Orbit-Raising and Maneuvering Propulsion: Research Status and Needs [1 ed.] 9781600863448, 9780915928828

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ORBIT-RAISING AND MANEUVERING PROPULSION: RESEARCH STATUS AND NEEDS Edited by Leonard H. Caveny Air Force Office of Scientific Research Boiling AFB Washington, D.C.

Volume 89 Progress in Astronautics and Aeronautics

Martin Summerfield, Series Editor-in-Chief Princeton Combustion Research Laboratories, Inc. Monmouth Junction, New Jersey

Published by the American Institute of Aeronautics and Astronautics, Inc. 1633 Broadway, New York, N.Y. 10019

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American Institute of Aeronautics and Astronautics, Inc. New York, New York Library of Congress Cataloging in Publication Data Main entry under title:

Orbit-raising and maneuvering propulsion. (Progress in astronautics and aeronautics; v. 89) Includes index: 1. Space vehicles—Propulsion systems. 2. Orbital transfer (Space flight) I. Caveny, Leonard H. II. American Institute of Aeronautics and Astronautics, Inc. III. Series. TL507.P75 vol. 89 629.1s [629.47'5] 83-26637 [TL782] ISBN 0-915928-82-5 Copyright © 1984 by American Institute of Aeronautics and Astronautics, Inc. All rights reserved. No part of this book may be reproduced in any form or oy any means, electronic or mechanical, including photocopying, by recording, or by any information or retrieval system, without permission in writing from the publisher.

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Progress in Astronautics and Aeronautics Series Editor-in-Chief

Martin Summerfield Princeton Combustion Research Laboratories, Inc.

Series Associate Editors Burton I. Edelson National A eronautics and Space Administration

Leroy S. Fletcher Texas A &M University

Alien E. Fuhs Naval Postgraduate School

J. Leith Potter Vanderbilt University

Norma J. Brennan Director, Editorial Department AIAA

Camille S. Koorey Series Managing Editor AIAA

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Table of Contents Preface. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . ix List of Series Volumes. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . x v Chapter I.

Laser and Solar Driven Propulsion . . . . . . . . . . . . . . . 1

Power-Beaming Technology for Laser Propulsion . . . . . . . . . . . . . . . . 3 L. N. Myrabo, The BDM Corporation, McLean, Va.

Potential of Advanced Solar Thermal Propulsion . . . . . . . . . . . . . . . 30 J. M. Shoji, Rockwell International Corporation, Canoga Park, Calif.

Laser Radiation to Supply Energy for Propulsion . . . . . . . . . . . . . . . 48 C. L. Merkle, The Pennsylvania State University, University Park, Pa.

Laser Energy Absorption in Gases: Research Problems . . . . . . . . . . . 73 N. H. Kemp, D. I. Rosen, and H. H. Legner, Physical Sciences, Inc., Woburn, Mass.

Repetitively Pulsed Laser Propulsion: Needed Research. . . . . . . . . . . 95 D. I. Rosen, A. A. Pirri, R. F. Weiss, and N. H. Kemp, Physical Sciences, Inc., Woburn, Mass.

Steady (Continuous Wave) Laser Propulsion: Research A r e a s . . . . . 109 N. H. Kemp and H. H. Legner, Physical Sciences, Inc., Woburn, Mass.

Laser Thermal P r o p u l s i o n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 129 D. Keefer, R. Elkins, and C. Peters, The University of Tennessee Space Institute, Tullahoma, Tenn., and L. Jones, NASA Marshall Space Flight Center, Huntsville, Ala.

Numerical Modeling of Laser Thermal Propulsion Flows . . . . . . . . 149 T. D. McCay, NASA Marshall Space Flight Center, Huntsville, Ala., and J. Thoenes, Lockheed Missile and Space Co., Inc., Huntsville, Ala.

Laser-Driven Repetitively-Pulsed MHD Generators: A Conceptual S t u d y . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167 C. D. Maxwell, STD Research Corporation, Arcadia, Calif., and L. N. Myrabo, The BDM Corporation, McLean, Va.

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Chapter II. Continuous Operation Electric T h r u s t e r s . . . . . . . . 201 A Comparison of Electric Propulsion Technologies for Orbit Transfer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203 R. L. Poeschel and J. Hyman, Hughes Research Laboratories, Malibu, Calif.

A Strategy for Electric Propulsion Development . . . . . . . . . . . . . . . 233 P. J. Turchi, R&D Associates, Arlington, Va.

Theoretical Modeling of the Voltage Characteristics of MPD Devices . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 241 R. M. Patrick and G. S. Janes, Avco Everett Research Laboratory, Inc., Everett, Mass.

Applied-Field Magnetoplasmadynamic Thrusters for Orbit-Raising M i s s i o n s . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 260 G. R. Seikel, SciTec, Inc., Cleveland, Ohio, T. M. York, Pennsylvania State University, University Park, Pa., and W. C. Condit, Westinghouse R&D Center, Pittsburgh, Pa.

Thrust for Interorbit Propulsion: A Question of Lifetime . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 287 S. B. Gabriel and D. Q. King, Jet Propulsion Laboratory, California Institute of Technology, Pasadena, Calif.

Electric Thruster Performance for Orbit-Raising and M a n e u v e r i n g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 303 H. R. Kaufman and R. S. Robinson, Colorado State University, Fort Col I ins, Colo.

Electric Thruster Capabilities for Orbit-Raising and Maneuvering Missions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 327 P. J. Wilbur, Colorado State University, Fort Collins, Colo.

Chapter III.

Pulsed Electric Thrusters . . . . . . . . . . . . . . . . . . . 343

Metallic Induction Reaction Engine . . . . . . . . . . . . . . . . . . . . . . . . . 345 P. Mongeau and H. Kolm, Massachusetts Institute of Technology, Cambridge, Mass.

Plasma-Surface Interactions for Electromagnetic P r o p u l s i o n . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 362 C. L. Dailey, TRW Space and Technology Group, Redondo Beach, Calif., and R. H. Lovberg, University of California at San Diego, La Jolla, Calif.

Deflagration Plasma T h r u s t e r . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 371 D. Y. Cheng and C. N. Chang, International Power Technology, Inc., Sunny vale, Calif.

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Configurations, Materials, and Performance Considerations for Railguns in Space Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . 385 A. C. Buckingham and R. S. Hawke, Lawrence Livermore National Laboratory, University of California, Livermore, Calif.

Chapter IV. Nuclear Propulsion . . . . . . . . . . . . . . . . . . . . . . . 403 Nuclear Reactor Sources for Space Prime Propulsion and P o w e r . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 405 L. N. Myrabo, The BDM Corporation, McLean, Va., and J. R. Powell, Brookhaven National Laboratory, Upton, N.Y.

Nuclear Space Power Systems for Orbit-Raising and M a n e u v e r i n g . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 425 D. Buden and J. A. Sullivan, Los Alamos National Laboratory, Los Alamos, N.M.

Ultra-Performance Closed-Cycle Gas Core Reactors for Orbit Raising . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 460 R. J. Rosa, Montana State University, Bozeman, Mont. and L. N. Myrabo, The BDM Corporation, McLean, Va.

Space Nuclear Multi-Mode R e a c t o r s . . . . . . . . . . . . . . . . . . . . . . . . . 477 L. N. Myrabo, The BDM Corporation, McLean, Va.

Particle Bed Reactors for Space Power and Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 495 J. R. Powell and T. E. Botts, Brookhaven National Laboratory, Upton, N. Y.

Rotating Bed Reactor: Research and Development Issues . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 526 O. C. Jones Jr., Rensselaer Polytechnic Institute, Troy, N.Y.

Nuclear Electric Propulsion (NEP) Spacecraft Configuration Study . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 544 P. W. Garrison, K. T. Nock, and R. Jones, California Institute of Technology, Jet Propulsion Laboratory, Pasadena, Calif.

Chapter V. Advanced Chemical Propulsion . . . . . . . . . . . . . . 557 Advanced Liquid Propellant Systems for Chemical Propulsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 559 J. Q. Weber, Rockwell International Corporation, Canoga Park, Calif.

Author Index for Volume 89 . . . . . . . . . . . . . . . . . . . . . . . . . . . 569

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Preface Advanced primary propulsion for orbit transfer periodically receives attention, but invariably the propulsion systems chosen have been adaptations or extensions of conventional liquid- and solid-rocket technology. The advent of reusable launch vehicle capability for low Earth orbit creates new opportunities for advanced propulsion for interorbit transfer. For example, 75% of the mass delivered to low Earth orbit may be the chemical propulsion system required to raise the other 25% (i.e., the active payload) to geosynchronous Earth orbit; nonconventional propulsion offers the promise of reversing this ratio of propulsion and payload masses. Prior to the present considerations, the advanced propulsion concepts could have been placed into several categories, e.g., 1) Sufficient onboard electrical power did not exist. 2) Solutions to indentified problems could not be assured. 3) Potential for space contamination was not understood. 4) An important enabling technology was lacking. 5) Knowledge of the concept was narrowly held, thus it escaped attention. 6) Overall system reliability could not be assured. The dominant consideration in previous years was that the missions could be performed using conventional chemical propulsion. Consequently, major initiatives to provide technology and to overcome specific barriers were not pursued. Chapter I treats several aspects of beamed energy propulsion, i.e., power sources and transmission, radiation absorption in flowing media, and thruster configurations. The immediate advantage of beamed energy is that high specific impulses are obtained by using low molecular weight working fluids heated by external power sources to temperatures greatly in excess of combustion gas temperatures. The premise of this chapter is that suitable megawatt laser sources will be available and justified for applications other than propulsion. The initial paper by Myrabo provides both background and new insights into power beaming, transmission and collection efficiencies, and thruster configurations. He supports the thesis that ground-based free-electron lasers operating at suitable short wavelengths will enable attractive transmission efficiencies and compact collection optics. Conversely, a scenario for using orbiting laser systems for beam energy propulsion is considered to be premature. The discussion of solar thermal propulsion by Shoji focuses on the

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nearer term approach of using onboard solar collectors as the heat energy source. The thoughtful approaches to solar radiation absorption in flows provide a transition to the challenges of the higher power density, laser supported plasmas. Merkle introduces the temperature and wavelength dependencies of laser supported hydrogen plasmas in terms of axial flows. The next three papers, integrating contributions of Kemp, Legner, Pirri, Rosen, and Weiss, deal with the physics of laser-induced breakdown and beamed energy absorption/emission for both continuous and pulsed energy beams; quantitative understanding of these processes is essential to predicting and interpreting plasma initiation, stability, and flow in thruster configurations. They address the realities of such processes as flowfield confinement, chamber heating, and seeding to promote plasma initiation. The paper of Keefer, El kins, Peters, and Jones provides additional insights into questions associated with the control of plasma position, flow stability, and mixing with buffer gases. McCay and Thoenes present a strong case for developing an analytical model that comprehensively includes the nonequilibrium interactions of the multidimensional radiation and flowfield processes. In the last paper of the beamed energy chapter, Maxwell and Myrabo describe the first of several dual-mode concepts (i.e., providing for both electrical power and propulsion) presented in this volume. Their concepts include bold propositions to use laserinduced detonation and blast waves to drive MHD generators and propulsive expansions of working fluids. Chapter II, the first chapter on electric propulsion, emphasizes concepts which lend themselves to sustained operation at megawatt power levels, possibly using clusters of thrusters. The authors explore the dual-mode premise of the megawatt power supply onboard for the main mission being available for propulsion power. Thus, propulsion does not take all of the mass penalty for the power source. This represents a major departure from pulse-mode electric propulsion considered for station keeping or planetary probes, e.g., millisecond pulses driven by a capacitor bank charged by a dedicated kilowatt power source. When large total impulses are required, thruster mass is small compared to the fuel mass; thus, a premium is placed on increasing fuel efficiencies, and the low power densities of the thrusters become relatively unimportant. The overview paper by Poeschel and Hyman describes the capability of three categories of electric propulsion (electrostatic, electromagnetic, and electrothermal) and provides the reader with sufficient information to make comparisons using simplified expressions for mission and propulsion parameters. Turchi proposes, for the intermediate time frame, to take advantage of payload power supplies and to build on the success of the pulsed plasma thruster experience. Patrick and Janes describe the potential of two externally applied magnetic field

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thrusters and point to the timeliness of extending recent advances in plasma science to understand the phenomena that limit performance. Seikel, York, and Condit advance the argument that high instantaneous power levels and externally applied magnetic fields lead to higher thrusts and efficiencies. Electrode and insulator lifetimes have been identified as primary barriers to sustained, highpower density operation of magnetoplasmadynamic thrusters. Hypotheses for mass loss mechanisms and for conditions leading to the abusive environment are offered by Gabriel and King. Kaufman and Robinson use rather direct approaches to mission and thruster parameters to show the conditions under which arcjet thrusters and electrostatic thrusters offer distinct advantages. Their conclusions point to research directed at improving acceleration efficiency. Wilbur focuses attention on approaches which will improve the efficiency of ionizing the working fluids in electrostatic thrusters. Chapter III, the second chapter on electric propulsion, deals with the concepts which by their nature operate in the pulsed mode. Mongeau and Kolm describe processes for magnetic induction, acceleration, and vaporization of metal foils. Dailey and Lovberg address the extreme environment on the insulating surfaces of a pulsed inductive thruster; they stress the importance of understanding uv radiation damage to insulators and plasma formation in gases adjacent to insulator surfaces. Cheng and Chang discuss plasma acceleration mechanisms that tend to result in smaller temperature increases and larger momentum increases. Railguns accelerate small masses to high velocities; thus, it is appropriate to consider how they can be used as impulsive thrusters. Buckingham and Hawke set forth the operational principles and address the challenges of achieving systems that can withstand the rigors of repetitive operation and of avoiding ejecta which could damage other spacecraft. Nuclear propulsion, as discussed in Chapter IV, includes either direct heating of the working fluid (i.e., hydrogen) by fission reactors or nuclear electric power systems to drive electric thrusters. The papers emphasize the nuclear reactor successes as well as the challenges of operating in space. As in the case of the megawatt lasers for beamed energy, the decision to invest in space nuclear reactors will probably be dictated by requirements for the main mission. However, mission analyses clearly show that the existence of a megawatt reactor, as part of the mission payload, will enable heretofore unobtainable payload fractions, as a direct result of higher propulsion efficiencies, greatly reducing the fuel mass for orbit transfer. In the first paper, Myrabo and Powell review the progress of the U.S. space nuclear reactors program, which in effect stopped in 1972. The concepts covered range from solid core reactors, to fluidized bed reactors, to gaseous fission reactors. In the

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second paper, Buden and Sullivan describe the long trip-time missions which are suited to nuclear-powered electric propulsion and point out that direct thruster nuclear rockets may not be cost effective for the shorter trip-time missions. Particularly significant is their summary of the readily achievable requirements for avoiding radiological risk to the Earth's populace and environment. Closedcycle megawatt power systems will be impractical unless major advances are made in lightweight waste-heat radiator systems. The paper introduces a series of imaginative and thought-provoking radiator concepts. The word ultra-performance in the title of the paper by Rosa and Myrabo is an appropriate adjective for several of their high-temperature (>4000 K) closed-cycle systems. If their goals can be realized, the resulting order-of-magnitude increases in kW/kg will enable electrical propulsion to rely on nuclear reactors even if high power is not required for the main mission. Myrabo offers several approaches to generating high peak power (direct thrust or electric) in open-cycle modes or auxiliary power in closedcycle modes. Powell and Botts describe the advances that have been made in defining both the rotating bed and fixed bed particle-fuel reactors. From the standpoint of direct thrust, such configurations may achieve important increases in working fluid temperature. Jones provides an in-depth examination of the issues relating to heat transfer and fluid mechanics of fluidized reactor beds. Garrison, Nock, and Jones introduce several aspects of achieving a spacecraft configuration that is compatible with the nuclear power system. The final paper in the volume is in the chapter on advanced chemical propulsion. The absence of other papers in this chapter reflects the relative technical maturity of chemical propulsion with respect to the types of propulsion addressed in previous chapters. However, continued advances in conventional solid and liquid space engines will insure that they will have significant roles in the 1990's. Weber's results show that important gains in both specific impulse and propellant density are possible through the use of fluorine, lithium, and hydrogen. The scope of the chapters and the focus of the papers were developed in two workshops held in Orlando, Fla., during January 1982. Each of the papers was peer reviewed by the authors contributing to the chapter. The reviews were for technical accuracy and clarity; no premium was placed on consensus views. Indeed, individual and bold positions were encouraged. An iterative review and revision process was carried out to achieve balanced coverage of the prospects and barriers to advanced space propulsion. Where appropriate, the authors included statements on research needs and approaches. The workshops established the following guidelines for the preparation of papers. Nonchemical propulsion concepts may have

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relied either on power sources onboard for the main mission or on power beamed from Earth or orbiting stations. The technology was to be consistent with optimistic projections for the 1995-2000 time frame. The papers were to provide broad coverage of the underlying technologies leading to descriptions of the technical and research issues. Where appropriate, performance projections were to be included. The emphasis was on establishing the conceptual and basic research aspects of processes, mechanisms, coupling, etc. The papers were not intended to provide solutions; rather they were to provide an introduction and perspective on the challenges. The issues were necessarily complex; controversial statements were expected. From the foregoing introduction, the reader may get the impression that the papers have a decided device orientation whereas the volume title promises research status and needs. This is not a contradiction. Describing the research needs in the context of intended applications often accelerates the communication process needed to establish new research directed at specific goals. In putting together the individual papers and chapters, one of the first obligations was to establish which concepts are of interest for the 1995-2000 time frame. This naturally leads to analyses of systems and devices. This open and effective advocacy is part of the recently revitalized national forum to clarify the issues and approaches which relate to major advances in space propulsion. As editor I am indebted to the expertise, dedication, and objectivity of the Editorial Committee and the authors, which for this volume are one and the same. They accepted the task of contributing to and reviewing a volume which carries two messages: 1) major advances in space propulsion are both obtainable and necessary, and 2) considerable progress is being made in defining the long-term research goals and approaches. Since most of the papers were not supported under active projects and since the invited paper process involved several iterations among the authors, to various degrees all of the authors used the editorial, illustration, and word processing services of the BDM Corporation. The sustained cooperation and editorial support of Joann S. Kieffer during the lengthy preparation and revision period is gratefully acknowledged. Leik N. Myrabo was instrumental in stimulating several of the papers in Chapter IV. I gratefully acknowledge the cooperation of Camille S. Koorey, Managing Editor of the Series, and Dr. Martin Summerfield, Editor-in-Chief of the AIAA Progress in Astronautics and Aeronautics Series. Leonard H. Caveny December 1983

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Progress in Astronautics and Aeronautics Volume Titles *1.

*2.

Volume Editors Martin Summerfield

Solid Propellant Rocket Research. 1960

Princeton University Loren E. Bollinger

Liquid Rockets and Propellants. 1960

The Ohio State University Martin Goldsmith The Rand Corporation Alexis W. Lemmon Jr.

Battelle Memorial Institute *3.

Nathan W. Snyder

Energy Conversion for Space Power. 1961

Institute for Defense Analyses

*4.

Space Power Systems. 1961

Nathan W. Snyder Institute for Defense Analyses

*5.

Electrostatic Propulsion. 1961

David B. Langmuir Space Technology Laboratories, Inc. Ernst Stuhlinger NASA George C. Marshall Space Flight Center J. M. Sellen Jr. Space Technology Laboratories, Inc.

*6.

Detonation and Two-Phase Flow. 1962

S. S. Penner California Institute of Technology F. A. Williams Harvard University

*7.

Hypersonic Flow Research. 1962

Frederick R. Riddell A VCO Corporation

*8.

Guidance and Control. 1962

Robert E. Roberson Consultant James S. Farrior

Lockheed Missiles and Space Company *9.

Ernst Stuhlinger NASA George C. Marshall Space Flight Center

Electric Propulsion Development. 1963

*Nowout of print. XV

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xvi *10. Technology of Lunar Exploration. 1963

Clifford I. Cummings and Harold R. Lawrence Jet Propulsion Laboratory

"11. Power Systems for Space Flight. 1963

Morris A. Zipkin and Russell N. Edwards General Electric Company

*12. lonization in HighTemperature Gases. 1963

Kurt E. Shuler, Editor National Bureau of Standards John B. Fenn, Associate Editor Princeton University

*13. Guidance and Control—II. 1964

Robert C. Langford General Precision Inc. Charles J. Mundo Institute of Naval Studies

. Celestial Mechanics and Astrodynamics. 1964

Victor G. Szebehely Yale University Observatory

*15. Heterogeneous Combustion. 1964

Hans G. Wolfhard Institute for Defense Analyses Irvin Glassman Princeton University Leon Green Jr. Air Force Systems Command

*16. Space Power Systems Engineering. 1966

George C. Szego Institute for Defense Analyses J. Edward Taylor TRW Inc.

*17. Methods in Astrodynamics and Celestial Mechanics. 1966

Raynor L. Duncombe U. S. Naval Observatory Victor G. Szebehely Yale University Observatory

*18. Thermophysics and Temperature Control of Spacecraft and Entry Vehicles. 1966

Gerhard B. Heller NASA George C. Marshall Space Flight Center

*19. Communication Satellite Systems Technology. 1966

Richard B. Mars ten Radio Corporation of America

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*20. Thermophysics of Spacecraft and Planetary Bodies: Radiation Properties of Solids and the Electromagnetic Radiation Environment in Space. 1967

Gerhard B. Heller NASA George C. Marshall Space Flight Center

*21. Thermal Design Principles of Spacecraft and Entry Bodies. 1969

Jerry T. Bevans TRW Systems

*22. Stratospheric Circulation. 1969

Willis L. Webb Atmospheric Sciences Laboratory, White Sands, and University of Texas at El Paso

*23. Thermophysics: Applications to Thermal Design of Spacecraft. 1970

Jerry T. Bevans TRW Systems

24. Heat Transfer and Spacecraft Thermal Control. 1971

John W. Lucas Jet Propulsion Laboratory

25. Communications Satellites for the 70's: Technology. 1971

Nathaniel E. Feldman

26. Communications Satellites for the 70's: Systems. 1971

Nathaniel E. Feldman The Rand Corporation Charles M. Kelly

The Rand Corporation Charles M. Kelly The Aerospace Corporation

The Aerospace Corporation 27. Thermospheric Circulation. 1972

Willis L. Webb Atmospheric Sciences Laboratory, White Sands, and University of Texas at El Paso

28. Thermal Characteristics of the Moon. 1972

John W. Lucas Jet Propulsion Laboratory

29. Fundamentals of Spacecraft Thermal Design. 1972

John W. Lucas Jet Propulsion Laboratory

30. Solar Activity Observations and Predictions. 1972

Patrick S. Mclntosh and Murray Dryer Environmental Research Laboratories, National Oceanic

and Atmospheric Administration

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xviii 31. Thermal Control and Radiation. 1973

Chang-Lin Tien University of California, Berkeley

32. Communications Satellite Systems. 1974

P. L. Bargellini COMSA T Laboratories

33. Communications Satellite Technology. 1974

P. L. Bargellini COMSA T Laboratories

34. Instrumentation for Airbreathing Propulsion. 1974

Alien E. Fuhs Naval Postgraduate School Marshall Kingery Arnold Engineering Development Center

35. Thermophysics and Spacecraft Thermal Control. 1974

Robert G. Her ing University of Iowa

36. Thermal Pollution Analysis. 1975

Joseph A. Schetz Virginia Polytechnic Institute

37. Aeroacoustics: Jet and Combustion Noise; Duct Acoustics. 1975

Henry T. Nagamatsu, Editor General Electric Research and Development Center Jack V. O'Keefe, Associate Editor The Boeing Company Ira R. Schwartz, Associate Editor NASA Ames Research Center

38. Aeroacoustics: Fan, STOL, and Boundary Layer Noise; Sonic Boom; Aeroacoustics Instrumentation. 1975

Henry T. Nagamatsu, Editor General Electric Research and Development Center Jack V. O'Keefe, Associate Editor The Boeing Company Ira R. Schwartz, Associate Editor NASA Ames Research Center

39. Heat Transfer with Thermal Control Applications. 1975

M. Michael Yovanovich University of Waterloo

40. Aerodynamics of Base Combustion. 1976

S. N. B. Murthy, Editor

Purdue University J. R. Osborn, Associate Editor Purdue University A. W. Barrows and J. R. Ward, Associate Editors Ballistics Research Laboratories

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xix 41. Communication Satellite Developments: Systems. 1976

Gilbert E. LaVean Defense Communications Engineering Center William G. Schmidt CML Satellite Corporation

42. Communication Satellite Developments: Technology. 1976

William G. Schmidt CML Satellite Corporation Gilbert E. LaVean Defense Communications Engineering Center

43. Aeroacoustics: Jet Noise, Combustion and Core Engine Noise. 1976

Ira R. Schwartz, Editor NASA Ames Research Center Henry T. Nagamatsu, Associate Editor General Electric Research and Development Center Warren C. Strahle, Associate Editor Georgia Institute of Technology

44. Aeroacoustics: Fan Noise and Control; Duct Acoustics; Rotor Noise. 1976

Ira R. Schwartz, Editor NASA Ames Research Center Henry T. Nagamatsu, Associate Editor General Electric Research and Development Center Warren C. Strahle, Associate Editor Georgia Institute of Technology

45. Aeroacoustics: STOL Noise; Airframe and Airfoil Noise. 1976

Ira R. Schwartz, Editor NASA Ames Research Center Henry T. Nagamatsu, Associate Editor General Electric Research and Development Center Warren C. Strahle, Associate Editor Georgia Institute of Technology

46. Aeroacoustics: Acoustic Wave Propagation; Aircraft Noise Prediction; Aeroacoustic Instrumentation. 1976

Ira R. Schwartz, Editor NASA Ames Research Center Henry T. Nagamatsu, Associate Editor General Electric Research and Development Center Warren C. Strahle, Associate Editor Georgia Institute of Technology

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47. Spacecraft Charging by Magnetospheric Plasmas. 1976

Alan Rosen TRW Inc.

