowing from turbopump assemblies (TPAs) cools the engine’s nozzle, pressure vessel, neutron reflector, and control drums and, in the process, picks up heat to drive the turbines. The turbine exhaust is then routed through the core support structure, internal radiation shield, and coolant channels in the reactor core’s fuel elements where it absorbs fission energy, is superheated to high temperature (2500–3000 K depending on fuel type), and then expanded out a high area ratio nozzle to generate thrust at twice the specific impulse (Isp) of today’s best chemical rockets. Controlling the NTP engine during its various operational phases (startup, full thrust, and shutdown) is accomplished by matching the TPA-supplied LH2 flow to the reactor power level. Multiple control drums, located in the reflector region surrounding the core, regulate the neutron population and reactor power level over the operational lifetime of the engine. NTP systems underwent extensive development and testing from 1955 to 1973. Smaller programs since that time have furthered the SOA. First-generation NTP systems are projected to provide a specific impulse in excess of 900 s at a high thrust-to-weight ratio of 3.5. More advanced systems promise even higher performance.

13.6.2

NUCLEAR FUELS

The primary materials challenge for NTP systems is developing the nuclear fuel used to directly heat hydrogen propellant. The fuel must be capable of operating at very high power density (1 MW/L) and very high temperature (.2500 K) for several hours (cumulative) in a flowing hydrogen environment. The fuel may also need to withstand two or more thermal cycles from ambient temperature to operating temperature and thermal transients approaching or exceeding 100 K/s during engine start. Increasing maximum fuel operating temperature, power density, or both could provide significant performance benefits to NTP systems. NTP fuel must also have adequate uranium loading to enable criticality (a selfsustained fission reaction) in a compact reactor core that simultaneously incorporates adequate hydrogen coolant channels for high thrust-to-weight operation. In addition, other isotopes in the fuel must have acceptable neutronic properties, principally a low neutron absorption cross section in the operating spectrum. A representative thermal neutron spectrum system could use composite fuel with a uranium carbide loading of 700–800 mg/cc and zirconium hydride (UZrH1.85) moderator incorporated into structural tie tubes [50]. A representative fast spectrum system could use a W-UO2 cermet with Gd2O3 added for chemical stabilization. A tungsten fraction of ,40% is desirable for cermet fuels to enable a compact engine. Several potential fuel forms have been developed for NTP systems. During the Rover/NERVA program [24], graphite matrix, composite, and carbide fuels were developed and tested to varying degrees. For total engine burn times of 2 h, graphite matrix fuel (pyrocarbon coated UC2 particles in a graphite substrate) was shown adequate for providing a bulk hydrogen exit temperature of 2500 K. Composite

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L. JOHNSON AND T. LOCKETT

fuel (a UC-ZrC dispersion in a graphite substrate) was shown adequate for providing a bulk hydrogen exit temperature in excess of 2600 K. Both of these fuel forms employed an NbC or ZrC coating to prevent corrosion. Initial testing was also performed on carbide fuels, which showed promise for enabling bulk hydrogen exit temperatures in excess of 3000 K. The brittle nature of pure carbide elements would need to be accommodated in fuel element and engine design. Tungsten uranium dioxide and tungsten uranium nitride cermet fuel was developed and tested during the 1960s [51]. In addition, modern fabrication techniques may further improve the viability and performance of cermet fuels [25, 52]. For total engine burn times of 2 h, cermet fuels appear capable of providing a bulk hydrogen exit temperature in excess of 2600 K. Cermet fuels also have excellent burn up and fission product retention capability and may be applicable to high-performance nuclear electric propulsion systems. Multiyear lifetimes at fuel temperatures in excess of 1800 K may be feasible. Fuel fractions in excess of 60% appear feasible while still maintaining full tungsten encapsulation of individual fuel particles. For solid core NTP systems, modern fabrication techniques may enable even higher performing systems. For example, options may exist for allowing UC2 or other fuel particles to run at or possibly beyond their melting point while still maintaining a solid overall core structure and adequate NTP engine lifetime [53]. It may also now be feasible to fabricate advanced, high-temperature carbides in the geometries needed to withstand very high power densities and operating temperatures. Candidate carbides include tantalum zirconium carbide and tantalum hafnium carbide (Tmelt 4488 K). Uranium could be added either directly in the carbide phase or as particles in the carbide matrix. These and other 21st-century technologies could potentially enable solid core nuclear thermal rockets with specific impulses in excess of 1000s. In the early 1970s, preliminary study was also given to potential liquid, gas, and plasma core nuclear thermal rockets. Additional details would need to be developed for the more promising concepts before specific materials needs could be identified. In general, high-temperature structural materials would be required. The ability to use specific isotopes in structural or other applications could also provide a significant benefit to advanced, high-performance NTP systems. Nonsolid-core NTP systems would also rely on advanced fuel and hydrogen flow control to maximize performance while ensuring adequate lifetime. Candidate high-temperature structural alloys include W-30%Mo. Propulsion-only NTP systems typically have very low fuel burn up: approximately 0.1%. Low burn up will facilitate NTP fuel development by reducing the effect of fission products on fuel element performance. Nonnuclear and shortduration nuclear tests could also be performed with nonradioactive fission product surrogates in the fuel to more accurately determine fuel behavior. Bimodal (propulsion and power generation) NTP systems would have higher fuel burn up, but potential issues would be mitigated by the relatively low power generation requirements (,50 kWe) anticipated for initial systems.

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

Fig. 13.12

721

NTREES schematic [54].

One approach to developing NTP fuels would involve developing and testing small samples of the desired fuel form and then fabricating representative fuel elements to be tested at temperature and at near-prototypic power density in a flowing hydrogen environment. Elements with adequate performance in nonnuclear screening tests would then progress to nuclear testing. Nuclear testing would ideally be performed at near-prototypic conditions, including temperature, hydrogen pressure, power density, fuel burn up, and operating time. Highly realistic nonnuclear testing of NTP fuel elements could potentially be performed at NASA’s Nuclear Thermal Rocket Element Environmental Simulator (NTREES). A schematic of the NTREES is shown in Fig. 13.12. Typical NTP fuel elements are 1.0–1.5 m in length. Initial testing could be performed on shorter fuel elements that would still allow hydrogen flow to fully develop and for directly useful data to be obtained. A 0.30 m cermet fuel element (with surrogate in place of uranium fuel) that includes a representative hydrogen flow channel geometry is shown in Fig. 13.13. The element was fabricated using modern cermet fabrication techniques. Fig. 13.13 Representative cermet fuel element (surrogate in place of uranium fuel) [55].

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L. JOHNSON AND T. LOCKETT

13.7 SOLAR SAIL PROPULSION Dave Edwards and Roy Young NASA Marshall Space Flight Center, Huntsville, Alabama

13.7.1

INTRODUCTION

Solar sail propulsion uses sunlight to propel vehicles through space by reflecting solar photons from a large (typically 20–100 m per side), mirror-like sail made of a lightweight, highly reflective material. The continuous photonic pressure provides propellantless thrust to perform a wide range of advanced maneuvers, such as to hover indefinitely at points in space, or conduct orbital plane changes more efficiently than conventional chemical propulsion. Eventually, a solar sail propulsion system could propel a space vehicle to speeds that are theoretically much faster than any present-day propulsion system. Because the sun supplies the necessary propulsive energy, solar sails require no onboard propellant, which can significantly increase useful payload mass. Practical concepts for solar sailing have existed for approximately 100 years, beginning with Konstantin Tsiolkovsky and Friedrich Tsander in the 1920s. A team at the Jet Propulsion Laboratory completed the first serious mission study in the late 1970s for a rendezvous with Halley’s Comet [56]. An effort by Colin McInnes in the 1990s and the publication of his Ph.D. dissertation as a textbook on solar sailing helped reinvigorate interest in solar sailing as a research topic [57]. In the early to mid-2000s, NASA’s In-Space Propulsion Technology Project made substantial progress in the development of solar sail propulsion systems. Two different 20-m solar sail systems were produced and successfully completed functional vacuum testing in the NASA Glenn Research Center Space Power Facility at Plum Brook Station, Ohio [58]. Solar sailing has been tested in space. In the summer of 2010, the Japanese Aerospace Exploration Agency launched a solar sail spacecraft named IKAROS in tandem with another mission to Venus. The sailcraft IKAROS (14 m by 14 m) is the first in-flight demonstration of solar sailing [59]. Although the effects of solar radiation pressure (SRP) are smaller on this sailcraft as compared with other concepts for solar sails, numerous program objectives have been achieved, including verifying SRP effects on the sail and performing in-flight guidance and navigation techniques using the solar sail. The force generated by reflection of the sun’s radiation represents a sailcraft’s thrust. The net force on a sailcraft surface is generally perpendicular to its reflective surface. Desired orbital maneuvers are conducted by pointing this thrust vector in a particular direction to alter the sailcraft’s velocity vector in such a way as to effect a desired change in this velocity (DV) over a given period (sometimes considerable) of time.

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

723

By tilting the sail surface such that the reflecting surface is no longer perpendicular to the incident radiation, the thrust vector will have a component that is along the sailcraft’s trajectory as well as a perpendicular component. This perpendicular component enables the sailcraft to make inclination changes as well as spiral-in/spiral-out trajectories.

13.7.2

SAIL MEMBRANES

The driving parameters for sails are reflectivity and mass. Sailcraft designers commonly achieve high reflectivity and low mass by depositing a thin, highly reflective coating, usually aluminum, on a thin polymeric membrane. The most common materials for this polymeric membrane are Kapton, Mylar, and derivatives of Kapton such as Colorless Polyimide 1 and polyethylene naphtalate. Techniques to manufacture these films in mass quantity in thin layers (i.e., 5–10 mm) have been developed, and enhanced techniques to further “etch” these films to generate uniform thickness of 2–3 mm are being developed. The L’Garde Company used 2.0 mm Mylar in the development of its sail. Although decreasing thickness reduces mass and thus enhances the characteristic acceleration, there are negative aspects that necessitate trades to be considered. As the thickness of these films decreases, material property thresholds of mechanical strength, temperature limits, and glass transition temperature decrease from their bulk property values. It has been well documented [60, 61] that the bulk value of the glass transition temperature (Tg) is dependent on film thickness, and such dependence increases with decreasing film thickness. The temperature of spacecraft surfaces is dependent on the absorbance and emittance of external surfaces, as shown by Eq. (13.1) [62]: T 4 ¼ ðas =11 þ 12 Þ  ðS=sÞ

ð13:1Þ

where

as ¼ 11 ¼ 12 ¼ S¼ s¼ T¼

solar absorbance front surface emittance back surface emittance solar constant (1366.1 Wm22) Stefan–Boltzmann constant (5.6704  1028 Wm22K24) temperature, K

Realistic sailcraft design is not this straightforward. As the polymeric substrate decreases in thickness, the thermo-optical properties change. This is due to the metalized coating having increased bias on the thermo-optical measurements as the polymeric layer thickness decreases. The effective emittance of the polymeric layer decreases with decreasing thickness. Table 13.3 lists values for thermo-optical properties for specific polymeric materials according to thickness, with and without an aluminum coating. Figure 13.14 shows the sail temperature for aluminum-coated Kapton using 2

724

TABLE 13.3 Material CP1 CP2 Kapton Kapton Mylar AL

6.36 12.7

Alpha

Emittance

Alpha with AL Coating

Emittance with AL Coating

0.072

0.194

0.106

0.03

263

0.07

0.243

0.088

0.03

209

0.23

0.24

500

360

0.46

0.86

254

80

0.14

0.28

254

80

0.19

0.77

254

80

2.03 127 3.8 127 0.08

0.02

Melt Temp, 8C

660

Glass Transition, 8C

L. JOHNSON AND T. LOCKETT

Mylar

Thickness, Micron

THERMO-OPTICAL PROPERTIES FOR SPECIFIC POLYMERIC MATERIALS [19–21]

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

Fig. 13.14 of NASA.)

725

Effective sail temperature as a function of solar distance in AU. (Image courtesy

thicknesses of Kapton, 127 mm and 2 mm, and two thicknesses of Mylar, 127 mm and 3.8 mm. As the polymeric layer thickness is decreased, the aluminum layer begins to influence the thermo-optical properties and hence alter the sail temperature. The characteristic acceleration of a sail can be calculated using Eq. (13.2), where the reflectivity R ¼ 0.9, the solar radiation pressure Ps ¼ 9.12 mN/m2, and ra is the sail areal density (defined as the sail density multiplied by the sail thickness): ac ¼ R Ps =ra

ð13:2Þ

This acceleration is good enough to generate a spacecraft velocity increase or DV 100 m/s per day at the Earth’s orbit; this sail could reach Mars within a year with a modest payload, but the acceleration is limited by the mass of the plastic substrate. Although the conventional sail offers potential for interplanetary flight, acceleration and the distance of closest solar approach are the key variables that determine if a solar sail can reach velocities of interest for interstellar travel. Usually, the characteristic acceleration, ac [i.e., the sail

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L. JOHNSON AND T. LOCKETT

acceleration at 1 AU (the Earth’s orbit)], is a good means of comparing one sail design to another. Combining Eq. (13.2) with an expansion of the expression for radiation pressure yields ac ¼ Ps Cos2 a½1 þ RA=ðmr2 Þ

ð13:3Þ

where A ¼ surface area of the sail a ¼ angle between the incident solar radiation and the reflective surface normal m ¼ sail mass r ¼ distance from the sun in AU Utilizing Eq. (13.3), the driving materials parameters for sailcraft designers to obtain optimum performance are R, reflectivity of the sail materials; m, mass of the sail materials; and A, the surface area of the sail. The reflectivity of a sail material is a primary factor in sail performance. Figure 13.15 shows the dependence of sailcraft acceleration on the sail material

Fig. 13.15 Solar sail characteristic acceleration dependence on sail material reflectivity and sail area. Reflectivity calculations were made using a sail area of 1026 m2. Sail area calculations were made using sail reflectivity of 1. (Image courtesy of NASA.)

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

727

Fig. 13.16 Ultimate tensile strength on selected polymeric materials as a function of increasing radiation dose. reflectivity and also on sail area. The calculations in Fig. 13.14 were based on a total sailcraft mass of 1000 Kg, a distance of 1 AU (astronomical unit) from the sun, and the sail surface normal parallel to the incident flux (i.e., a ¼ 0). Sailcraft designers have been challenged with developing lightweight materials and structures and with integrating guidance, navigation, and control systems, which use these first principles of photon pressure. Interest in solar sails has resulted in development of solar sail materials, structures, and deployment mechanisms. Many challenges confront sail materials technologists, and among these is the development of low areal density sails. Present SOA sails have an areal density ranging between 3 and 15 g/m2. The far-term goal of sail materials manufacturers is a sail material on the order of .01 g/m2 areal density. These low areal density, thin, polymeric materials are highly susceptible to low energy-charged particle radiation. Experiments have shown that ionizing radiation exposure has little effect on thermo-optical properties; however, radiation doses above 1 Giga rad (Grad) begin to decrease the ultimate strength of the polymeric material, as shown in Fig. 13.16.

13.7.3

ADVANCED SAIL MATERIALS

Conventional solar sails have been designed and built using thin films of aluminum, approximately 100 nm thick, deposited on a plastic substrate typically

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L. JOHNSON AND T. LOCKETT

0.9–25 mm thick. Aluminized Mylar or Kapton is commercially available and can be used for this purpose; for example, a 5 mm thick (0.2 mil) product that is 90% reflective and has an areal density of 7 g/m2 has an acceleration at 1 AU, excluding payload mass, of only 1.2 mm/s2. Eliminating the plastic substrate altogether would dramatically increase sail performance. Scaglione [63] describes two methods that may be applicable to an in-orbit removal of the plastic substrate: UV degradation and etching by atomic oxygen. If it were possible to remove the plastic substrate and retain the aluminum coating from commercially available products, the resulting sail acceleration will be over an order of magnitude higher than that of a conventional solar sail. Further orders of magnitude improvement are possible by reducing aluminum layer thickness and developing perforated sails, although manufacturing challenges are high. The following improvements in sail acceleration are possible by advanced development of manufacturing technologies: 25x improvement by removing the plastic substrate, leaving 100 nm Al layer; 300x improvement by reducing aluminum sail thickness to 4–5 nm; 500–5,000x improvement by perforating the aluminum sail, possible in the near term; and 10,000–100,000x improvement by doping carbon nanotubes, well into the 21st century. A conventional Kapton solar sail of only 40 m (0.040 km) on a side will weigh almost 10 kg and will have an acceleration at 1 AU of 1.4 mm/s2 if payload mass  sail mass. For a payload mass of 10 kg, the acceleration would only be reduced by a factor of two. An ultrathin all-aluminum sail only 4 nm thick and 1 km on a side also weighs 10 kg and is capable of carrying a nanosat class payload. Furthermore, it is capable of 200x higher accelerations. In the future, solar sails may eventually be constructed of carbon fibers or carbon nanotubes. Thermally, carbon–carbon can survive temperatures of 3000oC before outgassing becomes excessive. Structurally, carbon–carbon has continuous use temperatures exceeding 1600oC; furthermore, its structural properties remain relatively constant over this range, unlike the structural properties of metals. If carbon–carbon can be doped or coated with a high-temperature material to enhance its solar reflectivity, then a carbon sail can probably approach within four solar radii of the sun, enabling extremely high accelerations of 5 m/s2.

13.8 TETHER PROPULSION Rob Hoyt and Les Johnson NASA Marshall Space Flight Center, Huntsville, Alabama

13.8.1

INTRODUCTION

Electrodynamic tether (EDT) propulsion uses the force the Earth’s magnetic field exerts on a wire carrying an electrical current. The effect is the basis for electric motors and generators. The technology is relatively mature and has benefited from many component-level spaceflight tests over the last two decades.

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EDTs act as energy transducers. If electrical energy (e.g., solar generated) is used to force current flow against the electromotive force (emf) created across a conducting tether by its motion through the planetary magnetic field, the interaction of the tether current with the planetary field creates thrust. Conversely, if the motion-generated emf is allowed to drive current through the tether, mechanical energy from planetary rotation is converted into electrical energy, producing drag thrust. Because the system reacts against the planetary spin motion, it requires no propellant. The equivalent specific impulse can exceed 1  105 s, compared with 300 s for chemical thrusters and 5  103 s for solar-electric propulsion. EDT propulsion can be used for satellite deorbit, orbit transfer, orbit maintenance, electrodynamic orbit capture, and, because it works wherever there is a magnetic field and an ionosphere, planetary exploration missions. An EDT upper stage could be used as an orbit transfer vehicle (OTV) to move payloads within LEO. The OTV would rendezvous with the payload and maneuver it to a new orbital altitude or inclination without the use of propellant. The tug could then lower its orbit to rendezvous with the next payload and repeat the process. Conceivably, such a system could perform several orbital maneuvering assignments each year without resupply, making it a low-recurring-cost space asset. The environment of the Jovian system has properties that are particularly favorable for use of an EDT. Jupiter has a strong magnetic field, and the mass of the planet dictates high orbital velocities that, when combined with the planet’s rapid rotation rate, can produce very large relative velocities between the magnetic field and the spacecraft. In a circular orbit, tether propulsive forces are found to be as high as 50 N and power levels are as high as 1 megawatt. With current spacecraft being extremely limited in solar power generation because of their distance from the sun, this level of available power could enable a whole new suite of science instruments such as high-power radar.

13.8.2

MOMENTUM EXCHANGE

A spinning tether system can be used to boost payloads into higher orbits with a Hohmann-type transfer. A tether system would be anchored to a relatively large mass in LEO, awaiting rendezvous with a payload delivered to orbit. The uplifted payload meets with the tether facility, which then begins a slow spin-up using electrodynamic tethers (for propellantless operation) or another low-thrust, highspecific impulse thruster. At the proper moment and tether system orientation, the payload is released into a transfer orbit, potentially to geostationary transfer orbit or lunar transfer orbit. A network of such systems could be developed to “hand off’ a payload until it reaches the desired location. Following spin-up of the tether and satellite system, the payload is released at the local vertical. The satellite is injected into a higher orbit with perigee at the release location; the orbital tether platform is injected into a lower orbit with

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apogee at the release location. The satellite enters a geostationary transfer orbit trajectory and accomplishes the transfer in as little as 5–16 h, where the lower number applies to a single-stage system and the higher number to a two-stage system. The platform then reboosts to its operational altitude using electric thrusters. The system thus achieves transfer times comparable to a chemical upper stage with the efficiencies of electric propulsion.

13.8.3

HIGH-TENACITY FIBERS

In many space tether systems, the tether must incorporate high-tenacity materials to provide high tensile strength with minimal mass. The tenacity of the tether material is particularly important for rotating momentum-exchange tether systems, where the required mass of the tether scales as the exponent of the ratio of the square of the tether’s tip velocity to the specific tensile strength of the material. Candidate high-tenacity materials include ultra-high molecular weight polyethylene (UHMWPE), PBO (Zylon) polymer, PIPD (M5) polymer, and carbon nanotubes. Spectra and Dyneema are trade names for forms of ultra-high molecular weight polyethylene created by drawing polyethylene fibers in the solid state. The low van der Waals intermolecular forces in polyethylene enable the molecules to slip along each other easily, resulting in strong alignment of the parallel polymer chains. These fibers can achieve tensile strengths of up to 6 GPa and tensile moduli of about 124 GPa. The low chain-to-chain forces, however, allow the material to creep under load, making their use problematic for momentum-exchange tether systems, which may require portions of the tether to maintain very high loading levels for many months or years. Additionally, oriented polyethylene has very low adhesion susceptibility, making the application of protective coatings difficult or impossible without the use of surface treatments that break some of the polymer chains so as to provide chemical bonding sites. Zylon is the commercial trade name for polybenzobisoxazole (PBO) fibers produced by Toyobo of Japan. The PBO polymer was originally invented in the 1980s by the U.S. Air Force Research Laboratory and has been licensed for exclusive production by Toyobo for use in products such as bulletproof vests and composite armor. Like aramid fibers such as Kevlar, the PBO polymer achieves its strength through the alignment of rigid rod-like molecules. The PBO molecule, shown in Fig. 13.17, has a much stronger rod-like character than the aramids, and the polymer chain has no freedom to adopt folded configurations through rotation of chain bonds, therefore creating a very strong polymer alignment. The PBO molecule also is highly conjugated, meaning that its interatomic bonding is much stronger than found in polyethylene chains. It thus can achieve impressive tensile properties, with a tenacity of 5.5 GPa and an elastic modulus of 280 GPa. Its rigid polymer structure and weak van der Waals intermolecular interactions, however, result in relatively poor compressive strengths [65].

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

Fig. 13.17

731

Chemical structure of PBO [64].

Zylon also is susceptible to degradation due to UV exposure. Zylon is sold in two forms: Zylon-AS (“As Spun”), which has a bright golden color, and Zylon-HM (“High Modulus”), which undergoes a heat-treatment process to increase its tensile modulus and has a darker, almost bronze color. Magellan Systems International LLC, a majority-owned subsidiary of DuPont, is developing a process for fabricating yarns of polyf2,6-diimidazo[4,5-b:40 ,50 -e]pyridinylene-1,4(2,5-dihydroxy) phenyleneg (PIPD), marketed under the trade name M5. M5 is a fiber that was engineered by a team of scientists led by Dr. Doetze Sikkema while working for Akzo Nobel, a pharmaceuticals, coatings, and chemicals company headquartered in the Netherlands. The goal was to create a highstrength synthetic fiber suitable for use both as a ballistic fabric and a composites material. The result of this work was a unique rigid rod polymer that has hydrogen bonds in both the x and y directions (with z being the rod-like polymer chain direction). The bidirection hydrogen bonding connects the M5 molecules into a honeycomb network producing a well-oriented fiber with mechanical properties that are projected to be far superior to any other fiber on the market today. The bidirectional bonding also results in greater compressive strengths than PBO materials, on the order of 1.7 GPa. The transverse bonding in the M5 fiber is also expected to make it more amenable to application of coatings than Spectra or Zylon. Early in its development, researchers predicted that M5 could achieve tenacities exceeding those of UHMWPE and PBO. However, although limited quantities of M5 fiber have been delivered to U.S. Department of Defense customers under research contracts, M5 is not yet commercially available and the highest tenacity publicly released for experimental quantities of the fiber is 5 GPa [66]. Multiwall carbon nanotubes (MWCnTs) have the highest tensile strength of any material yet measured, with laboratory samples demonstrating strengths up to 63 GPa and theoretical strengths as high as 300 GPa [67]. However, producing nanotubes with sufficient length to be spun into useable yarns and achieving yarn

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tenacities even a small fraction of the single-nanotube value remains a significant challenge. An example of the SOA in nanotube yarns is Nanocomp Technologies, Inc., which has demonstrated fabrication of multi-kilometer lengths of yarn composed of spun millimeter-scale nanotubes, achieving tensile strengths of 3 GPa, or 0.15 N/dTex.

13.8.4

HIGH-TENSILE STRENGTH MATERIALS COMPARISON

The relevant material characteristics of PBO, UHMWPE, PIPD, and MWCnTs are summarized below in Table 13.4. In this table, we have converted the tensile and modulus data, usually expressed in GPa, into newtons per decitex to facilitate calculation of breaking strengths for the tethers.

13.8.5

TETHER CONDUCTORS

For electrodynamic tethers, high-conductivity materials such as aluminum and copper are preferred to minimize ohmic losses in the tether while also minimizing tether system mass. Aluminum wire or tape tends to be the favored conductor for most published designs for electrodynamic tethers due to its superior conductivity-per-mass. Aluminum has a resistivity of 27.4 nV . m, a density of 2700 kg/m3, and a “specific conductivity” of 13,500 m2/V . kg. Copper has a resistivity of 17.0 nV . m, a density of 8933 kg/m3, and a specific conductivity of 6585 m2/V . kg. Although its specific conductivity is half that of aluminum, the lower cost, higher flexibility, and better availability of fine-gauge copper wire can make copper competitive with aluminum from an overall system cost and performance perspective for certain applications. In applications requiring higher strength and flexibility than aluminum or copper wires can provide, metallization of high-strength polymer fibers can be used to produce conductive yarns with high tenacity, high conductivity, and excellent flexibility. Micro-Coax, Inc. produces Kevlar yarns plated with layers of nickel, copper, and silver under the trade name Aracon. The specific resistance of Aracon is typically on the order of four times that of copper wire, depending on coating materials and thickness, while providing a breaking strength roughly three times that of an equivalent copper wire. Syscom Advanced Materials, Inc., produces PBO fibers metalized with various combinations of nickel, copper, silver, and gold under the trade name AmberStrand.

13.8.6

COATINGS FOR ENVIRONMENTAL SURVIVABILITY

Tethers deployed in orbit will be subject to space environment effects (SEE) that could degrade their performance or limit their useful lifetime. In LEO, momentum-exchange tethers will be exposed to atomic oxygen (AO) that can cause erosion of high-strength polymers. AO is also a concern for electrodynamic

Material

Tenacity, N/dTex

CHARACTERISTICS OF HIGH-TENSILE-STRENGTH MATERIALS FOR SPACE TETHERS Modulus, N/dTex

Elongation at Break, %

Density, g/cm3

Degradation Temperature, 8C

UV Tolerance

Coating Acceptance

650

poor

moderate

PBO Zylonw AS w

Zylon HM

†

w

Spectra 2000

0.37

11.5

3.5

1.54

0.37

17.2

2.5

1.56

0.335

7.9

2.9

0.97

150

moderate

very poor

530

excellent

excellent

good

excellent

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

TABLE 13.4

w

M5

Experimental

.0.36

.22.3

1.5

1.7

Target

.0.61

.25.5

1.3

1.7

MWCnTs (SOA)

0.15

6

unknown

2.1

.450

†

Toyobo produces two forms of its Zylon PBO fiber: “As Spun” (AS) and “High Modulus” (HM).

