Impact Behavior of Fibre Reinforced Laminates: Fundamentals of Low Velocity Impact and Related Literature on FRP (Materials Horizons: From Nature to Nanomaterials) 9811694389, 9789811694387

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Materials Horizons: From Nature to Nanomaterials

Kalyan Kumar Singh Mahesh Shinde

Impact Behavior of Fibre Reinforced Laminates Fundamentals of Low Velocity Impact and Related Literature on FRP

Materials Horizons: From Nature to Nanomaterials Series Editor Vijay Kumar Thakur, School of Aerospace, Transport and Manufacturing, Cranfield University, Cranfield, UK

Materials are an indispensable part of human civilization since the inception of life on earth. With the passage of time, innumerable new materials have been explored as well as developed and the search for new innovative materials continues briskly. Keeping in mind the immense perspectives of various classes of materials, this series aims at providing a comprehensive collection of works across the breadth of materials research at cutting-edge interface of materials science with physics, chemistry, biology and engineering. This series covers a galaxy of materials ranging from natural materials to nanomaterials. Some of the topics include but not limited to: biological materials, biomimetic materials, ceramics, composites, coatings, functional materials, glasses, inorganic materials, inorganic-organic hybrids, metals, membranes, magnetic materials, manufacturing of materials, nanomaterials, organic materials and pigments to name a few. The series provides most timely and comprehensive information on advanced synthesis, processing, characterization, manufacturing and applications in a broad range of interdisciplinary fields in science, engineering and technology. This series accepts both authored and edited works, including textbooks, monographs, reference works, and professional books. The books in this series will provide a deep insight into the state-of-art of Materials Horizons and serve students, academic, government and industrial scientists involved in all aspects of materials research.

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Kalyan Kumar Singh · Mahesh Shinde

Impact Behavior of Fibre Reinforced Laminates Fundamentals of Low Velocity Impact and Related Literature on FRP

Kalyan Kumar Singh Department of Mechanical Engineering Indian Institute of Technology (Indian School of Mines) Dhanbad, India

Mahesh Shinde Department of Mechanical Engineering Indian Institute of Technology (Indian School of Mines) Dhanbad, India

ISSN 2524-5384 ISSN 2524-5392 (electronic) Materials Horizons: From Nature to Nanomaterials ISBN 978-981-16-9438-7 ISBN 978-981-16-9439-4 (eBook) https://doi.org/10.1007/978-981-16-9439-4 © The Editor(s) (if applicable) and The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 This work is subject to copyright. All rights are solely and exclusively licensed by the Publisher, whether the whole or part of the material is concerned, specifically the rights of translation, reprinting, reuse of illustrations, recitation, broadcasting, reproduction on microfilms or in any other physical way, and transmission or information storage and retrieval, electronic adaptation, computer software, or by similar or dissimilar methodology now known or hereafter developed. The use of general descriptive names, registered names, trademarks, service marks, etc. in this publication does not imply, even in the absence of a specific statement, that such names are exempt from the relevant protective laws and regulations and therefore free for general use. The publisher, the authors and the editors are safe to assume that the advice and information in this book are believed to be true and accurate at the date of publication. Neither the publisher nor the authors or the editors give a warranty, expressed or implied, with respect to the material contained herein or for any errors or omissions that may have been made. The publisher remains neutral with regard to jurisdictional claims in published maps and institutional affiliations. This Springer imprint is published by the registered company Springer Nature Singapore Pte Ltd. The registered company address is: 152 Beach Road, #21-01/04 Gateway East, Singapore 189721, Singapore

To Science and Technology To our Parents

Preface

We study the impact response behaviour of fibre reinforced polymer (FRP) materials. During our daily discussion related to our research always had difficulty in finding a single book related to our research because all the available books either completely dedicated to FRPs materials or mathematical modelling or impact loading. Thus, we decided to write a book which should cover all the basic and essential information about impact loading, particularly low velocity impact, FRP materials and different numerical modelling techniques. This book deals with FRP material response to impact loading, particularly low velocity impact. FRP materials are gradually replacing the conventional metals and alloys in the field of automobiles, aircraft, marine vehicles, defence protective armours and civil constructions due to their high specific strength, high specific stiffness and manipulation of the material properties in the required direction without compromising the strength. However, FRP materials have weak through-thickness strength properties due to which these FRP composites are vulnerable to impact loading particularly to low velocity impact because the low velocity impact creates barely visible impact damages (BVID) in FRP composites. These BVID are neither visible to naked eyes nor detectable to non-destructive methods during an inspection. Further, these BVID grow throughout the structure with time and severely reduces the residual compressive strength, which leads to premature sudden catastrophic failure of the structure. Consideration of impact and compression after impact criteria in consideration of the factor of safety during material design attained considerable attention after NASA’s Colombia space shuttle disaster which took all the onboard crew’s life, which happened almost 17 years back, but the understanding of the impact response of FRP laminates is still inadequate. Thus, we are working on to analyse the impact response of the FRP materials. The first half of this book covers the general information regarding fibre reinforced polymer composites, classic laminate theory, these composites testing methods under different loading conditions, low velocity impact and major mathematical models used in the numerical simulation, while the second half of the book covers the current research trends related to impact and previous literature works by various authors. vii

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Preface

Literature work includes GFRP, CFRP, hybrid FRP, different nanomaterial-doped FRP materials and other FRP composites (natural fibres). This book is useful for those who are interested in working with FRP materials under impact loading, those who wish to gain general information about FRP materials and those who want to study the basics of different testing methods associated with FRP materials. Overall, this book helps those who want to carry out their research in the field of FRP composites under low velocity impact. Dhanbad, India