48. Scientific Investigations on the Skylab Satellite. 1976

Marion I. Kent and Ernst Stuhlinger NASA George C. Marshall Space Flight Center Shi-Tsan Wu The University of Alabama

49. Radiative Transfer and Thermal Control. 1976

Allie M. Smith ARO Inc.

50. Exploration of the Outer Solar System. 1977

Eugene W. Greenstadt TRW Inc. Murray Dryer National Oceanic and Atmospheric Administration Devrie S. Intriligator University of Southern California

51. Rarefied Gas Dynamics, Parts I and II (two volumes). 1977

J. Leith Potter ARO Inc.

52. Materials Sciences in Space with Application to Space Processing. 1977

Leo Steg General Electric Company

53. Experimental Diagnostics in Gas Phase Combustion Systems. 1977

Ben T. Zinn, Editor

Georgia Institute of Technology Craig T. Bowman, Associate Editor Stanford University Daniel L. Hartley, Associate Editor Sandia Laboratories Edward W. Price, Associate Editor Georgia Institute of Technology James G. Skifstad, Associate Editor Purdue University

54. Satellite Communications: Future Systems. 1977

David Jarett TRW Inc.

55. Satellite Communications: Advanced Technologies. 1977

David Jarett TRW Inc.

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56. Thermophysics of Spacecraft and Outer Planet Entry Probes. 1977

Allie M. Smith ARO Inc.

57. Space-Based Manufacturing from Nonterrestrial Materials. 1977

Gerard K. O'Neill, Editor Princeton University Brian O'Leary, Assistant Editor Princeton University

58. Turbulent Combustion. 1978

Lawrence A. Kennedy State University of New York at Buffalo

59. Aerodynamic Heating and Thermal Protection Systems. 1978

Leroy S. Fletcher University of Virginia

60. Heat Transfer and Thermal Control Systems. 1978

Leroy S. Fletcher University of Virginia

61. Radiation Energy Conversion in Space. 1978

NASA Ames Research Center

62. Alternative Hydrocarbon Fuels: Combustion and Chemical Kinetics. 1978 63. Experimental Diagnostics in Combustion of Solids. 1978

Kenneth W. Billman

Craig T. Bowman Stanford University Jorgen Birkeland Department of Energy Thomas L. Boggs

Naval Weapons Center Ben T. Zinn Georgia Institute of Technology

64. Outer Planet Entry Heating and Thermal Protection. 1979

Raymond Viskanta Purdue University

65. Thermophysics and Thermal Control. 1979

Raymond Viskanta Purdue University

66. Interior Ballistics of Guns. 1979

Herman Krier University of Illinois at Urbana-Champaign Martin Summerfield New York University

67. Remote Sensing of Earth from Space: Role of "Smart Sensors." 1979

Roger A. Breckenridge NASA Langley Research Center

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XX11

68. Injection and Mixing in Turbulent Flow. 1980

Joseph A. Schetz Virginia Polytechnic Institute and State University

69. Entry Heating and Thermal Protection. 1980

Walter B. Olstad NASA Headquarters

70. Heat Transfer, Thermal Control, and Heat Pipes. 1980

Walter B. Olstad NASA Headquarters

71. Space Systems and Their Interactions with Earth's Space Environment. 1980

Henry B. Garrett and Charles P. Pike Hanscom Air Force Base

72. Viscous Flow Drag Reduction. 1980

Gary R. Hough Vought Advanced

Technology Center 73. Combustion Experiments in a Zero-Gravity Laboratory. 1981

Thomas H. Cochran NASA Lewis Research Center

74. Rarefied Gas Dynamics, Parts I and II (two volumes). 1981

Sam S. Fisher University of Virginia at Charlottesville

75. Gasdynamics of Detonations and Explosions. 1981

J. R. Bowen University of Wisconsin at Madison N. Manson Universite de Poitiers A. K. Oppenheim University of California at Berkeley R. I. Soloukhin Institute of Heat and Mass Transfer, BSSR Academy of Sciences

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XX111

76. Combustion in Reactive Systems. 1981

J. R. Bo wen University of Wisconsin at Madison N. Manson Universite de Poitiers A. K. Oppenheim University of California at Berkeley R. I. Soloukhin Institute of Heat and Mass Transfer, BSSR Academy of Sciences

77. Aerothermodynamics and Planetary Entry. 1981

A. L. Crosbie

78. Heat Transfer and Thermal Control. 1981

A. L. Crosbie University of Missouri-Rolla

79. Electric Propulsion and Its Applications to Space Missions. 1981

NASA Lewis Research Center

University of Missouri-Rolla

Robert C. Finke

80. Aero-Optical Phenomena. 1982

Keith G. Gilbert and Leonard J. Otten Air Force Weapons Laboratory

81. Transonic Aerodynamics. 1982

David Nixon Nielsen Engineering & Research, Inc.

82. Thermophysics of Atmospheric Entry. 1982

T. E. Horton

83. Spacecraft Radiative Transfer and Temperature Control. 1982

T. E. Horton

84. Liquid-Metal Flows and Magnetohydrodynamics. 1983

H. Branover Ben-Gurion University of the Negev

The University of Mississippi

The University of Mississippi

P. S. Lykoudis

Purdue University A. Yakhot Ben-Gurion University of the Negev

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XXIV

85. Entry Vehicle Heating and Thermal Protection Systems: Space Shuttle, Solar Star probe, Jupiter Galileo Probe. 1983

Paul E. Bauer McDonnell Douglas Astronautics Company Howard E. Collicott The Boeing Company

86. Spacecraft Thermal Control, Design, and Operation. 1983

Howard E. Collicott The Boeing Company Paul E. Bauer McDonnell Douglas Astronautics Company

87. Shock Waves, Explosions, and Detonations. 1983

J. R. Bowen University of Washington N. Manson Universite de Poitiers A. K. Oppenheim University of California at Berkeley R. I. Soloukhin Institute of Heat and Mass Transfer, BSSR Academy of Sciences

88. Flames, Lasers, and Reactive Systems. 1983

J. R. Bowen University of Washington N. Manson Universite de Poitiers A. K. Oppenheim University of California

at Berkeley R. I. Soloukhin Institute of Heat and Mass Transfer, BSSR Academy of Sciences 89. Orbit-Raising and Maneuvering Propulsion: Research Status and Needs. 1984

(Other volumes are planned.)

Leonard H. Caveny Air Force Office of Scientific Research, Boiling AFB

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Purchased from American Institute of Aeronautics and Astronautics

Power-Beaming Technology for Laser Propulsion Leik N. Myrabo*

The BDM Corporation McLean, Va. Abstract This introduction to beamed-energy propulsion addresses some of the background issues and technology relating to such items as laser power sources, mirrors, desired wavelength, pointing, and tracking. The laser thruster configurations are surveyed by other papers in the chapter. The laser energy absorption schemes suggested in the laser propulsion literature include: inverse bremsstrahlung, particulate and molecular seeding, and heatexchanger mechanisms. Both continuously-operating (CW) and repetitively-pulsed (RP) thrusters have been tested. Studies of candidate molecular absorbers indicate a clear preference for specific laser wavelengths. Present understanding does not permit the prediction of which thruster concept and associated wavelength/waveform will lead to superior performance with future full-scale engines. Furthermore, a given thruster concept may be ideal for some missions, but inferior for others. Also, orbit-raising missions with attendant ranges of 10^ km or more will dictate the use of very short wavelengths (e.g., from 0.35 to 2.5 ym), which may, to a large extent, predetermine the choice of thruster concept. Regardless of the long-term outcome Invited paper received February 1, 1983. Revision received July 29, 1983. Copyright © American Institute of Aeronautics and Astronautics, Inc., 1983. All rights reserved. ^Scientist, Advanced Technology Group; now Assistant Professor of Mechanical Engineering, Rensselaer Polytechnic Institute,. Troy, N.Y.

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of thruster technology, however, one revolutionary powerbeaming technology appears to hold much promise for high versatility in serving the widest variety of potential engine concepts and missions: the free-electron laser (PEL). Not only does the PEL have exceptional flexibility and the promise of high conversion efficiency, but it is being pursued aggressively for other military directed-energy applications. The embryonic field of laser propulsion stands to benefit greatly from the future availability of a powerful ground- or space-based PEL source with adequate laser power for orbit-raising missions, and superior pointing and tracking capabilities that exceed propulsion needs. Introduction

The concept of beaming power from a remote source directly to a spacecraft propulsion presents a revolutionary point of departure from conventional chemical and electric rocketry. Since the power source remains independent of the spacecraft, a beamed-energy propulsion system can simultaneously overcome two traditional limitations: specific impulse and specific power. Beamed-energy thrusters may attain specific impulses approaching the regime of electric propulsion with thrust/mass ratios more typical of chemical rockets. Thruster design can be optimized for simplicity and high performance because the remote energy source can be made relatively sophisticated, by comparison. Thruster Design Laser propulsion concepts can generally be placed into three distinct generic classes: laser-thermal, laser-electric, and combined laser-thermal/electric. Most of the current attention is focused on the laser-thermal variety which absorbs beamed radiation into solid, liquid, or gaseous propellants by mechanisms of (I) inverse bremsstrahlung, (2) continuum absorption, (3) molecular (line) absorption, (4) parttculate absorption, or (5) indirectly, by way of heat exchangers. Kemp, et.al.' present a review of the current state of knowledge of laser energy absorption in gas engines. Three laser thermal rocket engine concepts are shown in Fig. I. The first two concepts, Pig. IA and IB, absorb repetitively-pulsed (RP) radiation into solid and gaseous propellants, respectively. Rosen and Pirri et.al. 2 review the

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POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION -LASER

H2 + SEED CONDENSED FUEL , PREPARATION

- LSC-WAVE IGNITION FRONT STEADY FLOW THRUSTING HIGH FLUX LASER PULSE DETONATES VAPORS LASER PULSE

A) LASER-SUPPORTED DETONATION WAVE ON CONDENSED SURFACE

B) PULSE PRODUCES BLASE WAVE; ACOUSTIC COUPLED FUEL FLOW

C) LENS-WINDOW; LASER-SUPPORTED COMBUSTION WAVE HEATING

Fig. 1 Laser-thermal rocket engine concepts.

prospects for such RP laser propulsion engines. Figure 1C is a schematic diagram of a continuous wave (CW) laser-heated thruster concept which utilizes seeded H2 propellant. Kemp and Legner^ discuss the areas in which research is needed for such CW laser-heated rockets. In a subsequent paper, Keefer and Jones et.al.4 identify additional CW thrusters research issues related to the ignition of hydrogen plasmas, position of absorbing plasmas within a laser beam, and plasma stability. Next, McCay and Thoenes^ discuss the numerical modeling of CW laser thermal propulsion flows. Merkle^ examines the various losses which can be incurred in both CW and RP laser-heated thrusters. Laser-thermal thrusters can use one of two basic window concepts for transmission of radiant energy into internal engine chambers: a) material (solid) windows, or b) aerodynamic windows as discussed by Shoji7 For solid windows, current research is centered on developing low absorption, low reactivity materials, and in understanding the mechanisms of optical surface degradation. Aerodynamic windows are, of course, being developed for a number of laser generators themselves. In addition to other research areas, basic mechanisms of gas transport are being modeled to examine the limits on aero-window operation in the space environment.

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)

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Optical Train Design for OTV/Thruster

Mission and engine requirements strongly influence overall optical train design. Earth-based laser transmitters must be equipped with adaptive optics to successfully propagate the beam up through the atmosphere without suffering excessive distortion caused by turbulence and thermal blooming. Clearly, space-based sources will avoid all such requirements on the laser transmitter. The overall geometry, size, articulation, and, therefore, weight of the receiver optics (mounted to the laserpropelled spacecraft) are dictated by the choice of laser wavelength, distance between source and receiver, and basing of the source. Mirror elements in the train can be reflecting or refracting, simple or compound. In most cases, reflecting elements will require high reflectivity, multilayer films to reduce cooling loads and thermal-induced optical distortion. High power windows and refracting mirrors may need anti-reflection coatings to reduce losses. Depending upon the incident time-average flux, engine mirror elements can either be radiantly-cooled (e.g., large primary receiver mirrors) or actively-cooled by gaseous, liquid or perhaps cryogenic coolants. For certain spacecraft designs, optical elements may be rotated to reduce the time-average thermal load. Heavy mirror cooling loads might be met by a new class of advanced "heat-pipe" system with superior heat-transfer capabilities.8 Certain reflecting optical elements may avoid heavy thermal loads by using grazing-incidence optics to spread the laser beam over a much larger footprint. Strict control over the incident laser beam pulse energy, fluence, time-average flux, and beam spot location is necessary to avoid exceeding the damage limits of elements in the optical train, as well as cooling limits of internal engine chamber walls. Inaccuracies in pointing would cause the powerful beam to stray across "softer" spacecraft external skins, so fail-safe control systems must be developed to prevent self-inflicted damage. Finally, for certain laser propulsion engines requiring seeding, careful control should be maintained over: a) injection of solid or liquid alkali seed materials into internal engine chambers, and b) subsequent ejection of contaminated high velocity exhaust plumes—to prevent coating spacecraft optical surfaces with highly absorbent particles. Otherwise, these surfaces might become so

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POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

7

seriously degraded in reflectivity as to trigger their own destruction. Assessment of Laser Power-Beaming Status The following discussion explores the state of development of principal elements of laser power-beaming systems. Laser Device Technology The candidate concepts for laser propulsion engines will collectively demand a high degree of flexibility from any futuristic HEL device technology. Depending on the thruster concept, laser power-beaming stations must deliver either CW or RP waveforms. Furthermore, thruster design and mission requirements will certainly combine in unexpected ways to demand a wide variety of laser wavelengths. For this reason, laser device technologies which depend upon direct-pumping by chemical, nuclear, and solar sources will probably be eliminated from considerations. By the very nature of their design, these lasers impose severe restrictions upon the resulting wavelengths and waveforms which can be generated. The free-electron laser (FEL) is the only high-energy electric laser technology which holds great promise for such versatility. The FEL is undergoing extensive development at several laboratories^"^ for defense power-beaming roles—largely because of the excellent prospect for achieving high brightness at short wavelengths (e.g., 0.35 to 1.0 ym) with high "wall-plug" efficiency (e.g., 25 to 50%). A good review of the several possible paths to high efficiency, high power FELs is given in Ref. 13. Informed experts agree on the feasibility of developing efficient FELs at power levels which meet the needs of future laser propulsion orbitraising missions, given sufficient funding. HEL Mirror Technology

The ability of an HEL mirror to handle a given level of laser power depends upon its diameter, total reflectivity, and cooling method. At power levels of interest for FEL power-beaming to laser propulsion systems, actively-cooled bare metal mirrors will be unable to maintain the required figure of merit at all interesting wavelengths, due to excessive levels

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8

L.N. MYRABO

of thermal absorption. One exception is for the case of very large mirrors which dilute the energy fluence to a survivable value. In this instance, aluminum is best at 0.35 ym giving 92% reflectivity; silver is best at 1.0 ym, yielding 99% reflectivity. Mirror coatings to achieve reflectivities better than 99.5% appear feasible at all but the shortest visible wavelengths, where atomic dimensions become significant in determining the achievable uniformity of layer thickness. Multi-layer thin films will be necessary to achieve reflectances of 99.5-99.9%. Mirror cooling constraints can be very severe for the smaller high power mirrors in the optical train, but new innovations appear to be providing the needed advances in the state-of-the-art for all but the shortest visible wavelengths. A comprehensive review of HEL mirror coating and cooling technology can be found in Ref. 8. Shown in Fig. 2 are the limits on total incident laser power imposed by the properties of mirrors using representative cooling schemes: i.e., a) water, b) liquid-metal (NaK eutectic), and c) wick-type heat pipes. Boundaries are shown on 99% and 99.9% surface reflectivity. A line is also indicated for 92% reflective bare aluminum at 0.35 ym. The 99% figures represent bare metal at 1.0 ym. As indicated, a 5 m diameter mirror with 99.9% reflective coating and heat pipe cooling system can handle (at near normal incidence) a 50 MW laser beam; 10 M mirror - 200 MW; 20 m mirror - 800 MW.

Basing Considerations

The most expedient route for earliest deployment of an effective PEL power-beaming system for laser propulsion applications would be to base it on the ground, and use lightweight "relay" mirrors (placed in fixed orbits) to reflect the laser beam as shown in Fig. 3. Since the laser installation would be on the ground, scientists could devote their efforts to developing the basic device and optical train, without simultaneously having to worry about such potentially critical issues as the system mass, volume, overall dimensions, and survivability during launch. The electric power plant becomes a non-issue, since today's conventional city power plants reliably generate 100 MWe or more. Also, the laser device and optical train

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I 23 CD

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Fig. 2 Power handling capabilities of HEL mirrors.

CO

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ORBITAL MOTION

C BEACON >: TRANSMITTER AND " DATA LINK

Fig. 3 Free-electron laser autonomous power source. atmospheric extinction.

site with orbital relay mirror, and High elevation reduces losses due to

could be engineered for easy maintainability, initially using traditional practices. Ground-based laser (GBL) installations would be placed upon mountain tops in order to minimize beam distortion effects of the atmosphere. The use of PEL device technology would permit the laser wavelength to be selected for minimum atmospheric absorptivity; hence, thermal blooming could then be reduced to insignificance (if not entirely eliminated) by the proper choice of beam power and output projector diameter. The beam-spreading effects of atmospheric turbulence can be greatly reduced by using adaptive "uplink" projector optics. The simplest embodiment of this scheme employs a beacon attached to the space relay platform. The beacon transmits an initially undistorted spherical wavefront from the relay satellite toward the ground transmitter to diag-

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POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

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nose the effects of the atmosphere on the beam path. (The transmitter aperture would be shared by a sensing system operating between laser power pulses.) The sensing system would provide feedback correction signals for the adaptive elements to pre-bias outgoing high power wavefronts so that they would arrive at the relay satellite as a diffractionlimited beam. This concept is further described in Ref. 14. A full-scale power-beaming system for laser propulsion would involve, for example, 6 to 10 mountain top installations with perhaps 2 to 4 GLBs at each site. Installations would be placed in "sunbelt" sites carefully selected from long-term meteorological data—such that one site could always count on clear access links to space relays. The number of relay mirrors at moderate (e.g., 3000 to 5000 km) orbits would be selected to maintain ''uplink" zenith angles within ±60° from vertical. OTV spacecraft coming in low over the horizon would seldom be engaged directly, but would be engaged by way of the space relay mirror to minimize atmospheric attenuation. Space-based lasers (SBL) have the advantage of inherently avoiding the atmospheric turbulence correction problem, but face the severe restrictions of light-weight design engineering for the laser device, power supply, beam projector, and integrating spacecraft structure. Efficient packaging for shuttle boost to orbit, as well as methods of packaging for easy erection (once in space) are additional major considerations. The rejection of waste heat from both power supply and laser device (to space) becomes a problem of paramount importance. Therefore, these requirements may push the deployment of space-based freeelectron lasers for OTV power-beaming applications into the "far-term", possibly as late as the 2lst century. Figure 4 (adapted from Ref. I5) is a mapping of representative laser parameters into beam brightness, the best measure of overall power-beaming laser system capability. Since this quantity is measured at the final output mirror (or projector)—on a direct line-of-sight path to the OTV vehicle, the chart applies equally well to both space-based and ground-to-space relay laser systems. The map reveals the interplay of equivalent mirror (aperture) diameter, wavelength, pointing and tracking accuracy ("jitter"), and laser power. An expression for beam brightness (B) is given as:

B = (P/2TTOJ. )

L (W/sterradian)

(1)

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12

LN. MYRABO

Fig. 4

Mapping of laser device/system parameters (from Ref. 15).

where P is the laser power in watts, 07 is the total beam beam spread in radians, and L is the beam quality. As indicated in Fig. 4, a 50 MW beam of 0.35 ym radiation will give a brightness of 5XI0 2 ' w/sterradian with a 5 m aperture, a very low value of jitter and perfect beam quality. A 10 MW beam appears close to the minimum useful level for "near-term" orbit-raising propulsion applications. Adaptive "Uplink" Optics

Several approaches are being considered to address the "uplink" turbulence correction problem. These include piezo-electrically driven mirrors of both segmented and deformable membrane type, and also electro-optical devices (particularly non-linear phase conjugators) which potentially can generate the needed phase corrections with no moving parts.

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POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

13

Requirements for vigorous "uplink'' turbulence correction will require a frequency response of 3 kHz. In the near-term (5-IO yr), this requirement will probably dictate a segmented approach to construction of the phase adaptive mirror, rather than the deformable 'membrane approach. At the shortest wavelength (0.35 ym) and large zenith angles (e.g., 60°), the atmospheric coherence scale (ro) is 2-3 cm under worst conditions. This situation will require the transmitter primary mirror to perform the phase correction function, because thermal loading of thousands of tiny elements in a secondary mirror would pose unsolvable problems. For laser propulsion application, uplink projection of a several hundred megawatt beam at 0.35 ym requires roughly a 4 m primary aperture because of linear and non-linear atmospheric propagation limitations. Hence, the primary mirror must be assembled from ~4 x 10^ adaptive elements. Atmospheric Propagation Constraints

Any serious consideration of the Ground-Based Laser/ Space-Based Relay (GBL/SBR) concept must recognize that up-link propagation of the laser beam through the atmosphere will be a decisive factor in determining feasibility of the system. It appears, on the basis of the current perspective, that turbulence problems are solvable in much of the wavelength range from 0.35 to 2.4 ym, but that the longer

wavelengths are preferable because of much lower absorption and scattering losses, provided that other system issues do not dominate the propagation issue. Atmospheric

radiation

extinction

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short

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tabulated information is available for laser lines of interest

and

a

number

of

different

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T R = exp - U

a

+

°a

(2)

where k a and o a are the aerosol absorption and scattering coefficients, Km and o m are the corresponding values for

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14

LN. MYRABO

molecules. The transmittance for a given wavelength is best for clear days, i.e., low aerosol contamination. Any attempt to propagate a high power laser beam from the ground to space will be heavily influenced by the integrated properties of the atmosphere in the beam path. The height of the GBL ground site above sea level is the principal determinant of the overlying air mass and its molecular composition. The zenith angle (i.e., angle Z, from the point straight overhead) determines the total distance traversed through the atmosphere, which is proportional to secant Z to a very good approximation. To deliver a prescribed brightness to the relay, the laser power must increased as a direct function of zenith angle to make up for losses caused by the atmosphere. The required number of relay satellites also becomes a function of zenith angle, because one or more satellites must always be visible within the accessible cone of zenith angles. The required number of relay satellites can be reduced by placing them in higher orbits where each satellite moves slower and views more of the earth; but this drives up the mirror size, and hence the mass, of each relay device. Other atmospheric factors are involved with the design of the ground-based transmitter optics. The projector aperture must be large enough to prevent the power density in the beam from exceeding the threshold where non-linear effects (e.g., thermal blooming and laserinduced breakdown) take place. Such effects vary greatly with wavelength. Also, the size of the atmospheric coherence length, ro, varies with wavelength. The associated disturbances can be corrected with adaptive optics of either the continuous membrane or multisegmented types, so r c determines the size, and thus the required number, of the adaptive zones in the mirror surface. Uplink Propagation Calculation

In this section, the "uplink" entire atmospheric transmission problem is analyzed in sufficient detail to determine a good approximation for all of the required system parameters. The work was first done analytically, and then the results were checked by modifying the Air Force state-of-the-art propagation code, ESP4-SWL, to simulate the GBL/SBR problem for laser propulsion

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POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

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applications. The results by both methods are in excellent agreement. Figure 5 displays the typical atmospheric transmittance to space from 3 km mountain peaks as a function of wavelength for several zenith angles of interest. A zenith of 60° is equivalent to two vertical atmospheric pathlengths The overall atmospheric bandpass spectrum is adapted from astronomical data. Atmospheric transmittance to space is clearly a strong function of wavelength and zenith angle. Superimposed upon the plot are values of transmittance calculated by the ESP4-SWL code for various wavelengths and zenith angles of interest, with the following exemplary assumptions: Mid-latitude summer atmosphere, 10 m/sec crosswind to clear the beam, 3 km altitude groundbased laser site, and beam slew calculated for direct uplink to relay mirrors in orbit at 3000 km. Figure 6A displays this same ESP4-SWL data, cross-plotted to show more clearly the transmittance versus zenith angle as a function of wavelength. Figure 6b displays the values for uncorrected turbulence-induced beam spread calculated by the ESP4-SWL code. It is evident that a factor of 30 to 40 reduction in these values must be accomplished by the adaptive transmitter optics in order to prevent the uplink beam from being dominated in turbulence considerations alone. Presented in Fig. 7 are the values for the thermal blooming parameter, IREL (relative irradiance), calculated by the code. The results indicate that for wavelengths of 0.35, 0.459, 0.66567, and 1.06, thermal blooming probably can be completely corrected by adaptive optics at zenith angles out to 60° and power levels up to 350-400 MW. Figure 8 shows the anticipated values for the coherence length, r o , as a function of zenith angle and laser wavelength. The calculation was made with Fried's turbulence model assuming a C n ^ value typical of conditions at a 3 km altitude site. Using the Miller and Zriske profile, the turbulence coherence length can be expressed as

( 2 ) r o = 0.1847 ) —————^——— r°° , ^2 / x '

I /.- * c

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20

LN. MYRABO

appears adequate for reducing thermal blooming to within limits correctable by adaptive optics (assuming a 10 m/s cross-wind velocity to clear the beam), even for zenith angles approaching 60°. Simple calculations indicate that electrical air breakdown will not occur at these power levels, pulse repetition frequencies, and pulse durations of interest. No other non-linear atmospheric propagation effects are anticipated. Therefore, the problem reduces to the linear regime where the necessary relay satellite receiver aperture becomes a direct function of transmitter diameter, zenith angle, relay altitude, and, of course, laser wavelength. The "uplink" map in Fig. 9 presents the results of such a calculation. The following initial assumptions are made in the analysis: a) the laser device has a specific beam quality (here beam quality refers to the narrow-angle beam spread component, and not the wide angle component associated with distant side-lobes), b) the "uplink" zenith angle is set at zero degrees (i.e., straight up), c) the ground-based laser is placed at the top of a 3 km altitude mountain, d) the use of adaptive optics reduces turbulence-induced beam spread by a factor of 40 from the uncorrected case, and finally, e) the relay receiver aperture is sized to intercept 95% of the incident radiation. Assuming a Gaussian distribution to represent the received beam intensity, the relay extends only to the e"~2 intensity level. The minimum beam spot size projected by GBLs on space relay receptive optics is a large function of the beam spreading due to diffraction, atmospheric turbulence, thermal blooming, and jitter of the transmitting optics. The total beam spreading may be expressed as

„*-,,* +

^

+

3+

.1

where D, T, J, and B denote the contributions due to diffraction, turbulence, jitter, and thermal blooming. The beam spread component due to diffraction alone is given by:

OD = 0 . 4 5 X B / D

(7)

where X is the laser wavelength, B is the narrow-angle beam quality, and D is the diameter of the GBL transmitter aperture. Hence, the peak intensity lo projected upon the relay mirror is

Purchased from American Institute of Aeronautics and Astronautics TRANSMITTER APERTURE DIAMETER, (M) RELAY DIAMETER • RIM INTENSITY = e'2 lpEAK

FINAL BEAM QUALITY

• »

95% ENERGY CAPTURED ZENITH = 0°

0.45

I 23 DO

O frl

O DIFFRACTION LIMIT

•

BEAM QUAL. - 1 0

(JT. LOW POWER BEAM DIVERGENCE (/IRAD)

O

5

=0.45 A./D

TURBULENCE • MAUI, 3 KM ALT. • PHASE CORRECTED (FACTOR OF 20) • ZENITH = 0° 0.233 URAD

i \

=*-|

§ m 3)

3

Fig. 9

Mapping of "uplink" laser propagation parameters.