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tether systems, as it can degrade polymers used on insulated conductors as well as oxidize bare-metal surfaces used for collection of currents from the ionosphere. There are several potential means for protecting tether materials from the space environment. All of these methods involve incorporating materials into the tether that either are highly resistant to oxidation or readily oxidize to form a solid and durable oxide layer that resists further erosion of the material. These protective materials can either be incorporated directly into the tether material, before the polymer is spun into yarn, or applied as a surface coating. When applied as a coating, the treatment can be applied to each individual fiber or, potentially, as a layer encasing an entire braided construction, such as an overbraid or wrapped layer of AO-resistant material on top of a hightenacity yarn. Each approach has advantages and disadvantages. Direct inclusion of the AO-resistant material into the tether polymer could provide the most robust protection, preventing AO-erosion even if micrometeoroid orbit debris impacts or surface abrasion create a hole or crack in the fiber surface. This method also avoids problems associated with differential elongation between materials under load, which can cause coatings to crack. Direct inclusion, however, adds significant complexity to the chemistries involved in fabricating the high-tenacity fibers and has the potential to degrade the strength of the polymer itself. Applying the AO/UV protective material as a surface treatment may provide a means for mitigating degradation of the fiber’s tenacity. Applying the coating or overwrap to an entire yarn or braided bundle can minimize the mass impact of the treatment, but if a micrometeoroid orbit debris impact or other phenomenon causes a hole or crack in the AO-resistant outer layer, the polymer inside will be exposed and the AO flux might erode the rest of the hightenacity material. Furthermore, the outer coating must be designed not to crack or separate when the tether is put under high load or under significant bending stress. Applying the coating to each individual fiber increases the mass impact of the treatment but can provide more robustness against damage. Coating individual fibers, however, introduces the issue of adhesion between the fibers. If the coating binds the fibers together, it can significantly increase the stiffness of the yarn and the coating can crack or cause fiber failures under bending or loading. Adhesion of the coating to the fiber is also a significant concern. The elasticity of a coating must be equal to or greater than the fibers to minimize cracking or delamination of the coating from the fiber when it is bent or put under load. For all AO/UV protection methods, a key issue for applicability to space tethers is minimizing the net impact on the strength-to-weight characteristics of the tether. This is particularly true for momentum-exchange tethers, where the extreme dependence of the total tether mass on the tether material tenacity means that any degradation in polymer strength due to processing and any

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735

Fig. 13.18 The molecular and crystal structures of M5 [69]. additional non-load-bearing mass applied to the tether will have an exponential impact on the required tether mass. Deposition of thin layers of certain metals onto the surface of the highstrength polymer fibers may protect the polymers from AO degradation because the top layer of the metallic coating will oxidize and form a stable oxide surface that resists erosion. Measurements of AO effects on Aracon have shown that the Aracon fiber actually gains mass during AO exposure, indicating the formation of an oxide layer [68]. The drawback to metallization, however, is that the metal applied to the polymer fibers has a much higher molecular weight than the fiber, and so even very thin metallic coatings can significantly impact the strength-per-weight characteristic of the tether. Polyhedral oligomeric silsesquioxanes (POSS) are known to oxidize to silicate in an AO environment, forming a protective film that resists further erosion of the material. Two forms of POSS, monohydroxy-POSS and dihidroxyPOSS, are illustrated in Fig. 13.18. The POSS molecule has a cage-like structure of silicon and oxygen, surrounded by organic groups. When POSS is exposed to atomic oxygen, the organic material is degraded and a stable silica passivation layer forms, which resists further degradation. Testing of Kapton materials incorporating POSS in the LEO environment aboard the International Space Station has demonstrated that POSS can dramatically reduce the erosion of polymeric materials by AO [70]. Integrity Testing Laboratory, Inc., has developed a surface modification process called Photosil, which incorporates silicon-containing groups into the subsurface layers of the polymer structure, creating an organo-silicon surface with a thickness of up to 1 mm. The Photosil process has been shown to be effective in protecting cords constructed of Nomex fibers from AO erosion [71]. Triton Systems, Inc., has developed a polyarylene ether benzimidazole polymer with exceptional atomic oxygen resistant attributes, marketed under the trade name TOR (Triton Oxygen Resistant) polymer. The company has produced yarns composed of TOR fibers, and these AO-resistant yarns could be used as an overwrap around high-tenacity polymer yarns to enable them to survive the space environment. TOR can also be applied as a thin coating, adding as little as 2% to the lineal density of a yarn. AO exposure testing of TOR coatings on PBO yarns has shown that TOR coatings can significantly reduce the AO degradation of high-strength polymer yarns [72]. Triton Systems has also developed a conductive TOR coating for wires and other metal surfaces that may be useful for preserving the conductivity of bare wire anodes or passive collection surfaces in electrodynamic tether systems.

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13.9 ADVANCED PROPULSION TECHNOLOGIES Sonny White NASA Johnson Space Center, Houston, Texas

13.9.1

VARIABLE SPECIFIC IMPULSE MAGNETOPLASMA ROCKET [MARK CARTER]

The variable specific impulse magnetoplasma rocket (VASIMR) is a high-power electric space propulsion engine capable of Isp/thrust modulation at constant input power. Figure 13.19 depicts a graphical representation of the system. Plasma is produced in the helicon antenna stage, and the ions are energized azimuthally by the ion cyclotron resonance heating (ICRH). Axial momentum is obtained by adiabatic expansion of plasma in a magnetic nozzle. Thrust/specific impulse ratio control is achieved by partitioning radio frequency (RF) power and propellant flow between helicon and ICRH systems. VASIMR engines have been built and tested at the 10, 20, 50, 100, and 200 kW power levels. The most advanced VASIMR prototype to date is the VX-200 device that uses a 35 kW helicon source to generate argon plasma, a 170 kW ICRH section to heat plasma, and an adiabatic magnetic nozzle to produce exit plume. The VX-200 leverages commercially developed solid-state RF generators with efficiencies of 98% and specific mass less than 1 kg/kW. It uses a cryogen-free superconducting magnet to produce fields approaching 1 tesla (T) required for helicon and ICRH

Fig. 13.19

VASIMR schematic [73].

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737

sections. The VX-200 has been operated at power levels of up to 200 kW and with performance data estimates of 5,000 s Isp and 70% efficiency. Materials needs for the VASIMR include high-temperature high-current density superconducting (HTHCDS) wire in lengths of hundreds of meters or more, cryogenic epoxy formulations suitable for locking together HTHCDS wire in magnet configurations capable of surviving launch environments, high-temperature-compliant materials, metal-to-ceramic bonding/joining techniques, and high-temperature heat pipes. Some of the materials problems are simply that they are not easily manufactured in large quantities at reasonable cost. Solutions exist, but there is no continuing research to make the material readily available cheaply and in large quantity.

13.9.2

MAGNETOPLASMADYNAMIC THRUSTERS [JAMES POLK, JPL; DAN GOEBEL, JPL; RICHARD HOFER, JPL]

Magnetoplasmadynamic (MPD) thrusters in their basic form consist of a cylindrical cathode surrounded by a cylindrical anode. A high current arc is initiated between the cathode and anode, and the electrons collide with and ionize a propellant gas generating plasma. The current returning to the power supply from the cathode generates an azimuthal magnetic field between the cathode and anode. The radial electric field and azimuthal magnetic field generates a Lorentz force on the plasma, pushing it out the back of the thruster. See Fig. 13.20 for an illustration of the thruster concept and field interactions. For low-power applications, an external magnetic field can be added. MPD thrusters will facilitate very high power-processing capability in a small volume with high efficiency and have been demonstrated at 1 MW steady-state power levels. To date, MPD thrusters have demonstrated efficiencies over 50% at Isp greater than 10,000 s and thruster power levels of multi-MW. SOA laboratory model (TRL 3–4) lithium-applied field thrusters have demonstrated efficiencies greater than 50% at 4000 secs in a 200-kW device, and modeling indicates they can achieve over 60% efficiency at power levels of 250 kW and above. MPD thrusters rely on high current arc discharges and therefore present a particularly challenging environment for electrode materials. In addition, the most promising propellant for near-term applications is lithium, and so materials exposed to the liquid in the feed system and the plasma in the discharge must be compatible with this highly reactive alkali metal. The thermionic cathode must be capable of sustaining current densities of hundreds of amperes per cm2 for thousands of hours. The SOA cathode material is pure tungsten. Reductions in cathode work function have a large impact on the operating temperature, and so schemes such as including a low work function activator like barium in the propellant flow are very attractive. Submonolayer coverage of barium on the surface of the tungsten can reduce the temperature required for a given current density by hundreds of degrees Celsius. Alternate materials that combine very high melting temperatures, low evaporation rates,

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Fig. 13.20

Magnetoplasmadynamic thruster [74].

and low work function or alternate ways of supplying activator materials to reduce the work function are of interest. The anode is subject to very high heat loads, so it must be capable of operating at high temperature for thousands of hours and must incorporate active or passive cooling. The SOA anode material is also pure tungsten, although it is very difficult to fabricate in large sizes and becomes brittle after operating at high temperatures. Tungsten anodes have been fabricated by machining large billets and by chemical vapor deposition or plasma spraying onto carbon or molybdenum mandrels. Alternate materials with high melting temperatures but greater ductility, such as rhenium or tungsten–rhenium alloys, could significantly reduce fabrication cost and risk. Passive radiative cooling has been employed, in some cases with high-emittance coatings such as zirconium diboride. Other technologies such as

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

739

black rhenium (textured rhenium deposits that provide high emittance) could be applicable. Integrated high-temperature heat pipes could be used to couple anodes to larger radiator surfaces. For example, tungsten–lithium heat pipes are capable of operating at the required temperatures and heat fluxes, although fabrication in the required configuration has not been demonstrated. Other fabrication technologies and materials enabling active cooling of anodes could be of interest. Compatibility of insulators exposed to liquid lithium in flow system components such as electromagnetic pumps and flowmeters, or to the plasma in the discharge chamber, is a key materials issue. Boron nitride has been used in laboratory-model thrusters but probably does not have sufficient life. Materials such as hafnia and yttria have demonstrated good compatibility with lithium in nuclear applications but have not been demonstrated in the MPD thruster discharge environment. Lithium is an extremely strong reducing agent, and so oxides are particularly vulnerable and other types of insulators may be preferable.

13.9.3

PULSED INDUCTIVE THRUSTER [YIANGOS MIKELLIDES, JPL; DAN GOEBEL, JPL]

The pulsed inductive thruster (PIT) is an electromagnetic plasma accelerator that discharges a short, high-current pulse using energy stored in a bank of capacitors through an inductive winding that puffs propellant onto the surface. The high fields near the surface of the coil cause inductive breakdown of the propellant. An azimuthal current sheet forms with the current in the opposite direction, resulting in magnetic pressure between the coil and the current sheet. As a result, the current sheet and entrained propellant are accelerated away from the coil generating thrust. Figure 13.21 depicts a PIT test article. The PIT has demonstrated efficiency of greater than 50% and an Isp of 2000–9000 s in a single pulse. New pulsed inductive concepts have operated at much lower stored energy (100 J vs 2–4 kJ of the high-power PIT) and provides a .3 smaller thruster operating at 20–40 less energy than the larger variant. Because the PIT is an “electrode-less” thruster, the major material degradation issues are associated with the spark gap switches and the propellant valve (rather than erosion of thruster components by plasma bombardment). The new highpower PIT designs propose to replace spark gap switches with solid-state switches (thyristors) that do not erode; however, Fig. 13.21

PIT test article [75].

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L. JOHNSON AND T. LOCKETT

Fig. 13.22

Microwave thermal launch vehicle [76].

meeting the stringent requirements of this thruster for fast current rise and speed of turnon must be demonstrated in ground tests. Gas valve seat life is another challenge for longduration missions (e.g., requiring .108–109 pulses). Several lifeextending approaches have been proposed, among which is advanced seat materials such as polyetheretherketone (PEEK), which is a polycrystalline thermoplastic nearly as hard as metals and serviceable up to 5008F.

13.9.4

BEAMED ENERGY PROPULSION [CREON LEVIT, ARC; ERIC DAVIS, EARTHTECH]

Beamed energy propulsion uses laser or microwave energy from a ground- or space-based energy source and beams it to a launch or in-space vehicle that either converts the beamed energy to electric power for an electric thruster or uses the beam energy to heat a propellant, with the advantage of higher exit velocity of exhaust products than traditional chemical propulsion. Figure 13.22 depicts a notional microwave launch system. A significant challenge, unique to the RF instantiation of this technology, is the materials technology required to manufacture a large, thin, lightweight, millimeter-wave absorbent refractory heat exchanger. The heat exchanger must efficiently absorb at least 20 MW/m2 of millimeter wave radiation beamed to the vehicle from the ground, convert this to heat, and transfer the heat to propellant at an inlet pressure of about 100 bars and outlet temperature of about 2000 K during a short (100 s), high-acceleration powered ascent. Existing materials technologies developed for other applications are likely to be sufficient. For example, integral woven ceramic composite heat exchangers have already been demonstrated with appropriate strength, weight, dimensional, refractory, and chemical properties [77]. These 3-D-woven SiC or carbon fiberreinforced ceramic composites also have versatile capabilities for bonding and/ or attach to the underlying vehicle, minimizing the otherwise vexing issues of differential thermal expansion. Although neither bulk SiC nor C are good millimeter wave absorbers, doping, layering, or simply patterning the outer surface can create an efficient microwave/millimeter wave absorbent material over a broad range of frequencies and incidence angles [78, 79]. Increased system performance may be accomplished by substituting SiC with HfC, TaC, or HfC/ TaC to increase operating temperatures to 3000 K or above, resulting in a

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

741

high-thrust launch system with ISP approaching 1000 s and payload mass fractions over 20%. Laser propulsion (e.g., lightcraft) will benefit from higher power lasers and development of novel optics/tracking and pointing systems. Ablative propellants and propellant feed systems need further development and consideration. Similar to the RF version, laser propulsion systems also have to develop efficient means to capture and transform the beamed energy into propulsive energy.

13.9.5

VERY ADVANCED PROPULSION SYSTEMS

13.9.5.1

FUSION PROPULSION [JEFF GEORGE, JSC]

Fusion propulsion involves using fusion reactions to produce the energy required for the spacecraft propulsion. Fusion technology is promising but has not yet been demonstrated in the laboratory for propulsion applications. Many of the issues to be faced by future commercial fusion-driven power plants are similar to those required for spaceflight, and it is expected that space application of the technology will advance at a rate commensurate with power generation systems on the ground. Figure 13.23 depicts plasma contained by external fields in an attempt to reach the temperature and pressure necessary to initiate and sustain fusion reaction. Fusion propulsion can be implemented in an indirect manner where the fusion reactor produces electrical power that is in turn used by an electric thruster. It can be implemented in a more direct manner where the system converts the thermal/kinetic energy (from the fusion reaction) to the plasma to kinetic energy of the vehicle by directing part of the plasma out from the reaction chamber through a magnetic nozzle. This leaking plasma generates thrust and accelerates the spacecraft. Another manner of implementation considered is to use lasers to ignite a fission/fusion reaction in a fuel pellet. The reactants are used to generate a directional thrust by means of an ablative-tolerant pusher plate or magnetic nozzle. All instantiations will need strong magnetic fields to contain and/or direct the reactants. Most fuels considered for fusion reactions generate high neutron flux/ fluences and high-energy electromagnetic radiation. Fusion propulsion will benefit from improvements in high strength at temperature materials, materials compatible with high strength magnetic fields, high-temperature high current density superconducting wire Fig. 13.23

Plasma containment [80].

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L. JOHNSON AND T. LOCKETT

Fig. 13.24 Penn State artist’s concept of an antimatter-powered Mars ship with equipment and crew landers at the right, and the engine, with magnetic nozzles, at left [81]. in lengths of hundreds of meters or more, magnets capable of generating high field strengths without failing, and radiation-tolerant materials. 13.9.5.2

ANTIMATTER PROPULSION [BOB FRISBEE]

Antimatter propulsion is the most efficient form of energy production/release known and has the possibility of converting a large percentage (up to 75%) of fuel mass into propulsive energy by annihilation of atomic particles with their antiparticles. The fundamental physics of antimatter production and annihilation are known, but no viable application of the reactions has yet been proven or demonstrated for space propulsion. Figure 13.24 shows a notional antimatter powered rocket. Proton–antiproton annihilations are used to generate charged ions that are directed to produce thrust with a magnetic nozzle. The reaction will also generate electromagnetic radiation by-products in the form of gamma rays that are not affected by the magnetic field. Improvements in hightemperature, high current density superconducting wire that is gamma ray tolerant in long lengths will be beneficial. Reduction in the cross-sectional area of the magnetic system components will reduce the cross section to gamma rays and result in reduced shielding mass. Storage techniques for the antimatter use magnetic bottles; this area will also benefit from magnetic system technologies developments. High-temperature heat rejection systems improvements will be useful, such as coatings that improve high-temperature emissivity of radiators. Some matter–antimatter systems have been considered for interstellar missions. The resultant high speeds (nontrivial fraction of the speed of light) coupled with interstellar dust and debris necessitate that the spacecraft’s geometric layout be exceedingly long compared with its “cylindrical” cross section to reduce cross-sectional area in the ram direction. This layout will benefit from active structure control and very lightweight materials. Storage of the fuels will be in the form of liquid hydrogen in a conventional tank, and the antimatter may be stored in the form of

ADVANCED MATERIALS FOR IN-SPACE PROPULSION

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antihydrogen pellets contained in magnetic storage, which will also benefit from improvements in high-temperature, high current density superconducting wire. These storage vessels will be under extreme compressive loads, so that materials with improved buckling strength capability will be needed.

13.10 SUMMARY From the high-temperature environment of a nuclear thermal rocket to the highly reflective and lightweight thin film required for a solar sail, new and advanced materials are required. The many in-space propulsion technologies currently in use or being developed for use in the near future depend on new materials and improved performance of existing materials. Technology advances are happening at a breakneck pace, and much of what is discussed in this chapter may be somewhat obsolete by the time of publication. The authors hope the reader will be able to use the material provided in this chapter as a starting point, rather than an end point, for new and exciting research and development of the materials that will be needed for the next generation of in-space propulsion systems.

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[9] Goebel, D. M., and Schneider, A., “High Voltage Breakdown and Conditioning of Carbon and Molybdenum Electrodes,” IEEE Transactions on Plasma Science, Vol. 33, 2005, pp. 1136–1148. [10] Figure 13.3, Soulas, G. C., “Improving the Total Impulse Capability of the NSTAR Ion Thruster With Thick-Accelerator-Grid Ion Optics,” 27th International Electric Propulsion Conference [online], IEPC-2001-081, Publisher’s name and city/state, URL: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20020012005.pdf [11] Soulas, G. C., Foster, J. E., and Patterson, M. J., “Performance of Titanium Optics on a NASA 30 cm Ion Thruster,” 36th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit, AIAA-2000-3814, AIAA, Reston, VA, 2000. [12] Wallace, N., “Optimization and Assessment of the Total Impulse Capability of the T6 Ion Thruster,” 30th International Electric Propulsion Conference, IEPC-2007-231, 2007, IEPC, Florence, Italy. [13] Doerner, R., White, D., and Goebel, D. M., “Sputtering Yield Measurements during Low Energy Xenon Plasma Bombardment,” Journal of Applied Physics, Vol. 93, No. 9, 2003, pp. 5816–5823. [14] Snyder, J. S., Brophy, J. R., and Anderson, J. R., “Results of a 1000-Hour Wear Test of 30-cm Carbon-Carbon Ion Optics,” 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA-2005-4394, AIAA, Reston, VA, 2005. [15] Beattie, J. S., Snyder, J. S., and Shih, W., “Manufacturing of 57 cm Carbon–Carbon Composite Ion Optics for a 20-kW Class Ion Engine,” 41st AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA-2005-4411, AIAA, Reston, VA, 2005. [16] Haag, T., and Soulas, G., “Performance and Vibration of 30 cm Pyrolytic Ion Thruster Optics,” 39th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA-2003-4557, AIAA, Reston, VA, 2003. [17] Sengupta, A., Anderson, J. R., Garner, C., Brophy, J. R., deGroh, K., Banks, B., and Karniotis, T., “Deep Space 1 Flight Spare Ion Thruster 30,000-Hour Life Test,” Journal of Propulsion and Power, 0748-4658 Vol. 25, No. 1, 2009, pp. 105–117, doi: 10.2514/1.36549. [18] Goebel, D. M., Watkins, R. M., and Jameson, K., “LaB6 Hollow Cathodes for Ion and Hall Thrusters,” Journal of Propulsion and Power, Vol. 23, No. 3, 2007, pp. 527–528. [19] Mikellides, I., and Katz, I., “Wear Mechanisms in Electron Sources for Ion Propulsion, I: Neutralizer Hollow Cathode,” Journal of Propulsion and Power, Vol. 24, No. 4, 2008, pp. 866–879, doi: 10.2514/1.33462. [20] Mikellides, I., Katz, I., Goebel, D., and Jameson, K., “Wear Mechanisms in Electron Sources for Ion Propulsion, II: Discharge Hollow Cathode,” Journal of Propulsion and Power, Vol. 24, No. 4, 2008, pp. 855–865, doi: 10.2514/1.33461. [21] Figure 13.4, Herman, D. A., “Status of the NASA’s Evolutionary Xenon Thruster (NEXT) Long-Duration Test After 30,352 Hours of Operation,” 46th AIAA/ASME/ SAE/ASEE Joint Propulsion Conference and Exhibit [online], AIAA-2010-7112, AIAA, Nashville, TN, 2010, URL: https://archive.org/details/NASA_NTRS_ Archive_20110000522 [22] Goebel, D. M., Jameson, K., Katz, I., and Mikellides, I., “Potential Fluctuations and Energetic Ion Production in Hollow Cathode Discharges,” Physics of Plasmas, Vol. 14, 103508, 2007. [23] Goebel, D., and Katz, I., “Fundamentals of Electric Propulsion Ion and Hall Thrusters,” Jet Propulsion Laboratory Science and Technology Series, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, 2008, Chap. 7.

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[24] Seikel, G. R., and Reshoktko, E., “Hall Current Accelerator,” Bulletin of the American Physical Society, Ser. II, Vol. 7, June 1962, p. 414. [25] Morosov, A. I., and Solov’ev, L. S., “A Similarity Parameter in the Theory of Plasma Flow,” Soviet Physics, Vol. 10, No. 9, 1966. [26] Artsimovich, L. A., et al., “Development of the Stationary Plasma Thruster (SPT) and Its Test on Meteor Satellite,” Journal Kosmicheskie Issledovanija, Vol. 7, No. 3, 1974, pp. 451–468. [27] Martinez-Sanchez, M., and Pollard, J. E., “Spacecraft Electric Propulsion: An Overview,” Journal of Propulsion and Power, Vol. 14, No. 5, 1998, pp. 688–699. [28] Zhurin, V. V., Kaufman, H. R., and Robinson, R. S., “Physics of Closed Drift Thrusters,” Plasma Sources Science and Technology, Vol. 8, 1999, R1. [29] Camino, O., Alonso, M., Gestal, D., de Bruin, J., Rathsman, P., Kugelberg, J., Bodin, P., Ricken, S., Blake, R., Pardo Voss, P., and Stagnaro, L., “SMART-1 Operations Experience and Lessons Learnt,” IAC-06-B5.3.08, International Astronautical Congress, Valencia, Spain, Oct. 2006. [30] Kim, V., Kozlov, V., Skrylnikov, A., Veselovzorov, A., Hilleret, N., Henrist, B., Locke, S., and Fife, J. M., “Investigation of Operation and Characteristics of Small SPT with Discharge Chamber Walls Made of Different Ceramics,” 39th AIAA Joint Propulsion Conference, AIAA-2003-5002, AIAA, Reston, VA, 2003. [31] Gascon, N., Thomas, C.A., Hermann, W.A., and Capelli, M.A., “Operating Regimes of a Linear SPT with Low Secondary Electron-Induced Wall Conductivity,” 39th AIAA Joint Propulsion Conference, AIAA-2003-5156, AIAA, Reston, VA, 2003. [32] Ahedo, E., Gallardo, J. M., and Martinez-Sanchez, M., “Effects of Radial Plasma-Wall Interaction on the Hall Thruster Discharge,” Physics of Plasmas, Vol. 10, No. 8, Aug. 2003, pp. 3397–3409. [33] Kamhawi, H., Manzella, D., Pinero, L., Haag, T., Mathers, A., and Liles, H., “In-Space Propulsion High Voltage Hall Accelerator Development Project Overview,” 46th AIAA Joint Propulsion Conference, AIAA-2010-6860, AIAA, Reston, VA, 2010. [34] Mikellides, I. G., Katz, I., Hofer, R. R., Goebel, D. M., de Grys, K., and Mathers, A., “Magnetic Shielding of the Acceleration Channel Walls in a Long-Life Hall Thruster,” 46th AIAA Joint Propulsion Conference, AIAA-2010-6942, AIAA, Reston, VA, 2010. [35] Peterson, P., and Manzella, D., “Investigation of the Erosion Characteristics of a Laboratory Hall Thruster,” 39th Joint Propulsion Conference, AIAA-2003-5005, AIAA, Reston, VA, 2003. [36] Gascon, N., Dudeck, M., and Barral, S., “Wall Material Effects in Stationary Plasma Thrusters I: Parametric Studies of an SPT-100,” Physics of Plasmas, Vol. 10, No. 4123, 2003. [37] Jiles, D., Introduction to Magnetism and Magnetic Materials, London, UK, Chapman and Hall, 1991. [38] Bugrova, A. I., Lipatov, A. S., Morozov, A. I., and Solomatina, L. V., “Global Characteristics of an ATON Stationary Plasma Thruster Operating with Krypton and Xenon,” Plasma Physics Reports, Vol. 28, No. 12, 2002, pp. 1032–1037. [39] Linnell, J. A., and Gallimore, A. D., “Krypton Performance Optimization in High-Voltage Hall Thrusters,” Journal of Propulsion and Power, Vol. 22, 2006, pp. 921–925.