Dr. Kalyan Kumar Singh Mahesh Shinde

Acknowledgements

The present book is an effort to throw some light on ‘impact behaviour of FRP composites’. This work would not have been possible without the help of some people in professional life as well as personal life. With the deep sense of gratitude, we acknowledge the guidance received by Prof. Prashant Kumar, Former Professor of IIT Kanpur, and Dr. R. K. Singh, Former Scientist of DMSRDE, Kanpur, for their guidance to Dr. K. K. Singh during his Ph.D. work to develop the interest in the field of FRP composites, especially in impact analysis. Mr. Mahesh Shinde is presently pursuing his Ph.D. in the area of impact analysis of FRP composites under the supervision of Dr. K. K. Singh. We would like to thank Prof. Rajiv Shekhar, Director, Indian Institute of Technology (Indian School of Mines), Dhanbad, for providing the research facility in the Department of Mechanical Engineering. This research facility is the basic need for writing a book in this research area. We would also like to thank and show our gratitude to our parents and family members for their humble and extraordinary support. We are thankful to our research group Dr. Vinay Ugale, Dr. Nand Kishore Singh, Dr. Dheeraj Kumar, Dr. Prashant Rawat, Dr. Anand Gaurav, Mr. Touhid Alam Ansari, Mr. Ruchir Srivastava, Mr. Santosh Kumar, Mr. Raju Thakur and Miss Nisha Sharma. Finally, we thank Springer for accepting our proposal for this book. We would like to extend our sincere thanks to all the employees present at the Springer office for their direct or indirect support and coordination. Dhanbad, India

Dr. Kalyan Kumar Singh Mahesh Shinde

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Contents

1 Introduction to Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.1 Basics of Fibre Reinforced Polymer Composite Laminates . . . . . . 1.2 Classification of Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3 Different Types of Fibre Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3.1 Synthetic Fibres . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.3.2 Natural Fibres . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.4 Different Types of Polymer Materials . . . . . . . . . . . . . . . . . . . . . . . . 1.5 Different Types of Fibre Weaving Architecture . . . . . . . . . . . . . . . . 1.6 Different Types of Fibre Reinforced Polymer Composite Laminate Designs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7 Different Fabrication Methods of Fibre Reinforced Polymer Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.1 Wet Hand Lay-Up Technique . . . . . . . . . . . . . . . . . . . . . . . . 1.7.2 Vacuum Bagging Process . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.3 Spray Lay-Up Technique . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.4 Filament Winding Method . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.5 Pultrusion Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.7.6 Resin Transfer Moulding Process . . . . . . . . . . . . . . . . . . . . 1.7.7 Autoclave . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.8 Micro- and Macromechanics of Fibre Reinforced Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.8.1 Micromechanics of Lamina . . . . . . . . . . . . . . . . . . . . . . . . . 1.8.2 Macromechanics of Laminate . . . . . . . . . . . . . . . . . . . . . . . . 1.9 Fibre Reinforced Polymer Nanocomposite Laminates . . . . . . . . . . . 1.10 Advantages of Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.11 Application of Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

1 1 2 3 5 8 9 10 11 17 17 17 19 19 20 21 21 22 22 26 29 30 31

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1.12 General Failure Mechanism of Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.12.1 Fibre Damage Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . 1.12.2 Matrix Damage Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . 1.13 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1.14 Further Reading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2 Mechanical Testing of Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1 Mechanical Testing of Fibre Reinforced Polymer Composite Laminates According to ASTM Standards . . . . . . . . . . . . . . . . . . . . 2.1.1 Measurement of Constituent Content in FRP Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.2 Tensile Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.3 Compression Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.4 Flexural Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.5 Short Beam Shear Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.6 Impact Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.7 Compression After Impact Test . . . . . . . . . . . . . . . . . . . . . . 2.1.8 Double Cantilever Beam Test . . . . . . . . . . . . . . . . . . . . . . . . 2.1.9 End Notch Flexural Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.10 Mixed Mode Interlaminar Fracture Test . . . . . . . . . . . . . . . 2.1.11 Translaminar Fracture Toughness Test . . . . . . . . . . . . . . . . 2.1.12 Fatigue Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.13 Lap Shear Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.1.14 Quasi-Static Indentation Test . . . . . . . . . . . . . . . . . . . . . . . . 2.1.15 In-Plane Shear Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2 Failure Modes in Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.1 General Failure Modes Observed Under Tensile Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.2 General Failure Modes Observed Under Compression Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.3 General Failure Modes Observed Under Low Velocity Impact or Quasi-Static Indentation . . . . . . . . . . . . 2.2.4 General Failure Modes Observed Under Compression After Impact Loading . . . . . . . . . . . . . . . . . . . 2.2.5 General Failure Modes Observed in Specimen Under Out-Of-Plane Shear Loading . . . . . . . . . . . . . . . . . . 2.2.6 General Failure Modes Observed Under Shear Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

37 37 40 42 42 42 45 45 45 49 52 54 56 57 61 62 64 65 68 68 69 69 71 73 73 74 74 75 76 76

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2.2.7

General Failure Modes Observed in Adhesive FRP Composite Lap Joints . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.2.8 General Failure Modes Observed in Filled or Non-Filled Open-Hole Specimens Subjected Tensile or Compression . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.3 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2.4 Further Reading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

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78 79 79 80

3 Low Velocity Impact on Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 83 3.1 Basics of Impact Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 84 3.2 Classification of Impact Loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86 3.3 Parameters Influencing the Impact Performance of Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . 86 3.4 Damage Detection Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 88 3.5 Numerical Simulation of Fibre Reinforced Polymer Composite Laminates Subjected to Low Velocity Impact . . . . . . . . 96 3.5.1 Material Models . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 96 3.5.2 Contact Definition . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 99 3.6 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 3.7 Further Readings . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 105 4 Low Velocity Impact on Carbon Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4.1 Low Velocity Impact Test on Controlled Carbon Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . 4.2 Low Velocity Impact Test on Hybrid Carbon Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . 4.3 Low Velocity Impact Test on Nanomaterial-Doped Carbon Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . 4.4 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5 Low Velocity Impact Test on Glass Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.1 Low Velocity Impact Test on Controlled Glass Fibre-Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . 5.2 Low Velocity Impact Test on Hybrid Glass Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.3 Low Velocity Impact Test on Nanomaterial Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5.4 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