Purchased from American Institute of Aeronautics and Astronautics

22

LN. MYRABO

I0 = P T R / 2 ^ ( o R ) 2

(8)

where P is the total transmitted power at the GBL aperture, Tp is the transmissivity of the atmosphere, a is the onesigma beam spread angle, and R is the range between the GSL laser and relay satellite.

insights from Uplink Calculation Clearly, it is the shortest wavelengths and the largest zenith angles that place the most stringent requirements on uplink adaptive optics—and fix the minimum effective atmospheric coherence length, which in turn sets the maximum dimensions of the GBL adaptive transmitter elements themselves. For 0.35 ym radiation, the elements must be roughly 2-3 cm across to be effective in turbulence correction at zenith angles between 60 and 70°. Regarding the effects of reduced beam quality (narrow angle beam spread only), the results summarized above are not significantly changed for any zenith angles between 0 and 60° if the relay design requires the same aperture diameter for projector and receiver elements. Generally, the beam would have to be expanded anyway to fill the entire receiver aperture. Laser Power and Run Time Requirements: Near-Term OTV Missions Long-running high energy FELs provide an exciting option for laser propulsion systems to perform useful orbital transfer missions. Significant payloads (e.g., 3 X I04 kg) can be raised to long-term parking orbits using moderate size laser system (e.g., 20 MW) with run times of less than a day. There are many ways to group and plot intersecting system parameters to serve as mission analysis tools. The author generated "first-order" estimate charts in Figs. 10 and II (from Ref. 20) to show the logical relationships among all of the principal parameters of laser propulsion for orbitraising missions. The first chart (Fig. 10) relates laser power to achievable orbital height for specified performance of the laser propulsion engine. Figure II, the second, uses a plausible tug model to find the duration of operation to raise a given payload to a given orbital height using the laser power found from the first chart.

Purchased from American Institute of Aeronautics and Astronautics

POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

23

REQUIRED FOR ORBITAL TRANSFER (M/SEC)

Mllhoz I I I I Mlilio3 I

Z

-r-

W

^

Fig. 10 Mapping of from 185 km,

laser propulsion parameters for orbit-raising

In the upper right hand quadrant of Fig. 10, the facing graphic plots the key mission parameter of a given increase in orbital velocity (total Av) required to deliver any payload from a 185 kms orbit to any selected orbital altitude. The remaining curves represent parametric assumptions to describe particular prdpulsion system options that lead to required laser power (upper left hand quadrant of Fig. 10), The significant engine performance parameters are specific impulse, l s p / and the energy coupling coefficient, C, which relates to the conversion efficiency from collected laser power to rocket thrust (e.g., units of dynes/watt). The chosen combination of C and l s p defines the required fuel flow rate. An example of how to use this plot is shown for the mission of raising a 32 metric ton payload (approximate weight of the expended shuttle main tank) from I85 to 3000 km orbit using 3.6 metric tons of residual hydrogen

Purchased from American Institute of Aeronautics and Astronautics

24

LN. MYRABO

Fig. 11 Mapping of performance parameters for TRW laser-propelled orbital transfer vehicle.

and a family of tug-like propulsion systems weighing between I and 5 metric tons (i.e., a mass fraction of approximately 0.9). Exhaust velocity for this example is selected as 10,000 m/s, corresponding to a thrust of 23,000 N, and the coupling coefficient is chosen to be C = 12. If these assumptions comprise a valid propulsion system, then the total power required is approximately 200 MW. Stippled areas have been added to Fig. 10 to designate areas of validity of plausibility. The chart may not be accurate to within '0% for mass ratios lower than 0.9, because if the fuel mass is sufficiently larger, then it will affect optimum mission parameters (see Fig. 11). The other boundaries of the stippled areas indicate a plausible regime vis-a-vis achievable physics. To go further, we must adopt a mathematical model of the laser tugboat. The mass of the tug is primarily

Purchased from American Institute of Aeronautics and Astronautics

POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

25

related to the thrust, both because of the size of the engine and pumps and because of the required stress-bearing components of the system as a whole. (Interestingly, the laser light collectors will have the same diameter regardless of the thrust for a specified laser wavelength and range to the laser source.) TRW has examined a laser-propelled tug (pictured in Fig. 12, with the addition of a shuttle external tank for payload), which seems to have plausible and justifiable characteristics.21 For the purposes here, the author derived the following scaling relationship based upon the conceptual TRW tug design.

My = 0.386 T + 964

(9)

where My is the total mass of a tug having a laser propulsion engine which develops thrust, T, measured in newtons. Again, it is assumed that the fuel mass will, in general, be a small fraction of the tug plus payload mass (< 10% of the total). It is important to understand that the chart in Fig. II is "slaved" to that in Fig. 10. The same thrust, altitude, and laser power must be used here that were chosen on the previous chart. In addition, the laser engine conversion efficiency is closely related to the coupling coefficient on the previous chart for a given engine design. Fifty percent efficiency is regarded as a reasonable value for C = 10 and l s p = 1000. With these constraints, we can then find the total thrusting time to perform the OTV mission. The dashed line in Fig. 11 applies to the mission of raising the Space Shuttle main tank to a 3000 km orbit from 185 km. It can be seen that this mission can be accomplished in -3500 s of thrusting time with 200 MW of delivered laser power. Or, retracing all of the steps, we find that the same mission can be performed in -26,000 s (7.2 hr) with 20 MW of laser power. Summary and Conclusions

It is unnecessary for laser propulsion researchers to: a) constrain their thruster concepts to operate on the narrow range of wavelengths and waveforms available with current laser device technology, or b) worry about the future availability of a versatile full-scale, high-power laser to exactly suit the needs of these thrusters—in 10 to 20 years

Purchased from American Institute of Aeronautics and Astronautics

rO O)

Laser Device

Reactor Power Plant (Closed Cycle)

Recieving Optics

Laser Propulsion Engine

Fig. 12 TRW's laser-propelled OTV tug concept (with shuttle main ET tank for payload).

Purchased from American Institute of Aeronautics and Astronautics

POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION

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when the technology is finally mature. The principal reason is that an ideal potential source already exists in the embryonic stage: i.e., the free-electron laser. Furthermore, PEL technology is actively being pursued for potential use in antiballistic missile (ABM) applications, because of the FEL's exceptional flexibility and promise of short wavelength operation at very high efficiencies. Projected GBL power levels for "robust" ABM roles agree well with that required for laser propulsion engines in near-term orbit-raising missions: e.g., tens to hundreds of MW. Also, a parallel intense development in sophisticated acquisition, pointing and tracking (AP&T) systems is being funded for defense laser applications. The demands placed upon AP&T systems by laser propulsion applications are significantly less severe than that required for ABM missions. On a unit process basis, the development of GBL/SBR power beaming stations suitable for future OTV laser propulsion can be postulated in terms of reasonable extensions of current technology. The process of integrating essential components into a workable system is largely an engineering challenge, requiring no major technology breakthroughs. However, certain physical processes require increased understanding which will come from additional directed basic research. In summary, if adequate research support is applied, all indications suggest that the problem of developing OTV power-beaming technology is tractable.

Acknowledgement Some of the technical background and illustrations for this chapter have been derived from prior work by the author, and J.D.G. Rather (previously of The BDM Corporation, now with DESE Research and Technology, McLean, Va).

References 1

Kemp, N. H., Rosen, D. I., and Legner, H. H., "Laser Energy Absorption: Research Problems/' Orbit-Raising and Maneuvering Propulsion—Research Status and Needs. AIAA Progress in Astronautics and Aeronautics, Vol. 89, edited by L. H. Caveny, AIAA, New York, I983, pp. 73.

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28

LN. MYRABO

2

Rosen, D. I., Pirri, A. A., Weiss, R. F., and Kemp, N. H., "Repetitively Pulsed Laser Propulsion - Needed Research/' Orbit-Raising and Maneuvering Propulsion — Research Status and Needs, AIAA Progress in Astronautics and Aeronautics, Vol. 89, edited by L. H. Caveny, AIAA, New York, 1983. pp. 95.

N. H. and Legner, H. H., "CW Laser Propulsion: Research Areas/' Orbit-Raising and Maneuvering Propulsion — Research Status and Needs, AIAA Progress in Astronautics and Aeronautics, Vol. 89, edited by L. H. Caveny, AIAA, New York, 1983. pp. 109. ^Keefer, D., Elkins, R.r Peters, C., and Jones L., "Laser Thermal Propulsion," Orbits-Raising and Maneuvering Propulsion — Research Status and Needs, AIAA Progress in Astronautics and Aeronautics, Vol. 89, edited by L. H. Caveny, AIAA, New York, 1983. pp. 129. 5 McCay, T. D. and Thoenes, J., "Numerical Modeling of Laser Thermal Propulsion Flows," Orbit — Raising and Maneuvering Propulsion — Research Status and Needs, AIAA Progress in Astronautics and Aeronautics, Vol. 89, edited by L. H. Caveny, AIAA, New York, 1983. pp. 149.

^Merkle, C. L., "Laser Radiation to Supply Energy for Propulsion," Orbits-Raising and Maneuvering Propulsion — Research Status and Needs, AIAA Progress in Astronautics and Aeronautics, Vol. 89, edited by L. H. Caveny, AIAA, New York, 1983. pp. 48. ^Shoji, J. M., Orbit"- Raising Needs, AIAA edited by L. H. 8

"Potential of Advanced Solar Thermal Propulsion," and Maneuvering Propulsion — Research Status and Progress in Astronautics and Aeronautics, Vol. 89, Caveny, AIAA, New York, 1983. pp. 30.

Sparks, M., "Assessment of Materials and Coatings Requirements

for UV Laser Optics," BDM/W-81-447-TR, Contract No. MDA90381-C-0151, The BDM Corporation, McLean, VA, 7 August 1981.

^Smith, T. I. and Madey, J. M., "Realizable Free Electron Lasers," Stanford University High Energy Physics Laboratory, Stanford, CA, Publication HELPL-885, December 1980. 1u *Brau, C. A., "Free Electron Lasers," Los Alamos National Laboratory, Report LA-UR-80-2015, Los Alamos, NM, 1980. 11 Neil, V. K. and Prosnitz, D., "Theoretical Investigation of a 10 MM Free Electron Laser Amplifier Using a 50 MeV Multi-kA Electron Beam," Lawrence Livermore National Laboratory, Livermore, CA, Report UCRL-85653, 24 June 1981.

J. M., "Scaling Relations for the Power Output of Gain Expanded Storage Ring Free Electron Lasers," Stanford University High Energy Physics Laboratory Publication HEPL-853, Stanford, CA, June 1979. ^Rather, J. D. G., Myrabo, L. N., and Henderson, W. D., "FreeElectron Lasers in Space," Report No. BDM/W-1217-82-S, Contract

Purchased from American Institute of Aeronautics and Astronautics

POWER-BEAMING TECHNOLOGY FOR LASER PROPULSION No. DAAHOI-8D-C-I587, The BDM Corporation, McLean, VA, April 1982.

^Rather, J. D. G., and Myrabo, L. N. "Alternative Laser Initiatives in Space," Report No. BDM/W-2I80-82-S, Contract No. DAAHOI82-S, Contract No. DAAHOI-80-C-I587, The BDM Corporation, McLean, VA, 9 October 1981. ^Henderson, W. D., "Space-Based Lasers—Ultimate ABM System?," Astronautics and Aeronautics, Vol. 20, No. 4, May 1982, p. 44. 16 Elterman, L., "UV, Visible, and IR Attenuation for Altitudes to 50 km," Environmental Research Papers No. 285, AFCRL-68-0153, Air Force Cambridge Research Laboratories, Bedford, MA, April 1968. 17 Zeideres, G. W., Schaefer, R., and French, F. W., "Short Wavelength Laser Beam Propagation Analysis for BMD Applications," WJSASTR-78-ID6, W. J. Schafer Associates, Inc., Wakefield, MA 28 February 1978.

18

Scherfflins, J. H., and Lush, K. J., "ARIS Aperture Relay," LMSCL030203, Contract No. DAAK90-76-C-DADE, 30 July I977. ^Decker, L., Aspinwall, D., Pohle, D., Dotson, R., and Bartosewz, M., "Aperture Relay Experimental Definition," LMSC-L048510, Contract No. DASG6-78-V-OIOQ, Lockheed Palo Alto Research Laboratory, Palo Alto, CA, 28 August 1979. 20

Rather, J. D. G., Borgo, P. A., and Myrabo, L. N., "Laser Propulsion

Support Program," Report No. BDM/W-80-652-TR, prepared for NASA Marshall SFC by The BDM Corporation, McLean, VA, 1 November 1980.

^Huberman, M., et. al., "Investigation of Beamed Energy Concepts for Propulsion," AFRPL-TR-76-66, Volume I, TRW Defense and Space Systems Group, Redondo Beach, CA, October 1976.

29

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Potential of Advanced Solar Thermal Propulsion James M. Shoji* Rockwell International Corporation, Canoga Park, Calif. Abstract With the advent of the Space Transportation System (STS) and a manifest of low Earth orbit (LEO) to geosynchronous equatorial orbit (GEO) missions, solar thermal propulsion offers a potentially attractive, high-performance propulsion system. This propulsion approach is capable of delivered specific impulse values of 800 Ibf-s/lbm to 1200 Ibf-s/ibm with low to moderate thrust levels. A variety of concepts utilizing solar radiation to heat propellents to high temperatures have been proposed and conceptually evaluated. These concepts include both windowless and windowed approaches. The windowless approaches represent a simpler, near term design with the least technical risks, and therefore a windowless heat exchanger cavity absorber/thruster has been analyzed, designed, and is currently being fabricated for ground test. Since this concept utilizes indirect heating, the achievable engine performance (i.e., maximum propellant temperature) is limited by the wall material maximum temperature limit. Advanced windowed concepts using direct solar radiation absorption are capable of delivered specific impulses approaching 1200 Ibf-s/lbm compared to the 800 to 900 Ibf-s/lbm of the windowless concept. Promising advanced concepts include Invited paper received March 22, 1982; revision received March 11, 1983. Copyright © by American Institute of Aeronautics and Astronautics, Inc., 1983. All rights reserved. *Manager, Advanced Systems Analysis, Rocketdyne Division.

30

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POTENTIAL OF ADVANCED SOLAR THERMAL PROPULSION

31

the rotating bed and the aerowindow absorber/thruster concepts. The ultimate performance potential and operational feasibility are dependent on further subcomponent improvements. The research and technology issues associated with these advanced concepts are identified. Introduction The basic mechanism involved in solar thermal propulsion is heating of the propellant within an absorber and subsequent discharge through a nozzle. This absorber is a direct or an indirect heat exchanger, depending on the specific concept, in which solar radiation is focused. The heated propellant then flows through a conventional converging-diverging nozzle to produce thrust. One may ask: "Why solar thermal propulsion?" As shown in Fig. I, solar thermal propulsion provides specific impulses and thrust levels between that of a chemical and an electric propulsion system. Also this propulsion concept makes use of an available power source, the sun, and therefore the development of a power source is not required. A comparison of chemical, ion, and solar propulsion systems are presented in Table I (Ref. I) for a LEO-to-GEO one-way mission from the shuttle. As shown in Table I, a solar thermal propulsion system with a moderate delivered specific impulse (855I N-s/kg or 872 Ibf-s/lbm) can provide a 45% increase in payload over a conventional chemical sysSOLAR-HEATED

10

100

1000

SPECIFIC IMPULSE, SEC

10,000

Fig. I Propulsion system performance comparison.

Purchased from American Institute of Aeronautics and Astronautics

32

J.M. SHOJI Table I Propulsion system comparisons

Av, tt/s a c

Trip time13

'.P. * Mass fraction

0

e

Payload, lb

L02-LH2

Ion

Solar I

14,000 5 h 475 0.90 20,400

19,200

19,200 14 days 872 0.85 20,500

180 days 2940 0.68 44,000

Solar 2

15,750

40 days 872 0.85 29,000

a

Total stage velocity increment required for LEO-to-GEO transfer. Elapsed time of LEO to GEO mission. c Delivered engine specific impulse. d Usable stage propellant mass/loaded stage propellant mass + spacecraft hardware mass except payload). e Usable mass delivered to GEOb

MIRROR AXIS OF REVOLUTION MIRROR INFLATABLE 2 REQ"D iEGMENT OF PARABOLOID)

PARABOLOID

INTERIOR MIRROR SURFACE MINIMIZES VEHICLE DAMAGE\M \//A FROM MISFOCUSED SOLAR ENERGY —»"^\l/Ar

MIRROR SUPPORT (INFLATABLE CONE)

Fig. 2

PARABOLOID THIS PORTION OF CONE WIRE REINFORCED AND MIRROR COATED INSIDE

Inflatable cone/paraboloid collector.

tern with a significant reduction in trip time over an ion propulsion system. A representative solar thermal propulsion system' is illustrated in Fig. 2. The solar radiation is collected using two large inflated solar collectors which are segments of a paraboloid. The collectors consist of a transparent front cover with a reflecting paraboloid segment inner surface. In addition, the configuration incorporated a conical internally mirrored surface which is intended to prevent stray solar radiation from impinging on the vehicle and acts like a 'light funnel/' The unique feature of this Rockwell Satellite

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POTENTIAL OF ADVANCED SOLAR THERMAL PROPULSION

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Systems Division design is that the collector system can rotate about an axis perpendicular to the vehicle axis to track the sun irrespective of the direction of vehicle travel. The 73 m (240 ft) diameter configuration shown in Fig. 2 would provide approximately 890 N (200 Ib) of thrust. The solar absorber/thruster system is located at the focal point of the paraboloid. As shown in Fig. 3, several candidate propellants have been evaluated for application to solar thermal propulsion. These propellants were primarily considered as single or monopropellant as opposed to bipropellants, and included hydrazine (^H^, ammonia (IMHs), methane (CH^, and hydrogen (H2). In terms of theoretical specific impulse, the choice is obvious. Hydrogen provided the highest specific impulse even at lower temperatures and is the best coolant for convectively cooling absorber and thruster components. Therefore, hydrogen was selected. Candidate Concept Evaluation and Comparison Concept Description

For solar thermal propulsion, two basic approaches exist to heat the propellant. The first approach is indirect heating in which a physical wall exists between the solar radiation and the flowing propellant. The solar absorber in this approach is basically a radiant heat exchanger. Two concepts utilizing this approach are presented in Fig. 4. NOZZLE AREA RATIO ( € )

2000

Ho 400-n 200-

§01500

cc

O

111

X H

500

200 2000

4000 6000 8000 GAS TEMPERATURE, R

10,000

Fig. 3 Theoretical vacuum specific impulse variation.

Purchased from American Institute of Aeronautics and Astronautics

34

J.M. SHOJI

_

-J X \/

v

^=—7—^ \ HEAT EXCHANGER CAVITY (ABSORBER)

u

i -^»-^ > '

J

/

|

i

FOCUSED SOLAR RADIATION

x

FOCUSED _SOLAR RADIATION

^ ^

/V , v

SOLID

TRANSPIRATION COOLED LINER

HEAT EXCHANGER CAVITY (WINDOWLESS) WINDOWED HEAT EXCHANGER CAVITY

Fig. 4

Indirect solar absorption concepts.

The windowless heat exchanger cavity concept consists of a radiant cavity heat exchanger in which the focused solar radiation is accepted. The propellant flows through the walls of the heat exchanger to cool the walls and heat the propellant. This heated propellant then flows to a separate thruster. The windowed heat exchanger cavity concept consists of an integral absorber and thruster configuration in which the propellant is heated primarily as it flows through a transpiration cooled wall (Fig. 4). This w a f f is heated by the focused solar radiation entering the cavity through a solid window. Since the pressure drop through the transpiration cooled wall would be low, this concept would reduce the structural requirements of the high-temperature regions of the absorber. The basic disadvantage of indirect solar radiation absorption is that the maximum propellant temperature achieved is limited by the maximum allowable temperature of the wall material. The propellant cannot be heated to a temperature higher f han the material maximum allowable temperature. With current state-of-the-art materials, the maximum attainable propellant temperature using the indirect heating concepts is approximately 2778 K (5000 R). The second basic solar thermal propulsion approach is to directly heat the propellant through a window by introducing a seed or some other molecular constituent which absorbs solar radiation. The heat is transferred from the

Purchased from American Institute of Aeronautics and Astronautics

POTENTIAL OF ADVANCED SOLAR THERMAL PROPULSION

35

FOCUSED — — SOLAR RADIATION

Fig. 5

Direct solar absorption

concepts windowed molecular or participate concept (discharged seed).

absorbing media to the hydrogen through conduction, convection, and radiation. The concept illustrated in Fig. 5 incorporates a seedant/gas mixer prior to introduction into the absorber. Pure hydrogen is used to convectively cool the solid window and to prevent deposition of the absorbing media. In this concept, the absorbing media (typically high molecular weight) is discharged through the thruster. This discharged media results in a decrease in delivered specific impulse and could deposit on the solar collectors. In an attempt to alleviate both of the problems encountered with the previous concept, a concept proposed for nuclear reactors and modified for solar thermal propulsion is shown in Fig. 6. This concept incorporates a vortex flow to retain the absorbing media (seed). Gaseous propel-

FOCUSEO ^ _ _ _SOLAR ^" RADIATION

Fig. 6 Direct solar absorption concepts windowed vortex flow concept (retained seed).