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[40] Massey, D. R., King, L. B., and Makela, J. M., “Development of Direct Evaporation Bismuth Hall Thruster,” 44th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, AIAA Paper No. 2008-4520, July 2008. [41] Szabo, L, Pote, B., Paintal, S., Robin, M., Hillier, A., Branam, R. D., and Huffman, R. E., “Performance Evaluation of an Iodine-Vapor Hall Thruster,” Journal of Propulsion and Power, Vol. 28, No. 4, 2012, pp. 848–857. [42] Makela, J. M., Washeleski, R. L., Massey, D. R., King, L. B., and Hopkins, M. A., “Development of Magnesium and Zinc Hall-Effect Thruster,” Journal of Propulsion and Power, Vol. 26, No. 5, 2010, pp. 1029–1035. [43] Figure 13.5, Gerrish, H. P., Jr., “Solar Thermal Propulsion at MSFC,” [online] NASA, Huntsville, AL, 2016, pp. 4, URL: https://ntrs.nasa.gov/search.jsp?R¼20160003173 [44] Figure 13.6, NASA [online], URL: https://www.nasa.gov/centers/marshall/ multimedia/photogallery/photos/photogallery/astp/astpphotos.html. [Accessed on 5 March 2018] [45] Figure 13.7, Gerrish, H. P., Jr., “Solar Thermal Propulsion at MSFC,” [online] NASA, Huntsville, AL, 2016, pp. 4, URL: https://ntrs.nasa.gov/search.jsp?R¼20160003173 [46] Figure 13.8, Gerrish, H. P., Jr., “Solar Thermal Propulsion at MSFC,”[online] NASA, Huntsville, AL, 2016, pp. 4, URL: https://ntrs.nasa.gov/search. jsp?R¼20160003173 [47] Figure 13.9, Gerrish, H. P., Jr., “Solar Thermal Propulsion at MSFC,” [online] NASA, Huntsville, AL, 2016, pp. 4, URL: from https://ntrs.nasa.gov/search. jsp?R¼20160003173 [48] Figure 13.10, Gerrish, H. P., Jr., “Solar Thermal Propulsion at MSFC,”[online] NASA, Huntsville, AL, 2016, pp. 4, URL: https://ntrs.nasa.gov/search. jsp?R¼20160003173 [49] Figure 13.11, Houts, M. G., Borowski, S. K., George, J. A., Kim, T., Emrich, W. J., Hickman, R. R., Broadway, J. W., Gerrish, H. P., and Adams, R. B., “Nuclear Cryogenic Propulsion Stage,” Nuclear and Emerging Technologies for Space 2012, NETS 2012, NETS, The Woodlands, Texas, 2012, URL: https://ntrs.nasa.gov/ archive/nasa/casi.ntrs.nasa.gov/20120014197.pdf [50] Koenig, D. R., “Experience Gained from the Space Nuclear Rocket Program (Rover),” LA-10062-H, Los Alamos National Laboratory, Los Alamos, NM, May 1986. [51] Burkes, D. E., Wachs, D. M., Werner, J. E., and Howe, S. D., “An Overview of Current and Past W-UO2 Cermet Fuel Fabrication Technology,” INL/ CON-07-12232, Idaho National Laboratory, Idaho Falls, ID, 2007. [52] Hickman, R., Emrich, W., Litchford, R., Broadway, J., Schoenfeld, M., Houts, M., Martin, J., and Pearson, J.B., “Nuclear Thermal Propulsion (NTP) Fuel Element Development and Testing for Future Transportation Systems,” Proceedings of Nuclear and Emerging Technologies for Space 2011, Paper No. 3298, Nuclear and Emerging Technologies for Space, Albuquerque, NM, February 7–10, 2011. [53] Maxwell, J. L., et al., “A Novel Approach to Actinide Nanocomposite Fuels for Nuclear Thermal Rockets,” LA-CP 10-00759, Los Alamos National Laboratory, Los Alamos, NM, 2010. [54] Figure 13.12, Emrich, W. J., “Nuclear Thermal Rocket Element Environmental Simulator (NTREES) Upgrade Activities,” NASA Marshall Space Flight Center, Huntsville, AL, 2014, URL: https://archive.org/details/NASA_NTRS_ Archive_20140008773

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[55] Figure 13.13, Houts, M. G., Borowski, S. K., George, J. A., Kim, T., Emrich, W. J., Hickman, R. R., Broadway, J. W., Gerrish, H. P., and Adams, R. B., “Nuclear Cryogenic Propulsion Stage,” Nuclear and Emerging Technologies for Space 2012, NETS 2012, NETS, Woodland, Texas, 2012, URL: https://ntrs.nasa.gov/archive/ nasa/casi.ntrs.nasa.gov/20120014197.pdf [56] Friedman, L., “Solar Sailing: The Concept Made Realistic,” 16th AIAA Aerospace Sciences Meeting, AIAA-78-82, AIAA, New York, Jan. 1978. [57] McInnes, C., Solar Sailing: Technology, Dynamics, and Mission Applications, Springer-Verlag, Berlin, 1999. [58] Johnson, L., Young, R., and Montgomery, E., “Recent Advances in Solar Sail Propulsion Systems at NASA,” Acta Astronautica, Vol. 61, 2007, pp. 376–382. [59] Tsuda, Y., Mori, O., Funase, R., Sawada, H., Yamamoto, T., Saiki, T., Endo, T., and Kawaguchi, J., “Flight Status of IKAROS Deep Space Solar Sail Demonstrator,” 61st International Astronautical Congress, IAC-10-A3.6.8, Prague, Czech Republic, September 2010. [60] Fryer, D. S., Peters, R. D., Kim, E. U., Tomaszewski, E., de Pablo, J. J., and Nealy, P. F., “Dependence of the Glass Transition Temperature of Polymer Films on Interfacial Energy and Thickness,” Macromolecules, Vol. 34, 2001, pp. 5627–5634. [61] Kim, J. H., Jang, J., and Zin, W.-C., “Thickness Dependence of the Melting Temperature of Thin Polymer Films,” Macromolecule Rapid Communications, Vol. 22, 2001, pp. 386–389. [62] Rowe, W. M., Carroll, W.F., Ingham, J.D., Rowe, W.M., Sarbolouki, M.N., Marsh, H.E., Cuddihy, E., Hong, S.D., and Steurer, W., “Sail Film Materials and Supporting Structures for a Solar Sail: A Preliminary Design,” Vol. IV, JPL D-720-9, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, CA, Oct. 1978. [63] Scaglione, S., “The AURORA Project: Removal of Plastic Substrate to Achieve an all Metal Solar Sail,” Acta Astronautica, Vol. 44, Nos. 2–4, 1999, pp. 147–150. [64] Figure 13.17, Toyobo Co., Ltd. [online], URL: http://search.toyobo-global.com/ en_all/search.x?q¼zylon&x¼0&y¼0&ie¼UTF-8&page¼1. [Accessed on 5 March 2018] [65] Sikkema, D. J., Northolt, M. G., and Pourdeyhimi, B., “Assessment of New High-Performance Fibers for Advanced Applications,” MRS Bulletin, Vol. 28, No. 8, Aug. 2003, pp. 579–584. [66] Afshari, M., Sikkema, D. J., Lee, K., and Bogle, M., “High Performance Fibers Based on Rigid and Flexible Polymers,” Polymer Reviews, Vol. 48, 2008, pp. 230–274. [67] Yu, M. F., Lourie, O., Dyer, M. J., Moloni, K., Kelly, T. F., and Ruoff, R. S., “Strength and Breaking Mechanism of Multiwalled Carbon Nanotubes Under Tensile Load,” Science, Vol. 287, No. 5453, 2000, pp. 637–640. [68] Finckenor, M. M., Vaughn, J. A., and Watts, E., “Changes in Polymeric Tether Properties Due to Atomic Oxygen,” 41st Aerospace Sciences Meeting and Exhibit, AIAA Paper 2003-1084, Jan. 2003. [69] Figure 13.18, “M5 Fiber,” Wikipedia [online], URL: https://en.wikipedia.org/wiki/ M5_fiber [Accessed on 5 March 2018] [70] Tomczak, S. J., Wright, M. E., Guenthner, A. J., Pettys, B. J., Brunsvold, A. L., Knight, C., Minton, T. K., Vij, V., McGrath, L. M., and Mabry, J. M., “Space Survivability of Main-Chain and Side-Chain POSS-Kapton Polyimides,” AIP Conference Proceedings, Vol. 1087, 2009, pp. 505–518.

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[71] Gudimenko, Y., Ng, R., Kleiman, J.I., Iskanderova, Z.A., Tennyson, R.C., Hughes, P.C., Milligan, D., Grigorevski, A., Shuiski, M., Kiseleva, L., Edwards, D., and Finckenor, M., “Enhancement of Surface Durability of Space Materials and Structures in LEO Environment,” Ninth Symposium on Protection of Materials from Space Environment, European Space Agency, Noordwijk, Netherlands, 2003, pp. 95–106. [72] Hoyt, R. P., et al., “MXER Strength Tether Development,” Tethers Unlimited, Inc., NASA Contract NNM04AA40C, Oct. 2006. [73] Figure 13.19, Ad Astra Rocket Company [online], URL: http://www.adastrarocket. com/aarc/ImageGallery. [Accessed on 5 March 2018] [74] Figure 13.20, NASA [online], https://www.nasa.gov/centers/glenn/about/fs22grc. html. [Accessed on 5 March 2018] [75] Figure 13.21, Frisbee, R. H., and Mikellides, I. G., “The Nuclear-Electric Pulsed Inductive Thruster (NuPIT): Mission Analysis for Prometheus,” 41st AIAA/ASME/ SAE/ASEE Joint Propulsion Conference and Exhibit [online], AIAA-2005-3892, AIAA, Reston, VA, 2005, URL: https://trs.jpl.nasa.gov/handle/2014/38357 [76] Figure 13.22, Parker, K. G., and Culick, F. E. C., “Feasibility and Performance of the Microwave Thermal Rocket Launcher,” Beamed Energy Propulsion: Second International Symposium on Beamed Energy Propulsion [online], edited by K. Komurasaki, American Institute of Physics, New York, 2004, pp. 408, URL: https://www.semanticscholar.org/paper/Feasibility-and-Performance-of-theMicrowave-Therm-Parkin-Culick/987210c93533befecfee8d69dafa11344a21a95a [77] Marshall, D. B., and Cox, B. N., “Integral Textile Ceramic Structures,” Annual Review of Materials Research, Vol. 38, 2008, pp. 425–443. [78] Munk, B. A., Frequency Selective Surfaces: Theory and Design, Wiley, New York, 2000. [79] Mirotznik, M. S., Good, B. L., Ransom, P., Wikner, D., and Mait, J. N., “Broadband Antireflective Properties of Inverse Motheye Surfaces,” IEEE Transactions on Antennas and Propagation, Vol. 58, No. 9, 2010, pp. 2969–2980. [80] Figure 13.23, NASA [online], 1999, URL: https://science.nasa.gov/science-news/ science-at-nasa/1999/prop12apr99_1 [81] Figure 13.24, NASA [online], 1999, URL: https://science.nasa.gov/science-news/ science-at-nasa/1999/prop12apr99_1

CHAPTER 14

Materials for Power Systems in Space Exploration Ajay K. Misra NASA Glenn Research Center Cleveland, Ohio

14.1 INTRODUCTION As is the case with terrestrial power systems, materials play an important role in determining the performance of various space power systems. Depending on the mission application, a variety of power conversion and energy storage systems can be employed for space missions. The type of power system required for space missions is a function of the power level and duration of the mission, as shown in Fig. 14.1. For short durations and large power levels, chemical power conversion systems such as batteries and fuel cells are ideal. Solar power using photovoltaic cells is used for low to moderate power levels and for long-duration missions. For large power levels, nuclear power becomes the primary option. For longduration missions to outer planets where the intensity of sunlight is very low, nuclear power is the preferred option. For robotic missions to outer planets, which require relatively low power levels (on the order of hundreds of watts), radioisotope power systems have been used. For large power levels, such as those required for travel to Mars and surface power on Mars or the moon, nuclear fission would be required. Material systems are expected to be different for different power systems. This chapter provides an overview of materials used in various space power systems, current materials research activities to meet future space power needs, and long-term research needs to increase power density and energy density of future space power systems. Specifically, it focuses on material systems for the following power systems: solar, fuel cells, batteries, radioisotope power system (RPS), and nuclear fission.

14.2 SOLAR 14.2.1

PHOTOVOLTAIC CELL FUNDAMENTALS

Solar power results from the photovoltaic (PV) effect in which a PV cell converts sunlight into electricity. The PV cell is essentially a p-n junction semiconductor This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.

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Figure 14.1

Potential power systems for space application.

device. The first generation of materials for the PV cells is based on silicon (Si), which is still the most widely used material for terrestrial applications. The principles of the PV effect and the significance of various material properties affecting the performance of PV cells can be best explained by using Si as an example. Based on quantum theory, the energy levels in solids are grouped into “bands,” which are separated by an energy gap. For semiconductor materials like Si, two bands are of particular significance. These are the valence band and the conduction band, with an energy gap between both the bands, as shown schematically in Fig. 14.2. The valence band represents all allowable energies of valence electrons that are bound covalently to the neighboring atom. For Si, each atom shares each of its four valence electrons with each of the four other Si atoms. The conduction band represents allowable energies of electrons that have received some form of energy. The energy gap between the valence band and the conduction band is called bandgap, and the energy associated with the bandgap is called bandgap energy. For a semiconductor material to have electrical conduction, there must be electrons in the conduction band. This is possible if the atom receives energy greater Figure 14.2 Schematic of energy levels in a semiconductor material.

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than the bandgap energy, which can knock some electrons from the valence band to the conduction band. The PV cell is formed by combining a p-type semiconductor with an n-type semiconductor to create a p-n junction. For Si, an n-type semiconductor is created by doping Si with a phosphorous atom, which has one more valence electron than Si. Doping Si with P creates additional conduction electrons, which have a negative charge. A p-type semiconductor is formed by doping Si with a boron atom, which has one less valence electron than Si. This creates a hole in the Si lattice with a positive charge. When a p-n junction is created, electrons from the n-side migrate to the p-side in the vicinity of the junction and combine with holes in the p-side. Similarly, holes in the p-side migrate to the n-side in the vicinity of the junction and combine with the electrons. This results in a charge buildup in a small region adjacent to the p-n junction, with positive charge on the n-side and negative charge on the p-side. The charge buildup creates a potential barrier and voltage across the p-n junction, which can drive current flow. The basic principles are schematically shown in Fig. 14.3. When photons of sunlight strike a PV cell, only the photons with energy levels greater than or equal to the bandgap energy are able to free electrons from their atomic bonds. The release of an electron creates a hole in the Si atoms, resulting in creation of an electron–hole pair. The electrons of the electron–hole pair in the n-side of the p-n junction are repelled by the potential barrier near the p-n junction and move within the n-side. The holes of the electron–hole pair on the n-side migrate to the p-side and are free to move in the

Figure 14.3 Schematic showing a p-n junction with voltage generation due to charge buildup near interface.

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Figure 14.4

Working of the photovoltaic (PV) cell.

p-side. Similarly, holes of the electron–hole pair created on the p-side are repelled by the potential barrier near the p-n junction and move inside the p-side. The electrons of the electron–hole pair created on the p-side migrate to the n-side. When the n-type side and p-type side of the cell are connected by an external electric circuit, current flows through the circuit. Negative charges flow out of the electrode on the n-side through a load and then flow into the p-side, where they recombine with the holes in the electrode. The working of the PV cell is schematically shown in Fig. 14.4.

14.2.2

PV CELL MATERIALS USED FOR SPACE POWER

Solar cells have been extensively used and continue to be used to power various space exploration missions and space satellites. Solar power is currently the primary mode of power generation for the International Space Station. Various planetary missions, such as Mars PathFinder, Mars Exploration Rovers, Juno, and Dawn, have used solar power. Solar power is also the primary source of power for all satellites. The first generation of solar cells for powering spacecraft and satellites was based on Si, with efficiencies on the order of 15%–18%. The solar arrays on the International Space Station are built from Si cells. Today, the most common material system used for solar PV systems is multijunction cells based on III-V semiconductors [1], which are compounds combining Group 3 (Al, Ga, and In) and Group 5 (P, As, and Sb) elements of the periodic table. Multijunction cells offer many advantages for space applications, which include smaller size of solar arrays because of higher conversion efficiency (greater than 25%), high power-to-mass ratio, and resistance to space radiation exposure. The multijunction PV cells widely used for space applications today are triple-junction

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Figure 14.5

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Schematic of triple-junction cell.

GaInP/GaInAs/Ge cells, as shown in Fig. 14.5. There are excellent reviews describing various aspects of multijunction PV cells [2–5]. The following gives a brief overview of the principles behind the triple-junction GaInP/GaInAs/Ge cells used in various space missions today. The key to obtaining an efficient PV cell is to convert as much sunlight into electricity as possible. The bandgap energy of the semiconductor used in the PV cell material determines how much of the sun’s spectrum can be absorbed. Photons with less than the bandgap energy of the semiconductor material are not absorbed. Photons with too much energy compared to the bandgap energy are absorbed, but the excess energy is lost as heat. In other words, energy coming from the sunlight is either lost as heat or not converted if the photon is not perfectly matched to the bandgap energy of the absorbing semiconductor in the PV cell. Silicon, with band gap energy of 1.1 eV, can absorb only a limited portion of the incoming sunlight, as shown in Fig. 14.6. Multijunction cells achieve their high efficiency by combining several solar cells, or p-n junctions, in which each cell is composed of a different semiconductor material with different bandgap energy. The individual cells are positioned in optical series such that the top layer consists of material with the highest bandgap energy (with lower wavelength), which allows the top cell to absorb high-energy photons, while allowing the lower-energy photons to pass through. A material with slightly lower bandgap is then placed below the high bandgap junction to absorb photons with slightly less energy (longer wavelengths). The triple-junction GaInP (1.88 eV)/GaInAs (1.4 eV)/Ge (0.67 eV) cells currently used in space application can absorb a larger fraction of the incoming energy from sunlight, as shown in Fig. 14.7. Multijunction solar cells are typically fabricated by an epitaxial process in which each semiconductor layer is monolithically grown on the

Figure 14.6 Portion of sunlight effectively used to generate electricity for Si PV cell.

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Figure 14.7 Larger fraction of sunlight energy converted to electricity in triple junction cells. top of the other as one single piece by molecular organic chemical vapor deposition or molecular beam epitaxy. The transport of electrons between layers is accomplished using a tunnel junction, which is a stack of highly doped layers, producing effective potential barriers for both minority carriers. The various layers in a typical triple-junction solar cell [6] are shown in Fig. 14.8. Materials play a key role in the design of multijunction cells. It is important that multiple bandgap energy materials are available with the same or nearly the same atomic structure and lattice constant as that of the substrate material. Controlling the defect structure of each of the bandgap layers through doping and control of doping levels is important to achieve the desired properties. Controlling the defect structure of semiconductors is also important for designing the tunnel junction layer to reverse the n-p polarity between the individual cells so that current can be transmitted through the device.

14.2.3

MATERIALS FOR FUTURE PV SYSTEMS

Triple-junction GaInP/GaInAs/Ge PV cells that have been commercialized and are being used for space applications have beginning of life efficiencies on the order of 25%– 30%. Since commercialization of triple-junction cells for space applications, improvements in cell materials have resulted in Figure 14.8 Different layers in GalnP/ GalnAs/Ge triple-junction cell.

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

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efficiencies on the order of 40% [7]. Continuous increases in efficiency of multijunction PV cells are being achieved through materials advancements, particularly through modification of the defect chemistry and dopant levels in individual cell layers. The standard triple-junction solar cell used in space is Ga0.5In0.5P/ Ga0.99In0.01As/Ge. This materials combination has all materials with the same lattice constants. In other words, the materials combination is lattice matched, which results in low defect densities in the system. However, such lattice-matched materials combination puts constraints on the bandgap energies for the materials, which may not be optimum for achieving high conversion efficiency. If the latticematching constraint is removed, alternate bandgap materials can be selected to increase the conversion efficiency. This approach, called metamorphic multijunction cell design, matches the lattice constants for the top two layers but creates a lattice mismatch with the bottom layer. This approach enables the top two layers to have lower bandgaps than what is possible for Ga0.5In0.5P and Ga0.99In0.01As. Lower bandgaps are possible by increasing the In content in the top two layers. An example of a metamorphic triple-layer-junction is a combination of Ga0.35In0.65P/Ga0.83In0.17As/Ge or Ga0.35In0.65P/Ga0.83In0.17As/GaAs [8]. Because of lattice mismatch between Ga0.35In0.65P/Ga0.83In0.17As with Ge or GaAS, defects (such as dislocations) will be introduced in the PV cell. The defects are minimized by having a buffer layer between the bottom Ge and top two layers. The buffer layer is a region with a graded semiconductor composition and needs to be designed through a careful combination of semiconductor materials. Metamorphic multijunction cells have achieved over 40% conversion efficiencies [3]. Another approach for increasing the number of layers in the multijunction cell and incorporation of new bandgap materials [4, 9] is shown in Fig. 14.9. Some of the new semiconductor materials being developed for multijunction PV cells include GaInNAs and GaInNAsSb with 1 eV bandgap and reasonable lattice match with Ge. In the traditional GaInP/GaInAs/Ge triple-junction cell, there

Figure 14.9

Multijunction PV cells with more than three layers and new band gap materials.

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was a need to have a material in between GaINAs and Ge with an intermediate bandgap, and the new materials (GaInNAs and GaInNAsSb) satisfy that need. Similarly, use of a new semiconductor AlGaInP increases the top cell bandgap, resulting in an increase in conversion efficiency. Bandgap engineering will continue to increase the efficiency of multijunction solar cells. Revolutionary advances in the efficiency of solar cells are being envisioned using quantum dots [10]. Using quantum dots in a solar cell to create an intermediate band will allow the harvesting of a much larger portion of the available solar spectrum. Theoretical studies predict a potential efficiency of 63.2%, which is approximately a factor of 2 better than any state-of-the-art devices available today [11]. Quantum dots are nanoparticles of semiconducting materials with quantum optical properties that are absent in the bulk material. The first advantage of quantum dots is their tunable bandgap. It means that the wavelength at which they will absorb or emit radiation can be adjusted at will: the larger the size, the longer the wavelength of light absorbed. In principle, a mixture of quantum dots of different sizes can be used for harvesting the maximum proportion of the incident light. Whereas traditional semiconductors only produce one electron from each photon, nanometer-sized crystalline materials such as quantum dots produce more than one electron from each photon and are being developed as promising photovoltaic materials. An increase in the efficiency comes from quantum dots harvesting energy that would otherwise be lost as heat in conventional semiconductors. In quantum dot solar cells, the quantum dots are sandwiched in an intrinsic region between the PV solar cell’s ordinary p- and n-type regions, as shown in Fig. 14.10. Development of quantum-dot-based PV cells is an active area of research, and new discoveries will have a profound influence on the future of solar PV cells. Recently, researchers from the National Renewable Energy Laboratory (NREL) reported [12] the first solar cell that produces a photocurrent with external quantum efficiency greater than 100% when photoexcited with photons from the high-energy region of the solar spectrum. NREL researchers achieved the 114% external quantum efficiency with a layered cell consisting of antireflection-

Figure 14.10

Schematic of solar cell containing quantum dots.

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

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coated glass with a thin layer of a transparent conductor, a nanostructured zinc oxide layer, a quantum dot layer of lead selenide treated with ethanedithiol and hydrazine, and a thin layer of gold for the top electrode. In another research effort [13], it has been demonstrated that the quantum dots can be selectively doped to contain “a significant built-in charge” (dubbed “Q-BICs”), which forces electrons to keep bouncing around and thus minimizes recombination losses, resulting in a 45% increase in conversion efficiency. Although this research relied on an InAs/GaAs test device, it is also claimed that the technology can be applied to many different solar PV structures. Another materials development that will influence the future of solar cells for space application is in the area of thin and flexible solar cells. With efficiency of commercial III-V semiconductor space PV cells somewhere on the order of 28%, typical solar panels with honeycomb substrates can achieve a specific power of 150 W/kg and a power density of 350 W/m2 [14]. Considerable improvements in specific power and cost savings are possible if rigid solar panel materials and III-V solar cells can be replaced by blanket-like panel substrate and lightweight, flexible, multijunction solar cells. Initial research has shown [14] the feasibility of thin GaInP/GaINAs/Ge triple-junction cells with thicknesses in the range of 10–50 mm. Average conversion efficiency of 28% has been demonstrated for 50-mm-thin cells. The 50-mm cells can have specific power of 500 W/kg, and 10-mm or thinner cells can have a specific power of 2067 W/kg [14]. Recently, the concept of thin and flexible solar cells has been demonstrated with quantum dots [15] by fabricating thin-film InAs/GaAS quantum dot solar cells on mechanically flexible plastic films.

14.3 FUEL CELLS 14.3.1

BASIC PRINCIPLES

A fuel cell is an electrochemical device in which the energy of a chemical reaction can be used to produce electricity. A fuel cell operates as a continuous flow system to which reactants are continuously fed, and electricity is continuously withdrawn as a steady current. A schematic of one type of fuel cell operating with hydrogen and oxygen is shown in Fig. 14.11. One electrode is known as the fuel electrode (anode), and the other is known as the oxidant (e.g., oxygen) electrode (cathode). The reactions at different electrodes can be described as Anode: 2 H2 ¼ 4 Hþ þ 4e þ Cathode: O2 þ 4 H þ 4e ¼ 2H2 O Overall reaction: 2H2 þ O2 ¼ 2H2 O

(14:1) (14:2) (14:3)

As a system, the cell facilitates combustion of the fuel via facilitating electrochemical reactions at the anode and cathode.

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Figure 14.11

Schematic of a fuel cell.

The general design of most fuel cells is similar except for the electrolyte. Several different substances have been used as the electrolyte in fuel cells, each with their own advantages and disadvantages. The five main types of fuel cells as defined by their electrolyte are alkaline fuel cells (AFCs); solid polymer fuel cells, also known as proton exchange membrane (PEM) fuel cells; phosphoric acid fuel cells (PAFs); molten carbonate fuel cells (MCFCs); and solid oxide fuel cells (SOFCs). Alkaline and PEM fuel cells operate at lower temperature, whereas other fuel cells operate at higher temperature.

14.3.2

FUEL CELLS FOR SPACE APPLICATIONS

NASA first developed PEM fuel cells for the Gemini mission, but because PEM fuel cells had water-management problems, alkaline fuel cells were used for the Apollo mission and the space shuttle (Fig. 14.12). The alkaline fuel cell used on

a)

b)

Figure 14.12 Fuel cells used in space applications. (a) Alkaline fuel cell used in Apace Shuttle Orbiter, (b) PEM fuel cell used in Gemini spacecraft.

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759

the space shuttle is the only existing space-qualified fuel cell technology. This technology has a number of drawbacks. These include extensive operations requirement, limited life, high costs, and increasing obsolescence. For future space applications, PEM fuel cells promise to be more powerful, lighter, safer, simpler to operate, and more reliable. They will last longer, perform better, and may cost much less than current alkaline fuel cells. PEM fuel cells use hydrogen fuel and produce only water, which is so pure that NASA plans to use it as drinking water for spacecraft crews. There is currently significant commercial interest in PEM fuel cells for transportation applications. Although the basic PEM fuel cell technology that is being developed for transportation applications can be applied for space missions, there are some unique challenges for application of fuel cells in space. These more stringent space requirements include the operation with pure oxygen (instead of air) and the management of water in reduced- and zero-gravity environments. During the past several years, NASA has been advancing PEM fuel cell technology for space applications [16, 17]. SOFCs are being considered for power generation and for use in space because of their high efficiency, high power density, and extremely low pollution. They have an all-solid construction and can operate at high temperatures (600–10008C)—producing clean, efficient power from easy-to-transport fuels (such as liquid hydrocarbons) instead of pure hydrogen. SOFCs might find application for power generation in space vehicles using green propellants such as liquid methane and liquid oxygen. SOFCs are particularly attractive for generating power using propellants produced by in situ resource utilization technologies. Because of significant current interest in development of PEM and SOFC fuel cells for space applications, the material aspects of only these two types of fuel cells will be described here.

14.3.3

MATERIALS FOR PROTON EXCHANGE MEMBRANE FUEL CELL

A PEM fuel cell, schematically shown in Fig. 14.13, consists of a membrane electrode assembly (MEA) sandwiched between two flow field plates that are often mirrored to make a bipolar plate when cells are stacked in series to create a fuel cell stack. The MEA comprises a proton exchange membrane, catalyst layer, and gas diffusion layers (GDLs). The performance of the fuel cell is a strong function of the materials used in the proton exchange membrane, electrodes, and GDLs. The baseline materials for a PEM fuel cell [18] along with materials challenges are shown in Table 14.1. Future advances in PEM fuel cell performance and durability will be a strong function of the development of new material systems. Some of the materials development efforts include electrode catalysts with significant reduction in platinum (Pt) loading [19, 20], new membranes with higher conductivity and low swelling [21], and new bipolar plate materials, which include

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Figure 14.13 Schematic of single PEM fuel cell.

metals, carbon–carbon composites, and polymer composites with high electrical conductivity [22, 23].