107 107 128 134 141 142 149 149 169 181 183 183

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6 Low Velocity Impact Test on Other Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.1 Impact Test on Controlled Other Fibre Reinforced Polymer Composites Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.2 Low Velocity Impact Test on Other Hybrid Fibre Reinforced Polymer Composites Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6.3 Low Velocity Impact Impact Test on Nanomaterials Doped Other Fibre Reinforced Polymer Composites Laminates . . . . . . . . 6.4 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7 Low Velocity Impact Study on Symmetric and Asymmetric Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . 7.1 Introduction to Symmetric and Asymmetric Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.2 Low Velocity Impact Test on Symmetric and Asymmetric Composite Laminate Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7.3 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8 Compression After Impact on Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.1 Compression After Impact Test on Carbon Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.2 Compression After Impact Test on Glass Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.3 Compression After Impact Test on Other Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8.4 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9 Numerical Analysis of Low Velocity Impact and Compression After Impact on Fibre Reinforced Composite Laminates . . . . . . . . . . . 9.1 Numerical Analysis of Low Velocity Impact on Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . 9.2 Numerical Analysis of Compression After Impact on Fibre Reinforced Polymer Composite Laminates . . . . . . . . . . . . . . . . . . . . 9.3 Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . References . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . .

191 191 199 212 214 216 221 221 224 233 234 237 237 248 257 260 261 265 265 290 297 298

About the Authors

Dr. Kalyan Kumar Singh is an Associate Professor in the Department of Mechanical Engineering and presently officiating as Associate Dean (Monitoring and Review) at IIT (ISM) Dhanbad. He is with IIT (ISM) Dhanbad since August 2009. Previously, he has worked at NIT Jalandhar, HCST Mathura and IISc Bangalore in different capacities. He is selected “State Coordinator” for the state of Jharkhand by Ministry of Education, government of India for State specific perspective plan for technical education in India. He is the Founder Chairman of Indian Society for Advancement of Materials and Process Engineering (ISAMPE) Dhanbad chapter. He was also the Honorary Secretary of “The Institution of Engineers (India)” Dhanbad Local Centre during 2016–18. Dr. Singh has been working in the area of Mechanical Characterization of FRP composites for over 18 years. His research includes the impact, fracture, fatigue, tribology, failures and machining of FRP composites. Impact on the FRP composite remains as the dominant area of his research. He has guided 12 Ph.D. and 35 M.Tech. Students as sole guide or main guide. He has published more than 116 research papers in international journals only. Currently, he is supervising seven Ph.D. scholars in the field of FRP composites. Total value of the R&D projects and STP courses under him is approximately 100 lakhs. He is also a member of International Society of Impact Engineering USA. Dr. Singh has received N. K. Iyengar award in 2014 and Canara Bank Research Paper Publication award for two consecutive years 2016 and 2017 at IIT(ISM) Dhanbad. Mr. Mahesh Shinde is currently a Ph.D. student in the Department of Mechanical Engineering at Indian Institute of Technology (ISM) Dhanbad. He obtained his B.E. in Mechanical Engineering and M.Tech. (Computer Integrated Manufacturing) from Visvesvaraya Technological University, Karnataka in 2014 and 2016, respectively. During his M.Tech., he worked on copper nanocomposites to increase the strength and electrical conductivity of the material. He has published some good research papers in international journals and conferences. Currently, he is pursuing his research work under Dr. K. K. Singh on the impact damage behaviour of fibre reinforced nanocomposites.

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Chapter 1

Introduction to Fibre Reinforced Polymer Composite Laminates

Abstract Fibre reinforced polymer (FRP) composites are becoming well known and finding widespread applications in various industries due to their high specific stiffness and strength compared to conventional metals and their alloys. This chapter covers various aspects of FRP composite materials such as different methods used in classifying the FRP composites, different types of fibre and matrix materials used in fabricating the FRP composite laminates. Further, this chapter also covers different fibre weaving architectures which significantly affects the FRP composites strength. Followed by which a brief introduction is given on the laminate design techniques and various FRP composite laminates fabrication methods. Furthermore, this chapter also briefly covers micro- and macromechanics of FRP composite materials. Nowadays, the nanomaterials are extensively used in polymer composites to improve the mechanical properties of FRP composites. Thus, in this chapter a brief section is also dedicated to FRP nanocomposites. In the subsequent section of this chapter, advantages and applications of FRP materials are also briefly covered. This chapter finally ends with general failure mechanisms associated with FRP composites. Keywords FRP composites · Classification of composites · Fibre type · Polymer Type · Fibre weaving · Stacking sequence · FRP Fabrication · FRP mechanics · FRP nanocomposites · FRP advantages · FRP application · Failure mechanism

1.1 Basics of Fibre Reinforced Polymer Composite Laminates When two or more different materials mixed at a macroscopic level, the resulting third useful material is known as composite material (Fig. 1.1). Nevertheless, unlike alloys, the constituents of composite materials retain their original properties even after mixing and also after disintegrating. Further, when a load applied onto a composite material, initially the matrix takes the applied load and then transfers the applied load to the fibres. Here, fibre is the primary load-bearing member and matrix is just supporting the fibres.