Purchased from American Institute of Aeronautics and Astronautics

36

J.M. SHOJI

Fig. 7 Direct solar absorption concepts rotating bed concept retained seed.

lant is injected tangentially into the absorber to create a vortex action and the resultant centrifugal force in the rotating flow keeps the heavier seed near the walls of the absorber. The orifice-like entrance to the thruster prevents the loss of seedant. Analyses of this concept with spherical carbon particles led to the conclusion that from a solar radiation absorption standpoint, submicron particles were preferred and from the standpoint of particle retention, particles greater than 10 ym were preferred. Therefore, the particle size desired for heating was different than that desired for retaining the particles in the vortex flow. Another concept utilizing advanced nuclear reactor technology for space propulsion is illustrated in Fig. 7. This concept incorporated a rotating porous cylinder to create the centrifugal force required to retain the seed. The hydrogen flow is introduced radially in the absorber through the porous rotating cylinder. The seedant bed could be operated in a fluidized or packed bed mode. Cold tjow tests performed by Brookhaven National Laboratory for a nuclear space application indicated stable operation with uniform particle concentrations. The high temperatures achieved within the solar absorber will require hightemperature bearings and seals. Concept Comparison

The delivered performance and the technology unknowns were compared for each of the candidate con-

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POTENTIAL OF ADVANCED SOLAR THERMAL PROPULSION

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cepts. The delivered performance analyzed both the absorber and the thruster influences. Absorption, radiation, reradiation, and convective cooling were included in the absorber analysis. The thruster performance was determined using the one-dimensional equilibrium (ODE) theoretical vacuum specific impulse values and the simplified JANNAF procedure including nozzle geometric, boundary-layer, and reaction kinetic losses. The resulting delivered specific impulse and thrust are presented in Fig. 8 for the candidate concepts using a I00-to-l area ratio nozzle and two 30.5 m (IOO ft) diameter collectors, representing a fixed power input of approximately 1.5 mW. As shown in Fig. 8, the two indirect heated absorber/thruster concepts (windowless and windowed heat exchanger cavity) achieved a maximum specific impulse approaching 8826 N-s/kg (900 Ibf-s/lbm) at 2778 K (5000 R) hydrogen temperature. As previously discussed, this temperature is considered the maximum allowable temperature for current state-of-the-art materials. The thrust obtained from these concepts was in the ISO to 220 N (30 to 50 Ibf) thrust range. The windowed heat exchanger cavity concept resulted in a slightly lower delivered specific impulse due to the influence of window cooling. The absorber/thruster concepts using direct solar heating (rotating bed and vortex flow concepts) resulted in delivered specific impulses exceeding 9800 N-s/kg (IOOO Ibf s/lbm) and approaching 11,800 N-s/kg (1200 Ibf-s/lbm). Due to the constant available power from the two 30.5 m (IOO ft) ROTATING BED & VORTEX

FLOW CONCEPT \>>^2 laser (10.6 ym) and an e-beam pumped XeF laser (0.35 ym). While detailed comparisons of these models with experiments have yet to be made, the predictions of overall energy absorption and late time shock dynamics have been shown to be consistent with the available experimental data. In future studies, models similar to those developed in Ref. 9 should be used to predict the laser absorption dynamics that will occur for other interesting combinations of laser wavelength, laser pulse duration, propellent composition, and nozzle configuration. At the same time, welldesigned and carefully diagnosed experiments should be carried out. They will serve to test and validate the models, and to help point out where modeling improvements are most needed to further the understanding of laser energy absorption in pulsed laser propulsion. Mode-Media Interactions

In all studies to date, it has been assumed that the laser beam is unaffected by the absorbing gas, except for absorption reducing the beam intensity. In fact, the beam can interact with the absorbing medium in such a way as to lead to index of refraction changes. These changes can significantly alter the laser beam configuration, which can result in self-focusing or other mode-media interaction effects. These effects can in turn alter the absorption of laser energy. This possibility needs both theoretical and experimental study.

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LASER ENERGY ABSORPTION GASES

91

Conclusions

Laser energy absorption is a crucial element of any laser-powered gas engine, such as a CW or pulsed laserheated rocket. Absorption depends on the laser wavelength, the absorbing mechanism, the absorbing species, and (for pulsed applications) the laser pulse time. For CW applications, high temperature absorption can proceed by inverse bremsstrahlung as long as electrons are present. They can be provided by thermal ionization of the propellent gas (> 7000 K in hydrogen) or by the addition of easily ionized seedants {> 3000 K for cesium). To achieve this kind of absorption, the gas must be raised to the necessary temperature. One method for heating the gas is a laser-supported combustion (LSC) wave, which heats the gas by conduction and radiation from the hot absorption zone. A model of a one-dimensional LSC wave in pure hydrogen has been developed, showing peak temperatures near 20,000 K; and a simplified model with a two-dimensional temperature field has also been constructed. The addition of alkali metal seeds would greatly reduce the peak temperature and associated losses, and modeling of such seeded LSC waves is needed. The possibility of strong two-dimensional effects on both temperature and velocity exists, and needs to be assessed to ascertain the realism of the one-dimensional models. Experiments on LSC waves in candidate propellents are also needed to verify the models. A second method for heating the gas to ionization temperature is the use of molecular seedants which absorb by resonance processes. These molecules need to absorb at low temperature (~ 1000 K) and remain absorbing up to temperatures at which ionization may occur (~ 3000 K). Some candidate molecules have been studied, but no good absorber over this whole range has been identified for 10.6 ym radiation, for example. Experimental study of the absorption of candidate absorbers for various laser wavelengths are needed. Another important issue for molecular absorption is the "bleaching" phenomenon. At some laser intensity the gas will become transparent because the absorbing state population is depleted faster than it can be replenished by deactivation of the excited state population. Multiphoton absorption will ameliorate this effect though not

Purchased from American Institute of Aeronautics and Astronautics

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N.H. KEMP, D.I. ROSEN, AND H.H. LEGNER

completely. The laser intensity levels at which these phenomena occur are not known at high temperatures, and experimental studies are needed. The stability of the CW absorption process to perturbations in gas conditions is also an important issue. Some preliminary studies have indicated possible instabilities which may destroy the absorption. More definitive work is needed to specify the stable regions of absorption. For pulsed applications, laser breakdown thresholds are an important parameter. They are functions of the gas, the laser wavelength, and the laser pulse time. While reasonably well-known and understood for 10.6 ym radiation, there are serious discrepancies in the visible/UV wavelength region. Further experiments at candidate wavelengths in candidate gases are needed at various pulse times to develop a reliable and understood breakdown data base. The effects of possible mode-media interaction on laser absorption in gas engines are unknown, and an evaluation of importance of these phenomena is needed. References ^Kemp, N. H. and Lewis, P. F., "Laser-Heated Thruster Interim Report" NASA CR 161665 (PSI TR-205), Physical Sciences Inc., Andover, Mass., February 1980. ^Caledonia, G. E,, "Conversion of Laser Energy to Gas Kinetic Energy/' Journal of Energy, Vol. 1, March-April 1977, pp. 121-124. ^Kemp, N. H. and Krech, R. H., "Laser-Heated Thruster Final Report," NASA CR 161666 (PSI TR-220), Physical Sciences Inc., Andover, Mass., September 1980. E. R. and Krech,

R. H.,

"Absorptivity

of

Water

Vapor

for

10.6 ym Radiation," AIAA Journal, Vol. 20, June 1982, pp. 863-864. ^Fowler, M. C., "Measured Molecular Absorptivities for a Thruster," AIAA Journal, Vol. 19, August 1981, pp. 1009-1014. 6

Laser

Legner, H. H. and Douglas-Hamilton, D. H., "CW Laser Propulsion,"

Journal of Energy, Vol. 2, March-April 1978, pp. 85-94. 7 Kemp, N. H. and Root, R. G., "Analytical Study of Laser-Supported Combustion Waves in Hydrogen," NASA CR-135349, (PSI TR-97), Physical Sciences Inc., Andover, Mass., August 1977. See also: Journal of Energy, Vol. 3, January-February 1979, pp. 40-49.

^Keefer, D., Crowder, H., and Elkins, R., "A Two-Dimensional Model of the Hydrogen Plasma for a Laser-Powered Rocket," AIAA Paper 82-0404, January 1982.

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93

Wu, P. K. S. and Pirri, A. N., "Stability of Laser-Heated Flows/'

AtAA Journal, Vol. 14, March 1976, pp. 390-392. 1u

Baizer, Yu. P., "Heating of a Gas by a Powerful Light Pulse/' Soviet

Physics JETP, Vol. 21, 1965, pp. 1009-1017.

^Rosen, D. I., Kemp, N. H., Weyl, G., Nebolsine, P. E., and Kothandaraman, R., "Pulsed Laser Propulsion Studies Vol I, Thruster Physics and Performance," PSl TR-184, Physical Sciences Inc., Andover, Mass., 1982. 12 Jones, L. W. and Keefer, D. R., "The NASA Laser Propulsion Project: An Update," AtAA Paper 81-1248, June 1981. Published in Astronautics and Astronautics, Vol. 20, Sept. 1982, pp. 66-73.

^Jacob, J. H. and Moni, S. A., "Thermal Instabilities in High-Power Laser Discharges," Applied Physics Letters, Vol. 26, 15 January 1975, pp. 53-55. 14 Biblarz, 0. and Nelson, R. E., "Turbulence Effects on an Ambient Pressure Discharge," Journal of Applied Physics, Vol. 45, February

1974, pp. 633-637.

^Generaiov, N. A., Zimakov, V. P., Kozlov, G. I., Masyukov, V. A., and Raizer, Yu. P.; "Experimental Investigation of a Continuous Optical Discharge," Soviet Physics JETP, Vol. 34, April 1972, pp. 763-769.

fsr, D. R., Hendrickson, B. B., and Braerman, W. F., "Experimental Study of a Stationary Laser-Sustained Air Plasma," Journal of Applied Physics, Vol. 46, March 1975, pp. 1089-1093.

^Conrad, R. W., Roy, E. L., Pyles, C. F., and Mangum, D. W., "LaserSupported Combustion Wave Ignition in Hydrogen," Army Missile Command Technical Report RH-80-1, October 1979. 18

Weyl, G. M., Rosen, D. I., Wilson, J., and Seka, W., "Laser-Induced Physical Review A, Vol. 26,

Breakdown of Argon at 0.35 ym," August 1982, pp. 1164-1167.

^Buscher, H. T., Tomlinson, R. G., and Doman, E. K., "Frequency Dependence of Optically Induced Gas Breakdown," Physical Review Letters, Vol. 15, 1965, p. 847. 20

Alcock, S. J., Kato, K., and Richardson, M. C. "New Features of Laser-Induced Gas Breakdown in the Ultraviolet," Opticai Communications, Vol. 6, 1972, p. 342. i, A. N., "Theory of Momentum Transfer to a Surface with a High-Power Laser," The Physics of Fluids, Vol. 16, Sept. 1973, pp. 1435-1440. 22

Pirri,

A. N.,

Root,

R. G.,

and

Wu,

P. K. S.,

"Analytical

Laser/Material Interaction Investigations," TR-104, Physical Sciences Inc., Andover, Mass., Sept. 1977.

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N.H. KEMP, D.I. ROSEN, AND H.H. LEGNER 23Pirri, A. N., Root, R. G., and Wu, P. K. S., "Plasma Energy Transfer

to

Metal

Surfaces

Irradiated

by

Pulsed

Lasers/'

AIAA

Journal,

Vol. 16, 1978, pp. 1296-1304. 24

Root, R. G., Pirri, A. N., Wu, P. K. S., and Geiman, H., "Analyses of Laser-Target Interaction," TR-170A, Physical Sciences Inc., Andover, Mass., March 1979.

Purchased from American Institute of Aeronautics and Astronautics

Repetitively Pulsed Laser Propulsion: Needed Research David I. Rosen*, Anthony N. Pirri +

Robert F. Weiss,* and Nelson H. Kemp* Physical Sciences Inc., Woburn, Mass. Abstract This paper reviews the prospects for RP (repetitively pulsed) laser propulsion - a pulsed jet propulsion concept where the propellant energy is supplied by the absorption of short, repetitive laser pulses beamed to the thruster from a remote laser power station. This advanced propulsion concept promises several potential advantages for orbitraising and maneuvering missions including a remote power source, high specific impulse, high payload to total mass ratio (a consequence of the first two features}, and moderate to high thrust (limited primarily by the average laser power available). While analytical and experimental studies to date have made considerable progress toward establishing the feasibility of RP laser-heated thrusters, a number of critical technical issues still are unresolved. Introduction The concept of remotely powering a rocket thruster using a high power laser beam has been under consideration for nearly a decade.^ Studies have shown that this propulsion concept has several potential advantages including high invited paper received May 25, 1982; revision received Dec. 28, 1982. Copyright © by American Institute of Aeronautics and Astronautics, Inc., 1983. All rights reserved. * Principal Research Scientist. + Vice President of Defense Programs. ~~ President.

95

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specific impulse, moderate to high thrust (limited primarily by average laser power available), a remote power supply, and a high payload to total mass ratio. The combination of high specific impulse and high thrust is desirable for many space missions. Chemical propulsion is limited to specific impulses less than 500s because of upper limits to combustion temperatures. Electric propulsion is capable of high specific impulse, but is limited to low thrust by the necessity of carrying or generating electric power on board. Figure 1 illustrates thruster configurations for both continuous wave (CW) and repetitively pulsed (RP) laser propulsion schemes. In the CW thruster, CW laser energy is collected, focused, and absorbed in a heating chamber through which propellant gas continuously flows. The propellant gas is heated to very high temperatures (limited in practice by the amount of available laser energy and the maximum allowable temperature of the chamber walls) and allowed to expand through a nozzle converting thermal energy to thrust. The laser beam continuously provides energy to the propellant and the thrust remains constant in time. In this thruster the "combustion" side of the engine is REPETITIVELY PULSED LASER POWERED LASER-INDUCED BLAST WAVE

PROPELLANT STORAGE

PROPELLANT HEATING BY LASER ABSORPTION

SELF-FOCUSING NOZZLE

• CONTINUOUS WAVE LASER POWERED

EXHAUST COLLECTOR

PROPELLANT FEED

WINDOW

PROPELLANT HEATING

Fig. 1 Thruster configurations for laser propulsion.

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unconventional, but the sonic and supersonic regions of the rocket nozzle are standard. With the RP thruster, on the other hand, the subsonic flow is cold, and the propellent is "explosively" heated by bursts of laser energy that are focused by the supersonic nozzle. In this paper we will review the prospects for pulsed laser propulsion. In our discussion, we will describe the generic propulsion approach and its advantages, indicate the underlying technologies upon which it depends, and review the research accomplishments to date. We will then conclude with an assessment of the key research problems that remain to be solved if pulsed laser propulsion is to be viable by the 1990's. Early suggestions for CW thruster designs proposed an incident laser beam which entered the thruster from the nozzle exit and impinged upon a solid propellant in the throat region of the nozzle.^"5 The idea was to ablate the propellant and heat the ablated gas with the laser energy. However, it soon became apparent that laser absorption waves would form in the ablated gas, shielding the solid propellant from further laser heating and preventing addition production of ablated gas, until the initial gas was expelled out the nozzle. A similar effect occurs if the solid propellant is replaced by a flowing gas which absorbs the laser energy. One solution to this difficulty was the RP thruster, which uses short, spaced laser pulses to heat the gas, let it exhaust out the nozzle, and then repeat the cycle. 4 In the RP thruster concept shown in Fig. 2, parabolic nozzle walls focus the incoming beam to yield breakdown in a propellant at the focal region of the parabola. The resulting high pressure plasma is characteristic of detonation wave initiation by high power laser-induced breakdown. With a short laser pulse, the detonation wave quickly becomes a blast wave, which propagates to the nozzle exit plane converting the high pressure of the propellant gas behind it into a force on the nozzle wall. Propellant is fed to the focal region from a plenum chamber. The laserinduced blast wave stops the propellant flow through the throat until the pressure at the throat decays to the sonic pressure; then, the propellant flow restarts. The process is repeated with each successive laser pulse. In addition to circumventing the absorption decoupling problem associated with CW thrusters, the RP thruster concept presents other advantages. One is that the engine

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D.I. ROSEN, A.A. PIRRI, R.F. WEISS, AND N.H. KEMP

ALUMINUM PARABOLIC SHELL OF REVOLUTION

LASER INDUCED BREAKDOWN

PROPULSION SEQUENCE - COLD P R O P E L L A N T FLOWS T H R O U G H THROAT

- LASER E N E R G Y ABSORBED DOWNSTREAM AT FOCUS - SHOCKED GAS EXPANDS OUT NOZZLE

- S E Q U E N C E REPEATS: P R O P E L L A N T USE C O N T R O L L E D BY MASS FLOW AND LASER REPETITION R A T E Fig. 2

Repetitively pulsed laser powered thruster.

design is relatively simple, requiring no window and no external focusing optic. Another advantage is that the propellant requirements are rather flexible. Utilizing high intensity laser-heated breakdown to initiate the absorption process, pulsed laser-heated thrusters do not require propellants with any particular low temperature absorption properties. Impetus for research into RP thrusters has also come from developments in pulsed laser technology. Recently, significant advances have been made in the performance of RP laser devices.6'7 RP laser systems now exist or are under development that have average power outputs rivalling the larger CW laser devices. Furthermore, recent results indicate that pulsed laser propulsion missions may be compatible with the developing technology of high power, short wavelength RP lasers for space defense. The Underlying Technologies

The development of pulsed laser propulsion is dependent upon several underlying technologies.

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Laser Technology

RP laser devices are needed that have high average power, good beam quality, and long run times. Advances must also be made in the areas of beam propagation, pointing, and tracking. As mentioned above, many of these improvements are being pursued for space defense. Propellant Technology

There is a need for storable, low molecular weight propeliants which will undergo laser-induced breakdown, absorb the incident radiation strongly, and then allow efficient conversion of thermal energy to exhaust kinetic energy. Materials Technology

Chambers are needed that can withstand the inherent cyclical thermal and mechanical loadings. In addition, materials for the interior walls of the self-focusing nozzle must be found that are highly reflective to the incident laser radiation and will survive high temperatures. Flow Control or Valving Technology

To achieve efficient mass utilization, techniques are needed for high frequency valving. Valving schemes are required which are capable of large mass flow rates (>l kg/s) and which operate at frequencies > 10^ Hz and with characteristic opening and closing times 10"^ -* 10~3 s. The advances required in the above technologies may, in many cases, be achieved solely by good engineering design. However, in some cases, basic research will be needed. At the conclusion of this paper we will give our assessment of where basic research is needed. Research Accomplishments to Date The emphasis in pulsed laser propulsion research thus far has been on characterizing basic thruster design and performance. Toward this end, significant milestones have been achieved. Laboratory experiments in the United States and the Soviet Union suggest that high specific impulse and

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large thrust wilJ be feasible with RP thrusters; but scale-up from the laboratory to prototype must demonstrate this. Pulsed thruster designs are an outgrowth of laser/ plasma-interaction experiments using pulsed lasers. Evaporating a surface continuously with a laser does not yield high temperatures because the vaporization energy limits the temperature of the vapor to less than 4000 K. To raise the specific impulse, it is desirable to heat the vapor directly. However, stable heating of a gas to plasma temperatures involves either tailoring the spatial intensity distribution, controlling propellant flow rate, or proper propellant selection. To circumvent the stability problem, it is easiest to pulse the laser. Experimental studies of the interaction of high-intensity laser pulses with material surfaces^"^ have led to the simplest high specific impulse rocket concept. If the laser intensity rises high enough to break down the vapor or air above the surface, it will ignite a laser-supported detonation wave. A highly absorbing plasma with temperatures of 10,000-20,000 K propagates behind the detonation front. Proper containment and expansion of the plasma results in thrust at high specific impulse. The specific impulse achieved depends upon the laser intensity, propellant density, and the ability of the propellant to absorb laser radiation (determined by the ionization potential, molecular weight, and laser wavelength). Pirri and Weiss introduced the concept of the singlepulse laser-supported detonation wave rocket engine. 4 Subsequently, Douglas-Hamilton and colleagues proposed the double-pulse version of the detonation wave rocket shown schematically in Fig. 3.^ The first pulse determines the amount of propellant that will be heated by evaporating a controlled amount of solid or liquid, while the second pulse ignites the detonation wave plasma in the propeliant. Thus, two laser pulses yield one thrusting pulse, and the resulting plasma expands out a skirt nozzle before the next laser pulse. Calculations show that operating conditions for an 800 s specific impulse engine require evaporating propellant to obtain one-third the normal atmospheric density over a distance 7 cm from the base and a laser intensity of 2 x 10^ W/cm 2 for 10 ys. 1 1 An RP laser-powered rocket engine that operates with a self-regulating gaseous-propellant feed system has been tested at Physical Sciences Inc. 12 '^ Figure 2 shows a schematic diagram of the thruster that illustrates the

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S O L I D OR L I Q U I D P R O P E L L A N T SURFACE

EXPANSION SKIRT LOW FLUX L A S E R

PULSE

STEP I: A LOW FLUX LASER PULSE ABLATES A MEASURED AMOUNT OF PROPELLANT GAS FROM A SOLID PROPELLANT SURFACE OR FROM A LIQUID FILM ON A POROUS PLATE.

PLASMA ABSORPTION FRONT

H I G H FLUX LASER P U L S E STEP II: A HIGH FLUX PULSE IGNITES A LASER SUSTAINED DETONATION WAVE AT THE SURFACE. A FRONT OF HIGHLY ABSORBING PLASMA, WHICH SHIELDS THE SURFACE FROM FURTHER LASER RADIATION. PROPAGATES BACK TOWARDS THE LASER, HEATING THE GAS PROPELLANT TO « 20.000 °K. THE BEAM IS TURNED OFF WHEN THE FRONT HAS PASSED THROUGH MOST OF THE GAS. THE CYCLE IS REPEATED AFTER THE HOT GASES HAVE EXPANDED.

Fig. 3 Laser-sustained-detonation wave rocket concept - proposed by Avco Everett Research Laboratory.

propulsion sequence. Cold gaseous propellant enters the throat from a plenum and laser-induced gas breakdown takes place downstream of the throat. A parabolic nozzle focuses the incoming radiation to the propellant-breakdown intensity, and thrust results from propagation of the blast wave generated by the breakdown. Since the pressure behind the blast wave, as it starts to propagate down the nozzle, greatly exceeds the plenum pressure, the propellant feed stops until the blast wave pressure decreases below the plenum pressure. Thus, the propellant feed system becomes "self-regulating," providing fuel for the breakdown before the next laser pulse. Figure 4 is a photograph of one version of this laserpowered thruster. This version has a conical nozzle, since early experiments with it employed external focusing optics. With external optics the intensity in the breakdown region

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D.I. ROSEN, A.A. PIRRI, R.F. WEISS, AND N.H. KEMP

Fig. 4

Pulsed laser powered thruster with conical nozzle.

can be controlled to insure a strong blast wave. The fullscale thruster, if powered by a ground-based laser, requires parabolic nozzle focusing, since the beamed power density must be kept below atmospheric propagation limits. Experiments were performed with the thruster configurations shown in Figs. 2 and 4 using multiple 10.6 ym laser pulses of 10 J energy.^2,13 Figure 5 gives specific impulse and energy conversion efficiency (defined as the ratio of the exhaust kinetic energy per pulse to the laser energy per pulse) obtained with a self-focusing parabolic nozzle, argon and hydrogen propellants, and a background environment of 10~4 atm. The specific impulse for constant laser energy per pulse is controlled by the time between laser pulses and the propellant mass flow through the throat. It can be seen that a maximum specific impulse of 1000 s was obtained using hydrogen as the propellant. With argon, the maximum specific impulse from this small-scale thruster was 500 s. The physics of the thrusting process depends upon the background environment. At 1 atm, the nozzle entrains several shocks from successive laser pulses. Figure 6 gives experimental measurements obtained with helium propellant and a 1 atm background environment^ 2 '^ j^e theoretical lines presented in Fig. 6 correspond to differing values of the energy conversion efficiency.