14.3.4

MATERIALS FOR SOLID OXIDE FUEL CELL

Figure 14.14 shows schematically how a SOFC works. The cell is constructed with two porous electrodes that sandwich an electrolyte. Air flows along the cathode (which is therefore also called the “air electrode”). When an oxygen molecule contacts the cathode/electrolyte interface, it catalytically acquires four electrons from the cathode and splits into two oxygen ions. The oxygen ions diffuse into the electrolyte material and migrate to the other side of the cell where they encounter the anode (also called the “fuel electrode”). The oxygen ions encounter the fuel at the anode/electrolyte interface and react catalytically, giving off water, carbon dioxide, heat, and—most important—electrons. The electrons transport through the anode to the external circuit and back to the cathode, providing a source of useful electrical energy in an external circuit. Two commonly used design configurations for SOFCs include the tubular design (Fig. 14.15a) and the planar design (Fig. 14.15b). In the planar design, the components are assembled in flat stacks, with air and fuel flowing through channels built into the cathode and anode. The preferred planar design that is currently being developed is the anode-supported SOFC design (Fig. 14.16), in which the anode is the thickest element of the cell and provides mechanical support for the cell. The baseline materials along with materials requirements/challenges for various SOFC components [24–26] are shown in Table 14.2. Current commercial SOFC systems have low power densities, on the order of 0.2 kW/kg. Power systems for space applications will require significantly higher power density, on the order of 1 kW/kg or higher. Such increase in power density will require significant advancement in the materials used in SOFC components and innovative microstructures. Recent research [27] at NASA has shown that

BASELINE MATERIALS FOR PEM FUEL CELL COMPONENTS

Fuel cell Component

Material

Primary Function

Material Requirements

Proton exchange membrane

Nafion, a perflurosulfonic acid polymer made by Dupont

Allows protons to travel through but prohibits transport of electrons

.

Proton conductivity

.

Chemical stability

.

Mechanical stability

.

Low gas permeability

.

Low swelling

.

Proton conductivity

.

Electrical conductivity

.

Chemical stability

.

Mass transport

.

No Pt dissolution

.

Minimal Pt agglomeration

.

Corrosion resistance

.

No catalyst poisoning

Catalyst layer (active electrode)

Particles of platinum on carbon support, with trend toward using nanoparticles of platinum. Thin Nafion film binds C-supported catalyst particles.

Electrochemical reactions take place on the surface of the catalyst layer. The catalyst layer must also enable transport of protons from the membrane to the catalyst and transport of reactant and product gases to and from the catalyst layer.

761

(Continued )

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

TABLE 14.1

762

TABLE 14.1

BASELINE MATERIALS FOR PEM FUEL CELL COMPONENTS (Continued )

Fuel cell Component

Material

Primary Function

Gas diffusion layer (GDL)

Porous carbon paper or carbon cloth, might contain electrically conductive fibers

.

Ensures that reactants effectively diffuse to the catalyst layer

.

Must be an electrical conductor to transport electrons to and from the catalyst layer

Bipolar plate

Nonporous graphite plate

.

Assists in water management by allowing an appropriate amount of water to be held at the membrane for hydration

.

Primary function is to supply reactant gases to the gas diffusion layer

.

Must provide electrical connection between individual cells

Material Requirements .

Corrosion resistance

.

Good mass transport

.

Good mechanical stability

.

High electrical conductivity

.

Easy for gas flow

.

High electric conductivity

.

Low permeability to gases

.

Low weight High chemical stability and corrosion resistance Low thermal resistance

Must remove the water produced at the cathode effectively

.

Must be strong enough to support stack assembly

.

A. K. MISRA

.

.

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

Figure 14.14

a)

Schematic of a single SOFC.

b)

Figure 14.15

Schematic of SOFC stack: (a) tubular design and (b) planer design.

Figure 14.16

Anode-supported SOFC.

763

764

TABLE 14.2

BASELINE MATERIALS FOR SOFC COMPONENTS

Baseline Material

Primary Function

Material Requirements

Electrolyte

Yttria-stabilized zirconia (YSZ), Y2O3-ZrO2, dense

Transport of oxygen ions

.

High oxide ion conductivity

.

Low electronic conductivity

.

Chemical and thermal stability in air and fuel and chemical stability under oxygen potential gradient

.

Good mechanical strength

.

High electrochemical activity toward oxidation of fuel

.

Allows fuel and by-products to be delivered and removed from surface site

.

Provides path for electrons to be transported from electrolyte/anode reaction site to interconnect in SOFC stacks

.

Mixed ionic/electronic conduction with predominant electronic conduction for passage of electrons

.

No reaction with electrolyte

.

Coefficient of thermal expansion (CTE) match with other SOFC components

Anode

Ni/ytrria-stabilized zirconia (YSZ) cermet, 30% porous

Electrochemical reaction site for oxidation of fuel

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SOFC Component

Cathode

.

Ca-doped LaCrO3 (lanthanum chromite) for SOFCs operating at 10008C; Ni-based, Cr-based, and Fe-based alloys for SOFC systems operating below 8008C

.

17%–26% Cr ferritic stainless steel for lower temperatures

.

Oxide coatings for metals

Site for electrochemical reduction of oxygen

.

.

Provides electrical connection between anode of one individual cell to cathode of neighboring one Acts as physical barrier to protect air electrode material from reducing environment of fuel on the fuel electrode side, and equally prevents the fuel electrode material from contacting with oxidizing atmosphere of the oxidant electrode side

.

High electronic conduction under oxidizing condition

.

Matched coefficient of thermal expansion and chemical compatibility with electrolyte and interconnect material

.

Adequate porosity to allow gaseous diffusion to readily diffuse through cathode and cathode/electrolyte interface

.

Stability in oxidizing atmospheres at high temperatures during fabrication and operation

.

High catalytic activity for oxygen reduction

.

Excellent electrical conductivity with preferably nearly 100% electronic conduction

.

Adequate stability in terms of dimension, microstructure, chemistry, and phase at operating temperature in both reducing and oxidizing environments

.

Exceptionally low permeability for oxygen and hydrogen to minimize the direct combination of oxidant and fuel during operation

.

CTE compatibility with electrodes and electrolyte

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

Interconnect

Sr-doped LaMnO3 (LSM, lanthanum strontium manganate)

765

(Continued )

SOFC Component

Sealing material

BASELINE MATERIALS FOR SOFC COMPONENTS (Continued )

Baseline Material

.

.

Rigid seals: glass and glass-ceramic sealants, which are multicomponent silicates (e.g., BaO-Al2O3-SiO2) Compressive seals: metallic compressive seals based on silver and gold or corrugated C-shaped gaskets fabricated from superalloys, mica compressive seals

Primary Function

.

Ensures gas separation between both gas atmospheres (i.e., separating air on the cathode side and fuel gas on the anode side) and to provide electrical insulation between two successive interconnect plates

766

TABLE 14.2

Material Requirements .

No reaction or diffusion between interconnect and adjacent components

.

Fairly good thermal conductivity

.

Long-term stability under dual air/wet fuel atmospheres at operating temperature

.

Chemical stability with adjacent components

.

Electrical insulation between interconnect plates

.

Matching CTE with adjacent components

.

High bond strength to joined components or applied compressive load

.

Tolerance to thermomechanical and externally applied stress

A. K. MISRA

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

767

Figure 14.17 New bi-electrode SOFC cell concept with potential for >1 kW/kg power density.

microstructural engineering of the electrode pore structure through innovative processing techniques and elimination of the metallic interconnect offer the potential for significant increase in power density (Fig. 14.17). Materials advances will also be enabling for direct utilization of methane in the liquid methane–oxygen propulsion system using SOFCs to produce power, which will eliminate the need for an external reformer. The required materials advances include modifying the anode to make it more tolerant to carbon poisoning and residual sulfur.

14.4 BATTERIES Although Ni-Cd and Ni-H2 batteries have been used in the past for various space applications, lithium-ion batteries are currently used for space applications including satellites, Mars and lunar rovers, and many future space exploration missions. The key advantage of the Li-ion battery is significant weight reduction

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due to higher specific energy. The specific energy of the Li-ion battery is higher than 125 Wh/kg compared with 60 Wh/kg for Ni-H2 batteries, resulting in weight reduction of at least 50%. Because the Li-ion battery is the state of the art for current and future space application, this review will focus on material aspects of the Li-ion battery. A Li-ion battery (shown in Fig. 14.18) consists of a positive electrode (cathode), a negative electrode (anode), and an electrolyte separated by a microporous membrane (separator). The lithium ions transfer between the two electrodes through the electrolyte system, which is commonly a liquid solution containing a Li salt dissolved in a solvent. During the charging process, lithium ions are extracted from the cathode, transported through the electrolyte, and incorporated into the anode through a process called intercalation, in which Li atoms are incorporated into the carbon anode lattice structure. At the same time, electrons are liberated from the cathode and go through the external circuit to the anode. The reverse process takes place during the discharging process. The current collectors allow transport of electrons to and from the

Figure 14.18

Schematic of Li-ion battery.

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769

electrodes. Typical current collectors used are copper for the anode and aluminum for the cathode. The overall electrochemical performance of Li-ion batteries is a strong function of the properties of the anode and cathode materials, and advances in these materials are critical for achieving high specific energy for the batteries. The anode and cathode typically are complex composites containing active material along with polymeric binders to hold the powder structure together. In addition, the porous structure contains a conductive phase, such as carbon black, to provide electronic conductivity so that electrons can be transported to the active reaction site. Both the anode and cathode must have sufficient porosity to allow the liquid electrolyte to penetrate the powder structure so that the ions can reach the reaction sites. Some of the requirements for the anode and cathode [28, 29] are that 1) the material reacts with lithium in a reversible manner [i.e., there is no modification to the host structure during lithium insertion (intercalation) and removal (deintercalation)]; 2) the host structure must have a high chemical diffusion coefficient; 3) the material reacts with lithium very rapidly both on insertion and removal; 4) the material reacts with lithium with a high free energy of reaction; 5) the material is a good electronic conductor; and 6) the material is stable, environmentally benign, and inexpensive. The state-of-the-art commercial Li-ion batteries use carbon anodes, LiCoO2 cathodes, and LiPF6 dissolved in dimethyl or ethylene carbonate as the electrolyte. The chemical reactions for charging cycle are as follows: Positive electrode: Negative electrode: Overall reaction:

LiCoO2 ¼ Li1x CoO2 þ xLiþ þ xe C þ xLiþ þ xe ¼ CLix LiCoO2 þ C ¼ Li1x CoO2 þ CLix

(14:4) (14:5) (14:6)

The reverse reactions take place during the discharge cycle. The energy density of state-of-the-art commercial Li-ion batteries is on the order of 120 to 150 Wh/kg. Because weight is critical for space missions, future space applications will require Li-ion batteries with significantly higher energy density as well as the ability to operate safely over a wide temperature range. Advanced materials will be critical for achieving these goals, and a brief description of various advanced materials development efforts is provided in the following sections.

14.4.1

ADVANCED ANODE MATERIALS

Among potential anode materials [30], silicon has the highest theoretical specific capacity (Fig. 14.19), nearly 10 greater than the state-of-the-art carbon/graphite anode. Therefore, significant research is focused on using Si anodes [31]. However, there are major challenges with Si electrodes because of electrode degradation. This is due to the 400% volume change in Si during the insertion and extraction processes, which induces large stresses that result in cracking and

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Figure 14.19 Theoretical specific capacity of anode materials. pulverization of Si. This leads to loss of electrical contact and loss of capacity with increasing cycles. In addition to the generation of large stresses, there is the challenge related to the stability of the electrode-electrolyte interface. When the potential of the anode is below 1 V vs Li/ Liþ, the decomposition of the electrolyte forms a layer on the anode surface called solid-electrolyte interphase (SEI). The SEI stability at the interface between Si and the liquid electrolyte is a critical factor in obtaining long cycle life. However, because of the volume expansion during insertion and extraction processes, the SEI at the anode-electrolyte interface grows with time. Various degradation modes for a Si anode are shown in Fig. 14.20. Several approaches have been identified based on nanostructured Si materials to reduce or eliminate the adverse effect associated with large volume expansion for the Si anode. A few of these approaches are listed here: 1. New binders [32]: A new class of binders called polyacrylic acid (PAA) possessing certain mechanical properties has been developed for the Si anode. a)

b)

c)

Figure 14.20 Si electrode failure mechanisms: (a) material pulverization, (b) morphology and volume change of the entire Si electrode, and (c) continuous solid electrolyte interphase (SEI) growth. From [31].

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

771

The carbon-coated Si nanopowder anodes with PAA binder have shown excellent stability during the first few hundred cycles. 2. Porous Si/C Composite [33]: Preexisting pores in the carbon matrix can provide the volume needed for Si expansion and allow for fast transport of Li ions. Carbon allows for improved solid electrolyte interface formation, structural integrity, and high electrical conductivity. The concept requires 10–30 nm of Si for Li transport and porous C with interconnected porosity in the range of 34–102 nm, which poses significant fabrication challenges. A hierarchical bottoms-up approach has been developed (Fig. 14.21), in which annealed carbon black is coated with Si by chemical vapor deposition, and Si-coated carbon black particles are assembled into rigid spherical granules by a self-assembly process. 3. Si Nanowires [31]: A Si-nanowire array, a schematic of which is shown in Fig. 14.22, provides sufficient empty space between adjacent nanowires to accommodate the volume change during alloying and dealloying of Li. Each Si nanowire is connected to the metallic current collector, enabling robust electrical contact to be maintained. Nanostructured Si anodes are also more resistant to fracture than larger Si structures. The Si nanowire has been fabricated by vapor-liquid-solid and supercritical fluid-liquid-solid synthesis methods. 4. Hollow Si Nanostructures [31]: Compared to solid structures, hollow Si anode structures provide empty interior space for the volume expansion, with significantly lower diffusion-induced stress. Finite element modeling has shown a five-fold reduction in tensile stress for hollow spheres. Initial

a)

b)

c)

Figure 14.21 Schematic of Si-C nanocomposite granule formation through hierarchal bottom-up assembly: (a) annealed carbon black dendritic particles are (b) coated by Si nanoparticles and then (c) assembled into rigid spheres with open interconnected internal channels during C deposition. From [33].

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Figure 14.22 From [31].

Si nanowire anode accommodate strain due to large volume change in Si.

results show high specific capacity with long cycle life for hollow Si nanostructures. 5. Porous Si Nanowire [34–36]: Porous Si nanowire embedded in a graphene matrix is another approach for addressing the volume expansion problem associated with reaction with Li. Initially, the concept was demonstrated by fabrication of porous Si nanowire through direct etching of silicon wafer and using alginate as a binder [35]. A two-step scalable process has now been developed [34] in which silicon nanowires/nanoparticles are first doped with boron and then etched with an etchant to create a porous Si structure. Nanostructured Si anodes have demonstrated specific capacity in the range of 1000–1500 mAh/g with a 100- to 200-cycle life. Further improvements and engineering microstructures along with innovative fabrication approaches will lead to increased specific capacity and cycle life.

14.4.2

ADVANCED CATHODE MATERIALS

The cathode materials typically used in Li-ion batteries are lithiated metal oxides, LiMO2 (where M ¼ Co, Ni, Mn, or mixtures of these three elements), with LiCoO2 being the most widely used in commercial batteries today. The low specific capacity is attributed to the intrinsic structural instability of the material when more than half of the Li ions are extracted. This, along with toxicity and the cost associated with Co ions, has prompted the development of advanced cathode materials, a review of which is given by Xu and colleagues [36]. Materials of interest include Li-Ni-Co oxides and Li-Mn-Ni-Co oxides with varying chemistries. One chemistry that has received considerable attention recently is Li (Ni1/3Co1/3Mn1/3)O2, called the NMC cathode.

MATERIALS FOR POWER SYSTEMS IN SPACE EXPLORATION

773

Cathodes with solid solutions of NMC and LiMnO2 and a surface coating are promising to achieve high specific capacity [37]. Although Li-MN-Ni-Co-O compounds offer the potential of higher specific capacity, another class of cathode materials called olivines (e.g., LiFePO4), despite having low specific capacity, is attracting considerable interest because of its excellent electrochemical properties as well as low cost, nontoxicity, excellent thermal stability, and environmental friendliness. Table 14.3 shows chemistries of some of the advanced cathode materials along with their theoretical and practical specific capacities. As future space exploration missions demand Li-ion batteries with high energy density, development of cathode materials with high specific capacity cathodes remains a major challenge. Although advances are being made to develop layered complex oxides, the majority of such structures allow a maximum of one electron transfer per transition metal, which poses a fundamental barrier to increasing specific capacity. To circumvent the one electron transfer limit, new cathode materials such as metal fluorides [38] and Li2MnSiO4 silicates [39] are being developed. Metal fluorides can potentially offer three electron transfers per transition metal; however, this will require development of nanocomposites containing metal fluoride materials.

14.4.3

ADVANCED ELECTROLYTE MATERIALS

The electrolytes used in commercial Li-ion batteries are composed of LiPF6 dissolved in organic carbonates such as ethylene carbonate. This electrolyte has poor performance at low temperatures because of its high melting point. Many space exploration missions require battery performance over a wide temperature range, which could be as low as –1208C for Mars and as high as 4758C for Venus. The first generation of low-temperature electrolytes, developed by the Jet Propulsion Laboratory [40], consisted of LiPF6 dissolved in an equiproportion mixture of ethylene carbonate, diethyl carbonate, and dimethyl carbonate. Considerable enhancement in low-temperature performance was achieved with retention of 70% of the room temperature performance at –208C. Li-ion batteries for Mars exploration rovers Spirit and Opportunity were based on this low-temperature

TABLE 14.3

THEORETICAL AND PRACTICAL SPECIFIC CAPACITIES OF CATHODE MATERIALS

Cathode Material

Theoretical Specific Capacity, mAh/g

Practical Specific Capacity, mAh/g

LiCoO2

274

140

LiNi0.8Co0.2O2

274

165

Li(Ni1/3Co1/3Mn1/3)O2-Ni(NMC)O2

278

180–200

LiFePO4 (olivine)

170

160

774

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electrolyte. Subsequently, Jet Propulsion Laboratory researchers [41, 42] developed improved electrolytes with good performance up to –608C. The electrolytes consist of LiPF6 dissolved in mixtures of ethylene carbonate and ethyl methyl carbonate with and without ester solvents. Other low-temperature electrolyte development efforts [43] have been focused on lithium tetraflurooxalatophosphate(LiPF4(C2O4)-, or LiFOP-) and LiBF4-based salts. The liquid electrolytes currently employed in Li-ion batteries often involve safety issues because of their flammability, which can be mitigated by new materials along with improved designs. Gas generation in Li-ion cells, resulting from the decomposition of organic solvents under extreme conditions, affects safety. Gas production, if generated at sufficient pressure, will vent flammable solvent vapor into the surrounding environment, which can be quite explosive if there is an ignition source [44]. The flammability and the associated safety issues can be mitigated by the use of either flame-retardant additives or an inherently nonflammable electrolyte. Development efforts related to nonflammable electrolytes include ionic liquids [45] and solid electrolytes [39]. Recent development of the solid oxide [46] Li1þxþyAlxTi2–xSiyP3–yO12 with high Li-ion conductivity holds the promise for all solid-state Li-ion batteries in the future.

14.5 RADIOISOTOPE POWER SYSTEMS RPSs have been essential to the U.S. exploration of outer space [47]. The unique characteristics of these systems make them especially suited for environments in deep space where large solar arrays are not possible. So far, the United States has launched RPSs on 27 space systems, which range from navigational satellites to challenging outer-planet missions such as Pioneers 10 and 11, Voyagers 1 and 2, Galileo, Ulysses, Cassini, New Horizons Pluto, and the Mars Science Laboratory. All of these missions have used radioisotope thermoelectric generators (RTGs), shown in Fig. 14.23, that convert the heat generated by the natural decay of a radioisotope fuel (plutonium-238) into electricity through the use of thermoelectric couples.

14.5.1

THERMOELECTRIC PRINCIPLE

The thermoelectric couple, shown in Fig. 14.24, consists of p- and n-type semiconducting legs sandwiched between electrically insulating plates. The temperature difference between the top and bottom plate results in flow of electricity through the circuit. The efficiency of the thermoelectric element is largely determined by the properties of the p- and n-type leg materials and the temperature difference between the top and bottom plates. The thermoelectric efficiency of a material is quantified by a figure of merit, called zT, which contains three physical quantities: Seebeck coefficient (S), electrical conductivity (s), and thermal conductivity (k). The Seebeck coefficient tells how many volts per degree temperature

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a)

b)

Figure 14.23 RTGs flown in space missions: (a) general purpose heat source (GPHS-RTG) and (b) multimission (MMRTG).

difference are generated. The zT is related to these three parameters and absolute temperature (T ) by the equation zT ¼ S2 sT=k

(14:7)

The larger the value of zT, the greater the energy conversion efficiency of the material. The zT values of several n- and p-type thermoelectric materials are shown in Fig. 14.25.

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Figure 14.24

a)

Thermoelectric principle.

b)

Figure 14.25 Figure of merit (zT) for several thermoelectric materials: (a) n-type leg and (b) p-type leg.

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777

THERMOELECTRIC MATERIALS USED IN SPACE RADIOISOTOPE THERMOELECTRIC GENERATORS

The widely used RTG system for space exploration, general purpose heat source RPS (GPHS RTG), uses a SiGe thermoelectric element [48], the schematic of which is shown in Fig. 14.26. Two compositions of the SiGe alloy were used in the legs: 78 at.% Si for most of the length and a short segment of 63.5 at.% Si at the cold end of the couple. The purpose of using this segment with lower silicon content was to match the thermal expansion of bonded parts. The n-type material was doped with phosphorous and the p-type with boron. The SiGe couple was bonded to a cold stack assembly of tungsten, copper, and alumina parts that separated the electrical and thermal elements. On the hot side, the SiGe thermocouple was bonded to silicon–molybdenum (85 wt % Si) hot shoe. Except for the lower-Si segments and the hot shoes, a coating of silicon nitride (Si3N4) was applied to the thermocouple legs to retard Si sublimation. The design hot junction temperature was 10008C (hot-shoe temperature of 10358C) with a cold junction temperature of 3008C. Although SiGe thermocouples have been used for temperatures greater than 9008C, for mid-temperature applications (600–7008C), materials based on group-IV tellurides (such as PbTe, GeTe, or SnTe) are typically used [49]. These materials have peak zT values in the range of 0.6–0.8. One p-type telluride material, (GeTe)0.85(AgSbTe2)0.15, commonly referred to as TAGS, has a peak zT value greater than 1.2. The thermoelectric elements used for the multimission

Figure 14.26

SiGe thermocouples used in GPHS RTG.

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Figure 14.27 Thermocouple elements used in MMRTG flown on Mars Science Laboratory.

radioisotope generator (MMRTG) that is powering the Mars Science Laboratory consist of PbTe as the n-leg and a combination of TAGS and PbSnTe as the p-leg (Fig. 14.27). The hot side and cold side temperatures are 5388C and 2108C, respectively.

14.5.3

IMPACT OF MATERIALS ON FUTURE RADIOISOTOPE THERMOELECTRIC GENERATOR SYSTEMS FOR SPACE POWER

The RTG systems used so far for space exploration have only 7% thermal-to-electrical energy conversion efficiency, with power densities ranging

Figure 14.28 Advanced thermoelectric materials for achieving higher thermal-electric energy conversion efficiency.

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from 3 kW/kg for MMRTGs to 5 kW/kg for GPHS RTG. To reduce consumption of scarce radioisotope Pu-238 and mass of the RPS unit, future RTGs will require a thermoelectric system with higher thermal-to-electric energy conversion efficiency. Research and development efforts are ongoing [50] to develop thermoelectric systems with 10%–15% thermal-to-electric energy conversion efficiency. One system of interest, shown in Fig. 14.28, consists of a segmented thermoelectric couple comprising p-Zintl and n-lanthanum telluride in the hot end and alloyed CoxSby in the colder section. Several technical barriers must be overcome before such as system can be seriously considered for future RTG application. The primary barrier is achieving desired long life. Some of the life-limiting issues include sublimation of constituent elements of thermoelectric material in vacuum, stresses induced because of CTE mismatch between different elements of the thermoelectric assembly, and diffusion of elements between different thermoelectric segments. Nanostructured thermoelectric materials offer the potential to significantly increase the figure of merit (zT), as demonstrated by Wang et. al. [51] for nanostructured SiGe materials.

14.6 FISSION POWER FOR SPACE Fission power is being considered for two elements of future space exploration: one for nuclear-powered spacecraft for faster travel to interplanetary destinations, and the other for generating on-board power on spacecraft as well as surface power for long-duration stays on the moon and Mars. Nuclear-powered spacecraft will enable missions well beyond the capabilities of current chemical propulsion and significantly reduce travel time for interplanetary missions. Fission surface power will enable long-duration stays on the moon and Mars to support a human habitat. Generation of nuclear surface power on the moon and Mars requires a nuclear reactor for generating heat and a thermal-to-electric power conversion system (such as those based on thermoelectric conversion, Brayton, or Stirling cycles) for generating electricity. The only U.S. space nuclear reactor that has flown in space is SNAP-10 (SNAP refers to “system for nuclear auxiliary power”) in 1965 [52]. Also, 34 nuclear-powered Soviet spacecraft were launched between 1970 and 1989 [52]. To date, no nuclear thermal propulsion system has flown, although a system called NERVA (nuclear energy for rocket vehicle application) was built and tested with flight design components between 1955 and 1972 [53]. Recent efforts [54, 55] have focused on surface nuclear reactors for moon and Mars missions.

14.6.1

MAJOR COMPONENTS FOR SPACE FISSION REACTOR

Schematics of a space nuclear fission reactor (Fig. 14.29) show the essential elements of the basic reactor assembly, which consists of 1) a reactor core containing the nuclear fuel elements, 2) heat transfer fluid to transfer the heat from

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a)

b)

Figure 14.29 Cross section of space nuclear reactor: (a) schematic (modified from [54]) and (b) top view (modified from [54]). nuclear reaction to a power conversion or propelling system, 3) a reactor core vessel to maintain reactor pressure, 4) a reflector that reflects neutrons produced in the nuclear chain reaction back, and 5) a rotating control drum embedded in the reflector for controlling the rate of nuclear reaction. Materials requirements

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are different for different components of the reactor. The most demanding component is the reactor core containing the fuel elements, in which the materials are subjected to high temperature, stress, and radiation.