© The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_1

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1 Introduction to Fibre Reinforced Polymer …

Fig. 1.1 Cross-sectional view of a composite material

stacked laminate resin

fibre fibre-matrix interface matrix rich region reinforcement layers reinforcement

Constituents of fibre reinforced polymer composite laminates As already said, the FRP composite material consists of two distinct phases known as discontinuous phase called as reinforcement and a continuous phase known as the matrix. The reinforcement is always stronger and stiffer than the matrix, and it is the main load-carrying member of the constituent. While the matrix material helps in transferring the applied load to the fibres. The main function of the reinforcement is to carry the applied load. Other matrix functions are to keep the fibres in the desired position with intact, protect the fibres from the environment and avoid the abrasion of the fibres from each other. The main reason for the broad application of composite materials in various industries is high degree of freedom to a material designer to get the required mechanical properties in the desired direction along with high specific strength, high specific stiffness, high fatigue strength and good chemical inertness. Further, the high strengthto-weight ratio and high strength-to-stiffness ratio are primary required properties in a material for high-performing structural applications used in defence, aerospace and automobile. These FRP composite material properties not only provide high strength to these structures but also reduce the structural weight, which results in fuel efficiency which again helps in cost cutting.

1.2 Classification of Fibre Reinforced Polymer Composite Laminates Generally, there are two methods to classify the composite materials. The first classification method is based on the matrix materials. For example, if the composite matrix material is made of polymer material, then that composite material is termed as polymer matrix composites. Similarly, if the matrix material of a composite is made of metal and ceramic material, then the composite material is called as metal matrix and ceramic matrix composites, respectively. The second classification method is based on the reinforcement geometry. For example, if the reinforcement has high

1.2 Classification of Fibre Reinforced Polymer …

3

aspect ratio, then it is called as continuous reinforcement, and if the reinforcement has low aspect ratio, then it is called as short reinforcement. Further, if the reinforcement shape is similar to sphere with size being around the micrometres, then the reinforcement is called as particulate reinforcement. Furthermore, if the reinforcement has irregular shape with negligible thickness, then the reinforcement is called as whiskers or flake reinforcement. Apart from this general classification of composite materials, the FRP composite laminates are also classified based on the polymer matrix used and the reinforcement geometry. Based on the type of polymer material used, the polymer composite materials are classified thermoset polymer and thermoplastic polymer composite. The polymer composites are further classified based on the reinforcement materials such as fibre reinforced polymer composites and particulate reinforced polymer composites. Further, the fibre reinforced polymer composites are classified as long or continuous fibre reinforced polymer composites and short or discontinuous fibre reinforced polymer composites (Fig. 1.2a, b). These types of FRP composites are further classified based on the alignment of the fibre in the polymer matrix. If the reinforced fibres are oriented along the same direction, then the polymer composites are termed as aligned fibre reinforced polymer composites while if the reinforced fibre direction is random, then the polymer composites are termed as randomly oriented fibre reinforced polymer composites (Fig. 1.2c, d). The randomly oriented FRP composites nearly behave like isotropic material. Because in the case of randomly oriented FRP composites, it is assumed that the probability of fibre orientation in the matrix in all direction is approximately the same. Thus, it shows almost the same strength properties in all the directions. However, in the case of aligned FRP composites, the strength properties are maximum along the fibre direction thus assumed as highly anisotropic. The discontinuous fibre reinforced polymer composite is also termed as ‘chopped mat’. If the continuous fibres arranged in one direction, then it also termed as unidirectional FRP composites, whereas if the continuous fibres arranged or weaved into a fabric then it is termed as the woven FRP composites. The woven FRP composites are further classified based on their weaving structures which will be discussed in the coming section.

1.3 Different Types of Fibre Materials Fibres are matter with high aspect ratio where the diameter of the fibre is significantly small compared to its length. The length of the fibre is in few millimetres, while the diameter of the fibre is in microns. Synthetic and natural are the two major types of fibres available. Synthetic fibres are human-made such as glass, carbon and Kevlar. While, natural fibres are produced in nature and these are further classified as plant fibres, animal fibres and mineral fibres. Plant fibres include hemp, kenaf and flax. Whereas, animal fibres include collagen, spider silk and wool. Example of mineral fibre is asbestos. The natural fibres cannot compete with the synthetic fibres. However, the natural fibres can be recycled and cost-effective. The synthetic fibres can be further

4

1 Introduction to Fibre Reinforced Polymer …

continuous or long fibre

(a)

aligned continuous or long FRP laminates

(c)

randomly oriented continuous or long FRP laminates

discontinuous or short fibre

(b)

aligned discontinuous or short FRP laminates

(d)

randomly oriented discontinuous or short FRP laminates

Fig. 1.2 Classification of FRP composite a continuous aligned FRP composite, b short aligned FRP composite, c continuous randomly oriented FRP composite, d short randomly oriented FRP composite

classified into polymer synthetic fibres and non-polymer synthetic fibres. The first synthetic fibres made are polyamide or nylon in the year 1938. All the polymer-based synthetic fibres are made from polycondensation and/or polymerization process of polymer chains. There are different types of units used for defining the fibre. First is the ‘Denier’, and second is the ‘Tex’. Both the units are weight of the fibre as a function of fibre length. Denier is the one gram per nine kilometre, whereas the Tex is the one gram per kilometre. The fibres used in textiles are polyester, polyolefins (polypropylene and polyethylene), nylon and acrylic. All fibre surfaces are coated with compatible polymeric sizing and binders. The function of the sizing is to protect the fibre from damage that may occur during handling. The sizing refers to the thin polymer film made of a silane coupling agent which helps in adhesion or chemical bonding with the polymer matrix.