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104

D.I. ROSEN, A.A. PIRRI, R.F. WEISS, AND N.H. KEMP

propellant leads to the laser-powered pulsejet By replacing the plenum shown in Fig. 2 with an air inlet and diffuses we could design an engine that requires no propellant. Extensive single-pulse testing of this concept has been reported in the Soviet literature. 5/15,16 Typical results for the pressure impulse in the Soviet experiments include experimental values of up to 45 dyne-s/J for the ratio of the impulse to the laser energy^ - a somewhat higher value than early experiments gave with this concept. 4 All the thruster experiments described above have utilized 10.6 ym pulsed lasers. Recently, the scaling of thruster performance to visible/UV wavelengths has been addressed. Theoretical calculations have shown that although inverse bremsstrahlung heating of electrons by free-free absorption decreases significantly at these wavelengths, the absorption due to bound-free transitions (photoionization) increases. The net result is that laser-induced breakdown thresholds as well as absorption coefficients in the laser-produced plasmas are found to be within acceptable operating limits. ^ TO support the theoretical predictions, a single-pulse thruster experiment was carried out at 0.35 ym.18 With a conical nozzle and external focusing optics, argon and helium propellents were heated with approximately 5 J pulses of XeF laser energy. The measured strengths of the resulting blast waves indicated that energy conversion efficiencies (laser energy into blast wave energy) in excess of 50% were achieved for plenum delivery pressures 10 atm. Future Research Needs While promising results have been achieved with smallscale RP thrusters, they are still laboratory devices. Several issues must be resolved before pulsed laser propulsion concepts evolve into full scale rocket engines. Laser-induced Breakdown

The laser-induced formation of an optically absorbing plasma is an essential first step in the pulsed laser propulsion process. The threshold flux and pulse time required to ignite a plasma will in general depend upon laser wavelength, propetlant species, propellant density, and the presence of any impurities with low ionization potential. While there is a great deal of experimental data and much

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REPETITIVELY PULSED LASER PROPULSION

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theoretical understanding for laser-induced gas breakdown at infrared wavelengths (i.e., X > 1 ym), this is not the case in the visible/UV. Since visible/uv lasers are presently candidates for high power RP systems for space applications (due to their reduced diffraction-limited beam spreading over long ranges), further research in this wavelength regime is merited. Based on a review of previous investigations and on PSI's recent studies in this area, we believe that some of the key issues that need to be addressed are: • Pulse Duration Scaling: Much of the available breakdown data have been obtained for short pulse durations, i.e., ~ 10~8 s. Since efficient thruster performance will require longer pulses, i.e., > 10~6 s, it is essential that the scaling of breakdown thresholds with increasing pulse duration be determined. Experimental measurements should be performed with beams that are well characterized both temporally and spatially. To provide understanding of the data, theoretical modeling is also merited. Theoretical calculations can also be used to yield predictions in regimes not accessible to present laboratory lasers. • Effects of Particulates and Low lonization Potential Seed Species: Focusing requirements to achieve the threshold fluxes for propellant breakdown might be relaxed by adding a small percentage of particulates or low ionization potential seed gas. Investigations of such approaches could prove very valuable. • Effects of Self-Focusing and Other Nonlinear Processes: Previous studies 1 ^ have suggested that selffocusing may play a significant role in the laser-induced breakdown process at optical frequencies. Appropriate experiments and theoretical analyses should be performed to resolve this issue. Plasma Optical Absorption As with laser-induced breakdown, laser/plasma optical absorption at infrared wavelengths is, in general, well understood and predictable. At visible/UV frequencies, however, this is not the case. Several issues requiring further investigation include: • What is the degree of nonequilibrium that occurs between electrons and heavy particles, i.e., what is the true electron temperature? How does this temperature

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D.I. ROSEN, A.A. PIRRI, R.F. WEISS, AND N.H. KEMP

affect the competition between the formation of excited states by three-body recombination and their depletion by photoionization? • What are the appropriate cross-sections for multiphoton ionization? In argon, for example, at X = 0.35 ym, the two-photon ionization of Ar*(4s) is believed to play an important role; yet the rate constant for this process has only been crudely estimated. Quantum-mechanical perturbation calculations should be carried out to provide a better estimate of this rate. Radiation Losses/Thermal Protection The nozzle walls of a pulsed device will contain very hot gases and be subject to intense radiation. While it has been estimated that the fraction of energy lost to radiation in the laboratory-scale experiments was relatively small, no detailed radiation analyses have yet been performed. Studies to evaluate the radiation losses from a full-scale breakdown plasma are especially needed. The interior nozzle walls of the pulsed thruster will be required to maintain high reflectivity and satisfactory optical quality while subjected to significant thermal and mechanical stresses. Research is needed to provide an accurate assessment of these stresses as well as to arrive at material configurations that will satisfy the wall requirements. Chemical Nonequilibrium

The studies to date have not considered chemical nonequilibrium in the absorption or flow processes. Some molecular absorbers may not dissociate on the absorption time scale. Further, if high specific impulse is to be realized, it will be necessary to recover the chemical energy invested in internal degrees of freedom. Investigations should be performed, for example, into the degree of nonequilibrium that may occur in the acceleration of dissociated and ionized hydrogen. Propellant Technology

Research should be carried out in the area of novel storage techniques including the possible use of porous

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REPETITIVELY PULSED LASER PROPULSION

1 07

metal hydrides, or solid propellant formulations that could be "burned" to yield gaseous hydrogen as a major product. Pulsed Flow Control To achieve efficient mass utilization in pulsed thrusters, techniques for high frequency valving will be required. Valving concepts that allow large flow capacity are needed that can be operated at frequencies on the order of 10^ Hz and with opening and closing times of 10~4 •> 10"^ s. The degree of flow stability and uniformity that would result must also be evaluated. Beam Quality/Beam Propagation of High Energy RP Lasers Research needs to be performed on the inherent beam quality limitations of large Fresnel number, high gain RP lasers. Investigations to date on such systems have often revealed significant diffraction losses between the nearfield and far-field beam energy content. The origin of these losses must be understood if the output energy of future laser devices is to be delivered efficiently over long distances to designated receivers. For a ground-based laser, losses due to atmospheric propagation will also have to be understood and dealt with.

References 1 Kantrowitz, A., "The Relevance of Space/' Aeronautics, Vol. 9, No. 3, March 1971, pp. 34-35.

Astronautics

and

2

Kantrowitz, A., "Propulsion to Orbit by Ground-Based Lasers," Astronautics and Aeronautics, Vol. 10, No. 5, May 1972, p. 74. 3 Rom, F. E. and Putre, H. A., "Laser Propulsion," TM-X-2510, NASA, April 1972.

^Pirri, V N. and Weiss, No. 72-719, June 1972.

R. F.,

"Laser

Propulsion,"

AIAA

Paper

kin, F. V. and Prokhorov, A. M., "Uses of a Laser Energy Source in Producing a Reactive Thrust," Soviet Physics~-Uspekhi, Vol 19, No. 7, July 1976, pp. 561- 573, ^Bradford, R. S., Munch, J., Shackleford, W. L., and Schwartz, J., "TRW/Navy Fifty Liter Repetitively Pulsed Chemical Laser (RLCL-50)," Seventh Annual Tri-Service Chemical Laser Symposium, 25-28 October 1982, Kirtland AFB, N. Mex..

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7

McClure, J. D., et. a!.f "Development of a High Power Repetitively Pulsed DF Demonstrator Laser," Seventh Annual Tri-Service Chemical Laser Symposium, 25-28 October 1982, Kirtland AFB, N. Mex.. A. N., Schlier, R., and Northam, D., "Momentum Transfer and Plasma Formation Above a Surface With a High Power CC>2 Laser," Applied Physics Letters, Vol. 21, No. 3, August 1972, pp. 79-81. 9

Lowder, J. E., Lencioni, D. E., Hilton, T. W., and Hull, R. J., "HighEnergy Pulsed CC>2 Laser-Target Interaction in Air," Journal of Applied Physics, Vol. 44, No. 6, June 1973, pp. 2759-2762. 10 Hall, R. B., Maker, W. E., and Wei, P. S. P., "An Investigation of Laser-Supported Detonation Waves," AFWL-TR-73-28, June 1973.

^Douglas-Hamilton, D. H., Kantrowitz, A. R., and Reilly, D. A., "Laser Assisted Propulsion Research," Radiation Energy Conversion in Space, Progress in Astronautics and Aeronautics, Vol. 61, edited by K. W. Billman, AIAA, New York, 1978, pp. 271-278. 12

Nebolsine, P. E., Pirri, A. N., Goela, J. S., Simons, G. A., and Rosen, D. I.. "Pulsed Laser Propulsion," Physical Sciences Inc. TR-142; presented at the AIAA Conference on Fluid Dynamics of High-Power Lasers, Oct. 1978. See also Simons, G. A. and Pirri, A. N., "The Fluid Mechanics of Pulsed Laser Propulsion," AIAA Journal, Vol. 15, No. 6, June 1977, pp. 835-842. 13 Nebolsine, P. E., Pirri, A. N., Goela J. S., and Simons, G. A., "Pulsed Laser Propulsion," AIAA Journal, Vol. 19, No. 1, January 1981, pp. 127-128.

I^Pirri, A. N., Monsler, M. J., and Nebolsine, P. E., "Propulsion by Absorption of Laser Radiation," AIAA Journal, Vol. 12, No. 9, September 1974, pp. 1254-1261. I^Barchukov, A. I., Bunkin, F. V., Konov, V. I., and Prokhorov, A. M., "Laser Air-Jet Engine," JETP Letters, Vol. 23, No. 5, March 1976, pp. 213-215. ^Ageyev, V. P., Prokhorov, A. M., et at, Kvantovaya Elektronika, Vol. 4, No. 12, p. 2501 (in Russian); see Laser Focus, March 1978, p. 80. l, G. M. and Rosen, D. I., "Laser-Induced Breakdown of Argon at 0.35 ym," Physical pp. 1164-1167.

Review A,

Vol. 26,

No. 2,

August

1982,

18 Rosen, D. I., Kemp, N. H., and Weyl, G. M., "Pulsed Laser Propulsion Studies - Final Report," PSI TR-184, Physical Sciences Inc., Woburn, Mass., October 1982. 19 Alcock, A. J., DeMichelis, C., and Richardson, M. C., IEEE Journal of Quantum Electronics, Vol. QE-6, October 1970, p. 622.

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Steady (Continuous Wave) Laser Propulsion Research Areas Nelson H. Kemp* and Hartmut H. Legner* Physical Sciences Inc., Woburn, Mass. Abstract This paper discusses the areas in which research is needed for steady or continuous wave (CW) laser-heated rockets. These include gas confinement and losses, conversion of thermal to kinetic energy, and laser beam windows. (The area of absorption of laser energy is covered in a companion paper.) The quantitative results obtained to date in these areas are briefly discussed with some typical results presented. It is concluded that the concept of CW laserheated thrusters with l sp greater than 1000 s has so far shown no insuperable technical barriers. The heat loads are manageable, possible window concepts exist, and energy absorption into the gas appears to be feasible. The additional research areas which need study are then identified. Among these are the convective heating of walls in contact with hot (2 Laser Radiation," Applied Physics Letters, Vol. 22, May 1973, pp. 500-503. 16 Kemp, N. H. and Root, R. G., "Analytical Study of Laser Supported Combustion Waves in Hydrogen," PSI TR-97, NASA CR-135349, August 1977.

"^Appleton, J. P. and Bray, K. N. C., "The Conservation Equations for a Nonequilibrium Plasma," Journal of Fluid Mechanics, Vol. 20, Part 4, 1964, p. 659. r, y. P., "Subsonic Propagation of a Light Spark and Threshold Conditions for the Maintenance of Plasma by Radiation," Soviet Physics JETP, Vol. 31, Dec. 1970, p. 1148. ^Jackson, J. P. and Nielson, P. E., "Role of Radiative Transport in the Propagation of Laser Supported Combustion Waves," AIAA Journal, Vol. 12, Nov. 1974, p. 1498. J. H. and Keefer, D. R., "Two-Dimensional Generalization of Raizer's Analysis for the Subsonic Propagation of Laser Sparks," IEEE Transactions on Plasma Science. Vol. PS-2, Sept. 1974, p. 122.

O. C. and Thoenes, J., "Study of Laser Heated Propulsion Devices, Part 1," Final Report NASA Contract NAS8-33974, LMSCHREC-TR-D784744, April 1982. ^Robinson, E. G., Robinson, M. A., and Spadley, L. W., "GIM Code User's Manual," LMSC-HREC-TR-D784043, Jan. 1981. C., "Calculation of Radiative Properties of Nonequilibrium Hydrogen Plasma," Journal of Quantitative Spectroscopy & Radiative Transfer, Vol. 22, 1979, p. 101.

^Prandtl, L., "Ueber Fiussigkeitshewegung bei sehr kleiner Reibung," Verhandlung III, Intern. Math Kongress, Heidelberg, 1904.

Zs.

H., "Grenzschichten in Flussigkeiten mit kleiner Reibung," Math, u. Phys. Bd., Vol. 56, 1908, p. 1.

Purchased from American Institute of Aeronautics and Astronautics

Laser-Driven Repetitively-Pulsed MHD Generators A Conceptual Study Craig D. Maxwell"1"

STD Research Corporation, Arcadia, Calif. and Leik N. Myrabo*

The BDM Corporation, McLean, Va. Abstract A feasibility analysis is presented for laser-driven, repetitively-pulsed (RP) magnetohydrodynamic (XMHD) generators proposed for shared propulsion and electricpower roles in near Earth missions. Two basic power conversion approaches are analyzed in detail: 1) open-cycle H2 systems suitable for direct laser thermal RP rockets, and 2) closed-cycle monatomic gas systems compatible with RP electromagnetic accelerators. Although no such RP electric power generating devices currently exist, the analysis draws upon an extensive body of knowledge on analogous chemical explosive-driven magnetohydrodynamic generators. Finally, research and technology issues are identified. Introduction Described in Fig. 1 are the two approaches for shared electric power and propulsion explored in this paper. Both systems concepts would provide repetitively-pulsed (RP) Invited paper received February 1, 1983. Revision received July 29, 1983. Copyright © American Institute of Aeronautics and Astronautics, Inc., 1983. All rights reserved. + Vice President and MHD Project Manager. * Scientist, Advanced Technology Group; now Assistant Professor, Department of Mechanical Engineering, Rensselaer Polytechnic Institute, Troy, N.Y.

167

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168

c.D. MAXWELL AND LN. MYRABO A LASER THERMAL ROCKET CYCLE (HYDROGEN) RP PAYLOAD ELECTRIC POWER

RPLASER ENERGY

ROCKET EXHAUST

RADIATION SUPERHEATER

I |

SUPERSONIC NOZZLE

XMHD GENERATOR

b. LASER ELECTRIC ROCKET CYCLE (ARGON)

PUMP UJSPACE RADIATOR

2)

RP PAYLOAD ELECTRIC POWER

RADIATION SUPERHEATER

Fig. 1 Schematic diagrams of laser-driven space power supplies for shared propulsion and electric power roles.

electric power for use by the payload of an orbital transfer vehicle in near Earth missions. This XMHD generator concept offers some unique advantages as a very lightweight RP space power supply, especially in the multi-megawatt ranges. Figure 1a is a schematic diagram of an open-cycle space power plant which can generate either direct-laserheated rocket thrust, or pulsed electric power (i.e., when the XMHD generator field coils are energized). Gaseous H2

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is proposed as the open-cycle coolant/propellant which absorbs pulsed laser energy in a radiation superheater chamber before expanding into a combination XMHD generator/supersonic-nozzle. Whether producing electric power or rocket thrust, the heated H2 is finally ejected into space. Figure 1b shows an alternate closed-cycle space power plant scheme. Here a monatomic gas (e.g., argon) is proposed as the working fluid in a closed-cycle conversion scheme which rejects waste heat to space with a special high temperature radiator. As with the previous cycle, beamed RP laser power is first absorbed by the working gas in a radiation-superheater chamber; this high pressure, high temperature gas then expands through an XMHD generator, enters a space radiator and is finally recycled. The output pulsed electric power can be used either to 1) accelerate open-cycle propellant in an electric thruster or 2) energize the vehicle payload. As a thruster for orbital transfer vehicles, the power system in Fig. la should be able to produce rocket specific impulses of at least 1000 s, since this performance has already been demonstrated in pulsed laser propulsion experiments. Alternately, the options in Fig. 1 could enable specific impulses of 2000 to 5000 s or more, because electric thrusters are utilized. However, the overall efficiency of the latter cycle may be lower than the former, due to the additional laser-to-electric power conversion step. The analysis presented here assesses the feasibility of the XMHD generator concept itself, and generally ignores the thruster aspects. A reference design configuration for the concept is displayed in Fig. 2. The starting point for this analysis is the nominal system described in Refs. 1 and 2, for which the following parameters were used: 1) Working fluid: molecular hydrogen; ideal diatomic gas; ratio of specific heats = 1.4, Cp = 3.475 cal/g-K; initial temperature, 20.36 K. 2) Electrical conductivity = 1000 mho/m (obtained from nonequilibrium ionization of H2). 3) Mass of working fluid: 22.2 g (11.1 moles of H2) per pulse. 4) 5 MJ laser pulse. 5) XMHD generator dimensions: 0.1m X 1m X 1m. 6) Plenum chamber volume: 0.5 m^. 7) Magnetic induction: 2 T (up to 10 T). 8) MHD Generator output: 2.75 MJ(e) per pulse.

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170 •

C.D. MAXWELL AND LN. MYRABO PERSPECTIVE VIEW • LASER BEAM FOOT PRINT HIGH POWER LASER WINDOW TRANSPARENT "LIGHT BULB" RADIATOR ^ _ B

(END PLATE REMOVED FOR CLARITY)

PRIMARY MIRROR

7

GRAZING INCIDENCE SECONDARY MIRROR

x

MAGNETIC FIELD COIL

END VIEW

Fig. 2

Detailed configuration for reference open-cycle, H2 XMHD generator.

9) Enthalpy extraction: - 50%. 10) Specific energy extraction: 124 MJ/kg. These parameters are used as a reference to examine the physics of the electric-power generation process.

Relation to Previous Chemical Explosive MHD Work At the outset let it be said that MHD extraction of 50% of the total enthalpy of a working fluid is very far beyond the present state-of-the-art. The highest demonstrated enthalpy extraction ratios in any MHD generator experiment are those in seeded noble gas shock-tube experiments at General Electric^ and in the Netherlands'*, where levels approaching 25% were achieved with stagnation temperatures in the neighborhood of 4000 K. It is important

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171

to distinguish between the enthalpy extraction ratio (electrical power out/total power in) and the "turbine efficiency" (electrical power out/enthalpy change with isentropic expansion over the same pressure ratio). Thus, a turbine efficiency of 80% can be achieved easily in a generator that extracts only 20% of the total enthalpy of the working fluid. Nevertheless, it should be noted that advanced, high temperature turbines currently under development are able to demonstrate enthalpy extraction in the range of 40-50% from 2000-2500 K gaseous working fluids. Hence, the possibility of a unique approach to MHD generation yielding an enthalpy extraction of 50% cannot, from first principles, be ruled out. The projected specific energy extraction (2.75 MJ)/ (0.0222 kg) = 124 MJ/kg, is more than two orders of magnitude greater than the maximum demonstrated level of approximately 0.75 MJ/kg (Ref. 5, also with a top temperature of about 4000 K). However, this may not be an intrinsic limitation, because the total enthalpy of most of the working fluids considered to date was only typical of chemical levels, e.g., 4-6 MJ/kg. The total enthalpy of the working fluid is essentially unlimited with laser addition of energy. Finally, the highest demonstrated conversion efficiency of chemical energy to electrical energy in an explosive MHD generator has been approximately 5-6%.6 Again, as will be seen later, chemical-driven explosive generator experience does not necessarily place a limitation on laserdriven generators. Table 1 summarizes the results of the major chemicalexplosive MHD generator experiments to date. In Table 1, B is the applied magnetic induction, f-j is the percentage of the energy stored in the magnetic field Table 1 Experimental chemical-explosive MHD generator data Author

Ref.

Facility

B (T)

f

Jones Jones Jones Kirillin Kirillin

1, 8 1, 8 7, 8 9 9 9 10 6

2.5X2.5X46 cm3 2.5X10X46 cm3 20X15X78 cm3 VG-5 VG-8 VG-6 VG-10 X-MHD-14

1.7 2.3 2.8 1.35 0.8 2.5 5.3 3.5

30 35 40 30 40 50 50 54

Kirillin Asinovskiy

Bangerter

l' °/0

f

2' °/0

0.2 1.0 1.6 0.03 0.1 0.6 3.7 5.7

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C.D. MAXWELL AND LN. MYRABO

that was extracted as electricity, and \i is the percentage of the chemical energy released by the explosive that was extracted as electricity. More information on the above experiments is given in Refs. 11-17. Additional relevant citations to explosive MHD technology are given in Refs. 18-24. Among these, the work of Teno and Sonju at Avco Everett Research Laboratory^ W as primarily theoretical, although electrical

conductivities of explosion products were measured. Also worthy of mention is the work of Artec, Inc. on denseplasma explosive MHDJ9 in the latter work, extremely high power densities are achieved, but the conversion efficiency from chemical to

small.

electrical energy

is typically

The investigators who have achieved the highest levels of energy conversion are unanimous in stating that there is at present an inadequate understanding of the physical processes that control explosive generator performance (e.g., Refs. 6 and 10). Asinovskiy claims that the upper limit on

f-j is approximately 60%, while the practical upper limit for f2 is estimated to be approximately 6-9%. Bangerter indicates

that

extrapolated

Again,

it

is

to

his

conversion

approximately

emphasized

that

efficiencies

8%

these

chemically driven explosive generators.

with

could

be

optimization.

results

apply

to

Analysis of Reference Design (Open-Cycle H2) A

detailed

configuration

for

the

reference

XMHD

generator is presented in Fig. 2. Electromagnetic field coils are wrapped about the XMHD generator channel to produce a magnetic field vector in the direction indicated. As will be further described below, the pulsed laser energy enters a

high power material window located at the upstream end of

the plenum chamber, then impinges first upon the primary mirror and next upon the grazing-incident secondary mirror. After reflection off the secondary mirror, the radiation

emerges in a uniform, parallel beam to fill the entire crosssection of the MHD generator duct. The cycle begins when laser-induced electrical gas breakdown is triggered at the duct exit, thereby igniting a laser-supported detonation (LSD) wave.

Before proceeding, it is first necessary to distinguish

between the two different modes of MHD power generation

that have been postulated to occur in the reference open

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173

cycle configuration (Fig. 1a). In Mode 1, power generation takes place as the laser-supported detonation (LSD) wave passes back through the XMHD generator at high velocity. In Mode 2, the high temperature plasma remaining in the plenum chamber is vented through the generator. The power output time history is postulated to appear as in Fig. 3, in which a brief, intense burst of power is followed by a much longer duration, smaller amplitude power level with reversed voltage. Mode 1 was generally ignored in the earlier open-cycle laser-heated XMHD generator studiesj'2 because it was thought to contribute negligible amounts of pulsed power as compared with that of Mode 2. Figure 4 presents a schematic diagram of the power cycle for Mode 1. MHD generation requires a high conductivity working fluid moving at high velocity. As indicated, an LSD wave is ignited at the XMHD generator exhaust port, and propagates at a high velocity back through the generator. The point at which the LSD wave enters the plenum chamber signals the termination of Mode 1 and the beginning of Mode 2.

-1.0-

Fig. 3 XMHD H2 design).

generator

output waveform (reference open cycle,

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CD. MAXWELL AND LN. MYRABO

174 LASER BEAM , ON | | 1 1

1 1 1

\|-K

OPTICAL WINDOW ^ P L E N U M CHAMBER VOLUME 0 5 M' LASER INDUCED ELECTRICAL BREAKDOWN

NON-EQUILIBRIUM IONIZED PROPAGATING REGION

[•- 6 TO 10 CM CONDUCTING SLUG

END OF MODE 1 ELECTRIC POWER PULSE

LOCATION OF LSD WAVE AT END OF MODE 1

Fig. 4 Schematic diagram of power cycle for Mode 1, (open cycle, H2).

A schematic of the power cycle for Mode 2 is shown in Fig. 5. Mode 2 begins with the LSD wave reflecting off the secondary optics. Subsequently, the LSD absorption wave propagates through the plenum chamber, in the process creating a hot (10,000 to 15,000 K) high pressure hydrogen plasma. As indicated in Fig. 5, the laser pulse is terminated just before the LSD wave contacts the primary optic. Immediately, the LSD wave deteriorates to a strong blast

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LASER-DRIVEN PULSED MHD GENERATORS

175

LASER

HEL WINDOW

LSD W A V E R E F L E C T S OFF S E C O N D A R Y OPTICS S M A L L AMOUNT

OF GAS EJECTED

S E C O N D A R Y OPTICS

LASER BEAM

'

• MHD

GENERATOR

UNPROCESSED W O R K I N G

PROCESSED WORKING FLUID

ELECTRIC POWER

EJECTED GAS

LSD W A V E DEGENERATES TO STRONG BLAST W A V E

REFLECTED B L A S T W A V E

ELECTRIC POWER

Fig. 5 Schematic diagram of power cycle for Mode 2 (H2 gas, open cycle).

wave, reflects off the primary optics (i.e., the back end of the plenum chamber), and helps to expel all the hydrogen working fluid out through the XMHD generator—acting much like a piston. The MHD generator output voltage depends on the magnitude of the generated electric field (UB) and the

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1 76

C.D. MAXWELL AND LN. MYRABO

lateral dimensions of the cavity. In Mode 1, the conducting slug is moving with a high velocity to the left, whereas in Mode 2 the high pressure plenum plasma expands back out through the generator towards the right. Since the applied magnetic field is constant and U changes sign (i.e., goes through 0), the resultant output wave form will be as shown in Fig. 3. In Mode 1, additional analysis is required to establish whether the gas velocity behind the LSD wave is the same as the LSD wave velocity.