14.6.2

REACTOR CORE MATERIALS

The reactor core consists of tubes containing the nuclear fuel. The tubes containing fuels are known as fuel cladding and are typically made of metallic alloys. The rest of the reactor core structure (such as core internal support and spacers, ducting and piping for coolant flow, and the core containment vessel) is typically made from the same material as the fuel cladding. Reactor characteristics that influence materials selection include operating temperature, coolant, neutron flux, fuel type, and lifetime. The requirements for fuel cladding and other core structural materials include 1) stability under heat and radiation, 2) good mechanical strength, 3) chemical compatibility with the coolant, and 4) good heat transfer properties. The materials selection for core structural elements, including the fuel cladding, is a function of the operating temperature. A predominantly stainless steel construction would be adequate for lower operating temperatures on the order of 900 K. A refractory alloy of Nb, Ta, or Mo would be the material of choice for temperatures on the order of 1300 K. Ceramic composites would be likely candidates for temperatures beyond 1300 K. Busby and Leonard [56] provide an excellent review of materials for space fission reactors. Various alloys and composite materials that have been considered for space fission applications are shown in Table 14.4 [56]. 14.6.2.1

REFLECTOR

A reflector is a region of unfueled material surrounding the core. Its function is to scatter neutrons that leak from the core, thereby returning some of them back to the core. This design feature allows for a small core size. The reflector material must have low absorption, high reflection, radiation stability, and environmental resistance. The reflector material specified for most space reactors is Be or BeO, with Be being the heavier option. 14.6.2.2

CONTROL DRUM

Control drums embedded in the reflector are used to control the nuclear reaction. The cylindrical drums are typically made of the same material as that of the reflector (Be or BeO), with a portion of the drum consisting of boron carbide (B4C). The drums can be rotated so that either the B4C face or the Be (or BeO) face of the drum is oriented toward the reactor core. When the B4C side is rotated into place, neutrons are absorbed. When the Be (or BeO) side is rotated, neutrons are reflected back into the core. For surface power generation on the moon and Mars, a fission reactor combined with Stirling thermal-to-electric energy conversion is a viable option. The

782

TABLE 14.4 Alloy

CANDIDATE STRUCTURAL MATERIALS FOR SPACE FISSION REACTORS (FROM [53]) Density, g/cm3

Melting Point, K

Fe-18Cr-12Ni-2Mo Fe-15Cr-15Ni

7.92 7.95

1,670 1,665

Mature and proven alloy for nuclear reactors Lower swelling austenitic variant with TiC

Superalloys PE-16a Inc. 718b Inc. 617c

43Ni-17Cr-3Mo-Fe 53Ni-19Cr-2Mo-Fe 45Ni-22Cr-12Co-3Fe

8.00 8.19 8.36

1,610 1,570 1,620

Superalloy used in some reactors Commonly used high-strength alloy Better creep resistance than PE-16

Refractory alloys Nb-1Zr

Nb-1Zr

8.58

2,680

10.60 16.83 10.21 12.96

2,864 3,300 3,024 2,790

Commercial alloy, extensive database, considered in nearly all space reactor designs Greater strength and creep resistance than Nb-1Zr High-strength, commercial Ta alloy Commercially available, common Mo alloy Improved ductility over TZM

3.1

2,923

Stainless steels 316 SS

FS-85 ASTAR-811Cd TZMe Mo-41Re Composites SiC/SiC

Composition

Nb-1Zr-10W-26Ta Ta-8W-1Hf-2Re-C Mo-1Ti-0.5Zr Mo-41Re SiC fiber plus infiltrated

Remarks

Advanced composite considered for reactor applications

a

Nimonic PE-16 (Special Metals Corp.) is a Ni-based superalloy. Inconel 718 PE-16 (Special Metals Corp.) is a Ni-based superalloy. Inconel 617 (Special Metals Corp.) is a Ni-based superalloy. d ASTAR-811C is a precipitation-strengthened Ta-based alloy. e TZM Molybdenum (Ed Fagan, Inc.) is a Mo-based alloy. b

A. K. MISRA

c

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materials of construction for the thermal-to-electric conversion system must be durable under nuclear radiation conditions. The effect of radiation on various Stirling converter materials has been studied to identify materials susceptible to nuclear radiation [52]. Metallic components in Stirling converters are likely to experience minimal change in mechanical properties for fast-neutron fluence less than 1020 n/cm2. However, the radiation effects can impact the magnetic and electrical properties of metal at much lower fluence than is crucial for mechanical property integrity. The Stirling converter also contains many polymeric components used as bonds, seals, and insulation, which can be adversely affected by radiation doses as low as 1025 to 1026 rad.

14.7 CONCLUDING REMARKS Power systems are key components of spacecraft, and advanced materials will enable future space power systems. Currently, power systems constitute about 28% of the spacecraft mass, as schematically shown in Fig. 14.30. With the continuing trend toward reducing mass of the spacecraft to enable greater payload mass and reduce launch cost, dramatic increases in power density and energy density will be required for future space missions. For all power systems of interest for space applications, advanced materials will enable the achievement of high power density and energy density. Emerging technologies, including nanotechnology, computational design of materials, and innovative manufacturing concepts

Figure 14.30

Distribution of mass in a spacecraft.

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such as additive manufacturing, will have profound impact on the power density and energy density of future space power and energy conversion systems. Currently, the power systems are stand-alone units in a spacecraft. As more progress is made toward development of multifunctional materials and structures, it is expected that multifunctional structures with load bearing capability and power conversion or energy storage capability will be developed. Such multifunctional structures might include spacecraft structures with solar energy conversion capabilities and structural batteries where some of the battery components carry structural loads.

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INDEX Note: Page numbers followed by f or t (indicating figures and tables). Advanced Composite Cargo Aircraft (ACCA), 333 completed vertical tail, 335f composite fuselage, vertical tail, and fairings, 334f fuselage skins-mated, 335f Advanced electrolyte materials, 773– 774 Advanced Extremely High Frequency (AEHF), 709 Advanced nanoengineered materials, 275. See also Engineered materials aerospace commercial transport S-curve, 276f chemical processing, 290–291 future trends, 295– 296 mechanical processing, 291– 293 mixing mechanisms of co-rotating screws, 292f modeling of nanomaterials, 293– 295 multiscale modeling of structural materials, 294f nanoengineered multifunctional materials, 287– 289 nanoengineered structural materials, 277– 287 processing and manufacturing, 289– 293 solution intercalation, 290f Advanced propulsion technologies, 699, 736 beamed energy propulsion, 740–741 microwave thermal launch vehicle, 740f MPD thrusters, 737– 739 PIT, 739– 740 VASIMR, 736–737 Advanced sail materials, 727– 728 AEHF. See Advanced Extremely High Frequency

A-320neo, 591 A380 fuselage shells, 306 A4 (V2) turbopump, 678– 679 Ablative heat shields, 420, 536f Ablative nozzle liner, 629 Ablative-cooled thrust chambers, 702 ACC material. See Advanced carbon/ carbon material ACCA. See Advanced Composite Cargo Aircraft Acceleration grids, 705 – 707 ACDP. See Boeing Advance Composite Development Program ACEE. See Aircraft Energy Efficiency Program Acid pickling process, 89 Acreage, 439 Acreage TPS, 558 insulated structure, 559 – 560 internal insulation, 564 – 566 standoff TPS, 560 –564 structurally integrated TPS, 567 Adapters, 480 – 481 payload, 438 Addition polymerizations. See Chainreaction polymerization Additive manufacturing (AM), 392 additively manufactured composite processing tool, 394f ASTM process types, 393f Adhesion, 381 – 382 Adhesives, 427– 428, 446 ADI, 508 Adiabatic compression. See Rapid pressurization Advanced anode materials, 769 – 772 Advanced carbon/carbon material (ACC material), 543 Advanced cathode materials, 772 – 773

789

790

Aero engine materials. See also Composite materials General Electric aircraft engine business, 585 –586 historical perspective on aircraft engine evolution, 586 – 587 IFSD rate as function of time, 583f jet engine materials evolution, 587 – 603 materials processing, 603 – 605 passenger aircraft engines, 581 – 582 specific fuel consumption for engines, 584f T/W of military engines, 579 – 580f time series of key operating parameters, 581t titanium alloy fan disk for large turbofan engine, 582f trends and outlook, 605 –606 turbine engines, 580 – 581, 584 – 585 Aerobee, sounding rockets, 435 Aerodynamic factors, 383 Aeroengines, 41 Aerofairing, 438 Aerogel, 490 Aerojet, 709 HTPS, 709 Aeroshells, 558 load-bearing, 566 – 567 non-load-bearing, 438 Aerospace applications, 102, 215t Aerospace ceramic materials, 184 EBCs, 189, 191– 195 SiC/SiC ceramic matrix composites, 196 –200 thermal barrier coatings, 185 – 189, 190f Aerospace composites, polymers for, 146 history of resins, 146 – 147 reinforcements, 148 Aerospace industry, commercial, 344 Aerospace materials. See also Composite materials aerospace ceramic materials, 184 – 200 aluminum alloys, 13 – 38 copper alloys, 69 – 74 damage tolerance considerations for metallic materials, 75 – 84

INDEX

hydrogen embrittlement in metallic materials, 87– 111 material behavior in oxygen-rich environments, 113– 129 polymer matrix composites in aerospace structures, 158– 181 polymers and composites, 132– 156 steels, 55 – 61 superalloys, 62 – 67 titanium alloys, 41 – 54 Aerospace propulsion systems, 62 Aerospace Structural Metals Handbook, 442 Aerospace structures/systems, 277, 396 advances in, 1 polymer matrix composites in, 158– 181 Aerospace-grade superalloys, 64 AF1410 steel, 58 AFCs. See Alkaline fuel cells AFOSR. See U.S. Air Force Office of Scientific Research AFP machines. See Automated fiber placement machines AFRL. See Air Force Research Laboratory Age hardening alloys, 15 –16 Aging of alloys, 17 of polymers, 146 Air electrode, 760 Air Force, 358 Air Force Research Laboratory (AFRL), 333, 358 Air Force-developed Fe-Ni-Co type alloy steel, 58 Air-breathing hypersonic vehicles, TPS and hot structures for, 540– 541 Airbus, 306 Airbus/European Aeronautic Defence and Space (EADS), 277 astrium flexible external insulation, 560 Aircraft applications, 69 – 72 composite applications into commercial, 328f designers, 305 engine evolution, 586– 587

INDEX

fuselages, 340 propulsion systems, 1 turbine engine components, 60 turbine-powered military, 586 twin-engine, 581, 589 Aircraft Energy Efficiency Program (ACEE), 329 Airframes, 41 Airplanes, commercial, 310 AJ-10 Apollo service propulsion module engine nozzle and nozzle extension, 674f AKA Apollo 204, 114 Al-Al2O3 system, 21 Al-Li alloys. See Aluminum – lithium alloys Al-Si alloys. See Aluminum-silicon alloys Alclad 2xxx-T3 alloy, 26 Alkaline, 758 Alkaline fuel cells (AFCs), 758 Alliant Techsystems (ATK), 631 Alloy 2014, 466 Alloy 2024, 316 Alloy 2219, 465 Alloy 2224 extrusions, 317 Alloy 7079-T6 forgings, 316 Alloy 7178-T6 plate, 316 Alloying elements, 15 – 19 Alpha alloys, 46 Alumina (Al2O3), 244 Aluminum (Al), 42 – 43, 94, 158– 163, 612, 613t, 635, 686, 708 “airplane” with ceramic tiles and blankets, 543 Al-SiC lattices results generation, 270f bronzes, 69 “buy to fly” ratio, 358 honeycomb core, 484 MMC foams and lattices, 268 Aluminum alloys, 13, 94. See also Superalloys; Titanium alloys alloying elements and heat treatment, 16 –19 classification systems, 13– 15 DRA, 36 –38 high-temperature aluminum alloys, 34 –36

791

injector with baffles, 662f low-density, high-modulus Al-Li alloys, 30 – 34 properties, microstructure, and processing, 21 – 23 properties of cast aluminum alloys, 29 – 30 properties of wrought heat-treatable aluminum alloys, 23 –28 strengthening mechanisms, 19 – 21 temper designation systems, 15 2000 series of, 465 Aluminum hydride (AlH3), 613t Aluminum materials, 306, 314 advancing aluminum technology, 317f aluminum alloys and yield strength, 318f aluminum alloys evolution, 321f aluminum development, 314– 315 building block approach for airplane design, 315f current developments, 318– 321 development timeline, 316– 318 future aluminum potential, 321 new aluminum alloys, 316f strength-toughness combinations, 319f Aluminum oxide (Al2O3), 13, 596 Aluminum-based fiber metal laminates, 338 alternating layup of aluminum and glass-epoxy layers, 338f crack growth curves, 342f residual strength of GLARE and Al 2024-T6 rivet joints, 339f stress intensity factor reductions for patch repairs, 341f Aluminum-copper casting, 316 Aluminum-silicon alloys (Al-Si alloys), 29 Aluminum –lithium alloys (Al-Li alloys), 30 – 34, 33f, 320, 413, 463– 464 AM. See Additive manufacturing American F-86 Sabre jet, 586 American National Standards Institute (ANSI), 396

792

American Society for Testing and Materials (ASTM), 92 ASTM G 71 – 96, 380 ASTM-E-1559 standard, 409 ASTM-E-595 standard, 409 Ammonium nitrate (AN), 614, 614t Ammonium perchlorate (AP), 612, 613, 614t, 635 AN. See Ammonium nitrate Analysis models, 312 Anode, 764t anode-supported SOFC, 763f anode/propellant distributor, 712 ANSI. See American National Standards Institute Antimatter propulsion, 742– 743 Antimatter-powered Mars ship, 742f AO. See Atomic oxygen AP. See Ammonium perchlorate Apollo missions Apollo 1 mission, 114 Apollo 13 mission, 114, 430 Apollo Command Module heat shield, 488 –489 lunar dust experiment, 507f service module, 701 service propulsion module engine, 673f AR 70-38, 376 Aracon, 422 ARALL laminates, 338 – 342 Aramid, 148, 422 Arc-cast molybdenum, 706 Arc-jet coating evaluation of X-43 leading edges, 553 Archaic solid fuel mixtures, 611 Argon (Ar), 589 Artist’s conceptions of Solar Disc, 525f Artist’s rendition of propellant processing plant, 520f Asbestos, 621 Asbestos fiber, 620 Asbestos fiber NBR (ASNBR), 620, 621 Ashby-inspired log-log chart, 277 ASNBR. See Asbestos fiber NBR Assembly phase, 490 – 491

INDEX

ASTM. See American Society for Testing and Materials ATacMS. See U.S. Army Tactical Missile System ATJ, 631 usage in aerospace industry, 631– 632 ATK. See Alliant Techsystems Atlas launch vehicle, 469f Atlas missile, 470 Atlas turbopump, 680– 681, 681f ATLM. See Automated tape laying machines Atomic oxygen (AO), 404, 732 AO/UV protection methods, 734 Attribute limits, 216, 230 Austenitic stainless steels, 60, 102, 109– 110, 410, 441, 468, 470 screens, 661 Autoclave curing principle, 150f, 151f Autofrettage or sizing process, 479 Automated fiber placement machines (AFP machines), 356 Automated tape laying machines (ATLM), 356 Avcoat, 420 BAAM system. See Big Area Additive Manufacturing system Balloon tanks, 5, 469– 470 Bandgap, 750 Bandgap energy, 750 Barium-oxide dispenser cathodes, 707 Batteries, 749, 767 – 774. See also Fuel cells advanced anode materials, 769– 772 advanced cathode materials, 772–773 advanced electrolyte materials, 773– 774 Li-ion battery, 767– 768f bcc. See Body-centered cubic BCL. See Boron trichloride Beam, design requirements, 222t Beamed energy propulsion, 740– 741 Bearing alloys, 70 Bending-dominated foams elastic moduli of, 264 fracture toughness, 266– 267 strength of, 265 – 266

INDEX

Berylliosis, 613 Beryllium (Be), 29, 232, 449, 612, 613t, 706 Beryllium copper, 72 Beryllium hydride (BeH2), 613t Beryllium–aluminum alloy, 480 – 481 Beta-annealed heat treatment, 324 Bi-electrode SOFC cell concept, 767f Biaxial GLARE-3, 338 Big Area Additive Manufacturing system (BAAM system), 394 Binders, 611, 614 – 615t, 770 – 771 Binding agents, 611 Biocompatibility, 41 Bipolar plate, 762t Bipropellant systems, 642, 702 cryogenic, 701, 704 hypergolic, 701 Bismaleimides (BMIs), 136, 178 Bismuth, 712 – 713 Black Brant, sounding rockets, 435 Black rhenium, 739 Blade track and balance, 383 Blades, 220 – 222, 230 – 233 Blue Origin (commercial space launch companies), 649 BMC. See Bulk molding compound BMIs. See Bismaleimides BN. See Boron nitride Body-centered cubic (bcc), 43 – 44, 59 metals, 441 Boeing Boeing 707, 579 Boeing 777, 589 Boeing 787, 331, 331f conformal reusable insulation, 560 success, 310 Boeing Advance Composite Development Program (ACDP), 356 Bolted repairs, 366 Bonded repairs, 366 Booster, 435 SRBs, 609 Booster separation motor (BSM), 631 Boron (B), 148, 613t Boron nitride (BN), 196 –197, 711, 739

793

Boron trichloride (BCL), 148 BPR. See Bypass ratio BRI. See Bulk Resin Infusion Brown-out, 376 BSM. See Booster separation motor Buckling knockdown factors, 455 Buckling limited design, 247f “Buckypaper”, 284 Building block approach, 168 Bulk molding compound (BMC), 154 Bulk Resin Infusion (BRI), 336 Busek BHT-200, 709 Busek Company, 709 “Buy-to-fly” ratio, 395 Bypass ratio (BPR), 580 C/C. See Carbon/carbon C/SiC. See Carbon fibers in silicon carbide matrix C188 by Alcoa. See Alclad 2xxx-T3 alloy C47A-T851 Al-Li alloys, 320 CAI. See Composites affordability initiative; Compression after impact Calcium carbonate, 139 Calcium magnesium alumino-silicate (CMAS), 193 Candidate carbides, 720 Capability, 453 Capital costs, 517 CAPRI. See Controlled Atmospheric Pressure Resin Infusion Carbide fuels, 720 Carbide phase, 90 Carbon, 58 – 59 carbon-based polymers, 634– 635 carbon/epoxy, 158 Carbon cloth phenolic materials (CCP materials), 628, 630 Carbon dioxide (CO2), 519 Carbon fiber EPDM (CFEPDM), 620 Carbon fiber-reinforced plastics (CFRP), 49f, 279, 306, 329, 342 Carbon fibers, 327, 620 carbon fiber-based motor, 618 carbon fiber-reinforced composites, 12 composites, 158

794

Carbon fibers (Continued) candidate materials for use on launch vehicles, 160f crack blunting in composite, 162f crack growing in metallic structure, 162f examples of launch vehicles, 159f laminate cross-sectional schematic, 163f multidirectional composite and associated stress – strain curve, 161f stress – strain curves of Al/Li, 160f materials, 312 Carbon fibers in silicon carbide matrix (C/SiC), 420, 686 Carbon monoxide (CO), 513, 595 Carbon nanotubes (CNTs), 275, 390 sheet appliques on mission to Jupiter, 285 –286 Carbon/carbon (C/C), 482 combustion chamber test article, 537 Carboxyl group ( – COOH), 616 Carboxyl-terminated polybutadienes (CTPB), 615, 615t, 626 Cast alloys, 13, 14t Cast aluminum alloys, 13, 29 – 30 Cast turbine airfoil, 602f Casting alloys, 15 – 16 aluminum-copper, 316 conventional casting process, 65 die, 29 gravity, 513 – 514 investment, 29 sand, 29 techniques, 513 –514 Catalyst layer, 761t Cathodes, 707 – 708, 765t materials, 773 CCP materials. See Carbon cloth phenolic materials CE. See Cyanate ester Cell in low-density foam, 262f Cellular solids, 263 Cellular structures, 260, 262 – 263

INDEX

Centrifugal force (Cf), 370 Centrifugal processing, 293 Ceramic matrix composites (CMCs), 6, 184, 189 – 190, 237, 413, 531, 595, 605, 686 combustor liner, 596f General Electric and Materials Research & Design C/SiC body flap, 569, 570f hot structure control surface, 569 MT Aerospace C/SiC body flap, 569– 571 Ceramics, 184, 244 coatings, 405 composite system, 596 fibers, 198 materials, 11, 12, 237 CES hybrid synthesizer, 268– 269 CES Selector software application, 272 CES Selector software package, 233 CEV. See Crew exploration vehicle CEX. See Charge-exchange CFC. See Chlorofluorocarbon CFD. See Computational Fluid Dynamics CFEPDM. See Carbon fiber EPDM CFRP. See Carbon fiber-reinforced plastics; Continuous fiber reinforced polymer Chain-reaction polymerization, 140 Chamber core fairing design, 483 Charge-exchange (CEX), 706 Charged particle radiation, 404 Checkout phase, 490– 491 Chemical aging, 146 Chemical power conversion systems, 749 Chemical processing, 290– 291 Chemical propulsion, 700 materials requirement, 702– 704 pressure balance on chamber and nozzle wall, 700f pressures, temperatures, flow velocities and Mach number, 701f Chemical Systems Division (CSD), 631 Chemical vapor infiltration (CVI), 196– 197, 567

INDEX

Chemical-vapor deposition, 662 Chlorofluorocarbon (CFC), 486 – 487 Chopped fiber systems, 357 Clementine mission, 517 –518 Cloisite 30B, 389, 390 Cloisite 93A clay, 291 CMAS. See Calcium magnesium alumino-silicate CMCs. See Ceramic matrix composites CMH-17. See Composite Materials Handbook-17 CNTs. See Carbon nanotubes Co-bonded structure, 357 Co-cured skins, 370 Co-cured structure, 357 Coarse grain processing, 593 Coated materials, 260 Coating removal, 384 – 385 Coatings for environmental survivability, 732 –735 Coefficient of thermal expansion (CTE), 340, 427, 441 Cold hearth melting method, 589 Cold-wall magnetic, 514f Collected volatile condensable material (CVCM), 409 Colorless Polyimide 1, 723 Combustion, 643 chambers, 662 – 676 combustible materials, 121 –122 probability and severity for materials, 118 –120, 121t Combustor, 66, 595 – 596 Commercial purity (CP), 41 – 42 Component structural testing, 360 Composite field repair, 373 laminated skins, 357 laminates strength, 178 – 181 for launch vehicles, 163 – 168 manufacturing, 148 compression molding, 154 – 155 filament winding, 152 injection-molding, 155 liquid molding, 154 prepreg lay-up, 148, 150 – 152

795

pultrusion process, 156, 156f spray-up process, 153– 154 typical material product forms vs. process, 149t wet lay-up process, 152– 153 mechanical testing, 172 geometry dependence of compression test strength values, 176f open-hole specimens, 177f tensile strength of quasi-isotropic laminates, 174f test specimens, 178 testing notched specimens reduces variability, 173f transverse ply “failure stress” and thickness, 176f transverse tensile strength in realistic laminate, 175f process development, 356– 358 rotor blades, 369 structures, fatigue in, 170–172 tanks, 470– 475 Composite materials, 11, 209, 326. See Aerospace materials; also Aero engine materials advancing composite technology, 331f aerospace composite usage growth, 332f commercial application, 330– 333 composite applications into commercial aircraft, 328f composite matrix improvements, 337f development, 326–330 future composite potential, 336– 337 military application, 333– 336 NASA ACEE 737 carbon fiber stabilizer, 329f one-piece fuselage barrel section and CAB section, 332f Composite Materials Handbook-17 (CMH-17), 164 Composite overwrapped pressure vessels (COPVs), 451 Composites affordability initiative (CAI), 358– 359, 361f Compression after impact (CAI), 167f

796

Compression molding process, 154 –155, 155f Compressive strength, 265 – 266, 347 Compressor, 592 – 595 airfoils, 592 Computational Fluid Dynamics (CFD), 191 models of coated V-22 blade, 383 Concentrators, 717 – 718 Condenser region, 555 Conductive dissipative thermal control coatings, 416 Constitutive equation for creep deflection, 248 – 249 for creep fracture, 249 for relaxation, 249 –250 Constitutive material models, 239 Constitutive model, 241, 242t Constraints, 214t Contamination control, 403 Continuous fiber reinforced polymer (CFRP), 237 Continuous use temperature (CUT), 244 Continuum model, 260 Contour tape laminating machines (CTLMs), 356 Control configuration, 123 Control drum, 781 – 783 Control surfaces, 567 ceramic matrix composite hot structure control surface, 569 – 571 hybrid control surface, 571 – 573 image of hypersonic vehicle showing rudders, 568f insulated control surface, 568, 569f integrated fabrication, 571 Controlled Atmospheric Pressure Resin Infusion (CAPRI), 336 Controlling factors in hydrogen embrittlement hydrogen pressure, 107 – 108 operating temperature, 109 – 110 reduction in stress levels, 110 –111 Cooled cooling air, 605 Cooling methods, 643 Copolymer, 140, 141f

INDEX

Copper (Cu), 510, 686 alloys, 69, 101 aircraft applications, 69– 72 electronics and electrical applications, 72, 74 spacecraft applications, 72, 73t copper-based alloys, 69, 71t Copper chromite, 616 COPVs. See Composite overwrapped pressure vessels Core materials, 484– 485 Corrosion, 364–365 prevention, 312 processes, 75, 89 resistance and mechanical properties, 59 Corrosion-resistant steel (CRES), 59 – 60, 413, 478 Cost, 392 Counter-gravity containment melting, 514f CP. See Commercial purity Cracks, 371 blunting in composite, 162f extension, 77– 78 as function of stress cycles for metal, 445 growth, 161 in metallic structure, 162f rate, 78 –79, 106– 107 propagation, 405 “Crash” programs, 436– 437 Creep, design for, 246 candidate materials for deflectionlimited blade design, 258t case study, 255 constitutive equation for creep deflection, 248 – 249 for creep fracture, 249 for relaxation, 249–250 creep in design classes, 247f deflection-limited design, 250– 252 fan and turbine blades for gas turbines, 255– 256 fracture-limited design, 252– 253 materials selection, 246– 247 materials selection in creep regime, 250 relaxation-limited design, 253– 254

INDEX

selection chart for deflection-limited design, 256– 257f selection procedure, 254 – 255 Creep in design classes, 247f CRES. See Corrosion-resistant steel Crew exploration vehicle (CEV), 532– 533 Critical aerospace structure, 323 Critical stress intensity value, 79 – 80 “Cryo stir” procedure, 430 applications, 55 bipropellant systems, 701 cryogenic-based systems, 702 lines, 478 – 479 liquid, 700 – 704 Cryogens, 643 – 644 Cryopumping, 414 failure mode, 473 Crystalline lattice of graphite, 390 CSD. See Chemical Systems Division CTE. See Coefficient of thermal expansion CTF. See Cycles-to-failure CTLMs. See Contour tape laminating machines CTPB. See Carboxyl-terminated polybutadienes Curing process, 627 Curve shape, 371 – 372 CUT. See Continuous use temperature CVCM. See Collected volatile condensable material CVI. See Chemical vapor infiltration Cyanate ester (CE), 137 –138 Cyanoacrylates, 428 Cycles-to-failure (CTF), 105 – 106 Cyclic crack growth rates, 106 – 107, 107f Cytec FM-300 adhesives, 427 D-6AC steel alloy, 618 “D” spar, 369, 369f D6 alloys, 320, 321 Damage tolerance, 12 availability of damage-tolerance data, 84 behavior of metallic materials, 75 considerations for metallic materials, 75 crack growth rate, 78 – 79 fracture toughness, 76 –77

797

orientation, 83 – 84 processing, 82 – 83 properties of metallic materials, 81 service conditions, 81 – 82 stable crack extension, 77 – 78, 78f, 79f thickness, 83 transferability, 79 – 81 DARPA/Air Force Falcon HTV-2 vehicle, 566, 566f Dawn mission, 752 DC-3 airplane, 316 de Laval nozzle, 641 Debulking, 621 Deceleration grid, 706 Deep Space 1 probe, 424 Deflection-limited design, 250–252 Deformation, 15 Delamination, 36 Delft University of Technology (TU Delft), 338 Demonstrated limit load, 454 Density, 263 Density of states (DOS), 90 Department of Defense (DoD), 358 Department of Defense Rotor Blade Erosion Working Group, 375 Design limit load, 452 Design strength, 251– 252, 254 Determinant assembly, 355, 361 Devolatizing, 621 Die casting, 29 Dielectric properties, 268 Direct gain engines, 716 Direct melt intercalation into clay layers, 388 Direct metal laser sintering (DMLS), 695 Directed energy deposition, 395 Directionally solidified structure (DS structure), 65, 597 Discharge chamber, 708–709 Discharge channel walls, 710–711 Discontinuous reinforced metallic composites (DRX), 237 Discontinuously reinforced aluminum (DRA), 36– 38, 237 Dispersion hardening, 20 – 21