1.3 Different Types of Fibre Materials

5

1.3.1 Synthetic Fibres Carbon Fibre The carbon fibres are made from polyacrylonitrile (PAN). In the early days, the carbon fibres are synthesized from cellulose via pyrolysis process and used as light bulb filament. Carbon fibres are used in lightweight structural applications in industries such as aerospace, marine, automobiles and sports goods. Fishing rods, gold clubs, Airbus A380 and Boeing 787 are few examples which use carbon fibres. Both Boeing 787 and Airbus A380 use approximately 35 tonnes of carbon as reinforced polymer composites which reduces the overall weight of the aircraft and increases the aircraft carrier mileage. Carbon fibres are also used in windmills blades used for clean electricity generation. Use of carbon fibre reduces the overall weight of the structures which in turn also helps in reducing the emission of greenhouse gases. Carbon fibres are basically synthesized by two process. First process is by carbonizing precursor, and the second process is from hydrocarbon gas. There are three precursor methods to synthesize the carbon fibres: 1. PAN, 2. pitch and rayon-based carbon fibres. The carbon fibres have high strength, high modulus and high strength-to-weight ratio. Due to these properties, these materials mostly used in the lightweight and high-performing structures such as aerospace, space crafts, automobile and marine. These kinds of industries prefer lightweight materials with high strength to increase fuel efficiency. The main disadvantages associated with carbon fibres are high cost, poor performance under impact loading. The difference between graphite and carbon fibre is the presence of carbon content. Graphite fibre contains about 99% of carbon, whereas carbon fibre contains approximately 80–95% of carbon content. Presence of covalent bonding in the carbon fibres is the key to its strength. Among various methods, polyacrylonitrile is the most widely used process to fabricate carbon fibres. Glass Fibre Glass fibre is produced from silicate or sand. First, the glass fibre reinforced polymer composite was produced around 1935. Dale Kleist, Jack Thomas and Games Slayter were responsible for the modern fibre glass. Glass fibre is solidified in an amorphous arrangement. There are variety of glass fibres in the market. These glass fibres generally differ in the composition. Glass fibres are made of soda lime and silica. These fibres have low tensile strength along with poor chemical resistance under acidic environment. This type of glass fibres are used in thermal insulation application. There is wool glass fibre. The composition of wool fibre glass is the soda lime and silica along with 2–10% of B2 O3 . These fibre glass wools are used in thermal insulator and sound-absorbing applications. The other type of glass fibre is alkali-resistant (AR-glass fibre) glass fibres. These fibres are generally used in civil applications such as reinforcement in concrete. Most commonly used glass fibre type is E-glass (‘E’ means electrical resistivity) fibre. This type of glass fibre finds widespread application as reinforcement in polymer composites. Further, the E-glass fibres also used in electrical circuit board. The chemical attach resistance by this glass fibre is good. The composition of E-glass fibre is silica, calcia and

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1 Introduction to Fibre Reinforced Polymer …

alumina. Now, there are high-strength glass fibres known as ‘R’ and ‘S’ glass fibre. These fibres have high strength compared to all other types of glass fibres. Here ‘R’ and ‘S’ stand for strength. In these fibres, the major constituent material is the SiO2 which provides the strength to the fibres. These high-strength glass fibres are used in ballistics armours, wind turbine blades and gas-compressed gas tanks. These fibres are mostly preferred where high strength and good thermal durability are required. There is also a ‘D’ glass fibre which is used in electronic circuit. Here, ‘D’ stands for dielectric. Glass fibres are extensively used due to its low cost with reasonable strength along with its availability and variety as mentioned above. Nevertheless, the significant limitations of the glass fibres are poor abrasion resistance, poor performance in moisture containing environment. The thermal coefficient of expansion and strengthto-weight ratio is lower than that of the carbon fibres, and it performs better under impact loading compared to carbon fibres due to its better deformation. All types of glass fibres are fabricated by a standard method known as a molten-extrusion process. Variation in glass fibre properties obtained by varying the silica or other constituents ingredients to get desired property in the fibre. Aramid or Kevlar Fibre Aramid fibres are discovered in 1970. There are two different types: meta aramids and para aramids. The meta-aramid fibres are made from meta-phenylene isophthalamide, and the para-aramid fibres are made from para-phenylene terephthalamide. The difference between meta- and para-aramid fibre is the presence of –NH– and –CO–. Meta aramid fibre is simply called as aramid fibres, while the para aramid fibres are called as Kevlar fibres. Basically, in general the aramid fibres are nothing but the aromatic polyamides, which belongs to the family of nylon. These fibres are fabricated by polycondensation of diamines and halides at low temperature. There are two types of Kevlar fibres available known as Kevlar-29 and Kevlar-149. These fibres have a high impact resistance property and high compressive strength. Thus, these are usually used in bulletproof jackets. In general, these fibres are extensively used in protective gears. However, these are more expensive than glass fibres. Boron Fibre These fibres have high tensile strength, compressive strength and usually fabricated from chemical vapour deposition method. Their greater fibre diameter when compared to other fibres enables them to perform better under buckling loading. However, these fibres are costlier than carbon fibres. Nylon Fibre Nylon fibres are made from thermoplastics consisting of aliphatic polyamides with amide groups. These fibres are first introduced in 1939. It is used in many industrial applications due to their high strength, good abrasion resistance along with low moisture absorption. It is used in aircraft tyres as reinforcement, spinnaker sails, airbag sewing threads, sportswear and cloth for parachute. Nylon fibres are synthesized by