Mode 1 (Open-Cycle, H2 Device) A novel feature of Mode 1 operation is that the electrical generation is accomplished nearly simultaneously with the deposition of energy in the working fluid. This is much different than in conventional explosive MHD generators, where the energy is released at one instant, and the resulting slug of gas evolves in response to the MHD interactions with no means of making up momentum or energy losses by addition of more energy. The thermodynamic cycle for Mode 1 may perhaps be likened to that of a conventional multi-stage turbine with reheat between each stage, as portrayed schematically in Fig. 6. Conceptually, the XMHD generator concept could be envisioned as an infinite number of turbine stages running the entire length of the channel, with the addition of laser heat between stages. The high gas pressures created aft of the LSD absorption wave might be likened to the compressor function in this thermodynamic cycle. The velocity of the LSD wave is proportional to the cube root of the incident laser flux lo divided by ambient gas density p o in front of the wave. This relationship was first calculated by Raizer,25 and is given by:

where y is the ratio of specific heats. The pressure PDW behind the detonation wave is proportional to the product of the ambient gas density and the square of the detonation wave velocity (also by Raizer);25 P

DH ' "o

V

D/^+1>

The reference design assumed an incident laser flux density of 5 x 10? W/cm^ entering the XMHD generator and

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(ETC.) COMPRESSOR TURBINE (LSD WAVE)

TURBINE TURBINE

TURBINE TURBINE

TURBINE

- ANALOGY TO XMHD GENERATOR -

*Q=LASER HEATING

9

11 ,

(ROUGHLY 4000 K 60OO K)

MAXIMUM TEMPERATURE FOR NON REGENERATIVELY COOLED STRUCTURE. (1300 10,000 K above the ions and neutrals. The dependence of argon ionization upon temperature and pressure is shown in Fig. 14 (from Ref. 31). A sample calculation can more clearly demonstrate the LSD wave heating process. The internal energy (E|) of a mass (Am) of monatomic gas heated from initial temperature T o - to final temperature TO:? is approximately:

- T

= Am

\+

;, °

V ^T

(15)

where Cp is the coefficient of specific heat at a constant pressure, Myy is the molecular weight, N0 is Avagadro's number (6 X 10^3 atoms/mole), Bj is the degree of ionization, and C j is the ionization energy (15.75 eV for argon). The first quantity is the contribution from translation; the second from electronic excitation. To heat 100 g of argon from 500 to 30,000 K requires 1.5 MJ of energy. Ignoring for the moment the energy loss to ionization, and assuming ?2 = (T2/Ti) P-| we find that the static pressure of the gas at 30,000 K may be 62 atms. An additional 3.5 MJ is sufficient to ionize perhaps 93% of the 1.0-1

0.8-

z

o N

Z

o Li.

O LU

UJ

Q

0.2-

5,000

10,000

15,000

20,000

25,000

TEMPERATURE, K

Fig. 14 Dependence of ionization pressure, from Saha equation.

of

argon

on

temperature

and

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LASER-DRIVEN PULSED MHD GENERATORS

atoms.

1 93

Realistically, the Saha equation should be used to

determine the actual degree of ionization expected under these conditions, consistent with a total input laser pulse energy of 5 MJ. We now derive an expression to relate the pulsed laser beam parameters to the XMHD generator ("driver section") parameters, so that the detonation wave is able to process a!N the "driver section" gas exactly when the laser pulse terminates. Equation (1) gives the detonation wave velocity (VDW) as a function of laser flux (lo) and gas density (po). The total laser energy (Ep) deposited into the "driver section" gas can be related to the laser pulse duration (tp), and laser flux projected upon an LSD wave of crosssectional area (A), as follows:

Ep ' 'o * V where A

is

also the

cross-sectional

area

of

the

XMHD

generator (e.g., 0.1 m X 1.0 m for the reference design). Since the detonation wave must travel a distance equal to the length (L) of the XMHD generator channel during the laser pulse, we have:

tp We now solve for V

= L/V

(17)

and find that:

2 E V

DW

(T2

-

D

DW

For the argon design with a laser pulse energy of 5 MJ, Y = 1.67, p 0 = 1 kg/m3, A = 0.1 m2, 'and L = 1m, we determine that VDvv = 13-4 km/s' 'o = 66 - 9 MW/cm2, and t p = 74.8 us. (Note that Eq. (2) also tells us that the pressure behind the detonation wave will be 663 atm!) Assuming the total XMHD electrical pulse duration is 748 ys, we find from Eq.(3) that the current slug thickness (refer to Fig. 7) must be roughly 6 cm. If this can be accomplished the specific energy extraction will be 25.0 MJ/kg for the argon system. Table 5 presents the operating parameters for the argon design, differentiating between the Mode 1 and Mode 2 conditions.

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194

C.D. MAXWELL AND LN. MYRABO Table 5 Operating parameters for closed cycle XMHD

generator design (argon)

Mode 1a

Parameter Input laser energy (MJ) Conversion efficiency (K = 0.8), % Output electric energy (MJ) Mode 2 peak gas temperature (K) Mode 2 cut off temperature (K) Mode 2 peak gas pressure (atm) Working fluid conductivity (mho/m) Internal resistance (milli-ohm) Applied magnetic field (T) Magnetic Reynolds number Peak power density (GW/m^) Inductance (]aH) Short circuit current (MA) Open circuit voltage (kV) Peak load current (MA) Peak electrode current density (MA/m2) XMHD generator dimensions (m) Specific enthalphy extraction (MJ/kg) a D

Mode 2b 5.0 50 2.5 30,000 7500 66 5000

20.0

2.0 2.0

9.0

87.0 500

1.4 1.2

13.9 12.5 25.0

0.2 1.0

2.5 25.0 1 X 1 X 0.1 25.0

Assumes 10 cm slug length. Assumes 100 cm slug length.

Argon Gas Recirculation System A schematic diagram of the argon gas recirculation system is presented in Fig. 15. As mentioned earlier, the "compressor" function is provided by the -700 atm blast wave shed by^ the LSD wave at the end of Mode 1, which acts much like a piston to sweep the "used" argon (i.e., from the previous pulse) out of the diffuser/plenum section. As this blast wave reflects off the throat, peak overpressures may reach 1400 atm, forcing the gases out through a rotaryslit exhaust valve at the top of the primary mirror. At this point, the rotary "exhaust" valve closes, and this re-pressurized gas expands into an anechoic chamber designed to quickly damp out any residual acoustic noise caused by the motivating blast wave. Next, the gases pass through a low temperature "chiller" section (i.e., heat exchanger) which lowers the gas temperature to the range of 500-1000 K, and a pressure of perhaps 2 to 10 atm. At the appropriate moment in the engine cycle, the rotary "intake" valve opens, the flow chokes at the minimum section, and the XMHD generator channel is filled with a fresh load of argon at 1 atmosphere and 500 K (or perhaps

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HIGH TEMPERATURE "LIGHT BULB" RADIATOR

XMHD £ GENERATOR

\

\ \ Cm/

ACOUSTIC ABSORBER MATERIAL

COOLANT

ROTARY

VALVE

HIGH TEMPERATURE " "LIGHT BULB" RADIATOR

AUXILIARY LOW TEMPERATURE RADIATOR (500-1000°K)

Fig. 15 Argon gas recirculation system.

2 atm and 1000 K). Subsequently, the intake valve closes and a LSD wave is ignited at the far end of the XMHD generator channel. The wave then races across the channel and processes all the argon gas to 60 atm and 30,000 K in the next 75 microseconds. (Meanwhile the 700 atm blast wave is shed and pushes all the "spent" argon from the previous pulse out the rotary "exhaust" valve.) For the next 750 ys, this 60 atm driver gas expands into the plenum at supersonic velocities, suffering a large pressure drop through the 6-10cm electric discharge at the end of the XMHD generator channel. As mentioned above, the radiation emitted by the 30,000 K "driver gas" is transmitted directly out of the two actively-cooled refractory windows which serve as confining walls for the XMHD generator. (The other two orthogonal walls are the MHD generator electrodes, of course.) When these driver gases expand through the electric discharge into the diffuser/plenum section, they continue to radiate. This 0.5m3 cavity is also a one-sided "light bulb" radiator; the other major surface comprises the primary and secondary lenses. Here the gases radiatively cool to a temperature of perhaps 2000-3000 K, and await ejection by the next blast wave "piston".

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C.D. MAXWELL AND LN. MYRABO

Research and Technology Issues

A number of interesting research issues are associated with the 1) radiative production of nonequilibrium plasmas and 2) subsequent conversion to electric power. For example, to solve the problem of nonequilibrium absorption of laser energy, one must really use a computer code which solves the electron Boltzmann transport equation for the conditions of interest. This is because the electrons are not only out of equilibrium with the heavy particles, but will also have a non-Maxwellian velocity distribution. This solution would balance the monoenergetic electron excitation source (i.e., laser photons absorbed by inverse bremsstrahlung) and the various electron energy loss and energy redistribution mechanisms (i.e., electron neutral inelastic and electron-electron elastic collisions). One must first determine a region in which the electron temperature can be maintained significantly above that of the ions and neutrals, but where the energy absorbed by the electrons is equal to the energy transferred to the gas by the electrons. Next, one must show that the solution can be in steady state at the required choice of electron temperature and ionization level. The ultimate model would have to describe these processes in a good level of detail since the electron density, electron temperature, and energy absorbed are intimately related. As has been pointed out earlier, the success of the LSD-driven explosive MHD generator depends upon the creation of a relatively thick region of nonequilibrium, high conductivity working fluid. An adequate theory for the prediction of actual slug lengths, power densities, and other power generating characteristics (such as working fluid density/mass flow rate, incident laser flux, etc.) of a given device must couple solutions for strong interaction MHD flow, ionization/recombination kinetics, radiant energy transfer, and Ohm's law in a multicomponent, nonisothermal plasma (including ohmic dissipation, electron pressure and temperature gradients). Such solutions are available in existing computer models.32-37 However, it is doubtful that a reliable analytical expression could be developed without test and verification by computer solutions. Moreover, the existing computer solutions are capable of predicting the influence of a number of nonideal phenomena that can play an overriding role in determining MHD generator phenomena. Examples of such phenomena include

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plasma electrothermal stability, electrode voltage drops, effects of Hartmann layers, etc. The methods of Refs. 32-38 illustrate applications of such computer models to various related MHD problems over the past 15 years, albeit without the coupling of a solution for laser energy deposition in the plasma. The activation of a solution for the nonequilibrium inverse bremsstrahlung radiant energy absorption to the existing computer models in a onedimensional formulation would provide a first order analysis. Conclusions

The evolution of a concept for conversion of high power laser energy into electrical energy with a pulsed or explosive MHD generator has been described. The original concept for a component of a larger system has been refined to account for frozen losses and electron nonequilibrium effects with elementary calculations. Both open cycle, equilibrium systems with hydrogen working fluid and closedcycle, nonequilibrium systems with noble gas working fluids appear to be feasible. The noble gas system should have greater energy conversion efficiency due to reduced frozen losses. The addition of energy to the working fluid by absorption of laser energy introduces a novel aspect to previous explosive or pulsed MHD generator experiments: the possibility of adding energy to the working fluid simultaneously with the energy extraction process. The prospect exists for thereby overcoming some of the limiting processes (e.g., plasma instabilities) that govern MHD generator performance. Acknowledgement The results reported in this paper were supported by AFOSR/NA and DARPA/STO. The authors thank Dr. Manuel Martinez-Sanchez for his comments on ohmic heating.

References 1 Myrabo, L. N., "A Concept for Light-Powered Flight/' AIAA Paper No. 82-1214, June 1982. 2

Myrabo, L. N., "Solar-Powered Global Air Transportation/' AIAA Paper No. 78-689, April 1978.

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^Tate, E., Marston, C. H., and Zauderer, B., "Large Enthalpy Extraction Results in a Non-Equilibrium MHD Generator," Proceedings of the Sixth International Conference on MHD Electrical Power Generation, Washington, DC, June 1975, pp. 89-104. 4

Blom, J., Veefkind, A., Houben, J., and Rietjens, L, "High Power Density Experiments in the Eindhoven Shock Tunnel MHD Generator," Proceedings of the Sixth International Conference on MHD Electrical Power Generation, Washington, DC, June 1975, pp. 73-88. 5

Bangerter, C. D., et al., "Pulsed Magnetohydrodynamic Program," Final Technical Report, AFAPL-TR-76-34, July 1976.

^Bangerter, C. D., Hopkins, Generation A Progress International Conference on Generation, Washington, DC,

B. D., "Explosively Driven MHD Power Report/' Proceedings of the Sixth Magnetohydrodynamic Electrical Power June 9-13, 1975, p. 155-170.

7 Jones, M. S and McKinnon, C. N., "Explosive-Driven Linear MHD Generator," Conference on Megagauss Magnetic Field Generation by Explosives and Related Experiments," Frascati, Italy, September 1965.

^Jones, M. S., Bangerter, C. D., Peterson, A. N., and McKinnon, C. N., "Explosive Magnetohydrodynamics," AFAPL-TR-67-64, August 1967. 9

Kirillin/ V. A., Al'tov, V. A., Asinovskiy, E. I., Dremin, A. N. Dubovitskiy, F. I., Zenkevich, V. B., Kuznetsov, Yu. A., Lebedev, Ye. F., Savrov, S. D. and Sheindlin, A. Ye., "Explosively-Driven MHD Generator with a Superconducting Magnet System/' Doklady, Akademii Nauk SSR Vol. 185, No. 2, 1969, p. 316-319. 10 Asinovskiy, E. I., Lebedef, Ye. F., and Ostashev, V. Ye., "Investigation of Processes Determining the Efficiency of Energy Conversion in a Linear Explosive MHD Generator," 7th international Conference on MHD Electrical Power Generation, Cambridge, MA, June 16-20, 1980, Vol. I!, pp. 605-612. 11 Jones, M. S., Jr. and Blackman, V. H., "Parametric Studies of Explosive-Driven MHD Power Generators," Proceedings of the International Symposium on Magnetohydrodynamic Electrical Power Generation, Paris, July 1964, Paper 51. 12 Jones, M. S., Jr., Blackman, V. H., Brumfiefd, R. C., Evans, E. W., and McKinnon, C. N., "Research on the Physics of Pulsed MHD Generators," MHD Research, Inc., Report 646, Final Report under ARPA Contract Nonr-3859(00).

M. S., Jr., "The Use of an Explosively Driven MHD Generator for Laser Pumping," MHD Research, Inc., Report 672, a paper presented at The Second Laser Conference on Laser Technology, April 6-8, 1965.

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LASER-DRIVEN PULSED MHD GENERATORS

1 99

14 Jones, M. S., McKinnon, C. N., and Blackman, V. H., "Generation of Short-Duration Pulses in Linear MHD Generators/' 5th Symposium on Engineering Aspects of MHD, April 1964.

"^Asinovskiy, E. I., Kuznetsov, Yu. A., Lebedev, E. F., Maksimov, A. M.f and Ostashev, V. E., "Motion of a Plasma Driven by a NonConducting Piston in a Magnetic Field/' 6th International Conference on MHD Electrical Power Generation, June 1975, Washington, DC. ^Asinovskiy, E. I. and Ostashev, V. Ye., "Limiting Possibilities of a Pulsed MHD Power Generator to Generate Electrical Power into an Ohmic Load/' Teplofizika Vysokikh Temperature, Vol. 14, No. 5, 1976,

p. 1079-1082.

17 Bangerter, C. D., "Explosive Magnetohydrodynamic Program," Hercules Incorporated Systems Group, Technical Report AFAPLTR-73-16, May 1973.

"ISjeno, j. and Sonju, O.K., "Development of Explosively Driven MHD Generator for Short Pulse Aircraft High Power," Interim Technical Report No. 1, AVCO Research Laboratory, Inc., Technical Report AFAPL-TR-74-48, June 1974. 19 Baum, D. W., Gill, S. P., Shimmin, W. L, and Watson, J. D., "Dense Nonideal Plasma Research," Annual Report 130, Final report for Period 15 April 1979 to 14 March 1981 under US Navy Contract

No. N00014-78-C-0354, 30 April 1981.

H. T., and Smy, P. R., "Experiments on Power Generation from a Moving Plasma," Journal of Fluid Mechanics, Vol. 10, No. 1, 1961. 21

0stashev, V. Ye., Maksimov, A. M., Lebedev, Ye. F., Kuznetsov, Yu. A., and Davydon, A. N., "Non-stationary Interaction of Plasma Flow Behind a Strong Shockwave with a Magnetic Field," Teplofizika Vysokikh Temperature, Vol. 13, No. 1, 1975, p. 110-115. ^Bichenkov, E. I., "Explosive Generators," Soviet Physics - Doklady, Vol. 12, No. 6, December 1967, p. 567-569, translated from Doklady Akademii Nauk SSSR, Vol. 174, No. 4, p. 779-782, June 1967.

23|3urenin, YU. A. and Shvetsov, G. A., "Energetic Characteristics of Pulsed MHD Systems," Fizika Goreniya i Vsryva, Vol. 11, No. 3, May-June 1975, p. 433-437. 24

Brumfield, R. C., Evans, E. W., and McKinnon, C. N., "Pulsed MHD Power Generation/' 4th Symposium on Engineering Aspects of MHD, April 1963.

25Raizer, Y. P., "Heating of a Gas by a Powerful Light Pulse," Journal of Experimental and Theoretical Physics, Vol. 48, May 1965, p. 1508-1519; also, Soviet Physics JETP, Vol. 21, 1965, pp. 1009ff. 2

^Spitzer, L., Jr., Interscience Tracts Publishers, 1962.

"Physics of Fully Ionized Gases/' Number 3, on Physics and Astronomy, Interscience

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c.D. MAXWELL AND LN. MYRABO

2

^McWhirter, R. W. P., "Rates of Recombination Plasmas/' Nature, Vol. 190, p. 902, June 1961.

in

Hydrogenic

28

Bates, D. R., Kingston, A. E., and McWhirter, R. W. P., "Recombination Processes Between Electrons and Atomic Ions/' I. "Optically Thin Plasmas," Proceedings of the Royal Society of London, Ser. Vol. 267, p. 297, II. "Optically Thick Plasmas," Proc. Roy. Soc. A., Vol. 270, p. 155, 1962.

29park, C., "Calculation of Radiative Properties of Nonequilibrium Hydrogen Plasma," Journal of Quantitative Spectroscopy and Radiative Transfer, Vol. 22, pp. 101-112, 1979. ^Argyropoulos, Q. S. and Casteel, M. A., "Tables of Interaction Parameters for Computation of Ohm's Law Coefficients in Various Gases," J. Applied Physics., 4I, 10, pp. 4162-4165, September 1970. 31

Drellishak, K. S., Knopp, C. K., and Cambel, A. B., "Partition Functions and Thermodynamic Properties of Argon Plasma," Physics

of Fluids, Ser. 6, Vol. 9, p. 1280, 1963.

32Oliver, D. A., Swean, T. F., Bangerter, C. D., Maxwell, C. D., Demetriades, S. T., "A Computer Study of High Magnetic Reynolds Number MHD Channel Flow," STD Research Corp., Final Report under Contract No. N00014-77-C-0574, Work Unit NR 099-415, October 31, 1980.

33oiiver, D. A., Swean, T. F., Markham, D. M., and Demetriades, S. T., "Strong Interaction Magnetogasdynamics of Shock-Generated Plasma," AIAA Paper No. 80-0027, January 1980. •^Oliver, D. A., Swean, T. F., Markham, D. M., Maxwell, C. D., and Demetriades, S. T., "High Magnetic Reynolds Number and Strong Interaction Phenomena in MHD Channel Flows," Proceedings of the 7th International Conference on MHD Electrical Power Generation, June 16-20, 1980 Vol. II, pp. 565-572.

^Argyropoulos, G. S., Demetriades, S. T., and Kendig, A. P., "Current Distribution in Nonequilibrium J X B Devices," Journal of Applied Physics, Vol. 38, No. 13, December 1967. •^Lackner, K., Argyropoulos, G. S., and Demetriades, S. T., "Relaxation Effects in J X B Devices," AIAA Journal, Vol. 6, No. 5, May 1968, pp. 949-951. ^^Demetriades, S. T., Oliver, D. A., Swean, T. F., and Maxwell, C. D., "On the Magnetoaerothermal Instability," Paper No. 81-0248, January 1981. S. T., Argyropoulos, G. S., and Maxwell, C. D., "Progress in Analytical Modeling of MHD Power Generators/' Proceedings of the 12th Symposium on Engineering Aspects of Magnetohydrodynamics, Argonne National Laboratory, III., March 1972.

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A Comparison of Electric Propulsion Technologies for Orbit Transfer Robert L. Poeschel* and Jay Hyman + Hughes Research Laboratories Malibu, Calif. Abstract Electric propulsion technologies offer an opportunity for increasing the mass of spacecraft payioads by providing the propulsion requirements with significantly less propellent than chemical propulsion technologies. Propellant exhaust velocity (or specific impulse) can be increased by factors ranging from 2 to 30 using electric propulsion technologies thereby enabling a significant reduction in the mass of the propulsion subsystem that is required to perform a given mission. For a given spacecraft mass, the realization of this mass benefit depends on thrust system efficiency and thrusting strategy (orbit transfer time). This paper outlines the basic relationships that determine the tractability for achieving appreciable mass benefits in orbit transfer missions and then compares and discusses the capabilities and challenges for several types of electric propulsion technology that are available for meeting these requirements. We conclude that electric propulsion can increase payload transfer from LEO to GEO by a factor of 2 as compared with chemical propulsion if long transfer times (>60 days) are acceptable. If the payload is itself a power source, this benefit can be achieved with significantly shorter transfer times (10 to 20 days). Invited paper received June 23, 1982. Revision received A p r i l 25, 1983- Copyright © 1984 by R. L. Poeschel. P u b l i s h e d by the American I n s t i t u t e of Aeronautics and Astronautics, Inc., w i t h permission. *Asst. Manager, Plasma Physics Department. +Manager, Plasma Physics Department.

203

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R.L POESCHEL AND J. HYMAN

Nomenclature F go lsp Mo Mp m

= = = = = = = =

Pb Pj n t

= = =

v Av

= =

a

=

a

=

pp ps ay

= =

= =

~ = =

thrust produced by thrust subsystem gravitational constant specific impulse of thrust subsystem initial mass of the spacecraft mass of the propellant for achieving mission mass of the payload mass of the thrust subsystem mass flow rate of propellant expelled during thrusting power in exhaust beam of thruster(s) power supplied to the thrust subsystem time required to perform a propulsion maneuver (such as orbit transfer) average velocity of the exhaust propellant atoms increment in velocity associated with an orbit maneuver (such as orbit transfer) specific mass of the power processing hardware specific mass of the power source specific mass of the thruster and mounting structure specific mass of the thermal control hardware required to remove waste heat from any part of the thrust subsystem specific mass of the thrust subsystem specific energy parameter for the thrust subsystem efficiency of the power processing electronics efficiency of the thruster(s) efficiency of the thrust subsystem Introduction

Electric propulsion technologies can be divided into three generic types corresponding to the acceleration process employed to expel the propellent. These acceleration processes can be categorized as electrostatic, electromagnetic, and electrothermal. Several detailed analyses have been performed^~3 t h at project electric propulsion system capabilities for specific mission scenarios and which compare electric-thruster capabilities for the same technologies addressed in this paper. Our purpose here is to obtain a simplified analysis to illustrate the direct relationship between the thrust system efficiency, the thrust system

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ELECTRIC PROPULSION TECHNOLOGIES

205

specific mass, and the thrusting time that is required to achieve an overall performance objective. Although the performance projections obtained in this manner are not expected to be as accurate as those using more exact analyses, the implications on the requirements for thrust system technologies are not expected to be altered. In this paper we are concerned, primarily, with comparing the capability of the electric propulsion technologies now under consideration for meeting these implied system requirements.

Fundamental Relationships and Characteristics

Electric thrusters can be thought of as transducers that convert energy from an electric power source into directed kinetic energy of the propellant on an atomic basis. Examination of the fundamental relationships among the power required, the thrust produced, and the subsequent spacecraft acceleration provides an insight into the propulsion-system requirements and mission constraints to be expected in utilization of electric propulsion. For comparing electric thruster technologies, we present a brief discussion of the operating principles and development status of each generic type of electric thruster. In each case, the technology limitations are identified and prospects are assessed for achieving performance approaching the fundamental limits. Whereas component and total system efficiencies comprise the primary performance limitation, the nature of the limitation is different for each generic technology, as is the nature of the challenge to approach the limit. In electric thrusters, as in chemical rockets, the thrust or force F, produced by the device is related to the exhaust propellant parameters by the expression

F = m v

(1)

where m is the propellant flow and v is the average velocity of the propellant particles. The rate at which kinetic energy leaves the vehicle with the propellant exhaust is written as

2

1 P

b

=

Z

*

V

(2)

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206

R.L POESCHEL AND J. HYMAN

For electric thrusters, the efficiency of the thrust system, rijs/ is defined as the ratio of P^ to the electric power input, P-m, supplied to the thrust system by the power source

_ pb = ~

n

is

in

which can be written

nTS

-2

=

—,——

TS

The specific defined as

impulse,

2m P

lsp, of

(4)

in

a rocket type

thruster

is

(5)

where go is the gravitation constant (9.8 m/s 2 ). can be re-written as

,

F 9

0 'sp

Equation (3)

(6)

2 P

or

in

" g—r o sp y

Equation (7} illustrates a fundamental relationship between thrust and electrical power for electric propulsion systems. It should be noted that the inefficiencies in the thrust system associated with power matching and distribution are combined in r i y , so that

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ELECTRIC PROPULSION TECHNOLOGIES

207

TITS = ripp n T

{8)

where ripp is the power processor efficiency and nj is the thruster efficiency. In order to examine the mass benefit of electric propulsion for orbit transfer maneuvers in a more-or-less general manner, a simplified set of equations has been formulated to show first order effects. The initial mass of the vehicle Mo is defined as follows:

Mp where Mp[_ is the payload mass, Mp is the propellant mass, and Mys is the thrust system mass, including the power source. The variable of interest here is the ratio Mp[_/M o , the payload mass fraction, for an orbit maneuver that is characterized by a velocity increment Av, that has been derived using rigorous orbital mechanics. For simplicity, we assume that

Av

~ Ft/M

o

where F is the thrust produced and t is the orbit transfer time. Equation (10) is conservative in that the vehicle mass may be appreciably less than M o at the end of the thrusting period. On the other hand, the effects of atmospheric drag, solar pressure, etc., have been totally neglected. Having specified a value of Av, we calculate Mp using the familiar rocket equation

M

o

and Mjg using MT,

IS

5

QTCP.