798

Displacement-limited design, 247f DMLS. See Direct metal laser sintering DoD. See Department of Defense Dopant oxides, 187 –188 Dornier 328 regional jet, 333, 334f DOS. See Density of states DRA. See Discontinuously reinforced aluminum DRX. See Discontinuous reinforced metallic composites Dry hydrogen gas environment, 94 Dry reinforcement, 152 – 153 DS structure. See Directionally solidified structure Dual-mode system, 701 Ductility, 12 Ducts, lines and, 693– 695 Durability, 443 – 446 Duralumin, 316 Dutch Space hot rudder designed for X-38, 572f Dutch space X-38 hybrid control surface, 572 –573 Dynamic modulus, 245 Dynamic pressure, 532 Dyneema, 730 E-Beam, 336 E-glass, 327 EADS. See Airbus/European Aeronautic Defence and Space Earth-based personnel, 524 Earth-derived models, 525 Earth-launched solar power satellites, 522, 524 –526 EB. See Electron beam EB-PVD. See Electron beam-physical vapor deposition EBCs. See Environmental barrier coatings ECLSS. See Environmental Control and Life Support System Economic analysis methods, 516 Economic basis development for materials, 515 –526 degree of criticality, 521 – 522

INDEX

lunar oxygen production for spacecraft propellant, 516– 519 Mars in situ propellant production, 519– 521 outposts vs. settlements, 521– 526 processing space materials on grand scale, 522– 524 in situ products, 515–516 EDT propulsion. See Electrodynamic tether propulsion EFH. See Extra full-hard 8009-T6/T851, 320 EL. See Elongation Elastic moduli, 12 of bending-dominated foams, 264 of DRA materials, 37 of stretch-dominated lattices, 264– 265 Elastic stability limit, 483 Elastic-plastic fracture mechanics, 76 Elasticity (E), 104 Elastomers, 138, 446 Electric propulsion, 699 Electrical applications, 72, 74 Electrical grounding design for multilayer insulation blanket, 418f Electro-slag remelting (ESR), 604 Electrodynamic tether propulsion (EDT propulsion), 728, 729 Electrolyte, 764t advanced electrolyte materials, 773– 774 system, 768 Electromagnetic interference (EMI), 132, 386 Electron beam (EB), 493 Electron beam-physical vapor deposition (EB-PVD), 186 Electron spectroscopy for chemical analysis (ESCA), 411 Electronics and electrical applications, 72, 74 Electroplating process, 88, 89 Electrostatic acceleration grids, 705 Elongation (EL), 91 – 92, 144 EMI. See Electromagnetic interference Emissivity, 548, 549f Empennage, 308

INDEX

Enabling Propulsion Materials Program (EPM Program), 192, 196 – 197 Energy bandgap, 750 to earth and space habitation, 522 –524 levels in semiconductor material, 750f solar, 713 thermal activation, 110 Engineered materials, 256. See also Advanced nanoengineered materials aluminum MMC foams and lattices, 268 CES hybrid synthesizer, 268 – 269f criteria of excellence, 260 – 262 density, 263 electrical properties, 268 fracture toughness, 266 –267 holes in material-property space, 260 hybrid materials, 256 – 259f mechanical properties, 263 –265 synthesizing properties of cellular structures, 262– 263 thermal properties, 267– 268 yield strength, flexural strength, and compressive strength, 265 –266 Entire heat pipe, 557 Environmental barrier coatings (EBCs), 189 EBC mechanical stability safe design approach, 195f NASA EBC technology for SiC/SiC ceramic matrix composites, 194t recession determined under high gas velocity, 192f SiC/SiC recession rates, 193f surface recession of SiO2 scales, 191f Environmental Control and Life Support System (ECLSS), 412 Environmental durability, 573 – 574 Environmental exposure, 380 – 381 Environmental factors, 655 Environmental regulations, 449 Environmental survivability, coatings for, 732 –735 EPDM. See Ethylene-propylene dimer monomer

799

EPM Program. See Enabling Propulsion Materials Program Epoxy, 136, 137t, 138t Equiaxed grain structure (EQ grain structure), 597 “Equivalent isotropic” properties, 464 Erosion, 629 performance, 624 “pit-and-groves”, 706 rain, 380– 381 sand, 375– 380 surface, 624f Erosion coatings, 381 adhesion, 381– 382 aerodynamic factors, 383 EMI, 386 environmental design considerations, 385– 386 on-aircraft repairs of polyurethane, 384f repair, 383– 384 removal, 384–385 rotor blade stripped of coating, 385f strain compatibility, 382 substrate compatibility, 382 weight and span moment, 382– 383 ESCA. See Electron spectroscopy for chemical analysis ESR. See Electro-slag remelting ET. See External Tank Ethylene, 139– 140, 635f Ethylene carbonate, 773 Ethylene-propylene dimer monomer (EPDM), 619, 620 ETOPS. See Extended twin-engine operations EVAs. See Extravehicular activities Even larger five-segment reusable solid rocket motor (RSRMV), 609 Evolutionary laws, 241 EWR 127-1 system, 450fn Expander cycle, 649– 650 Expansion nozzles, 662, 668 AJ-10 Apollo service propulsion module engine nozzle, 674f Apollo service propulsion module engine, 673f

800

Expansion nozzles (Continued) assembly of Volvo Aero tubular nozzle, 671f brazing F-1 nozzle, 671f comparison of welded sheet metal nozzle to tubular nozzle, 667f completed Volvo Aero tubular nozzle, 672f Conical vs. contoured nozzle, 669f cross section of small ablative reactioncontrol engine, 675f F-1 rocket engine, 670f Fastrac engine hot-fire test at NASA Marshall Space Flight Center, 676f NAA (Rocketdyne) A6 rocket engine, 666f replica of Goddard’s first combustion chamber and nozzle, 664f RL-10 engine with extendable carbon – carbon nozzle extension, 675f Russian combustion chamber and nozzle, 668f Russian RD107 engine showing copper alloy inner liner, 669f small thruster with beryllium combustion chamber, 674f tubular designs, 663 –664 V-2 combustion chamber and nozzle, 676f Volvo laser-welded sandwich expansion nozzle, 673f Volvo welded square tube nozzle design, 672f Expendable rocket system, 677 Exploration systems. See also Power systems materials for costs of space infrastructure, 505 economic basis development for materials processing in space, 515 –526 technical basis development for materials processing in space, 506 –515 Extended twin-engine operations (ETOPS), 582

INDEX

External Tank (ET), 440 Extra full-hard (EFH), 469 Extravehicular activities (EVAs), 415, 428 Extreme temperature and moisture, 364 F-14, 333 F-15, 333, 355, 358 F-16, 48, 320, 333, 358 F-18, 358, 360 F-22 program, 168, 276, 333, 358 F1 turbopump (Mark 10), 681– 682 FAA. See Federal Aviation Administration Fabrication, 573 phase, 490– 491 processes, 148, 662 Face center cubic (fcc), 59, 62 metals, 441 Facesheet-stabilized aluminum honeycomb, 484 Failure limited design, 247f FAJ. See Floor-mounted assembly jigs Fan, 587– 592 FARs. See Federal Aviation Regulations Fastrac engine hot-fire test at NASA Marshall Space Flight Center, 676f Fastrac program. See MC-I program Fatigue, 312, 371– 373 in composite structures, 170– 172 crack growth rate, 22, 24f behavior of 2090 alloy vs. 2124 and 7150 alloys, 32f crack growth resistance, 26 strength, 22 fcc. See Face center cubic Federal Aviation Administration (FAA), 170, 394, 441, 582 Federal Aviation Regulations (FARs), 311 Feedlines, 478– 479 FEI. See Flexible external insulation Ferric nitrate, 702 Ferric oxide, 616 Fiber glass-reinforced plastics, 142 Fiber metal laminates (FML), 338, 342 aluminum-based, 338– 342 evolution, 338 titanium-based, 342– 344

INDEX

Fiber-reinforced plastics (FRP), 277, 279 Fiber-reinforced thermoset molding, 142 Fiberglass, 148 Figure of merit (zT), 774, 775, 776f, 779 Filament winding (FW), 138, 152 Filament-wound composite motor cases, 618 Fillers, 138 –139, 140t Film cooling, 536 – 537, 600 Financial viability, 519 Finer grains, 20 Finite element analysis model for reusable solid rocket motor, 623f Fire safety in oxygen systems design, 120 control configuration, 123 importance of appropriate cleaning, 127 –128 maximize use of oxygen-compatible materials, 122 minimize ignition sources and mechanisms, 123 – 127 minimize oxygen and combustible materials, 121 – 122 oxygen compatibility assessment, 128 –129 Fission power for space, 779 major components for space fission reactor, 779– 781 reactor core materials, 781 – 783 Flammability spacecraft, 408 – 412 Flexible external insulation (FEI), 560 Flexural modulus, 146 Flexural strength, 265 – 266 Flight thrusters, 708 Floor-mounted assembly jigs (FAJ), 355 Fluidity, 29 Fluorine (Fl), 404, 635, 636t Fluorohectorite, 389 FML. See Fiber metal laminates Foams, 262 – 263, 446 Forging billet, 604 4340 M alloy, 322 4340 steel, 55 Fracture, 144 control, 403, 405 mechanics, 227

801

properties, 83 – 84, 104– 105 toughness, 76 – 77, 266– 267, 308 Fracture toughness– strength relationship, 21 – 22 Fracture-critical components, 445–446 Fracture-limited design, 252– 253 Free variables, 214, 214t, 216 “Fresh Look Study”, 524 Friction stir welding (FSW), 353, 467 FRP. See Fiber-reinforced plastics FSW. See Friction stir welding Fuel, 612– 613 chemical structure, 635f cladding, 781 electrode, 757, 760 metal fuels in composite propellants, 613t slosh, 233 sources, 611 tank of space shuttle fabricated from Al-Li based 2195 alloy, 33f Fuel cells, 749, 757, 758f materials for PEM fuel cell, 759– 760f materials for SOFCs, 760– 767 for space applications, 758 Function, 213–214t component, 218– 219 functional requirements, 214 nozzle, 628– 629 Fuselage, 308 structural requirements, 350 allowable operating stress based on axial residual strength, 352f aluminum monocoque critical failure modes, 351f composite fuselage barrel, 353f critical considerations for fuselage design, 350t Fusion propulsion, 741– 742 Future composite potential, 336 – 337 Fuzzy-fiber composite, 287 FW. See Filament winding GAG. See Ground air ground Galvanized steel, 262 GAP. See Glycidal azide polymer

802

Gas diffusion layer (GDL), 762t Gas generator (GG), 643, 652, 653 cycle, 649 – 650 Gas tungsten-arc welding (GTAW), 466 –467 Gas turbine engines, 1, 579 Gas valve seat life, 740 Gas-tungsten arc (GTA), 670 Gaseous hydrogen, 105 – 106 Gaseous oxygen environment, 120 GCP. See Glass cloth phenolic GDL. See Gas diffusion layer Geared turbo fan engine (GTF engine), 591 GECCP. See General Electric Energy Ceramic Composite Products General Electric, 394 General Electric and Materials Research & Design C/SiC body flap, 569 individual parts to assemble control surface, 570f mechanically assembled body flap, 570f General Electric Energy Ceramic Composite Products (GECCP), 569 General purpose heat source-RTG (GPHS-RTG), 775, 777 SiGe thermocouples used in, 777f GEnx LPT, 603f GEO. See Geosynchronous Earth orbit Geometric throat, 628 Geosynchronous Earth orbit (GEO), 713 GFRPs. See Glass fiber-reinforced plastics GG. See Gas generator GLARE. See Glass-Fiber Reinforced Aluminum Glaser-era satellite, 524 Glass cloth phenolic (GCP), 630 Glass fiber, 327 Glass fiber-reinforced plastics (GFRPs), 283 Glass transition temperature (Tg), 619 Glass-Fiber Reinforced Aluminum (GLARE), 306, 338 –342 patch repairs, 341 Global Hawk, 360 Glycidal azide polymer (GAP), 615t GNSs. See Graphite nanosheets

INDEX

Gold, 510 GP zones. See Guinier-Preston zones GPHS-RTG. See General purpose heat source-RTG Grain size strengthening, 18 –20 GRANTA MI, 272 Grapheme sheets, 74 Graphene, 295 Graphite, 148, 708 Graphite nanosheets (GNSs), 390 Gravity casting, 513 – 514 Gravity Probe-B, 417 GRC NEXT thruster, 705 Gridded ion thrusters, 704 materials requirements, 705– 709 NSTAR DCA assembly, 708f Gross takeoff weight (GTOW), 540 Ground air ground (GAG), 313 Ground testing, 701 Grounding, 363– 364 GTA. See Gas-tungsten arc GTAW. See Gas tungsten-arc welding GTF engine. See Geared turbo fan engine GTOW. See Gross takeoff weight Guinier-Preston zones (GP zones), 17 Guncotton. See Nitrocellulose HADOS. See Hydrogen alters density of states Hafnia, 739 Hafnium (Hf), 547 Hail and incidental impact, 363 Hall propulsion systems, 709 materials requirements, 710– 713 properties of selected HET propellants, 713t Hall Thruster Propulsion System (HTPS), 709 Hall thrusters, 710 Hall-effect thrusters (HETs), 709 Halo effect, 375 Hard alpha defect, 588 HAZ. See Heat affected zones HCF. See High-cycle fatigue HDI. See High-density inclusion Heat affected zones (HAZ), 60

INDEX

Heat flux, 533, 541, 545 Heat load, 533 Heat pipe, 555 Heat sink structure, 534 – 535, 535f Heat treatment, 15, 103 Al-Cu phase diagram with markings, 18f principal aluminum alloys, 17f solid-state precipitation sequence for aluminum alloys, 19t Heat-of-compression. See Rapid pressurization Heat-pipe-cooled leading edges, 554 – 557 Heat-pipe-cooled wing leading edge, 535, 536f HEDE model. See Hydrogen enhanced decohesion model HEE. See Hydrogen environment embrittlement Helium (He), 91 HELP model. See Hydrogen enhanced localized plasticity model Hematite (Fe2O3), 511, 512 Hereditary deformation model, 241 HETs. See Hall-effect thrusters Hierarchical bottoms-up approach, 771 High area ratio nozzle, 702 High bypass turbofan engine, 195f High fracture toughness steels, 58 High modulus organic fibers, 336 High molecular weight (HMW), 388 High strength low alloy (HSLA), 322 High thermal gradients, 553 High Velocity Air Fuel (HVAF), 382 High Velocity Oxygen Fuel (HVOF), 382 High-conductivity materials, 732 High-cycle fatigue (HCF), 105 – 106 High-density inclusion (HDI), 588, 589 High-energy-density weld processes, 495 High-frequency vibrations, 660 High-pressure fuel turbopump (HPFTP), 685 High-pressure mechanical impact tester, 116f High-pressure oxidizer turbopump (HPOTP), 684 – 685

803

High-pressure turbine (HPT), 195, 596– 602 High-shear mixing, 292 High-speed machining, 355–356 High-strength steels, 477 High-temperature aluminum alloys, 34 – 36, 37t coatings, 547 operation, 715 SiC/SiC CMM SiC-based fiber creep properties, 199t High-temperature high-current density superconducting (HTHCDS), 737 High-temperature reusable surface insulation (HRSI), 419 High-tenacity fibers, 730– 732 High-tensile strength materials comparison, 732, 733t HIPed. See Hot isostatically pressed HiPerCom SiC/SiC CMCs, 198 HMW. See High molecular weight Hohmann-type transfer, 729 Hollow cathodes, 707, 710– 711 Hollow Si nanostructures, 771– 772 Homologous temperature, 238 Honeycomb structures, repairs for, 366 Hooke’s law, 239 Hot bond repairs, 366 Hot rolling produces plate, 53 Hot streaks, 596 Hot structures, 531, 535f for air-breathing hypersonic vehicles, 540– 541 components, 543 acreage TPS and aeroshells, 558– 567 control surfaces, 567– 573 leading edges, 544– 558, 559f “Hot-wet” combination, 342 HPFTP. See High-pressure fuel turbopump HPOTP. See High-pressure oxidizer turbopump HPT. See High-pressure turbine HRE. See Hydrogen reaction embrittlement

804

HRH-10 Nomex honeycomb from Hexcel, 335 HRSI. See High-temperature reusable surface insulation HSLA. See High strength low alloy HTHCDS. See High-temperature highcurrent density superconducting HTPB. See Hydroxyl-terminated polybutadiene HTPS. See Hall Thruster Propulsion System HVAF. See High Velocity Air Fuel HVOF. See High Velocity Oxygen Fuel Hybrid composites, 340 Hybrid control surface, 571 dutch space X-38 hybrid control surface, 572 –573 X-43 hybrid control surface, 571 – 572 Hybrid or nanoengineered materials, 280, 286 –287 Hybrid rocket platforms, 612 propulsion systems, 634 – 637 Hydraulic actuation, 495 Hydrazine (N2H4), 637t, 644, 701, 702 Hydrochloric acid (HCl), 512, 613 Hydrogen (H), 44, 637t effects on mechanical properties, 103 crack growth rate, 106 – 107 fracture properties, 104 – 105 low-cycle and high-cycle fatigue, 105 –106 tensile properties, 104 materials selection guide for hydrogen compatibility, 91 material database for HEE, 94, 95 –97t, 98 –100t materials screening methods, 91 –94 observations for metallic materials, 94 aluminum and aluminum alloys, 94 copper and copper alloys, 101 nickel and nickel-based alloys, 101 selecting proper materials, 103 steels, 102 superalloys, 102

INDEX

titanium and titanium alloys, 101– 102 pressure, 107– 109 turbopumps, 682– 684 Hydrogen alters density of states (HADOS), 90 Hydrogen embrittlement, 72 classification, 87 controlling factors in, 107– 111 hydrogen-embrittlement-resistant alloys, 655 in metallic materials, 87 HEE, 88– 89 HRE, 89 – 90 IHE, 89 mechanisms, 90 Hydrogen enhanced decohesion model (HEDE model), 91 Hydrogen enhanced localized plasticity model (HELP model), 91 Hydrogen environment embrittlement (HEE), 87, 88f, 88 – 89 material database for, 94, 95– 97t, 98 – 100t Hydrogen peroxide (H2O2), 635, 636t, 642 Hydrogen reaction embrittlement (HRE), 87 – 90, 88f Hydrogen-fueled rocket engine, 644 Hydrogen-fueled spacecraft, 72 Hydroxyl group (– OH), 616 Hydroxyl-terminated polybutadiene (HTPB), 615, 615t, 635 Hyper X Mach 10 vehicle, 553 Hypergolic bipropellant systems, 701 Hypergolic chemical propulsion systems, 702 Hypersonic vehicles with elevons and rudders, 567, 568f heating on reference 1-ft-diameter sphere, 533f highlighting multiple disciplines, 540f rocket and air-breathing propulsion systems, 534 technical challenges, 573 environmental durability, 573–574 fabrication, 573

INDEX

thermal management, 534 – 538 thermal protection systems for rocket-launch vehicles, 538 –539 thermal – structural challenges, 541 –543, 544f TPS and hot structure for air-breathing, 540 –541 components, 543 – 573 TSTO concept of operations, 532 Hypervelocity, 421 ICBM. See Inter-continental ballistic missile Ice protection systems (IPSs), 288, 288f Icing, 385 –386 ICRH. See Ion cyclotron resonance heating IFSD. See In-flight engine shutdown Igniters, 633 – 634 Ignition sources and mechanisms, minimize, 123 – 127 IHE. See Internal hydrogen embrittlement ILSS. See Interlaminar shear strength IMI. See Internal multiscreen insulation Impeller, 690 In situ polymerization, 290 –291 In situ propellant production, 515– 526 In Situ Resource Utilization (ISRU), 520 In-flight engine shutdown (IFSD), 583f In-service requirements, 362. See also Manufacturing requirements corrosion, 364 – 365 extreme temperature and moisture, 364 hail and incidental impact, 363 lightning and grounding, 363 – 364 repair, 365– 367 In-space propulsion, 1, 699, 700 advanced propulsion technologies, 736 –743 chemical propulsion, 700 – 704 gridded ion thrusters, 704 – 709 hall propulsion systems, 709 –713 liquid propellant combinations for, 703t NTP, 718– 721 solar sail propulsion, 722 – 728 STP, 713– 718 tether propulsion, 728 – 735

805

Inertial upper stage (IUS), 477 “Infinitely fast” modulus, 245 “Infinitely slow” modulus, 245 Inflatable concentrators, 717 Information management system, 272 Ingot metallurgy methods, 592 Ingot-based alloy, 28 Inhibitor, 626– 627 Injection-molding process, 155, 155f Injectors, 658 aluminum alloy injector with baffles, 662f Baffled, concentric-ring injector for Atlas launch vehicle, 661f coaxial injector face of J-2 engine, 663f concentric-ring injector, 660f cutaway of double-wall V-2 combustion chamber, 658f of V-2 burner cup and brass LOX injector head, 659f of V-2 combustion chamber head, 659f 3-D printed rocket injector, 663f Inner branch. See Payload attach fitting Inner mold line, 439 Insulated control surface, 568, 569f Insulated structure, 559 Boeing conformal reusable insulation, 560 EADS astrium flexible external insulation, 560 NASA advanced tile, 559– 560 Insulation materials, 619 cure, 622 devolatizing and debulking, 621 installation, 621 insulation at reusable solid rocket motor joint, 620f insulation composite materials, 620– 621 mechanical performance, 622–623 nozzle and aft case, 626f performance parameters, 622 processing, 621 RSRM components, 619f

806

Insulation materials (Continued) solid rocket motor configuration, 620f testing, 622 thermal performance, 623 – 626 Integrated fabrication, 571 Integrated powerhead demonstrator (IPD), 686 Integrity Testing Laboratory, Inc., 735 Inter-continental ballistic missile (ICBM), 469, 646 – 647 Intercalation process, 768 Interconnect, 765t Interlaminar shear strength (ILSS), 342 Interlaminar stresses, 391 Intermediate range ballistic missiles (IRBMs), 646 – 647 Internal hydrogen embrittlement (IHE), 87 –88, 88f, 89 Internal insulation, 564 MT Aerospace IMI, 565 Steve Miller and Associates Research Foundation OFI, 565 – 566 Internal multiscreen insulation (IMI), 565 International Alloy Designation System, 13 –14 International Space Station (ISS), 403, 467, 506, 554, 752 Cupola with open shutters, 422f program, 412 International Standards Organization (ISO), 76 Interstitial stabilized inclusions, 604 Intertanks, 438 Intertanks, 480 – 481 Intralaminar stresses, 391 Investment cast Ti frames, 591f Investment cast turbine airfoils, 601f Investment casting, 29 Ion cyclotron resonance heating (ICRH), 736 Ion thrusters, 704 IPD. See Integrated powerhead demonstrator IPSs. See Ice protection systems IRBMs. See Intermediate range ballistic missiles

INDEX

Iron, 511 Iron containing ilmenite (FeTiO3), 512 Iron oxide (FeO), 512 Iron-chromium-nickel (Fe-Cr-Ni), 59f Iron-nickel-chromium alloys, 60 ISO. See International Standards Organization Isostatic Molded ATJ Graphite, 631 Isostatic Molded ATJ Graphite. See ATJ Isothermal forging technique, 64 Isotropic behavior, 464 ISRU. See In Situ Resource Utilization ISS. See International Space Station Iteration, 210 IUS. See Inertial upper stage J2-X engine, 702 DMLS gas generator discharge duct, 695f engine, 694f JARs. See Joint Aviation Regulations Jet engines, 587 combustor, 595– 596 compressor, 592 – 595 fan, 587– 592 HPT, 596– 602 LPT, 602–603 material evolution, 587 Jet Propulsion Laboratory (JPL), 521, 773, 774 Joint Aviation Regulations (JARs), 311 JPL. See Jet Propulsion Laboratory JT-3C engines, 579 Juno, 752 Juno spacecraft, CNT sheet application areas on, 285, 285f Kapton, 723 Keeper electrode, 708 Kevlar, 730 Kevlarw fiber, 620 Kevlarw polymers, 619 Kevlarw fiber EPDM (KFEPDM), 620 “Kissing” bond, 366 Krumbein number, 376 Krypton, 712– 713

INDEX

L’Garde Company, 723 Large 777-class fan blades, 590f Larson-Miller plot for CVI-MI SiC/SiC CMCs, 200f Laser propulsion, 741 Lattices, 262– 263 Launch vehicle (LV), 158, 435 Aerobee sounding rocket, 436 basic material characteristics, 440 candidate materials for, 160f composites for, 163 –168 CTE, 441 durability and reusability, 443 – 446 feedlines, small lines, and pressure vessels, 478 – 479 manufacturing considerations, 490 manufacturing planning and execution, 491 –492 mechanical assembly processes, 495 –496 welding, 492 – 495 MMPDS, 442 orbital vehicles launched over two recent periods, 437t pressurized structure, 458– 477 rational methods of materials selection, 447 specialized materials, 446– 447 structural design and requirements, 447 contractual requirements, 448 – 449 laws and regulations, 449 level of conservatism, 457 pitfalls, controversies, and engineering judgment, 455 – 457 range safety, 449 – 451, 450t structural qualification, 452 –455 verification and qualification, 451 –452 structures, 438 – 440 thermal protection and insulation, 485 –490 unpressurized structure, 479 – 485 Laws and regulations, 449 LBB. See Leak-before-burst LCBT. See Low Cost Boost Technology LCF. See Low-cycle fatigue

807

LCROSS experiment, 518 LDEF satellite. See Long Duration Exposure Facility satellite Leading Edge Aviation Propulsion (LEAP), 394 Leading edges, 544 actively cooled leading edges, 557– 558, 559f effect of chord-wise position on heat flux, 546f heat-pipe-cooled leading edges, 554– 557 leading-edge thermal management options, 546f passive leading edges, 546–554 effect of radius on heat flux, 545f Leak-before-break, 227, 237 Leak-before-burst (LBB), 451, 463 LEAP. See Leading Edge Aviation Propulsion LEM. See Lunar Excursion Module LEO. See Low Earth orbit LH2. See Liquid hydrogen Li-ion battery. See Lithium-ion battery Li2MnSiO4 silicates, 773 LiFOP-based salts, 774 Light, design requirements, 222t Lightning, 363– 364, 385– 386 Lightweight (LWT), 465 nonmetallic materials, 11 Line of sight (LOS), 404 Linear-elastic fracture behavior, 81 mechanics, 76, 77 Liner testing, 626–627 Lines and ducts, 693–695 Liquid cryogenic, 700– 704 engines, 641 hydrocarbons, 759 lubricant, 424 molding process, 154 polymer, 292 storable, 700– 704 Liquid hydrogen (LH2), 421, 459, 517, 718

808

Liquid oxygen (LOX), 459, 646– 647, 517, 701 Liquid oxygen/gaseous oxygen (LOX/GOX), 405 Liquid propulsion systems, materials for general design considerations for materials selection, 649 – 654 lines and ducts, 693– 695 liquid rocket engines, 642 – 649 materials, 654 – 656 thrust chamber materials, 656 – 676 turbopump materials, 676 – 690 valves, 691 – 693 Liquid rocket engines, 437, 641 – 642 historical perspectives, 644 – 649 operational factors, 642 propellant selection, 643 – 644 Lithiated metal oxides (LiMO2), 772 Lithium, 739 Lithium tetraflurooxalatophosphatebased salts (LiPF4(C2O4)-based salts), 774 Lithium-ion battery (Li-ion battery), 767 –768f LMNS. See Lockheed Martin’s Nanosystems LMW. See Low molecular weight Load enhancement factors, 371 Load-bearing aeroshell, 558, 566 – 567 Lockheed Martin Aeronautics Company, 333 Lockheed Martin’s F-35, 279 Lockheed Martin’s Nanosystems (LMNS), 277 Long Duration Exposure Facility satellite (LDEF satellite), 414 Loral Space Systems, 709 LOS. See Line of sight Lost wax, 65 Low Cost Boost Technology (LCBT), 685 –686 Low Earth orbit (LEO), 516, 713 Low Earth Orbit Spacecraft Charging Design Handbook, 404 Low molecular weight (LMW), 389 Low Z materials, 420