1.3 Different Types of Fibre Materials

7

two methods. First method involved reaction between diamine and a dibasic acid. In the second method, the amino acid monomer chain is opened which contains lactam ring with amine and acid groups. Examples of nylon fibre produced by first method are nylon 6.6 and nylon 1.6. Here the first digit of the number indicates the number of carbon atoms present in the diamine, while the second digit of the number represents the dibasic acid content. Similarly, the nylon fibre produced by the second method includes nylon 6 and nylon 12. The number at the end represents the carbon atoms number in the lactam ring. Paul Schlack invented the nylon 6 in the year 1939. Dupont scientist Wallace Carothers produced the Nylon 6.6 in the year of 1934. Nylon fibres show good chemical resistance and thermal properties. Polyester Fibre Polyester fibres are the synthetic fibre made of aromatic carboxylic acid. These fibres are produced in 1951 by DuPont. Here, the polyester fibres are produced by using ethylene glycol and terephthalic acid. These fibres are mostly used in textiles. Further, the polyester fibres show good wear resistance, dimensional stability and low moisture absorption. Polypropylene and polyethylene Fibres Both polypropylene and polyethylene are members of polyolefins. These fibres also show good strength, better abrasion resistance and low moisture absorption along with low cost and ease of fabrication. Polypropylene fibres are produced by propylene polymerization. These fibres have good tensile strength and comparatively low strain to failure. These fibres are now losing their popularity due to increase in cure oil cost. Polyethylene fibres have high strength and modulus. There are different kinds based on the molecular weight of the polyethylene. For example, ultrahigh-molecular-weight polyethylene (UHMWPE) fibres, high-density polyethylene (HDPE) fibres, high-modulus polyethylene (HMPE) fibres and highperformance polyethylene (HPPE) fibres are produced from ‘gel-spinning’ process. Dyneema and Spectra are the trade names of the polyethylene fibres used by Nippon Dyneema, Japan and Honeywell Speciality Materials, USA, respectively. Acrylic Fibres Acrylic fibres are made from acrylonitrile monomer chains. Acrylic fibres are the precursors used for fabricating carbon fibre synthesis. The carbon fibre quality is controlled by the nitrile group and oxygen atoms along the fibre axis. Oxidation, stabilization and carbonization are the three main stages of the carbon fibre synthesis. Here, the stabilization stage involves exothermic reaction. The entire temperature range maintained during the stabilization process is from 200 to 320 °C. Oxidation process conducted at 220 °C, where the fibre tow is heated for several hours in air to add oxygen to the fibre structure. After oxidation process, the fibres are subjected to carbonization process conducted at 1500 °C. Unlike oxidation process, the carbonization of fibre is conducted at an inert atmosphere. Followed by carbonization, the fibres are subjected to graphitization. In this state, the fibre again heated to 2500 °C. In this stage, the percentage of carbon content is decided. If the carbon content is nearly

8

1 Introduction to Fibre Reinforced Polymer …

90%, then the fibre is called carbon fibre while if the carbon content is 100%, then the fibre is called graphitization.

1.3.2 Natural Fibres These types of fibres are found in nature. Growing concern for the environment is increasing the demand for natural fibres. The natural fibres are biodegradable, abundant and cost-effective. However, before using as reinforcement the natural fibres are subjected to surface treatment. Further, the natural fibres still do not meet the strength properties of the synthetic fibres. Furthermore, the natural fibres are made of cellulose and lignin which are hydrophilic in nature and thus absorb the moisture which reduces the strength in composites. There are a variety of natural fibres and classified based on their origin. If the fibre is from the plant, then they can be called as plant fibres. Correspondingly, there are also animal and mineral fibres. Flax fibre is one of the examples of natural fibre and obtained from seed crop plants. Hemp fibres are obtained from the cannabis plant. Sisal fibres are commonly obtained from cactus-type plants which contain needle-like structures. Animal Fibres Wool is also a fibre obtained from animals such as sheep. The Keratin is the major ingredient of the wool fibre. The chemical structure of the wool is helical unlike the folded molecular chains. Silk fibre is mainly obtained from silk worms/moths and spiders. Silk is a protein secreted by the insects. Both the fibres’ tensile strength is high. These fibres are mainly made of amino acid sequence. Collagen fibres are found in the muscles of the vertebrates. Plant Fibres Cotton Fibre Cotton fibres are commonly used in textiles. It is majorly made of long chains of carbohydrate molecules or polysaccharides. The convoluted arrangement of the fibrils and twisted-ribbon shape along the length of the fibre provides the cotton its tensile strength. Hemp Fibre Hemp fibre is obtained from hemp plant. It is the stem part of the plant. Agave Americana fibre is obtained from the leaves of the Agave American plant. The tensile strength of the hemp fibre falls in the range of 340–527 MPa. Basalt Fibre The basalt fibres are obtained from melting basaltic rock which is also known as mineral wool. Basalt fibres are soluble in lung fluids and can cause lung infection

1.3 Different Types of Fibre Materials

9

resulting in long lung irritation. Basalt fibres are initially used as high-temperature thermal insulation. Until recently, these fibres are also gaining lot of attention due their high tensile strength and finding widespread implications in structures.

1.4 Different Types of Polymer Materials The polymer resins are long chain of repeating units of monomers. Broadly, the polymer matrix materials are classified as thermoset and thermoplastic polymers. Further, the polymer matrix material can also be classified as synthetic and natural biopolymers. The synthetic polymer materials are manmade, while the natural polymers are directly obtained from the nature. Polymer matrix materials can also be classified based on the polymer molecular chain clustering or structure such as linear, branched and networked/crosslinked polymer chain materials (Fig. 1.3). Linear chain polymer materials have less strength and stiffness. Branched polymers are stronger than linear chain polymer materials, and these material structures of networking are rearranged upon application of heating. Thermoplastics materials come under this category of materials. Networked/cross-linked chain materials are stronger and stiffer than the other two types, and the chain structure of these materials cannot be rearranged, unlike the branched chain polymers. Thermosetting materials come under this category of polymer matrix materials. Thermoset Polymer Thermosetting polymer materials harden on heating. These materials have better strength, excellent thermal stability but cannot be recycled. These materials consist of covalent bonding and, hence possesses better temperature and environmental resistance. These materials also have cross-linking polymer chain networking which gives better hardness, strength and brittleness properties to the composites. Examples are polyether-ether-ketone, polyethylene and polyester.