TS in

02)

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208

R.L POESCHEL AND J. HYMAN

where 075 is the specific mass of the thrust system expressed in kg/kW and Pj n is determined by combining Eqs. (7) and (10) to o b t a i n

P

(13)

in

TS

Equation (13) can be exploited to illustrate the magnitude of power required for orbit transfer of a vehicle from low Earth orbit (LEO) to geosynchronous Earth orbit (GEO) using electric propulsion. Let us assume Mo = 35,000 kg (shuttle capacity projected for 1990-2000 time frame) and set the velocity increment to a value Av = 6000 m/s (the value required for a LEO to GEO transfer); the transfer time can then be taken as a parameter. For a perfect (HTS = 100%) electric propulsion system, Fig. 1 shows the input power required as a function of specific impulse. It is noteworthy that hundreds of kilowatts of electric power are required to transfer a full Shuttle load from LEO to GEO, independent of the electric propulsion technology used. Furthermore, the shorter the transfer time, the greater the power that is required. In^ fact, to achieve an appreciable advantage in specific impulse over chemical propulsion, megawatts of electric power are required. The payload mass fraction for a spacecraft using electric propulsion is derived by combining Eqs. (9), (11), (12), and (13) to obtain M

PL = exp M o

- Av 9 I ^o sp

- 5 x 10 ~^

a

n TC TS

TS

V g

t

I P

(14)

where the units of the velocity increment Av are m/s, the specific mass of the thrust system 0,73 is in kg/kW, the specific impulse l s p is in sec, and the time t is in sec. It should be noted that OLJ$ is the sum total of all thrust-system hardware including the power source or

(15) Q

TS '-

a

PS

+ G

PP

+ a

T

+ Q

Th

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209

ELECTRIC PROPULSION TECHNOLOGIES

3 -

cc

LU £

O Q_ h-

2 -

Q_

D

O

1 —

1000

2000

3000

4000

5000

6000

7000

SPECIFIC IMPULSE, ISR, s Fig. I Input power required for a perfect (Tlf$ = 1} electrically propelled spacecraft to raise 35,000 kg from LEO to GEO.

where ap5 is the specific mass of the power source, a pp is the specific mass of the power processor, ay is the specific mass of the thruster and associated structure, and aj^ is the specific mass of any thermal control hardware that is necessary to remove heat generated by inefficiencies of the electric propulsion hardware. We can simplify Eq. (14) by expressing t in hours and defining a specific energy parameter 875 where

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210

R.L POESCHELAND J. HYMAN Q C

TS

'-

r

TS

'TS

t

kg ' kW-hr

(16)

Substituting Eq. (16) into Eq. (14) we can write an expression for the payload fraction as a straightforward function of mission input parameters. Mp,

1.39 x 10

- exp

-7

:

TS 9C

I

(17) Av

Let us now calculate the values of payload fraction which can be achieved using electric propulsion under a variety of different mission constraints. The most optimistic case occurs when the payload is itself a power source; this is represented by setting Mp|_ = 0 (see Eq. (9)). For an idea! thrust subsystem with rijg = 1 and by setting QJS = 5 kg/kW (an optimistic value), Eq. (17) can be used to determine the orbit-transfer time as a function of the electric propulsion system lsp. Figure 2 shows that the minimum transfer time of about 3 days occurs at a minimum value of specific impulse. However, Eqs. (12) and (13) = 5 kg/kW 12

T

C/)




e— «^

^

5

^

"O

^-

^

"c D c? — ' ^ c:

>

to

*CL

1

!

13

"C

"C

D

if

C! c

£ Q

rc

-^-

i

0

~

fC

c c

J

00

C r E f C

^- c c c c c

.r *-

Z_

fb ^

r- CL c:

^I

CJrc

^ c C o

SI

^-v

O

*-

!Z

P

r- c

(U

| 11 ; III 1

h-

c

OJ

3 = c •-

2

>

C O

M- JZ

O

) e T e > 1), the arc impedance is increased.

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VOLTAGE CHARACTERISTICS OF MPD DEVICES

245

WATER COOLED ANODE - WATER COOLED CATHODE

B C A T M O D E APEX = 1 2 5 0 GAUSS AT IB = 100 AMPS

Fig. I Cross section view of the Magnetic Annular Arc showing the ionization region and magnetic field configuration.

This increased plasma impedance reduces the effect of electrode voltage drops and leads to a more efficient arc operation. (2) The discharge is stable when u) e T e > 1 and the arc is uniformly distributed throughout the channel. (3) Momentum can be added to the plasma by j x B forces in both azimutha! and axial directions. The azimuthally directed energy can be recovered in an expansion nozzle in th form of directed axial velocity. That is, as the rotating gas expands in the nozzle, the rotational velocity decreases (to conserve angular momentum) and this energy is transformed into an increased axial—i.e., thrust-velocity. (4) The magnetic field is aligned essentially parallel or at small angles to the channel walls, and since the oo e i e for the electrons is large compared to unity, the azimuthal (Hall) currents are the largest component of the current in the nozzle. The j x B forces are directed away from the channel walls. Therefore, both acceleration and containment is achieved in this device with plasma conditions corresponding to a relatively low degree of ionization and low electrical conductivity.

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246

R.M. PATRICK AND G.S. JANES

The total thrust developed by several variation of MAARC accelerator B field geometries was measured and reported in Ref. 3. This was done by measuring the local dynamic pressure with a cooled thrust plate and integrating these local measurements to obtain the total thrust. Further, the total mass flow and power input were measured. From these measurements one can compute the specific impulse, ISp,

Isp = F/mg

(1)

and the overall thrust efficiency

(2)

Varan

Some of the data obtained with MAARC accelerators designed for space plasma propulsion applications is shown in Fig. 2 from Ref. 3. The power level varied from 10 to T O O kW and several propellents were studied. It can be seen that interesting thrust efficiencies were obtained with this type of device, especially using hydrogen as a propellant. Further the thermal efficiency was measured and reported in Ref. 5. This was done by measuring the total

H 2 W I T H MEDIUM SIZED COILS

* WATER COOLED • HEAT SINK

2000

3000

4000

8000

Up (SEC)

Fig. 2 Overall thrust efficiency plotted vs the measured specific impulse, l sp . (Open symbols: data taken with a relatively large radius bias field coil. Solid symbols: hydrogen propellant at 0.05 g/s.)

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VOLTAGE CHARACTERISTICS OF MPD DEVICES

power input, IV, and the anode and cathode heating.

247

The

cathode heating was also measured and was always less than 10% of the anode heat transfer; so one may write for the thermal efficiency,

n

n

_ IV - Q anode

th ~

IV

(3)

where Qanode ' s the an°de heat transfer. The quantity Q-anode ' s closely related to the product of the total current, I, times the voltage drop associated with the current passage into the anode. This leads to an investigation of the voltage characteristics of MAARC devices described in the next section. Voltage as a Function of Magnetic Field Intensity in Magnetic Annular Arc Devices The most clearly defined scaling seen in the data on MAARC's was the parametric variation of the voltage of the arc. For a given B-field, electrode configuration, and feed gas, the arc voltage was essentially a constant over a wide range of power. That is, as the power was increased, the current increased, but the voltage was practically unchanged. But the voltage did depend upon the B-field and the type of gas. For a given gas, the voltage varied linearly with magnetic field, a scaling which has been observed in several studies of similar devices designed for different applications.^'^ The operating voltage of a magnetic annular arc (MAARC) is shown in Fig. 3. It is apparent that for several propellants the voltage is linearly dependent on the applied

magnetic bias field strength and the slope of the dependence is a function of the propellent. The arc voltage can be

described as V = V o + U C B£ where V 0 is a voltage drop associated with electrodes, B is the applied field strength, £

is nearly the cathode-anode spacing, and U c is a "cntical velocity'' associated with the propellant. The values of Uc for several propellants are given in Table 1. The value of U c can be characterized by

2

c

dissociation

ionization

^4^

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248

R.M. PATRICK AND G.S. JANES

60

80

100

B I A S M A G N E T I C F I E L D COIL C U R R E N T

Fig. 3 Magnetic annular arc voltage as a function of applied bias field coil current; 100 A corresponds to 1250 G at the cathode tip. Table 1 Critical speeds and electrode voltages Propellant

Uc

Vo

H

Hydrogen Nitrogen Argon Ammonia

55 16 8.7 26

45 30 15 36

1 1 1 1

where m is the ion mass and E d i s s o c j a t j o n and E i o n i z a t i o n are the energies necessary to dissociate the molecules and produce a single ionization. A theoretical description of U c has not been developed to date. The operating voltage is important in determining the efficiency of these devices used for space propulsion. This is true for several reasons. First the ratio between the electrode drop to the arc voltage V 0 /U C B£ should be minimized to produce small electrode losses for high efficiency. Second, the location of the arc currents in the nozzle region which produce accelerating body forces on the propellant is a function of the operating voltage. Third, the scaling of these devices as a function of mass flow, power level, and

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VOLTAGE CHARACTERISTICS OF MPD DEVICES

249

propellant choice must be carried out with understanding to enable one to achieve high efficiency, long life operation for space propulsion applications. A number of models have been proposed to explain the critical velocity. Some of the analyses which were prompted by astrophysical studies, and by colliding plasma experiments are reviewed in Ref, 11. Although some of the suggested explanations may apply to some of those situations, they either require that that gas be fully ionized, which is not the case in a MAARC, or else propose that the similarity between the measured E-ftetd and U C B was somewhat of a coincidence, which is quite unlikely to account for the scaling over orders of magnitude which was seen in the MAARC. Explanations have also been prompted by the MAARC results. Unfortunately, however, these analyses have generally not taken proper account of the collisional coupling between ions and neutrals. This has led to considerable confusion, a situation which has been exacerbated by reports of confusing and even contradictory experimental results. Observations of MPD thrusters operating with no net mass flow throughput have been reported.^ A model based on the concept that the ionized portion of the propellant is decoupled from the nonionized portion was developed to explain these results. 13,14 jhis model was made somewhat credible because the crude theory 1 5 for collisionfree discharges does predict the critical voltage for this type of coaxial discharge, i.e., radial currents in an axial B-field. The model^ assumes that the ionized portion of the propellant passes through the accelerator and does not share its momentum with the rest of the propellant. A critical mass flow is defined as that mass flow for which each propellant molecule is singly ionized and the total power delivered to the accelerator corresponds to this molecular flow times the ionization and dissociation energy per particle plus an equal thrust power. This would imply that the maximum thrust efficiency is 50%, i.e., that the input power is divided equally between thrust and ionization plus dissociation. A mass flow larger than this value corresponds to lower thrust efficiency and lower mass flows might correspond to higher thrust efficiencies, i.e., for low mass flows there is excess power to ionize and accelerate in unequal amounts. The two possible modes of MPD arc jet operation are m > m^ . or m < m^ . where m is the mass flow and mnu . is the^mass flow defined above. The model mm

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250

P.M. PATRICK AND G.S. JANES

assumes that when m > m^ . the excess flow is bypassed mm by the discharge. This cannot be the case for a large class of experiments reported for many MPD arc jets. In particular, the MAARC accelerator data described above are such that the mean free path for momentum sharing collisions is a very small fraction of the electrode spacing and a much smaller fraction of the measured axial length of the current distribution. As another example, Connolly, et al.J^ report spectroscopic measurements of exhaust velocity which show that, even at lower densities, there is acceleration of the neutral gas. Therefore a model must be developed which explains the observed voltage dependence for a collision-dominated accelerator as well as a collision-free device and takes into account the fact that nearly all of the plasma propellant is accelerated uniformly. Description of the Low Density Hall Accelerator

We now turn to low density Hall accelerators, a class of steady-state magnetic plasma accelerators which are geometrically akin to the magnetic annular arc but operate fully ionized at lower densities, magnetic fields, and power levels. Typical operation parameters are given in Table 2. In this device, the electric field is applied in the direction of plasma acceleration. Figure 4 illustrates the D.C. plasma accelerator studied at AERl.^7 Similar studies have been made by Seikel and ReshotkoJ 8 by Lary, Meyerand, and Salz, 1 9 and by Hess, et al. 20 Table 2 Typical experimental conditions

Gas Axial electric field, Ex Mean radial magnetic field, BR Electron density, n e Electron temperature, Te Final ion velocity Acceleration length Annulus radii Ion gyro radius Ion mean free path Electron gyro radius Debye length

Argon 20 V/cm 500 G lO^-IO 1 Vcm"3 10-30 eV 2-4x10^cm/s 13 cm 4 cm and 10 cm 30 cm > 30 cm 0.02 cm 0.01 cm

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VOLTAGE CHARACTERISTICS OF MPD DEVICES

251

Fig. 4 Schematic diagram illustrating the basic geometry of the E.M. region D.C. plasma acceleration experiment.

This is an E.M. region accelerator in which both the ion cyclotron radius and ion mean free path are large relative to the apparatus dimensions. (In the MAARC, where the ion gyroradius is smaller than the arc length, the ions tend to drift at the E/B velocity, the velocity per-

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252

P.M. PATRICK AND G.S. JANES

pendicular to E and B which is characteristic of any plasma, and arises because the gyrating particles move faster on the "downhill" side of each magnetic orbit. In the Hall accelerator, the ion gyroradius is so large that the ions cannot make a gyro-orbit within the machine and hence do

not attain the E/B speed.) Thus, the ions are freely accelerated in the axial direction by the electric field, since they

are relatively unaffected by either collisions or magnetic fields. The hot cathode at the exit serves to release electrons which both neutralize the outgoing plasma and maintain the axial voltage gradient. The electrons, which are many times lighter than the ions, have cyclotron radii which are small relative to the apparatus size. Thus, ideally, the electrons will be tightly coupled to the magnetic field lines. They will drift circumferentially at right angles to both the electric and magnetic fields at a velocity such that E + (v x B) = 0. This circumferential electron drift constitutes an electric current, J Q , which interacts with the radial magnetic field and thereby satisfies the requirements for overall momentum conservation. Detailed local time dependent measurements of ion flux momentum flow, electron temperature, and plasma potential verified most of this model and demonstrated the device's promise for propulsion. The hot probe plasma potential measurements plotted in Fig. 4 indicate that the magnetic field lines were indeed equipotentials. This verified the utility of the applied magnetic field in allowing us to maintain an accelerating electric field across the plasma. Figure 5 shows the axial ion mass flux, pv, for a typical set of experimental conditions. The uniform magnetic field extends from the anode at x = 23 to 10 cm. The decrease in ion flux beyond x = 10 cm is attributed to expansion. Langmuir probe studies indicated electron temperature of the order of 10 eV in the accelerating region, 3 or 4 eV downstream of the magnetic field, and as high as 20 to 30 eV in the ionizing region near the anode. Note that in Fig. 5 the ion flux increases to nearly its peak vaiue in the first few centimeters. This distance is comparable with the mean free path for ionization of an argon atom moving at room-temperature velocity in the presence of 25 eV electrons at a density of 4 x 10 10 cm"3. The ion is unlikely to become doubly ionized since the acceleration time (about 7 ys) is more than an order of magnitude shorter than the time corresponding to the ionization mean free path.

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VOLTAGE CHARACTERISTICS OF MPD DEVICES

253

IB = 200 AMPS

m

= .002

I0 = S A M P S I = 10 AMPS V = 400 VOLTS MAY 1963

X CM———

Fig. 5 Plot illustrating axial variation in electron density/ momentum flux, ion flux, and mean ion velocity for the annular E.M. region D.C. experiment. Momentum flux and ion flux were measured directly and utilized to compute the other two quantities.

The distribution of momentum flux, pv^, shown in Fig. 5 was obtained from measurements of the force on a small quartz plate. The data on ion flux and on momentum flux were combined to obtain the estimates of mean ion velocity and electron density that are shown in Fig. 5. (The pv 2 results assume that because of the long mean free path for ion-neutral collisions, the contribution of neutrals to the momentum flux measurements is negligible.) It is particularly important to observe that the ion velocity increases continuously throughout the accelerator up to a value consistent with the measurements of the integrated axial electrostatic field gradients. Furthermore, inductive measurements of the d e c r e a s e in magnetic flux were in rough agreement with the foregoing probe determination of electron density and electric fields. Measurements of the total integrated ion flux shown in Fig. 6 indicate that this quantity saturates at a value l max which corresponds to 100% single ionization of the total

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254

P.M. PATRICK AND G.S. JANES 10.0

7L

O

2.0

8

TOTAL

12

PLASMA

16

20

24

CURRENT, amps

Fig. 6 Total ion current Ij vs total current lt for three different mass flow rates. The measurements were made in a plane 7 cm below the anode. The total electron current, le, is It - Ij. m is the mass flow rate, lo is the ion current equivalent to m, lmax is the maximum ion current.

plus some reionization contribution from mass flux backstreaming of recombined neutrals. The desirable operating point is at the knee of the current where essentially all of the propellant has been accelerated and the total current is approximately twice the ion current. This suggests that the electric efficiency (see Fig. 5) is about 50%. At this operating point, the overall voltage is invariant to mass flow rate but increases with increasing B-field. Spoke and lonization Phenomena in the Low Density Hall Accelerator This device is also dominated by ionization phenomena; however, in the absence of collisions, these

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255

VOLTAGE CHARACTERISTICS OF MPD DEVICES

effects are somewhat different. Unlike the higher density MAARC device, the Hail accelerator exhibited a tilted rotating spoke phenomenon consisting of a region of increased ion density which originated near the anode and propagated axially with the ion particle velocity. Figure 7 illustrates a remarkable correlation between the previously referenced "critical ionization" velocity and the rotational velocity of this spoke which is highly suggestive of a strong but as yet unexplained coupling between spoke formation and ionization phenomenon. This hypothesis is further strengthened by time correlated current observations from a small isolated section of the anode indicating that most of the ionization actually occurs within this spoke which 1.0

0.9

0.8

: 0.7

0.6

5

0.5

o CRIT. XENON

0.4

0.3

0.2

O.I

L

10 RADIUS , CM. AXIS

Fig. 7 Comparison of spoke velocity vs radius with the "critical" velocity, vcr|t, for three different gases (V c r j t = 2 U|e/mj, where U| is the first ionization potential of the gas and mj is the ion mass).

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P.M. PATRICK AND G.S. JANES

completes a full revolution in the time required for a cold neutral atom to travel about a centimeter. Subsequent studies indicated that this spoke phenomenon controls the electronic component of the axial current and thus has a profound influence on the basic electrical efficiency of the device. The observed axial electron currents which cross magnetic field lines (see Fig. 6) are much too large to be explained by collisional diffusion theory. This theory predicts an axial electron current, j ez , given

The data (see Fig. 6) suggest that oo e i e has an effective value, (u)eie)eff, of about 3 which is at least two orders of magnitude smaller than expected. This phenomenon has also been observed in other devices but in the present device it is predicted by crossed field, electron drift theory in the presence of phase correlated nonuniformities in both the electron density and circumferential electric fields. The probe studies displayed in Fig. 8 illustrate a significant phase correlation between E o and n e which allows the electrons to drift toward the anode within the spoke where the electron density is higher. A quantitative examination of this effect predicts approximately the correct net average electron current, but there are several gaps in our understanding which preclude the development of scaling relations. One of these centers around the question of why the spoke has the character and "critical velocity" which just allows for the requisite ionization and for u)eie 3-10. A second related gap centers around the question of how the electron drift energy, originally invested in low frequency ion acoustic waves, evolves into an elevated electron temperature which is just adequate to ionize all the incoming atoms.

Summary The recent development of analytical techniques which predict both the plasma chemical kinetics of such discharges and the plasma dynamic stability of the resulting plasma flow has now made it possible to model such partially ionized, magnetized discharges. The authors

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257

VOLTAGE CHARACTERISTICS OF MPD DEVICES

P O T E N T I A L DIFFERENCE*!

(-20° SEPARATION)/ °

ION DENSITY PROBE

ION DENSITY

POTENTIAL PROBES

POTENTIAL DIFFERENCE \ (0° SEPARATION)/

ION DENSITY

POTENTIAL DIFFERENCE \ (+20° SEPARATION)/

ION DENSITY

Fig. 8 Measurements illustrating correlation between the fluctuating azimuthal electric fields and the density fluctuations. The azimuthal electric field is determined from the difference signal between two

heated potential probes arranged with azimuthal separation as shown in the sketches. The density measurements were obtained with a negatively biased Langmuir probe located as shown in the sketches.

believe that it should be possible to explain the observed behavior by taking into account the velocity distribution of the electrons and various rate constants for ionization, excitation, charge exchange, and j x B forces and that sufficient experimental evidence exists which can be used to check the analytical results. Two major unexplained phenomena have been briefly described. The successful modeling of the "critical voltage" controlled by ionization processes in D.C. magnetic arcs and the anomalous electron diffusion in low density accelerators would allow the design of more efficient electric propulsion devices for space flight orbit changing applications. It is suggested that an effort should be made to better understand these two basic phenomena.

References ''Boyle, M. J., Clark, K. E., and Jahn, R. G., "Flowfield Characteristics and Performance Limitations of Quasi-Steady Magnetoplasmadynamic Accelerators," AIAA Journal Vol. 14, July 1976, pp. 955-962. A. C.,

John,

R. R.,

Garrison,

R. L,

and

"Performance of Quasi-Steady MPD Thrusters at AIAA Journal, Vol. 10, February 1972, pp. 121-122.

Libby,

High

D. R.,

Powers,"

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258 ^Patrick,

R.M. PATRICK AND G.S. JANES R. M.

and

Schneiderman,

A. M.,

"Performance

Characteristics of a Magnetic Annular Arc," AIAA Journal, Vol. 4, February 1966, pp. 283-290.

4pugh, E. R., Patrick, R. M., and Schneiderman, A. M., "HighPressure High-Enthalpy Test Facility, " AiAA Journal. Vol. 9, February 1, pp. 200-204. ^Schneiderman, A. M. and Patrick, Thermal Efficiency of the Magnetic Vol. 4, October 1966, pp. 1836-1838.

R. M., "Optimization of the Annular Arc," AIAA Journal,

6

Janes, G. S. and Lowder, R. S., "Anomalous Electron Diffusion and Ion Acceleration in a Low-Density Plasma," Physics of Fluids, Vol. 9, June 1966, pp. 1115-1123. ^Alfven, H. On the Origin of the Solar System, Greenwood Press, Westport, Conn., 1954.

^Alfven, H., "Collision Between a Nonionized Gas and a Magnetized Plasma," Reviews of Modern Physics, Vol. 32, October 1960, pp. 710-713. 9

Danielsson, L. "Review of the Critical Velocity of Gas-Plasma Interactions,, I. Experimental Observations," Astrophysics and Space Science, Vol 24, 1973, pp. 459.

10 Anderson, O., Baker, W. R., Bratenahl, A., Furth, H. P., and Kunkel, W. B., "Hydromagnetic Capacitor," Journal of Applied Physics,

Vol. 30, February 1959, pp. 188-196.

^Sherman, J. C., "Review of the Critical Velocity of Gas-Plasma Interactions, II. Theory," Astrophysics and Space Science, Vol. 24, 1973, pp. 487. R. R. and Bennett, S., "Recent Advances in Electrothermal and Hybrid Electrothermal-Electromagnetic Propulsion," AFOSR Fourth Symposium on Advanced Propulsion Concepts, Palo Alto, Calif., April 1965. ^Bennett, S., John, R. R., Enos, G., and Tuchman, A., "Experimental Investigation of MPD Arc Jet," AIAA Fifth Electric Propulsion Conference, San Diego, Calif., March 1966. 14 Moore, R. A., Cann, G. L., and Gallagher, L. R., "High Specific Impulse Thermal Arc Jet Thruster Technology," Technical Report, AFAPL-TR-65-48, Part I, June 1965.

^Petschek, H. E., comment in the discussion following Ref. 8. ^Connolfy, D. J., Sovfe, R. J., and Seikel, G. R., "Performance and Diagnostics of a Water-Cooled Magnetoplasmadynamic Arc Thruster," NASA Technical Note, NASA TN D-5836, 1970.

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VOLTAGE CHARACTERISTICS OF MPD DEVICES

259

17 Janes, G. S. and Dotson, J., "Experimental Studies of Oscillations and Accompanying Anomalous Electron Diffusion Occurring in D.C. Low Density Hall Type Crossed Field Plasma Accelerators," Proceedings of Fifth Symposium of the Engineering Aspects of Magnetohydrodynamics, MIT, Cambridge, Mass., April 1-2, 1964,

18 Seikel, G. R. and Reshotko, E., "Hall Current Ion Accelerator," Bulletin of the American Physical Society, Series II, Vol. 7, p. 414, June 1962.