INDEX

Low-carbon iron-nickel lath martensite, 58 – 59 Low-cycle fatigue (LCF), 105– 106, 593 Low-density, high-modulus Al-Li alloys, 30 – 34 Low-pressure turbine (LPT), 49 – 50, 591, 602– 603 LOX. See Liquid oxygen LOX/GOX. See Liquid oxygen/gaseous oxygen LPT. See Low-pressure turbine Lubricants, 424– 427 Lunar Excursion Module (LEM), 647 Lunar geology, 507 Lunar oxygen production for spacecraft propellant, 516– 519 Lunar Prospector orbiter, 518 LV. See Launch vehicle LWT. See Lightweight 3M 9406PC adhesives, 428 3M 966 adhesives, 428 3M 9703 adhesives, 428 3M AF-191 film adhesives, 427 M5 fiber, 731, 735f M-5TM high modulus organic fibers, 336 Mach 0.3 particulate erosion, 189 Machinability, 440 Magnesium (Mg), 613, 613t, 712– 713 Magnesium fluoride, 423 Magnetic circuit elements, 711–712 Magnetic containment melting (MCM), 513 Magnetic layer thrusters (MLTs), 709 Magnetic storage, 743 Magnetite (Fe3O4), 511 Magnetoplasmadynamic thrusters (MPD thrusters), 737– 739 Manufacturability, 405 factors to consider for, 406 – 408t Manufacturing requirements, 354. See also In-service requirements assembly, 359– 362 automated layup methods, 356f

INDEX

automated splice fastening eliminates variability of hand installations, 360f CAI, 358– 359 composite process development, 356 –358 conventional 777 skin, stringer and frame construction, 362f determinant assembly, 355 future monolithic frame and shear tie construction, 362f high-speed machining, 355 – 356 machining of large wing panels, 355f roll forming, 354 – 355 MAPO. See Tris[1-(methyl) aziridinyl] phosphine oxide MAPTIS. See Materials and Processes Technical Information System Maraging steels, 55, 58 – 59 Margin of safety, 453 Mars Exploration Rovers, 752 Mars in situ propellant production, 519 –521 Mars PathFinder mission, 752 Mars Phoenix Mission Lander arm camera, 521f Marshall Space Flight Center (MSFC), 117, 118f Martian atmosphere, 519 Mass Loss, 379 Material Data Management Consortium, 272 Material(s), 579, 749, 754 data life cycle, 271f, 271 – 272 extrusion systems, 394 for future PV systems, 754 – 757 for high-temperature applications, 238 constitutive models and material parameters, 242t hereditary material behavior, 240f homologous temperature, 238 – 239 maximum and minimum service temperatures, 241 – 246 Norton– Bailey creep model, 240 –241 index, 217, 219f, 230, 231t

809

information, 211, 271f performance indices, 218 for bending at room temperature, 222– 223 blades, 220– 222 component function, 218– 219 pressure vessels, 224– 228 turbine jet engines, 221f for vibration at room temperature, 223– 224 processing, 603– 605 economic basis development for, 515– 526 technical basis development for, 506– 515 property charts, 228–230 modulus vs. density property chart, 229f removal processes, 212 selection advanced selection, 238 for aerospace systems, 209 core of materials selection, 211–212 data issues, 269– 272 design, 209– 210, 246– 256 engineered materials, 256– 269 life-cycle cost, 210 materials for high-temperature applications, 238–246 rational methods, 447 strategy, 213 systematic approach to materials selection, 212– 238 systems, 605 Materials and Processes Technical Information System (MAPTIS), 129, 409, 625 Materials on International Space Station Experiment (MISSE), 412 Materials Research & Design (MR&D), 569 Matrix, 146– 147 alloys, 36 matrix-dominated properties, 278 matrix-reinforcement interface, 36 MC-I program, 664

810

MCFCs. See Molten carbonate fuel cells MCM. See Magnetic containment melting MD simulations. See Molecular dynamics simulations MDA. See Methylene dianiline Mechanical assembly processes, 495 – 496 Mechanical fastening systems, 495 – 496 Mechanical impact test, 115f, 115 –116, 116f Mechanical processing, 291– 293 Mechanical testing of composites, 172 –178 Melt infiltration (MI), 200 Melt intercalation approach, 291, 292, 388 Melt processable rubbers (MPRs), 138 Melting techniques, 513 Membrane, 439 Mercaptyl group (– SH), 616 Mercury, 518 Metal matrix composite (MMC), 268, 413, 543, 686 Metallic Materials Properties Development and Standardization (MMPDS), 413, 441, 442 Metallic/metal(s), 229, 612 fluorides, 773 fuels in composite propellants, 613t materials, 11, 55, 237 damage tolerance considerations for, 75 –84 hydrogen embrittlement in, 87 – 111 ignition data for, 117 – 118 trim tabs, 369 Metastable beta alloys, 44, 324 Meteoroid/orbital debris shielding, 421 –422, 428 Methane (CH4), 513, 701 Methyl methacrylate, 635f Methylene dianiline (MDA), 136 – 137 MI. See Melt infiltration Micro-truss structure and unit cell, 263f Micromechanical model, 260 Micrometeoroids, 404 Microstructure, 21 – 23, 315 Microwave ion thrusters, 708 Mid molecular weight (MMW), 388

INDEX

MIL-HDBK-17, 413 MIL-HDBK-23, 413 MIL-HDBK-310, 376 MIL-STD-3033, 376 MIL-STD-810, 376 MILHDBK-5, 413 Military rotor blades, 369 MISSE. See Materials on International Space Station Experiment Missiles, 41 MLI blankets. See Multilayer insulation blankets MLTs. See Magnetic layer thrusters MMC. See Metal matrix composite MMH. See Monomethylhydrazine MMPDS. See Metallic Materials Properties Development and Standardization MMRTG. See Multimission RTG MMW. See Mid molecular weight Mo-Re. See Molybdenum-Rhenium Modal, 614 Model 377 Stratocruiser, 328 Modern aerospace systems, 1 Modulus, 146 Molecular dynamics simulations (MD simulations), 294 Molten carbonate fuel cells (MCFCs), 758 Molybdenum disulfide, 425 grids, 706 keeper electrode, 708 Molybdenum-Rhenium (Mo-Re), 555– 556, 556f Momentum exchange, 729– 730 tether systems, 730 Monocrystals (MX), 598 Monolithic ceramics, 184 Monolithic structure, 337 Monomer, 290 Monomethylhydrazine (MMH), 637t, 642, 701, 702 Monopropellant engines, 642 MPD thrusters. See Magnetoplasmadynamic thrusters MPRs. See Melt processable rubbers MR&D. See Materials Research & Design

INDEX

MSFC. See Marshall Space Flight Center MSFC-SPEC-1443, 411 MSFC-STD-3029, 413 MT Aerospace C/SiC body flap, 569 – 571 MT Aerospace IMI, 565 MTM-45 out-of-autoclave all-composite, 333 Multi-walled carbon nanotubes (MWCNTs), 391, 731 Multidirectional composite and associated stress – strain curve, 161f Multijunction PV cells, 755f Multijunction solar cells, 424, 753 Multilayer high-temperature ceramics, 420 Multilayer insulation blankets (MLI blankets), 416, 565 Multimission RTG (MMRTG), 775 thermocouple elements used in, 778f Multiple business case scenarios, 386 Multiscale modeling techniques, 293 –295 Multitechnique, 293 – 295 Multiwall nanotubes (MWNTs), 291 MWCNTs. See Multi-walled carbon nanotubes MWNTs. See Multiwall nanotubes MX. See Monocrystals Mylar, 723 N2O4. See Nitrogen tetra oxide NAA. See North American Aviation NAA A6/A7 engine, 665 – 666 Nanoclays, 389, 390 Nanocomp Technologies, Inc., 732 Nanocomposites, 280 – 284, 387, 389 Nanoengineered multifunctional materials, 287 – 289 Nanoengineered structural materials, 277 exemplary architecture types, 280f hybrid or nanoengineered materials, 286 –287 nanocomposites, 280 – 284 nanostructure to aerospace structures, 286f nanostructured fibers and sheets, 284 –286

811

scale and organization in CNT and carbon fiber composites, 279f specific stiffness and strength for comparing metal alloys, composites, and CNTs, 278f Nanographite layers, 390 Nanomaterials, 276, 280 modeling, 293– 295 Nanometer (nm), 387 nanometer-sized crystalline materials, 756 Nanoparticle-enhanced adhesives and composite matrices, 281– 282 Nanoporous material (NPM), 565 Nanoscale structures, 387 Nanosheets. See Nanographite layers Nanostructured fibers and sheets, 280, 284– 286 Nanostructured Si anodes, 772 Nanotechnology, 387– 392 NASA ACEE 727 elevator, 330 NASA ACEE 737 carbon fiber stabilizer, 329f NASA ACEE 737 spoilers, 330, 356 NASA advanced tile, 559– 560 NASA RP-1390 satellite, 416 NASA-led in-house design, 686 NASA-related oxygen events, 113 NASA-STD-6001 Test, 409 required and supplemental tests for each material use, 410– 411t NASP program. See National Aerospace Plane program National Aerospace Plane program (NASP program), 70, 472, 555– 556 National Renewable Energy Laboratory (NREL), 756 Natural flake graphite, 390 Natural rubbers (NR), 138 Natural vibration modes of beams, 224f NBR. See Nitrile butadiene rubber NDE. See Nondestructive evaluation NDI. See Nondestructive inspection Near-beta alloys, 324 Negligible damage, 365 NEPE. See Nitrate ester polyether

812

NERVA. See Nuclear energy for rocket vehicle application Net energy flow on passive leading edges, 546, 547f Neutralizer cathodes, 707 NH4ClO4. See Ammonium perchlorate NH4NO3. See Ammonium nitrate Nickel (Ni), 423, 587 Ni-base superalloys, 62, 592 Ni-Cd batteries, 767 Ni-H2 batteries, 767 and nickel-based alloys, 101 nickel-chromium steels, 55, 60 nickel-copper alloys, 690 substrates, 379 Nitrate ester polyether (NEPE), 615t Nitrile butadiene rubber (NBR), 619, 620, 626 Nitrocellulose, 611 Nitrogen tetra oxide (NTO), 635, 636t, 642, 644, 701, 702 Nitronium perchlorate (NO2ClO4), 614t NMC cathode, 772 “No bleed” process, 330 Nomex, 422 Non-heat-treatable aluminum alloys, 20 Nondestructive evaluation (NDE), 82, 405 Nondestructive inspection (NDI), 312, 344, 370 Nonmetallic materials, 11, 409, 414, 629 Nonsolid-core NTP systems, 720 Nonstructural adhesives, 427 Nonterrestrial materials, considerations for using, 508 North American Aviation (NAA), 645 Norton– Bailey creep model, 240 – 241 Nose cones, 438, 481– 484 Nose fairing structure, 438 Notched tensile strength (NTS), 91 –92, 93t Nozzle(s), 627, 662– 676 and aft case, 626f ATJ usage in aerospace industry, 631 –632 data sets, 632 – 633 extension, 656, 670, 674f

INDEX

function, 628– 629 materials and processes, 629–631 NPM. See Nanoporous material NR. See Natural rubbers NREL. See National Renewable Energy Laboratory NSTAR grids, 706 NTO. See Nitrogen tetra oxide NTP. See Nuclear thermal propulsion NTREES. See Nuclear Thermal Rocket Element Environmental Simulator NTS. See Notched tensile strength Nuclear energy for rocket vehicle application (NERVA), 779 Nuclear fission, 749 Nuclear fuels, 719– 721 Nuclear power, 749 Nuclear thermal propulsion (NTP), 718. See also Solar thermal propulsion (STP) cermet fuel element, 721f NTREES, 718f, 721f nuclear fuels, 719– 721 Nuclear Thermal Rocket Element Environmental Simulator (NTREES), 718f, 721 Nuclear-powered spacecraft, 779 Nylon, 12, 132 O/F. See Oxidizer-to-fuel ratio O’Neill –Glaser financial model, 524 OA condition. See Overaged condition Objectives, 214t, 214– 216 OCA. See Oxygen compatibility assessment OEMs. See Original equipment manufacturers Offgassing considerations of spacecraft, 408– 412 OFHC. See Oxygen-free high thermal conductivity OFI. See Opacified fibrous insulation OHC. See Open-hole compression Olivines, 773 OMMT. See Organic montmorillonite OMS. See Orbital maneuvering system

INDEX

Onboard propellant, 722 Opacified fibrous insulation (OFI), 565 Open Architecture, 333 Open-hole compression (OHC), 172 Operating temperature, 109 –110 Operations costs, 517 OPR. See Overall pressure ratio Optical materials, 423 Optical solar reflector (OSR), 423 Optimized O’Neill – Glaser model, 526 Optimum design, 218 Orbit mechanics, 520 Orbital maneuvering system (OMS), 674 Orbital Sciences, commercial space launch companies, 649 Orbital transfer vehicle (OTV), 517, 729 Orbiter, 435 Organic carbonates, 773 Organic montmorillonite (OMMT), 290f Organic-based materials, 405 Original equipment manufacturers (OEMs), 277 Orion capsule, 536 Orion crew module, 701 Orion’s service module, 701 Orowan strengthening mechanism, 20 Osmium, 423 OSR. See Optical solar reflector OTV. See Orbital transfer vehicle Outer branch. See Payload fairing Outer mold line, 439 Outposts, 521 – 526 Overaged condition (OA condition), 20 Overall pressure ratio (OPR), 580 Oxidant electrode, 757 Oxidation, 550 comparison of active and passive, 551f temperature vs. time and flow enthalpy vs. time, 552f Oxides of nitrogen (NOx), 595 Oxidizer, 611, 613 – 614, 634, 635, 690 for hybrid propulsion systems, 636t Oxidizer-to-fuel ratio (O/F), 634 Oxidizing agent, 611 Oxygen (O2), 113, 636t

813

combustion probability and severity for materials, 118– 120 fire safety in oxygen systems design, 120– 128 hazard examples, 113 Apollo 1, 114 Apollo 13, 114 fire triangle, 114f space shuttle extravehicular mobility unit, 114 material behavior in oxygen-rich environments, 113 maximize use of oxygen-compatible materials, 122 oxygen-ignition-resistant alloys, 655 testing of materials for oxygen compatibility, 114 ignition data for metallic materials, 117– 118, 119t mechanical impact test, 115f, 115– 116, 116f promoted ignition test, 116– 117, 117f Oxygen compatibility assessment (OCA), 118, 128 – 129 Oxygen-free high thermal conductivity (OFHC), 660 p-n junction, 751f P-static. See Precipitation-static P/M approach. See Powder metallurgy approach PA. See Polyamide PA condition. See Peak aged condition PAA. See Polyacrylic acid PAEK. See Polyaryletherketone PAFs. See Phosphoric acid fuel cells PAN. See Polyacrylonitrile Parametric economic model, 516–517 Passive leading edges, 546 emissivity, 548, 549f high-temperature coatings, 547 net energy flow on passive leading edges, 547f oxidation, 550– 551, 552f recombination (catalytic) efficiency, 548– 549

814

Passive leading edges (Continued) Shuttle Orbiter wing-leading-edge repair, 554 thermal conductivity, 547 –548 thermal expansion, 549 – 550 X-43 leading edge, 551– 554 Passive Optical Sample Assembly– I experiment (POSA – I experiment), 411 Passive radiative cooling, 738 Passive thermal control coatings, 416 of spacecraft, 414 Passive thermal management, 534– 535 Payload adapter, 438 fairings, 438, 481 – 484 fitting, 438 shroud structure, 438 Payload attach fitting, 438 Payload fairing, 438 PB. See Polybutadiene PBAA. See Polybutadiene acrylic acid PBAN. See Polybutadiene acrylonitrile PBI. See Polybenzimidazole PBI fiber-filled NBR (PBINBR), 620 PBO fibers. See Polybenzobisoxazole fibers PC. See Polycarbonate PE. See Polyethylene Peak aged condition (PA condition), 20 Peak heat release rate (PHRR), 389 PEEK. See Poly(ether ketone) PEI. See Polyetherimide PEM fuel cells. See Proton exchange membrane fuel cells PEPU. See Polyether polyurethane Performance index, 218 for bending at room temperature, 222 –223 for thermal and mechanical components of cryogenic storage tanks, 226t for vibration at room temperature, 223 –224 PET. See Polyethylene terephthalate PH. See Precipitation hardening Phenolic resin, 630

INDEX

Phenolic triazine (PT), 137–138 Phenolic-impregnated carbon ablator (PICA), 420 Phosphoric acid fuel cells (PAFs), 758 Photosil process, 735 Photovoltaic cell (PV cell), 749– 752f materials for future PV systems, 754– 757 materials used for space power, 752– 754 PHRR. See Peak heat release rate Physical aging, 146 PI. See Polyimide Pi joints, 360, 361f PICA. See Phenolic-impregnated carbon ablator PIT. See Pulsed inductive thruster “Pit-and-groves” erosion, 706 Plane strain threshold stress intensity factor ratio, 104 Planetary missions, 752 Plasma, 404 containment, 741f Plasma spray-physical vapor deposition (PS-PVD), 186 Plastic media blast (PMB), 384 Plasticizers, 616 Plate, design requirements, 222t Platinum liners, 702 PLMs. See Product life management tools Plutonium-238, 774 PMB. See Plastic media blast PMC. See Polymer matrix composites PMM. See Polymethyl-methacrylate PNC. See Polymer nanocomposite PNP. See Probability of no penetration POLIPOL 701, 291 POLIYA 420, 291 Poly(ether ketone) (PEEK), 147, 740 Poly(p-phenylene sulfide) (PPS), 147 Poly[styrene-b-(ethylene-cobutylene)-b-styrene] triblock copolymer (SEBS), 292 Polyacrylic acid (PAA), 770 Polyacrylonitrile (PAN), 327, 329 Polyamide (PA), 12, 132

INDEX

PA11 powder, 293 PA6, 387 Polyaryletherketone (PAEK), 393 Polybenzimidazole (PBI), 619, 620 Polybenzobisoxazole fibers (PBO fibers), 730, 731f Polybutadiene (PB), 634, 635f Polybutadiene acrylic acid (PBAA), 615t Polybutadiene acrylonitrile (PBAN), 612, 615, 615t Polybutadiene– acrylic acid– acrylonitrile terpolymer. See Polybutadiene acrylonitrile (PBAN) Polycarbonate (PC), 12, 132 Polyester and vinyl esters, 134 Polyether polyurethane (PEPU), 615t Polyetherimide (PEI), 147, 394 Polyethylene (PE), 12, 132, 420, 634, 635f Polyethylene naphtalate, 723 Polyethylene terephthalate (PET), 468 Polyhedral oligomeric silsesquioxanes (POSS), 735 Polyimide (PI), 137, 147, 428 Polymer matrix composites (PMC), 158, 232, 279, 568, 585, 605 in aerospace structures, 158 alternate method for generating allowables, 168 –170 carbon fiber composites vs. aluminum, 158 – 163 composites for launch vehicles, 163 –168 fatigue in composite structures, 170 –172 mechanical testing of composites, 172 –178 strength of composite laminates, 178 –181 Polymer nanocomposite (PNC), 279 Polymeric/polymers, 211, 246, 404 for aerospace composites, 146 – 148 aging, 146 and composites, 12, 132 elastomers, 138 erosion coatings, 374 fiber, 285

815

fillers, 138– 139, 140t manufacturing of composites, 148– 156 materials, 11 matrix of conventional composites, 471 mechanical behavior elongation, 144 fracture, 144 modulus, 146 strength, 143– 144 stress – strain behavior, 142– 143, 143f, 144f, 145t toughness, 146 processing, 141 fiber-reinforced thermoset molding, 142 processing operations, 141 – 142 structure and synthesis, 139– 141 thermoplastics, 132–134, 135t thermosets, 134–138 types, 132 Polymerization, 140f or cure, 134 process, 615 reaction, 290 Polymethyl-methacrylate (PMM), 634, 635f Polystyrene (PS), 12, 132 Polysulfide (PS), 615t Polytetra-fluoro ethylene (PTFE), 133 Polyureas, 137 Polyurethanes, 381, 382, 428 tapes, 374 Polyvinyl chloride (PVC), 12, 132 Poppet valve, 693f Porous Si nanowire, 772 Porous Si/C Composite, 771 POSA –I experiment. See Passive Optical Sample Assembly– I experiment POSS. See Polyhedral oligomeric silsesquioxanes Potassium nitrate (KNO3), 611, 614t Potassium perchlorate (KClO4), 614t Powder Bed Fusion, 394 Powder metallurgy approach (P/M approach), 28, 36, 64, 105– 106, 592 Powder process, conventional, 36

816

Powder-to-powder blending, 293 Power systems, 1. See also Exploration systems materials for, 749 batteries, 767 – 774 distribution of mass in spacecraft, 783f fission power for space, 779 – 783 fuel cells, 757 – 767 radioisotope power systems, 774– 779 solar power, 749 – 757 PPS. See Poly(p-phenylene sulfide) Pratt and Whitney JT-3C engine, 597 “Preburner”, 652 Precipitation, 315 Precipitation hardening (PH), 20, 41 – 42 stainless steels, 60 Precipitation-static (P-static), 385 – 386 PreMax High-Shear Mixer, 293f Prepreg, 327 hand lay-up process, 150f lay-up, 148 aerospace industry, 151 – 152 principle of autoclave curing, 150f traditional lay-up and autoclave cure process, 151f Prepregger, 327 Pressure sensitive adhesives (PSA), 374 Pressure vessels, 224 –228, 233 –238, 478, 479 Pressure-fed propulsion system, 459 –460 Pressure-fed rocket engines, 650f Pressure-sensitive adhesives, 428 Pressurized structure, 458 balloon tanks, 469 – 470 composite tanks, 470 – 475 equivalent axial loads, 461 fracture failure mode, 462 LBB criterion, 463 solid rocket motor cases, 475 – 477 Space Shuttle external LOX tank, 459 stable metal tanks, 463– 469 ullage pressure, 460 Primary fiber, 620 Primary flight-loaded structure, 313 “Pristine allowables” generation, 169

INDEX

Probabilistic approach, 458 Probability of no penetration (PNP), 422 Process variability, 396 Process window, 604 Processing space materials on grand scale, 522– 524 Product development team, 1 Product life cycle, 271f Product life management tools (PLMs), 272 Promoted ignition test, 116– 117, 117f Propellant(s), 616, 712– 713 boil-off, 702 grain, 610– 612 pumping elements, 689– 690 selection, 643– 644 tank barrels and domes, 464– 465 valves, 691 Propulsion system, 536– 537 Propulsion-only NTP systems, 720 Proton exchange membrane fuel cells (PEM fuel cells), 758, 759, 761t baseline materials for, 761– 762t materials for, 759– 760f PS. See Polystyrene; Polysulfide PS-PVD. See Plasma spray-physical vapor deposition PSA. See Pressure sensitive adhesives PT. See Phenolic triazine PTFE. See Polytetra-fluoro ethylene Pulsed inductive thruster (PIT), 739– 740 Pultruded “I” stiffener, 357, 357f Pultrusion process, 156, 156f, 357 Pump-fed designs, 704 Pure aluminum, 314 PV cell. See Photovoltaic cell PVC. See Polyvinyl chloride Pyrogen igniter components, 633f Pyrolysis, 629 Quality criteria, 509t Quantitative NDE, 82 Quantum dots, 756 Quantum theory, 750 Quantum-dot-based PV cells development, 756

INDEX

Quantum-level computing, 295 Quartz glass, 327 Quasi-isotropic laminates, 174f, 178 Quench, 17 RA. See Reduction of area Radiation equilibrium temperature, 548, 549f Radiation shielding, 420 –421 Radiative-cooled chambers, 702 Radio frequency (RF), 736 ion thrusters, 708 Radiography, 370 Radioisotope power systems (RPSs), 749, 774. See also Solar power thermoelectric principle, 774 – 776f Radioisotope thermoelectric generators (RTGs), 774, 775f materials impact on future RTG systems for space power, 778 – 779 thermoelectric materials used in space, 777 –778 Rain erosion, 380 – 381 Rapid pressurization, 125 ignition from rapidly compressing oxygen, 126f Rare earth oxides (RE oxides), 186 – 187 Rational methods of materials selection, 447 “Rau equations”, 668 – 669 RCC. See Reinforced carbon – carbon RCS. See Reaction control system RE oxides. See Rare earth oxides Reaction control system (RCS), 647, 674, 700, 701 Reactor core materials, 781 candidate structural materials for space fission reactors, 782t control drum, 781 – 783 reflector, 781 Recombination (catalytic) efficiency, 548 –549 Redstone engine A6/A7 turbopump, 679 –680 Reduction in stress levels, 110– 111 Reduction of area (RA), 91 – 92

817

Reflectivity of sail material, 726– 727 Reflector, 781 Reinforced carbon– carbon (RCC), 420, 539 Reinforcements, 36 – 37, 148 Reinspecting structure, 444 Relaxation, constitutive equation for, 249– 250 Relaxation-limited design, 247f, 253– 254 Repair methods, 362, 365– 367, 383– 384 Repairable damage, 365 Resin transfer molding (RTM), 138, 154f, 471 Resins, 146– 147 film infusion, 336 Resistance welding processes, 493 Resistivity, 268 Reusability, 443–446 Reusable Launch Vehicle (RLV), 686 Reusable materials, 418– 419 Reusable solid rocket motors (RSRMs), 609. See also Solid rocket motors (SRMs) components, 619f configuration, 618 finite element analysis model for, 623f insulation at, 620f nozzle, 628f “Reverse hybrids”, 635 RF. See Radio frequency Rhenium (Re), 599, 716 Rigid concentrators, 717 Rigimesh, commercial product, 661 RL-10 engine with extendable carbon – carbon nozzle extension, 675f RL10A turbopump assembly, 683f RLV. See Reusable Launch Vehicle Robert Goddard’s Turbopumps, 678 Robotic spray application, 383 Robotic welding, 493 Rocket engine(s), 641 systems, 677 thrust chamber, 657f Rocket-based vehicles, 540 Rocket-launch vehicles, 538– 539 Rocketdyne, 686 Roll forming, 354– 355

818

Roll milling, 291, 291f Rotor blade, 368 allowables, 370 – 371 “D” spar example from composite, 369f damaged by hostile fire, 372f erosion coatings, 373 Air Force Research Laboratory whirling arm rain erosion rig, 380f operational and environmental considerations, 381 rain erosion and environmental exposure, 380 –381 rotorcraft in Southwest Asia, 374f sand erosion, 375 – 380 titanium sparking due to sand erosion, 375f Rotor grade, 604 Rotorcraft, 368. See also Subsonic aircraft materials development composite field repair, 373 design and integration considerations for erosion coatings, 381 –386 fatigue, 371 – 373 preliminary considerations, 368 Rover/NERVA program, 719 RPSs. See Radioisotope power systems RSRMs. See Reusable solid rocket motors RSRMV. See Even larger five-segment reusable solid rocket motor RTGs. See Radioisotope thermoelectric generators RTM. See Resin transfer molding Rubber insulator, 619 Rule of thumb, 383 Russian MIG-15, 586 Russian satellite Kosmos-1275, 421 “S-405” catalyst, 702 S-T ductility. See Short transverse ductility Safe and arm device (S&A device), 633 Safe-life analysis, 444 Sail areal density, 725 Sail membranes, 723 – 727 Sailcraft designers, 727 IKAROS, 722