Linear chain

Branched chain

Networked chain

Fig. 1.3 Schematic representation of molecular arrangement in the polymer matrix

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1 Introduction to Fibre Reinforced Polymer …

Thermoplastic Polymer Thermoplastic matrix materials are those which soften upon heating. These materials have low strength and thermal stability compared to thermosetting materials, and these materials can be recycled. These materials consist of branched or linear polymer molecular chains. A drastic change in material properties is observed above its glass transition temperature, and hence, its brittleness and strength can be controlled by controlling the glass transition temperature by adding plasticizers. These kinds of resins include polyethylene-ether-ether-ketone (PEEK), polypropylene and polycarbonate. Epoxy Polymer Epoxy is the widely used thermoset matrix material due to its high mechanical strength for a low cost. It has low molecular weight with epoxide groups, i.e. one oxygen atom and two carbon atoms. Presence of polar hydroxyl groups and ether bonds helps epoxy to adhere with several materials. Epoxies synthesized by using epichlorohydrin with bisphenol-A. It has better moisture, chemical and temperature resistance properties with a temperature range up to 175 °C. Polyester Polymer Polyesters are a long unsaturated linear chain of carbon atoms with a double bond. It decomposes when heated, i.e. thermoset polymer. The condensation reaction between glycol and dibasic acid is used to synthesize polyester. Commonly used glycols are ethylene and dibasic maleic acid. These are non-polar, saturated, high-molecularweight hydrocarbon thermoplastics and synthesized by mixing the same polymers of ethylene. Polyesters are usually solid in state and cure exothermically. It is inexpensive, lightweight, resistant to environmental exposures with a temperature range up to 100 °C. These polymers have low strength, hardness and rigidity, and high ductility, high impact strength and low friction. Also, they exhibit good chemical, moisture and electrical resistance.

1.5 Different Types of Fibre Weaving Architecture Weaving pattern of the plies in FRP laminate plays one of the significant role in imparting properties to the laminates. Based on this, they broadly classified into two categories, i.e. unidirectional and bidirectional. In unidirectional weaving pattern, the fibres weaved in one direction only that is either along ‘x’ or ‘y’ direction. The strength properties are high along the fibre direction and low in the perpendicular direction that is same in both ‘y’ and ‘z’ directions. However, in the case of bidirectional weaving pattern, the in-plane fibre properties are same along both ‘x’ and ‘y’ directions. The bidirectional weaving pattern is further subdivided into three major categories such as plain woven, satin woven and twill woven fabric (Fig. 1.4). In plain weaving pattern, the fibres weaved in a criss-cross pattern, i.e. warp and weft weaving. The satin

1.5 Different Types of Fibre Weaving Architecture

plain weave

twill weave

11

satin weave

Fig. 1.4 Schematic representation of different weaving patterns

weaving pattern is made by passing one weft thread over more than one warp thread and one warp thread over more than one weft thread, and this process continued till the symmetric pattern along the diagonal achieved. Nevertheless, in case of the twill weave pattern, the weaving is carried out by passing four or more weft yarns over one weft yarn and vice versa. The end pattern is similar to a small structure floating in a river at a regular distance.

1.6 Different Types of Fibre Reinforced Polymer Composite Laminate Designs Single fabric and resin constitute a lamina or ply. The collection of these laminas or plies is called as a laminate. These laminates may consist of any number of laminas and can be oriented at any angle depending upon the required properties. Based on the fibre direction, the laminates are classified as unidirectional laminate and bidirectional laminate. Based on fibre orientation about the mid-plane, the laminates are classified as symmetric, asymmetric, antisymmetric, orthotropic and quasi-isotropic. Unidirectional Laminate In this type of laminate, the fibres in a ply oriented along one direction (Fig. 1.5). Here, the strength properties are high along the fibre direction while weak along the transverse direction. Strength properties along both ‘y’ and ‘z’ directions are the same. Bidirectional Laminate Here, laminate made of bidirectional woven fabric, i.e. in-plane properties of the fabric, is the same along ‘x’ and ‘y’ directions (Fig. 1.6).

12

1 Introduction to Fibre Reinforced Polymer … z x

laminate y unidirectional lamina

plies or laminas fibre yarn

Fig. 1.5 Schematic representation of a unidirectional FRP laminate z

bidirectional woven lamina tow cross over

Fig. 1.6 Schematic representation of bidirectional woven FRP laminate

Symmetric Laminate In a symmetric laminate, the orientation of the plies about the midplane (an imaginary plane that more or less divides the laminate into two equal halves) is same in terms of ply thickness, orientation and properties (Fig. 1.7). In general, it is represented by [81 /P2 //P2 /81 ]nS where ‘n’ and ‘s’ denote the number of plies and symmetry of the laminate. In symmetric laminate, the bending–extension coupling matrix is always zero. For symmetric laminate, the bending stiffness matrix (Bij ) present in classic laminate theory is equal to zero. Antisymmetric Laminate An antisymmetric laminate is symmetric about the midplane, but the orientation of the plies is precisely the mirror image of the corresponding ply about the midplane

1.6 Different Types of Fibre Reinforced Polymer …

13

z

θ1 Ѳ2 Ѳ2

imaginary mid-plane

θ1 ‘z‘ direction

Fig. 1.7 Schematic representation of symmetric laminate

(Fig. 1.8). The fabric thickness and properties remain the same, but the fibre/ply orientation will be exactly opposite to that of the corresponding ply about the midplane. The antisymmetric laminate is represented by [81 /82 //−82 /−81 ]nS where ‘n’ and ‘S’ represent the number of plies and symmetricity of the laminate, respectively. For this kind of laminate design, A16 = A26 = D16 = D26 = 0 (where A16 and A26 —represent the shear extension coupling; D16 and D26 —represent the bend twist coupling). Asymmetric Laminate In this type of laminate design, either the ply orientation or fibre thickness or ply properties may vary (Fig. 1.9). There is no plane of symmetry about the midplane. In general, the asymmetric laminate design represented by [81 /82 /83 /….]n where ‘n’ represents the number of plies present in that particulate laminate. z