E. C., Meyerand, Jr., R. G., and Salz, F., "Fluctuations in a Gyro-Dominated Plasma," Sixth International Conference on lonization Phenomena in Gases, Paris, Session Vc, Paper 14, July 8-13, 1963. 20

Burlock, J., Brockman, P., Hess, R. V., and Sidney, B. D., "Hall Currents and Oscillations for Steady Low Pressure Discharges Crossed with Magnetic Fields. I. Experiments — II. Theory," Presented at the Fifth Annual Meeting of the American Physical Society, Division of Plasma Physics, November 6-9, 1963.

Purchased from American Institute of Aeronautics and Astronautics

Applied-Field Magnetoplasmadynamic Thrusters for Orbit-Raising Missions George R. Seikel* SeiTec, Inc., Cleveland, Ohio Thomas M. York +

Pennsylvania State University, University Park, Pa. and William C. Condit* Westinghouse R&D Center, Pittsburgh, Pa. Abstract A review of the results of previous NASA major magnetoplasmadynamic (MPD) arc thruster experiments is presented. MPD performance is shown to improve with both increasing power level and with increasing strength of an applied axial magnetic field. The discussion points out how further studies utilizing advanced diagnostics should significantly improve the physical understanding of MPD arc thrusters with external fields. Rejection of anode heat losses limits the average power capability of an MPD arc thruster. Operation of a radiation-cooled thruster utilizing a IT superconducting magnet achieved 34% efficiency at a specific impulse of 2500 s at a power level of 25 kW. The higher performance potential from much higher power operation with stronger magnetic fields may only be achievable with practical radiation-cooled thrusters if operation is quasi-steady. The resulting potential performance improvements coupled with state-of-the-art power storage technology could make pulsed quasi-steady MPD arc thrusters Invited paper received March -20, 1982; revision received October 25, 1982. Copyright © by American Institute of Aeronautics and Astronautics, Inc., 1983- All rights reserved. (Limited copyright.) "President. +Professor of Aerospace Engineering jfSenior Scientist.

260

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APPLIED-FIELD MAGNETOPLASMADYNAMIC THRUSTERS

extremely missions.

attractive

candidates

for

future

261

orbit-raising

introduction An important space mission will be the transfer of large payloads from low Earth orbit (LEO) to geosynchronous orbit (GEO), where the payload has for its main mission either considerable onboard power and/or the power conditioning equipment to provide large high power pulses. Numerous examinations^"" 4 over the past two decades of the optimum specific impulse for such missions have shown that electromagnetic thrusters are very attractive if high performance can be obtained with a compact, lightweight thruster system (thruster plus necessary power conditioning, etc.). From the late fifties to the mid-sixties, a large variety of electromagnetic thrusters were investigated.5~7 By the mid-sixties, the majority of the technology effort was focused on the applied-field MPD arc thruster which had emerged as the most attractive contender.^ By the early seventies, reduction in the advanced space technology budget prevented continuation of MPD thruster engineering technology contracts and large facility testing at NASA Lewis Research Center was terminated. Research in the United States was conducted on applied-field MPD arc thrusters in a number of laboratories including teams led by the first author and R. V. Hess at NASA Lewis and Langley Research Centers, respectively; R. M. Patrick and R. R. John at AVCO Everett and Wilmington, respectively; A. C. Ducati at Giannini; G. L. Cann at Electro-Optical Systems; D. W. Esker at McDonnell Douglas Research Laboratories; and T. F. Stratton at Los Alamos Scientific Laboratory. Most of the important contributions of these teams are well reviewed, 5 ~8 Qf particular significance was the demonstration that very low (less than 10~4 torr) facility pressure was required to obtain reliable performance measurements. Only performance testing of condensab'e propellants, such as alkaline metals, was shown to be possible in small test facilities. Alkaline metal MPD arcs showed a potential for excellent performance, because of the low energy required to ionize the propellent Alkaline metal propellant, however, poses a potential environmental hazard to a spacecraft because of the possibility of propellant condensation on the spacecraft. Reliable performance testing of steady-state noncondens-

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262

G.R. SEIKEL, T.M. YORK, AND W.C. CONDIT

able propeliant MPD arcs was shown to require the large NASA Lewis Research Center facilities. Results of these previous noncondensable propeliant applied-field MPD arc efforts are reviewed, impact of recent developments in energy storage technology is discussed, and physics issues directly bearing on MPD thruster performance and operating life which need additional investigation are presented.

MPD Acceleration Mechanisms

Figure 1 illustrates the current and magnetic field distribution in a MPD arc thruster. The applied arc current between the anode and cathode flows radially with some pluming into the exhaust. The applied magnetic field diverges to form what can be termed a magnetic nozzle. In MPD A R C T H R U S T M E C H A N I S M S 1. P R E S S U R E , p ( N / m 2 ) , ON P H Y S I C A L N O Z Z L E 2. M A G N E T I C BLOWING, JrB 3. M A G N E T I C N O Z Z L E , jQB D r A. EXPANSION 1) RECOVERY OF T H E R M A L E N E R G Y , j 6« 3p/8r 2) R E C O V E R Y OF SWIRL E N E R G Y , v B. D R I F T C U R R E N T , j

D

« j & r

LI

« £ x B

W H E R E v IS VELOCITY ( m / s ) AND E IS ELECTRIC FIELD ( V / m ) . MAGNETIC FLUX DENSITY, B (TESLA): APPLIED SELF I N D U C E D

ARC CURRENTS, j ( A / m 2 ) : APPLIED INDUCED

J Fig. 1 MPD mechanisms.

arc thruster

currents,

magnetic fields, and thrust

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APPLIED-FIELD MAGNETOPLASMADYNAMIC THRUSTERS

263

addition to the applied magnetic field there is an azimuthal magnetic field which is caused by the applied arc current itself. In addition to the applied arc current there is also an induced azimuthal current. With this geometry and these currents and fields, there are three basically different thrust mechanisms simultaneously present. The first is simply the ordinary gas dynamic pressure on the physical anode nozzle. The second is the direct plasma acceleration which is termed magnetic blowing and is the result of the interaction of the radial current and the azimuthal magnetic field. The third category is the magnetic nozzle plasma acceleration processes which result from the interaction of the azimuthal current with the radial field of the magnetic nozzle. As indicated, there are two major subcategories of magnetic nozzle processes. The first subcategory, the processes termed "expansion" in a magnetic nozzle, is analogous to expansion processes in an ordinary physical nozzle. In the conversion of thermal energy into directed energy in a magnetic nozzle, the radial plasma pressure gradient gives rise to a diamagnetic azimuthal current. With regard to the second subcategory of expansion processes, the interaction of the applied arc current and applied magnetic field causes rotational or swirl energy to be added to the plasma. As the exhaust expands along the magnetic field lines, it moves to larger radii, to conserve angular momentum its azimuthal velocity decreases, and just as in a physical nozzle, the swirl energy is converted into directed energy, i.e., energy in the direction of the thruster's axis. Associated with this conversion is an azimuthal current driven by the ~\T X If electric field. Thus associated with expansion processes in a magnetic nozzle are two mechanisms that cause azimuthal current to be induced in the magnetic nozzle. A third alternative mechanism results from the drift or Hall current which is caused by the interaction of the applied electric and magnetic fields and flows in the -TxT? direction. For any given MPD arc experiment, the thruster mass flow, current, and magnetic field can be adjusted so as to make any one of the mechanisms discussed important. Conversely, to determine how a given MPD arc thruster is producing thrust requires an assessment of the relative importance of the possible thrust mechanisms. In the case where the magnetic nozzle thrust mechanisms dominate, that is, when the applied magnetic pressure (B^/2y o/ where

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264

G.R. SEIKEL, T.M. YORK, AND W.C. CONDIT

y o is the permeability of free space) is much greater than the plasma pressure and the physical nozzle diverges much more rapidly than the applied field, one can view the MPD arc thruster as an arc jet in which the physical nozzle has been replaced by a magnetic nozzle so that the gasdynamic heat-transfer losses of the physical nozzle are avoided.

High-Power MPD Thruster Performance Over the years the basic configuration of an MPD thruster has been investigated from 0.1-40 kW with steadystate radiation-cooled electrodes, to over 100 kW with steady-state water-cooled electrodes, and to many megawatts in pulsed experiments with and without external fields. In most steady-state experiments at least a modest external field was used to spin the arc. This insures time averaged azimuthal symmetry of the arc attachment on the anode and avoids potential anode destruction which could result from filamented arc attachment at one anode location. Figure 2 illustrates the most highly developed highpower steady-state radiation-cooled MPD thruster developed for NASA. This X-7 thruster was designed and constructed by McDonnell Douglas^ for the NASA Lewis Research Center (LeRC) under the management of the first author. The contractor tested and delivered to NASA the thruster with a water-cooled magnet. The design located the magnet upstream of the electrode structure and incorporated radiation shields and insulation to minimize the heat flux to the magnet from the high-temperature radiation-cooled electrodes. Testing by McDonnell-Douglas included an over 500 hour life test of the X-7 thruster with the solid-conical cathode illustrated in Fig. 2. Cathode erosion and erosion of the adjacent insulator during the 500 hour test indicated the need for better and more positive control of cathode current attachment. Subsequently, a hollow-cathode was developed^ for this thruster which increased arc voltage and improved performance; results of a 10 hour test indicated that it may have potential for long-life operation. Performance testing of the X-7 MPD thruster was conducted by NASA LeRC. 1 1 Connolly, Bishop, and Seikel 1 2 subsequently integrated the X-7 electrode assembly, as shown in Fig. 2, with both a permanent magnet and a superconducting magnet and measured thruster perform-

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APPLIED-FIELD MAGNETOPLASMADYNAMIC THRUSTERS

265

ance; testing included evaluation of a hollow-cathode. Performance testing was conducted in the 15 ft diameter NASA LeRC vacuum facility,"^ which is capable of maintaining sufficiently low pressure, less than 1CT4 torr, for meaningful thrust measurements. Connolly, Sovie, and SeikeP 4 summarize most of the NASA LeRC steady-state watercooled MPD thruster testing, exhaust beam diagnostics, and evaluation of effects of test facility pressure. Figure 3 summarizes the efficiency as a function of power for several thrusters^J1,14 operating on NH3 propellant and with modest applied-fields, 0.1-0.15 T. Efficiency is the ratio of thrust power, T2/2m, as determined from measured thrust (on a thrust stand) and mass flow (by calibrated choked orifice) divided by the arc electrode input

ANODE RETAINING SPRINGS TUNGSTEN NOZZLE GRAPHITE RADIATOR

PROPELLANT

RADIATION SHIELDSINSULATORS

-TUNGSTEN CATHODE

(a) Electromagnet thruster

SECONDARY FLOWS i

SOLID CONICAL CATHODE

DUAL FLOW HOLLOW CATHODE

(b) Cathode geometries

Fig.

2 30 kW MPD arc thruster.

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266

G.R. SEIKEL, T.M. YORK, AND W.C. CONDIT -RADIATION SHIELDS

—CATHODE

ANODE — TUNGSTEN NOZZLE ^-GRAPHITE RADIATOR

RETAINING SPRINGS-^

2.5 CM

(c) Permanent magnet thruster ANODE ^-RETAINING SPRINGS

T-HELIUM FILL & VENT LINE

\ \ '/-\

SUPERCONDUCTING MAGNET

'••-";--

VACUUM D E W A R — ^ ^ COIL — - - _ _ _ _ _ SUPER INSULATION—-^_J^

LHe DEWAR — ----MAGNET CORE—^

NOZZLE

PROPELLANT— ^^]

— CATHODE

MAGNET SUPPORTS-

V-GRAPHITE RADIATOR

2.5 CM

(d) Superconducting magnet thruster Fig. 2 (cont.) 30 kW MPD arc thruster.

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APPLIED-FIELD MAGNETOPLASMADYNAMIC THRUSTERS

267

SPECIFIC IMPULSE, SEC

——— 2400

- - - 1500

30

RADIATION COOLED THRUSTER HOLLOW CATHODE

EFFICIENCY,

CONICAL CATHODE

„

20

25

50 ARC POWER, kW

H20 COOLED THRUSTER

75

Fig. 3 Performance of various NH3 MPD thrusters with 0.1-0.15

tesia applied magnetic fields.

power. Magnet power is not included; for permanent or superconducting magnets it would be negligible. Figure 3 shows that the performance of all the MPD thrusters improves with increasing power level. Radiationcooled thrusters with high-temperature anodes have higher performance than water-cooled cold anode thrusters, presumably a result of the difference in the anode voltage drops. The high impedance hollow-cathode radiation-cooled thruster out performs the conical-cathode X-7 due to the reduction in the ratio of anode losses to total power input. The power lost to the anode is directly proportional to the current to the anode; therefore, by increasing arc voltage at constant power, the anode current is decreased and anode losses are decreased. Results shown in Fig. 4 of the thrust measurements, energy balance measurements, and spectroscopic diagnostics measurements in the exhaust of a water-cooled MPD thruster indicated one reason performance is increased with power level. From thrust and energy balance measurements, an upper limit can be calculated for the possible fraction of the propellant that is ionized.^ This is determined from the power per unit mass flow available to produce ions and the energy required for ion production. This fraction ionized limit, which is a frozen flow energy loss and not recoverable, decreases with increasing mass flow or power (power is proportional to mass flow at a given specific impulse). Also shown in Fig. 4 are results of spectroscopic measurements^ which illustrate that, even at low mass flows, the thruster's exhaust contained high velocity neutral propellant atoms; however, there is some

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268

G.R. SEIKEL, T.M. YORK, AND W.C. CONDIT H20 COOLED NH3 THRUSTER MASS FLOW, g/SEC

0.02 60 MAX POSSIBLE 40 % IONIZATION

20

500

(a)

1000 1500 2000 SPECIFIC IMPULSE, SEC

2500

Maximum fraction ionized

30x10

AXIAL VELOCITY, v, M/SEC

10-

0 1 2 3 A X I A L D I S T A N C E FROM ANODE EXIT PLANE,

4 IN.

(b) Axial velocity versus axial distance, m = 0.02 g/s,l s p = 1765. Fig. 4 Power balance considerations and spectroscopic diagnostics in exhaust of water cooled NH3 MPD thruster.

slip between the ionized and neutral atoms and an associated loss of effective energy. Figure 5 shows the variation of efficiency as a f unction of specific impulse for the superconducting magnet MPD thruster in Fig. 2 operating at 25 kW with argon propellent.^ 2 Efficiency increases significantly with magnetic field strength. Efficiency exceeds that assumed for MPD thrusters in recent missions studies, 4 and further increases would be extrapolated with higher fields which are well within the state of technology. Field strength in Fig. 4

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APPLIED-FIELD MAGNETOPLASMADYNAMIC THRUSTERS

269

MAGNETIC FIELD TESLA

1000

2000 SPECIFIC IMPULSE, SEC

Fig. 5 Thrust efficiency for superconducting magnet MPD thruster.

Is measured at the centerline of the cathode tip. It is significant to note that boil-off of He from such a superconducting magnet is negligible in mass flow when compared to that of the thruster. The operating power of a radiation-cooled MPD thruster is limited by the heat rejection capability of the anode. Figure 6 shows the thermal design calculation for a 16 kW heat rejection by the X-7 thruster anode. For this thruster, which has an anode diameter slightly less than 2 cm, the tungsten in the anode arc attachment region reaches a temperature of almost 3600 K. There is a 1400 to 1500 K temperature drop in conducting the anode heat loss the few centimeters radially outward to the graphite radiator flange. The graphite radiation flange has a temperature variation of 800 K between its inner and outer radii. Although slight increases in the power handling capability of the anode assembly could be made by extending the tungsten to a larger radius, this would significantly increase the weight of the anode and its difficulty of fabrication. Practically, to significantly increase the power leve* of the anode requires increasing the inner anode diameter. The effects on performance of this type scaling have not been adequately explored; arc voltage should increase due to the longer path length, but discharge density and the associated improvement in the coupling between neutrals and ions would not increase as it does with increasing power at constant specific impulse in a fixed anode diameter thruster.

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270

G.R. SEIKEU T.M. YORK, AND W.C. CONDIT

1439

1433

1370

1364

Fig. 6 Thermal design for 16 kW heat rejection by anode of 30 kW radiation-cooled MPD thruster.

Very High-Power MPD Thruster Performance Operation of quasi-steady megawatt MPD arc thrusters has been investigated at a number of laboratories. Of particular interest are the experiments conducted with a thruster equipped with a superconducting magnet to supply an external magnetic field 15; fields from 0 to 2 T and currents of up to 20 kA were investigated. The thruster was energized by a 10 kJ capacitor bank that was crowbarred at peak current to provide a monotonic decaying current for 0.500 ms.

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APPLIED-FIELD MAGNETOPLASMADYNAMIC THRUSTERS

271

Results presented in Figs. 7 and 8 include current, I, and applied magnetic field, B, as standard measured variables. However, the evaluation of the thrust was carried out by application of a novel pressure diagnostic developed by T. M. York. From the impact pressure measurements, and simultaneous Thomson scattering measurements of density, the flow velocity can be calculated. Further, when the impact pressure measurements are integrated across the diameter, thrust is uniquely calculated. From Fig. 7 it can be seen that thrust increases with both applied-field and instantaneous current. For the zero applied-field case, the anticipated thrust can be directly calculated from the Fig. 7 Thrust of quasi-steady MW MDP thruster as a function of current for various applied magnetic fields. Theory is for the zero applied-field case.

O 0 EXPERIMENTAL^ THEORETICAL

100 80 60 40 •z.

fc

20

10

10 INSTANTANEOUS CURRENT, I, RA

20

30

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272

G.R. SEIKEL, T.M. YORK, AND W.C. CONDIT 3.0X106

2.0

1.0

1.0 AUXILIARY MAGNETIC FIELD, T

2.0

Fig. 8 MW MPD thruster, thrust and power as a function of applied magnetic field strength for current of 8 kA.

current and estimated cathode and anode attachment diameters'^; this calculation is also indicated in Fig. 7. For the higher applied fields (2 T), measure thrust is a full order of magnitude higher than the thrust calculated for the blowing mechanism (6=0, self-field); therefore, at these large applied fields the self-field MPD arc blowing thrust mechanism is insignificant. Figure 8 more clearly shows that for a given current, power increases linearly with applied-field but the thrust increases more rapidly with the magnetic field strength. Thus, it can be seen clearly that efficiency which is proportional to thrust squared divided by power is a very rapidly increasing function of the magnetic field strength. Unfortunately, in these crowbarred very high-power experiments, the variation of power with time and the uncertainly in mass flow prohibited obtaining a definitive value of efficiency. By the time a well-controlled delay line power supply''' 7 was added to these experiments, investigation of this device as a thruster had been terminated and no additional thrust measurements were made. The results obtained by Michels and YorkJ^ however, clearly demonstrate that the performance of a thruster at high power levels is significantly improved, just as it is for lower power thrusters by the addition of a strong applied magnetic field. MPD Thruster Physics Issues

The physics issues, which must be addressed before complete comparison of experiments with analytical/com-

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putationai models of MPD arc thrusters can be made, are grouped into two general categories: 1) Those in the relative high pressure, a number of torr, arc chamber. 2) Those in the low pressure exhaust of the thruster

Phenomena in Arc Chamber Probably the most critical issue in obtaining high performance with MPD arc thrusters for sustained periods of time is the control of cathode arc attachment. This problem is illustrated in Fig. 9 which is from the 500 hour duration test of the X-7 thruster.9 Each of the photos in Fig. 9 is labeled by the time in hours after start of the test and the operating voltage of the arc. As can be seen, at 530 hours, current attachment is at the cathode tip and the arc is running in what would be termed a high voltage mode.

t = 530.3 V= 52.3

544.5 V=43.0

t = 544

V= 48.7

t = 553.5 V= 37.0

Fig. 9 Variation of MPD cathode current attachment and arc voltage with time in hours.

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As time proceeds, current attachment gradually migrates away from the cathode tip. At 553 hours, no current is attached at the cathode tip and the arc has shifted to a low voltage mode of operation. At this point the test was terminated because of the rapid erosion to the thruster's boron nitrite insulator which would result from the current attachment in its vicinity. Use of hollow cathodes in MPD thrusters is one promising approach to better control the cathode current attachment. A significant amount of research and development must be accomplished, however, before reliable highpower thruster performance can be obtained for the long duration that will be required for orbit-raising missions.2~4 Higher performance was obtained using a hollow cathode configuration similar to that shown in Fig. 2(b). To maintain current attachment in the hollow cathode, all of the propellant was required to flow through the cathode. When secondary flow was introduced, current attachment moved to the outside surface of the cathode. A second aspect of the current flow in MPD arc thrusters is illustrated in Fig. 10. Under most operating conditions in steady-state experiments and in some quasisteady experiments current flow in the arc has been shown to not be uniform, but to be concentrated in a spoke or number of spokes. With an applied axial magnetic field, the spoke rotates in the E X B direction. The resulting high frequency rotation, hundreds of kHz, prevents destruction of the anode due to localized current attachment. Spoke frequencies appear to correlate well with an electrothermal DIRECTION OF ROTATION-,

SIDE VIEW Fig. 10

FRONT VIEW

Rotating current spoke in MPD thruster.

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instability theory.^ The existence of current spokes permits more rapid radial diffusion of current across the magnetic field than would be calculated from classical steady-state diffusion theory; thus, an analysis of MPD arc thrusters must provide for inclusion of such effects. Past experiments only identified these spokes near their onset condition where they are discrete. The recent developments of instrumentation and powerful digital correlation techniques permit definitive study of turbulent diffusion across the fields of these thrusters. Such diffusion results from a correlation in the density and^ ejectric field fluctuations which leads to a net radial iTxlf drift of the electrons across the magnetic field. A third important phenomenon in the arc chamber is illustrated by Fig. II. The thrust-to-power ratio, which is proportional to the efficiency divided by the specific impulse, is shown as a function of an applied magnetic field shape parameter. This shape parameter is the derivative with axial distance at the cathode tip of the log of the applied magnetic field. As indicated, performance is optimized with a slightly diverging magnetic field at the cathode tip; however, a rapid diverging field is very detrimental to performance. It is not surprising that in thrusters where the magnetic nozzle thrust mechanisms dominate, the AMMONIA PROPELLANT

3. OxlO"

4

AMMONIA FLOW RATE,

THRUST TO

POWER RATIO, N/KW

FIELD SHAPE PARAMETER, CM" Fig. 11 MPD thrust to power ratio as a function of applied magnetic field shape.

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shape of the applied-field is important. Improved understanding of the effects of applied magnetic field shape, not only in the arc chamber but also in the low pressure exhaust, could be significant in improving performance.

Phenomena in Thruster Exhaust While considerable empirical testing 1 ^ and computational modeling work^O has been done to understand the performance of MPD arcs, there has not been a completely successful effort to define the mechanisms active in even self-magnetic field thrusters. The details of the particle interactions with the magnetic fields exhibit complex behavior. One specific series of experiments^*! carried out by York on a quasi-steady MPD clearly demonstrated the extent to which fields (and currents) are blown downstream in a self-field MPD. The acceleration chamber had a 12.7 cm diameter and was 4.1 cm deep, the conical-tipped thoriated tungsten cathode had a 1.0 cm diameter and was 2.5 cm long, and the aluminum anode had a 6.4 cm diameter orifice. The test gas was He and operation was at a total current of 14 kA. A magnetic probe was placed in the exhaust to measure enclosed current, i.e., current flow within a given radius. The results are shown in Fig. 12 for propellant mass flows of 1, 2, and 4 g/s. Currents and fields in this case extend significant distances downstream. With a larger anode orifice of 10.2 cm in diameter, the downstream current blowing is even considerably greater. Such behavior indicates that simple models which presume major interactions only within short distances from the electrodes may be inherently in error. Devices utilizing applied magnetic (nozzle) fields have received less attention. Of particular interest here is the quasi-steady megawatt MPD arc thruster^ previously discussed. In experiments conducted with this thruster, the local total pressure, equal to the momentum flux, was mapped by a unique technique.22 Radial profiles of the momentum flux at different currents and field conditions showed strong variations in acceleration dynamics^ as indicated in Fig. 13. Specifically, two points are worth noting. First, it can be seen that there is a sharply defined "blast front" at higher values of applied-fields; this front was found to be comprised of mostly neutral atoms swept in front of the current layer. Second, there is a pronounced increase in signal at larger radii and higher fields, thus

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I6r

12 -

8 -

0 ANODE FACE

5

10 AXIAL DISTANCE,cm

15

CURRENT PROFILES Fig. 12 Measured current extension into the exhaust of a selffield MW MPD thruster. Current enclosed by a probe of 2 1/2 cm radius located on the thruster axis is shown as a function of probe axial location relative to the anode face, for various propellant flows. Anode radius is 3.2 cm.

indicating the lateral deflection of a larger proportion of the exhaust. Such effects are not yet understood, but while they are probably detrimental, they do offer a potential for further improvements in thruster performance, if controlled. These profiles of momentum flux were combined with Thomson scatting measurements of electron density, n e / and electron temperature, Te, in the plasma region to provide indications of flow velocity. The possibility of refining such measurements can provide unique tools to unravel this complicated exhaust structure. Some studies of the detailed processes occurring during the expansion of plasma along confining magnetic field lines have been carried out in theta pinch experiments. 23 Although 6-pinch (J0B 2 ) interactions are definably simpler than plasma interactions in an appliedfield MPD thruster, the application of these new diagnostic techniques coupled with improved models should aid

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-^

u, .