INDEX

Salt fog, 381 Sand casting, 29 Sand erosion, 375, 380 eroded main rotor blade tip cap, 376f plot of particle angularity vs. aspect ratio for foundry sand, 378f for golf course sand, 378f for Krumbein scale, 377f for Kuwait sand, 379f Sandwich structures, 260 SARJ. See Solar Alpha Rotary Joint Saturn V launch vehicle with Apollo payload, 439f Saturn V – Apollo launch vehicle, 648 SBR. See Styrene butadiene copolymer rubber SCARLET. See Solar concentrator array with refractive linear element technology SCC. See Stress corrosion cracking SCP. See Silica cloth phenolic SCPs. See Spacecraft control processors SCRIMP process. See Seemann Composites Resin Infusion Molding Process Seal materials, 427 Sealing material, 766t SEBS. See Poly[styrene-b-(ethylene-cobutylene)-b-styrene] triblock copolymer SEE. See Space environment effects Seebeck coefficient (S), 774 Seemann Composites Resin Infusion Molding Process (SCRIMP process), 359 SEI. See Solid-electrolyte interphase Selection procedure, 254– 255 Semicrystalline thermoplastics, 134 rubbery behavior of amorphous phase, 138 Semipassive thermal management, 535– 536 SEMs. See Standard electronic modules Service conditions, 81 –82 Settlements, 521–526 SEUs. See Single-event upsets

INDEX

7000-series aluminum alloys, 480 7050-T74 alloy, 27 7055-T77 alloy, 28, 318 7150 alloy, 317 7150-T61 plate and extrusions, 27 – 28 737 HSLA flap tracks, 322f 737– 300 production program, 356 7xxx series alloys, 28, 319 SFC. See Specific fuel consumption SFEPDM. See Silica-filled EPDM Shaft horsepower (SHP), 591 Sharp leading edges impact, 545 Shearing and bypassing of precipitates by dislocation, 21f Sheet, 53 Sheet molding compound (SMC), 154 Shelf life of components, 405, 408 Shell Buckling Knockdown Factors, 455 Short transverse ductility (S-T ductility), 83 SHP. See Shaft horsepower Shroud, 481 Shuttle Orbiter wing-leading-edge repair, 554 Shuttle SRB structures, 610 Shuttle– Mir Program, 411 SHyFE. See Sustained Hypersonic Flight Experiment SiC. See Silicon carbide SiGe thermocouples in GPHS RTG, 777, 777f Silica, 550 ceramic tiles, 419 fiber, 620 Silica cloth phenolic (SCP), 631 Silica-filled EPDM (SFEPDM), 620 Silicon (Si), 750, 753 electrode failure mechanisms, 770f nanowires, 771 anode accommodate strain, 772f Silicon carbide (SiC), 184, 244, 543, 550 nanocomposite granule formation, 771f SiC/SiC ceramic matrix composites, 196 –200 creep and fatigue durability, 202f SiC/SiC CMC architecture designs, 201f

819

Silicon melt infiltration (SMI0), 197 Silicon nitride (Si3N4), 777 Silicone, 415 outgassing, 412 Silumin, 678 – 679 Silver, 510 Single crystal (SX), 65 Single load enhancement factor, 371 Single PEM fuel cell, 760f Single-event upsets (SEUs), 404 Single-stage magnetic layer thrusters, 709 Single-stage-to-orbit (SSTO), 470– 471, 531 Single-walled nanotubes (SWNTs), 390 6XXX alloys, 319 Skin-stringer aluminum aircraft structure, 538– 539 Skirts, 438, 480– 481 Skylab orbital laboratory, 507f SLA-561V, 420 Slow strain rates testing (SSR testing), 87, 92 SLS. See Space Launch System SLWT. See Super-lightweight tank SMARF. See Steve Miller and Associates Research Foundation SMC. See Sheet molding compound SNAP. See System for nuclear auxiliary power Snecma Propulsion Solide (SPS), 562, 562f SOA anode material, 738 SOA HETs. See State-of-the-art HETs Sodium nitrate (NaNO3), 614t SOFCs. See Solid oxide fuel cells SOHO. See Solar and Heliospheric Observatory Solar Alpha Rotary Joint (SARJ), 425– 426 damage in, 426f Solar and Heliospheric Observatory (SOHO), 423 Solar array materials, 423– 424 Solar cell containing quantum dots, 756f Solar concentrator array with refractive linear element technology (SCARLET), 424 Solar Disc, 525

820

Solar energy, 713 Solar power, 749. See also Radioisotope power systems (RPSs) materials for future PV systems, 754 –757 PV cell, 749– 752f materials used for space power, 752 –754 Solar power satellite (SPS), 524 Solar radiation pressure (SRP), 722 Solar sail propulsion, 722. See also Tether propulsion advanced sail materials, 727 – 728 effective sail temperature, 725f sail membranes, 723– 727 solar sail characteristic acceleration dependence, 726f thermo-optical properties for specific polymeric materials, 724t ultimate tensile strength, 727f Solar thermal propulsion (STP), 713. See also Nuclear thermal propulsion (NTP) materials requirements, 714 –718 melting point of materials, 715f operating principle, 714f tungsten and rhenium direct gain, 717f tungsten solar thermal propulsion engine components, 716f upper-stage concept, 714f Solder-welding, 668 Solid fuel(s), 610 burning process, 612 Solid loading, 615 Solid lubricants, 425 Solid materials, 132 Solid metallic nozzles and nozzle extensions, 672 Solid motors, 475 –476 Solid oxide fuel cells (SOFCs), 758, 759 anode-supported SOFC, 763f baseline materials for, 764 – 766t baseline materials for PEM fuel cell components, 761 – 762t bi-electrode SOFC cell concept, 767f materials for, 760

INDEX

single, 763f stack, 763f Solid oxidizer, 635 Solid polymer fuel cells, 758 Solid rocket boosters (SRBs), 609 Solid rocket engines, 609 binder, 614– 615t fuel, 612– 613 hybrid rocket propulsion systems, 634– 637 igniters, 633–634 insulation materials, 619– 626 liner and inhibitor, 626 – 627 minor ingredients, 615– 617f motor cases, 618– 619 nozzle, 627– 633 oxidizers, 613– 614 propellant grain, 610– 612 solid rocket components, 610 – 617 space shuttle’s reusable solid rocket motor, 610f Solid rocket motors (SRMs), 7, 437, 609, 641– 642 cases, 475– 477 components in, 609– 610 configuration, 620f Solid solution strengthening, 19 –20 Solid-electrolyte interphase (SEI), 770 Solid-forged Ti fan blades, 590f Solid-state precipitation sequence for aluminum alloys, 19t Solidification processing, 513 casting techniques, 513–514 material properties, 514– 515 melting techniques, 513 Solution treatment, 17 Southern Research Institute (SRI), 632 Soviet/Russian Buran Orbiter, 443 Soviet/Russian Energia vehicle, 436 Space base, 521 components for space fission reactor, 779 –781 debris impact on shuttle window, 421f particles, 404

INDEX

fuel cells for space applications, 758 missions, 749 PV cell materials used for space power, 752 –754 radiation, 420 resources, 522 structures/components/vessels, 225f Space environment effects (SEE), 732 “Space junk”, 435 Space Launch System (SLS), 609 Space Shuttle, 439 –440, 446 external LOX tank, 459 external tank, 496 extravehicular mobility unit, 114 reusable solid rocket motor, 610f system studies, 477 Space shuttle main engine (SSME), 60, 72, 536, 537f, 647 – 648, 684 alternate turbopump, 684 – 685 turbopump, 684, 685f Space Shuttle Orbiter, 435, 443, 488, 488f, 532 –533, 534f, 538 – 539, 558 airframes, 443 – 444 composites flown in, 164t nominal maximum temperatures on surface, 539f Space Shuttle’s Orbital Maneuvering System, 701 Space-derived model, 526 Space-derived solar power satellites, 524 –526 Spacecraft, 1, 41 applications, 72 room-temperature properties of current and potential rocket, 73t lunar oxygen production for spacecraft propellant, 516 – 519 materials adhesives, 427 – 428 defects in material, 429 – 430 deficiencies in design, 428 flammability, toxicity, and offgassing considerations, 408 –412 improper materials selection, 429 improper processing, 430 inadequate service, 430 – 431

821

inappropriate assembly, 430 lessons learned, 428 lubricants, 424– 427 meteoroid/orbital debris shielding, 421– 422 optical materials, 423 radiation shielding, 420 – 421 seal materials, 427 solar array materials, 423– 424 space environment, 404– 405 spacecraft design and materials requirements, 405–408 structural materials, 412– 414 thermal control materials, 414– 418 thermal protection materials, 418– 420 Spacecraft control processors (SCPs), 429 Spacewalks, 145 SpaceX (commercial space launch companies), 649 Span moment, 382– 383 Spar, 369 Specific fuel consumption (SFC), 582 Specific heat, 267 Spectra, 730 SPF. See Superplastic forming Spinning tether system, 729 “Split mission”, 519 Spray-up process, 153f, 153– 154 Sprayable polyurethanes, 384 SPS. See Snecma Propulsion Solide; Solar power satellite; Suspension plasma spray SRBs. See Solid rocket boosters SRI. See Southern Research Institute SRMs. See Solid rocket motors SRP. See Solar radiation pressure SS alloys. See Stainless steel alloys SSME. See Space shuttle main engine SSR testing. See Slow strain rates testing SSTO. See Single-stage-to-orbit Stable crack extension, 77 –78 stable crack growth, 79f unstable fracture, 78f Stable metal tanks, 463–469 Staged combustion cycle, 649– 650

822

Stagnation region aft, 546 Stainless steel alloys (SS alloys), 118, 322 Standard electronic modules (SEMs), 69 Standoff TPS, 560 Snecma Propulsion Solide CMC TPS array, 564f SPS CMC shingle TPS, 562f SPS CMC TPS array for mechanical, dynamic, thermal, 563f SPS shingle TPS attachment mechanism, 563f structural layout of X-33 RLV, 561f X-33 metallic TPS illustrating stand-off TPS attachment, 561f Stardust spacecraft, 411 State-of-the-art aircraft structures, 358 State-of-the-art HETs (SOA HETs), 709 –710 Static-dissipative thermal control coatings, 416 Stationary plasma thrusters (SPTs). See Single-stage magnetic layer thrusters “Steel balloon” design, 458 –459 Steel(s), 55, 56 – 57t, 102 corrosion-resistant, 59 – 60 high fracture toughness, 58 maraging, 58 –59 nickel – chromium, 60 and titanium evolution, 321 current developments, 324 – 325 development, 322 development timeline, 322 – 324 historical titanium alloy improvement, 324f strength toughness goals for high steel alloys, 323f ultrahigh-strength, 55, 58 Step-reaction polymerization, 140 Steve Miller and Associates Research Foundation (SMARF), 565 OFI, 565 – 566 Stiff tie, design requirements, 222t Stiffeners, 464 Stiffness, 455 and strength of foam, 487

INDEX

Stirling converter, 783 Storable propellants, 643 STP. See Solar thermal propulsion Strain compatibility, 382 hardened alloys, 15 –16 hardening, 20 induced porosity, 605 Strength of bending-dominated foams, 265– 266 of stretch-dominated lattices, 266 Strength-to-weight ratio basis, 41 – 42, 412 Strengthening mechanisms, 19, 50 – 51 dispersion hardening, 20 – 21 grain size strengthening, 20 precipitation hardening, 20 solid solution strengthening, 19 – 20 work (strain) hardening, 20 Stress corrosion cracking (SCC), 55, 87, 316 Stress levels, reduction in, 110 – 111 Stress– strain behavior, 142– 143, 143f, 144f room-temperature mechanical properties, 145t Stress– strain curve of Al/Li, 160f of multidirectional laminate, 179, 179f Stretch-dominated lattices elastic moduli, 264– 265 fracture toughness, 267 strength, 266 Structural adhesives, 427 Structural elements, 214 Structural requirements, 344 critical considerations for wing design, 347t critical material design properties, 345– 346t criticality of wing failure modes, 348f fuselage, 350– 354 lower wing panel material comparisons, 349f relationships, 348f upper wing panel strength evolution, 349f wing, 344– 350

INDEX

Structurally integrated TPS, 567, 567f STS-114 and tile insulation system, 559f STS-61-A Spacelab D-1 mission, 409 Styrene butadiene copolymer rubber (SBR), 138 Styrene-free resin, 291 Subsonic aircraft materials development, 305 enabling materials, structures, and manufacturing processes, 386 AM, 392 – 396 nanotechnology, 387 –392 promise of true multi-functional engineered materials, 388f relative structural cost and weight for material systems, 386f integrated design/manufacture/ material development, 307f structural materials on selected Boeing commercial aircraft, 307f Subsonic and supersonic fixed wing aircraft, 308. See also Subsonic aircraft materials development composites material/process checklist, 311t design criteria, 309, 309f evolution of aluminum materials, 314 –321 of composite materials, 326 –337 of fiber metal laminates, 338– 344 of steel and titanium, 321 – 325 of structural requirements, 344 – 354 in-service requirements, 362 – 367 manufacturing requirements, 354 – 362 metals material/process checklist, 310t preliminary considerations, 308 – 309 qualification of new designs, 311 – 314 of new materials, 310 Substrate compatibility, 382 Sun Tower, 525 Super-lightweight tank (SLWT), 413 Superalloys, 62, 102. See also Aluminum alloys; Titanium alloys metallurgy, 62 – 64 processing, 64 – 66

823

properties and applications, 66 –67, 67t Superplastic forming (SPF), 325 Supplier pre-impregnated fabrics, 328 Surface erosion, 624f Surface preparation techniques, 111 Surface recession mechanism, 191 Surface-driven convection, 506 Suspension plasma spray (SPS), 186 Sustained Hypersonic Flight Experiment (SHyFE), 566– 567, 567f SWNTs. See Single-walled nanotubes SX. See Single crystal Synthetic rubbers, 138 System for nuclear auxiliary power (SNAP), 779 Systematic approach to materials selection, 212 blades, 230–233 material choices for case studies, 230 material performance indices, 218– 228 material property charts, 228–230 methodology, 213 pressure vessels, 233– 238 ranking, 216– 217 research, 217 screening/rejection, 216 specific cultural constraints, 217 translation, 213– 216 universe of materials, 212– 213 T/W. See Thrust to weight T41. See Turbine inlet temperatures T77 temper, 317, 318, 319 Tail of airplane. See Empennage Tailoring, 337 Talc, filler, 139 TALs. See Two-stage-anode layer thrusters Tank system parameters relative to flight duration, 226t Tantalum hafnium carbide, 720 Tantalum zirconium carbide, 720 Tape wrapping, 630 TBC. See Thermal barrier coatings TBCC. See Turbine-based combined cycle TCP. See Topologically close packed

824

Technical basis development for materials processing, 506 considerations for using nonterrestrial materials, 508 hierarchy of materials, 508 – 510 materials availability and extraction, 510 –513 materials science and processing in space, 506 – 508 solidification processing, 513 – 515 Technology readiness levels (TRLs), 3, 277, 562 Teflon, 404, 702 Temper designation system, 15 Tensile modulus, 143 Tensile strength (TS), 12, 143, 144f of aluminum alloys, 465 – 466 of quasi-isotropic laminates, 174f Tensile yield strength (TYS), 48 Tether conductors, 732 Tether propulsion, 728. See also Solar sail propulsion coatings for environmental survivability, 732 –735 high-tenacity fibers, 730– 732 high-tensile strength materials comparison, 732 momentum exchange, 729– 730 tether conductors, 732 Tethered satellite system, 430 TGO. See Thermally grown oxide Thermal activation energy, 110 Thermal barrier coatings (TBC), 65 – 66, 185 –189, 190f, 601 Thermal conductivity, 267, 547 –548, 566 plain weaves and photographs of arc-jet tested leading edges, 548f Thermal control materials, 414 – 418 Thermal cycling effect, 340, 404 Thermal environments, 485 –486 Thermal expansion, 267, 549 – 550 Thermal management, 534 active, 536 – 538 passive, 534 – 535 semipassive, 535 – 536 Thermal management, active, 536 – 538

INDEX

film cooling and drawing of hypersonic vehicle, 537f transpiration cooling and C/C cooled combustion chamber test, 538f for vehicles, 538f Thermal performance analysis, 623 – 626 Thermal protection and insulation, 485– 490 materials, 418– 420 system used on X-33, 539 Thermal protection systems (TPSs), 6, 486, 531 and hot structure components, 543 acreage TPS and aeroshells, 558– 567 control surfaces, 567– 573 leading edges, 544– 558, 559f and hot structures for air-breathing hypersonic vehicles, 540– 541 for rocket-launch vehicles, 538– 539 Thermal soak back temperatures, 702 Thermal storage engines, 716–717 Thermal– chemical treatments, 198 Thermally grown oxide (TGO), 195– 196 Thermal– structural challenges, 541– 543 ceramic matrix composite with fibers inside of matrix, 544f Thermionic cathode, 737 Thermoelectric materials used in space RTG, 777–778 Thermoelectric principle, 774 – 776f Thermomechanical processing (TMP), 44, 55, 317 sequences, 53f Thermoplastic elastomers (TPEs), 138 Thermoplastic vulcanizate (TPV), 138 Thermoplastic(s), 132– 134 material, 140 polymer densities and melting and transition temperatures, 135t polymers, 147 Thermosets, 134 BMIs, 136– 137 CE and PT, 137– 138

INDEX

epoxy, 136, 137t, 138t PIs, 137 polyester and vinyl esters, 134 polyureas, 137 thermoplastic and thermosetting resin characteristics, 136t Thermosetting matrix, 134 polymers, 147 Thinky ARE-310 mixer, 293 Thixoformed A356. 0-T6 inner turbo frame for Airbus aircraft, 32f 3-D printed liquid hydrogen turbopump, 688f 3D printing. See Additive manufacturing (AM) Three-spool architecture, 592 Threshold pressure, 117 – 118 Threshold stress (k), 245 Thrust chamber materials, 656 combustion chambers, nozzles, and expansion nozzles, 662 – 676 F-1 rocket engine, 657f injectors, 658 – 662, 663f rocket engine thrust chamber, 657f pressure, 437 Thrust to weight (T/W), 579 of military engines, 580f Ti matrix composites (TMCs), 48 Ti-based intermetallic compound, 602 TiGr. See Titanium Graphite Time-dependent processes, 458 “Tin disease”, 429 “Tin pest”, 429 Tin plating, 429 “Tin whiskers”, 429 Titanium (Ti), 41, 101 – 102, 581, 613t Ti-10V-2Fe-3Al, 324 Ti-15V-3Cr-3Al-3Sn, 324 –325 Ti-5Al-5V-5Mo-3Cr, 324 Ti-6Al-4V, 702 TiAl, 602, 603 Titanium alloys, 41, 101 – 102, 244, 477, 702. See also Aluminum alloys; Superalloys

825

characteristics of Ti alloys compared with structural metallic materials, 42t classification, 43 commercial Ti alloys, 47t fan disk for large turbofan engine, 582f influence of microstructural parameters on mechanical properties, 46t metastable beta, 45 processing of titanium alloys, 51 – 54 properties of Ti and structural metal alloys, 43f pseudo-binary section of b isomorphous phase diagram, 44f specific tensile strength, 49f strengthening mechanisms, 50 – 51 typical microstructural classes of Ti alloys, 46f TYS, 48 Titanium Graphite (TiGr), 306, 342– 344 Titanium-oxide (TiO2), 101–102 TMCs. See Ti matrix composites TML. See Total mass loss TMP. See Thermomechanical processing Topologically close packed (TCP), 63 TOR polymer. See Triton Oxygen Resistant polymer Total mass loss (TML), 409 Toughened uni-piece reinforced oxidation resistance composite (TUFROC), 559– 560 Toughened unipiece fibrous insulation tiles (TUFI tiles), 419 Toughness, 146 Toxicity of spacecraft, 408– 412 TPAs. See Turbopump assemblies TPEs. See Thermoplastic elastomers TPSs. See Thermal protection systems TPV. See Thermoplastic vulcanizate Traditional lay-up and autoclave cure process, 151f Transferability, 79 – 81 Transformation process, 134 “Trapping-mechanism” model, 109

826

Triple-junction GaInP/GaInAs/Ge cells, 753, 753f layers in, 754f Tris[1-(methyl) aziridinyl] phosphine oxide (MAPO), 616 Triton Oxygen Resistant polymer (TOR polymer), 735 Triton Systems, Inc., 735 TRLs. See Technology readiness levels TS. See Tensile strength TS-1792 material, 632 TS-5245 material, 632 TSTO. See Two-stage-to-orbit TU Delft. See Delft University of Technology TUFI tiles. See Toughened unipiece fibrous insulation tiles TUFROC. See Toughened uni-piece reinforced oxidation resistance composite Tungsten (W), 244, 588 anodes, 738 tungsten – lithium heat pipes, 739 uranium dioxide cermet fuel, 720 uranium nitride cermet fuel, 720 Turbine blade, 255f fan and turbine blades for gas turbines, 255 – 256 drive fluid, 676 – 676 methods and components, 687 –689 engines, 66 inlet temperatures, 597 jet engines, 221f turbine-driven pumps, 676 turbine-powered military aircraft, 586 Turbine-based combined cycle (TBCC), 532 Turbofan engine, 579, 580f Turbopump assemblies (TPAs), 719 Turbopump materials, 676 propellant pumping elements, 689 – 690 sample turbopump configurations, 678 A4 (V2) turbopump, 678 – 679

INDEX

Atlas turbopump, 680–681, 681f F1 turbopump (Mark 10), 681– 682 hydrogen turbopumps (RL-10 and Mark 15), 682–684 Redstone engine A6/A7 turbopump, 679– 680 Robert Goddard’s turbopumps, 678 SSME alternate turbopump, 684– 685 SSME turbopump, 684, 685f technology development turbopumps, 685– 687, 688f turbine drive methods and components, 687– 689 Twaron, 422 Twin-engine aircraft, 581, 589 2091-T851 alloys, 320 2324-T39 alloy plate, 317 2XXX alloys, 26f, 319 Two-stage-anode layer thrusters (TALs), 709 Two-stage-to-orbit (TSTO), 531, 532f Type I chromic acid, 415 Type I defects. See Hard alpha defect Type II sulfuric acid, 415 Type III hard anodize, 415 TYS. See Tensile yield strength U.S. Air Force Office of Scientific Research (AFOSR), 276 U.S. Air Force Research Laboratory, 730 U.S. Air Force study, 306 U.S. Army Tactical Missile System (ATacMS), 632 UA condition. See Underaged condition UAS. See Unmanned aircraft system UAV/S. See Unmanned aerial vehicles/ systems UCAR. See Union Carbide UD tape prepreg. See Unidirectional tape prepreg UDMH. See Unsymmetrical dimethyl hydrazine UHMWPE. See Ultra-high molecular weight polyethylene Ultimate load (UL), 166

INDEX

Ultimate tensile strength (UTS), 104, 593 Ultra-high molecular weight polyethylene (UHMWPE), 730 Ultrahigh-strength steels, 55, 58 Ultraviolet (UV), 139 light exposure, 313 radiation, 718 solar radiation effects, 381 Un-tempered martensite, 322 – 323 Unbonded solid lubricants, 425 Underaged condition (UA condition), 20 Uniaxial Hooke’s law, 239 Unidirectional tape prepreg (UD tape prepreg), 327 Unified viscoplastic GVIPS model, 241 Union Carbide (UCAR), 631 Universe of materials, 212 – 213 University of Arizona Phoenix Lander mission, 521 Unmanned aerial vehicles/systems (UAV/S), 275 – 276 Unmanned aircraft system (UAS), 288 Unpressurized structure, 479 core materials, 484 –485 intertanks, skirts, adapters, 480 – 481 payload fairings and nose cones, 481 –484 Unsymmetrical dimethyl hydrazine (UDMH), 642 Uranium, 720 UTS. See Ultimate tensile strength UV. See Ultraviolet Vacuum arc remelting (VAR), 51, 58, 64 Vacuum Assisted Resin Transfer Molding (VaRTM), 154, 336, 358, 359f Vacuum induction melting (VIM), 64 Vacuum induction melting followed by vacuum arc remelting (VIM-VAR), 58 Valves, 691 – 693 Vanguard launch vehicle, 646 – 647 Vapor grown carbon nanofibers (VGNF), 391 VAR. See Vacuum arc remelting

827

Variable specific impulse magnetoplasma rocket (VASIMR), 736 –737 Variable-polarity plasma arc (VPPA), 466– 467 VaRTM. See Vacuum Assisted Resin Transfer Molding VASIMR. See Variable specific impulse magnetoplasma rocket Venting analysis, 428 Verification and qualification, 451– 452 Very advanced propulsion systems, 741. See also Advanced propulsion technologies antimatter propulsion, 742– 743 fusion propulsion, 741–742 Penn State artist’s concept, 742f plasma containment, 741f VGNF. See Vapor grown carbon nanofibers VIM. See Vacuum induction melting VIM-VAR. See Vacuum induction melting followed by vacuum arc remelting Vinyl-ester-epoxy resin, 291 Virgin Orbit, commercial space launch companies, 649 Volume Loss, 379 Volvo laser-welded sandwich expansion nozzle, 673f Volvo welded square tube nozzle design, 672f VPPA. See Variable-polarity plasma arc VX-200 device, 736 Weight, 382– 383, 386 Weldability, 440 Weldalite, 32 Welding aluminum for LV tanks and structures, 492 primary assembly method, 492– 495 Western turbine engine producers, 585 Wet lay-up process, 152–153, 153f Whipple shield design, 422 Window materials, 423 Wing structural requirements, 344– 350 Work hardening, 20

828

Wrought alloys, 13, 14t Wrought aluminum alloys, 13 Wrought heat-treatable aluminum alloys, 23 –28 Wrought processing, 64 X-33 cryogenic tank after failure, 414f liquid hydrogen tank, 472, 473f thermal protection system used on, 539 X-37B orbital test vehicle, 419 – 420 X-41 hybrid control surface, 571 – 572 leading edge, 551 – 554

INDEX

Xenon, 712–713 XLR43-NA-1 engine turbopump, 680f Y966. See 3M 966 adhesives Yield strength (YS), 104, 265– 266 Yield-before-break, 227, 237 Young’s modulus (E), 209, 447 Yttria, 739 Zirconia (ZrO2), 186– 187 Zirconium (Zr), 547, 613t Zylon, 730, 731 Zylon-AS, 731 Zylon-HM, 731

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ABOUT THE BOOK Aerospace Materials and Applications addresses materials selection and use in aircraft, spacecraft, launch vehicles, propulsion systems, and power systems. Advances in aerospace systems are strongly dependent on advances in materials and processing technologies. In the past 100 years of powered flight, aircraft structures have evolved around advances in materials that are lighter and stronger. Aircraft propulsion systems are constantly striving to become more fuel efficient via reductions in mass and improved capability for materials to operate at higher temperatures for longer periods. Gas turbine engines that power modern aircraft are being designed to run at higher pressures and temperatures to generate more thrust per pound of engine mass. Similar considerations apply to rocket engines where the power densities are much higher. Spacecraft are designed to operate in the harsh radiation environments of outer space. In-space propulsion and power systems are key components of spacecraft and advanced materials enable these systems. Hence it is important for aerospace systems designers to have a good understanding of how specific materials will perform in their systems. Aerospace Materials and Applications clearly shows the preferred approach to selecting materials given the unique requirements for design and construction of aerospace systems. Coverage includes: Aerospace Materials Characteristics Materials Selection for Aerospace Systems Advanced Nanoengineered Materials Subsonic Aircraft Materials Development Materials for Spacecraft Materials for Launch Vehicle Structures Materials for Exploration Systems

Thermal Protection Systems and Hot Structures for Hypersonic Vehicles Aero Engine Materials Materials for Solid Rocket Engines Materials for Liquid Propulsion Systems Advanced Materials for In-Space Propulsion Materials in Power Systems in Space Exploration

ABOUT THE EDITOR BILIYAR N. BHAT holds a Ph.D. in materials science and engineering from the University of Minnesota and a master’s in management of technology from Massachusetts Institute of Technology. He is an Aerospace Materials Engineer at the NASA Marshall Space Flight Center. An Associate Fellow of AIAA, Bhat led development of Aerospace Materials and Applications in conjunction with the AIAA Materials Technical Committee.

I S B N 978-1-62410-488-6

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