θ2 θ1 –θ1 –θ2 Z direction

Fig. 1.8 Schematic representation of antisymmetric FRP laminate design

imaginary mid plane

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1 Introduction to Fibre Reinforced Polymer … z

θ1 θ2 θ3 θ4

Fig. 1.9 Schematic representation of asymmetric FRP laminate design

Cross-ply Laminate In this type of laminate, the fibre or plies are orientated along 0° and 90° alternatively (Fig. 1.10) two differently oriented angle plies arranged alternatively. Generally, this type of laminate design represented by [0°/90°/0°/90°/…..]n where ‘n’ denotes the number of plies present in that laminate. Most commonly used cross-ply fibre orientation changes either to 0° or to 90°, but the plies’ properties and thickness remain the same. In this laminate, the design does not show any kind of coupling between in-plane extensional and the shear response which means that the terms A16 z

900 00 900 00

Fig. 1.10 Schematic representation of cross ply FRP laminate design

1.6 Different Types of Fibre Reinforced Polymer …

15

z

–450 +450 900 900

Imaginary mid plane

+450 z direction

–450

Fig. 1.11 Schematic representation of specially orthotropic FRP laminate design

= A26 = B16 = B26 = D16 = D26 = 0 for any kind of cross-ply laminate (where A16 and A26 are shear extensional coupling). Specially Orthotropic Laminate The laminate design in which A16 and A26 (where A16 and A26 are shear extensional coupling) are equal to zero, then the laminate design is called as specially orthotropic laminate (Fig. 1.11). Cross-ply, angle ply and antisymmetric laminate designs come under this category. For cross-ply to be a specially orthotropic laminate, there is no restriction that the corresponding pairing ply thickness needs to be same. Because in case of cross-ply laminate design, the Q 16 and Q 26 terms are equal to zero. However, for angle ply and antisymmetric ply laminate designs, the pairing angle plies’ thickness needs to be same to become specially orthotropic laminate. Quasi-Isotropic Laminate In quasi-isotropic laminate, the material properties such as stiffness and elastic constants are identical in specific directions, i.e. the extensional matrix is the same (Fig. 1.12). It means that in quasi-isotropic laminate design, the extensional stiffness matrix behaviour is similar to isotropic material. Here, A11 = A22 , A16 = A26 = 0 12 . To form a quasi-isotropic laminate, minimum of three plies of and A66 = A11 −A 2 equal thickness are needed. Further, the ply orientation of the adjacent ply in this type of laminate is π/n, where n is the total number of plies used in that laminate. The bending and bending–extension coupling matrices may or may not behave like isotropic material. Examples for quasi-isotropic laminates are [0°/ + 60°/−60°]S and [0°/45°/−45°/90°]S .

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1 Introduction to Fibre Reinforced Polymer …

900 + 600 – 600

Fig. 1.12 Schematic representation of quasi-isotropic FRP laminate design

Balanced Laminate In this type of laminates, the fibre/ply orientation is opposite with respect to the position of that ply about the midplane (Fig. 1.13). Here, also the fibre thickness and properties remain the same except the fibre/ply orientation. In general, it is represented by [81 /82 /83 …….//−81 /−82 /−83 …]n . Hybrid Laminate This type of laminate mainly consists of two different types of plies made of different materials and exhibits different materials properties (Fig. 1.14). This type of laminate is usually denoted by starting letter of its name and with fibre orientation as a subscript. Example [C 0 /G90 /Gr45 ………]n where n, C, G, Gr indicate the number of plies in that laminate in carbon ply, glass ply and graphite ply, respectively. z

300 600 – 300 – 600

Fig. 1.13 Schematic representation of balanced FRP laminate design

1.7 Different Fabrication Methods of Fibre Reinforced …

17

Glass fibre Carbon fibre Graphite fibre

Fig. 1.14 Schematic representation of hybrid FRP laminate design

1.7 Different Fabrication Methods of Fibre Reinforced Polymer Laminates The manufacturing process is one of the crucial parameters in deciding the laminate properties. There are various processes available to fabricate the FRP composites, and depending upon the size of the component, structural strength required, and application of the structural component, the fabrication process is selected. Moreover, the selection of the fabrication process also depends on the matrix and fibre material used. In this chapter, several fabrication methods are explained.

1.7.1 Wet Hand Lay-Up Technique This fabrication process is simple and cost-effective. In this method, the dry fibre fabric is wetted with resin by a soft brush, and extra resin and air bubbles entrapped between the plies are squeezed out using steel roller (Fig. 1.15). This procedure is continued until the required design for the laminate is achieved. Before using the stainless steel roller, it is cleaned with acetone, after that heated to remove the contaminating particles. However, the effectiveness and efficiency of removal of air bubbles and extra resin are not sufficient. This process becomes difficult in fabricating complex parts but useful for manufacturing large-sized parts.

1.7.2 Vacuum Bagging Process This method is an improvised version of hand lay-up process (Fig. 1.16). Here the removal of entrapped air and extra resin is extracted by roller followed by vacuum suction (Fig. 1.17). The vacuum provides extra benefit in the removal of extra air

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1 Introduction to Fibre Reinforced Polymer …

Resin

stainless steel roller for squeezing

wet laminate

Fabric Fig. 1.15 Schematic representation of wet hand lay-up process resin

stainless steel roller for squeezing extra resin

vacuum bagging set-up

wet laminate

fabric

Fig. 1.16 Schematic representation of vacuum bagging procedure ejector film FRP laminate perforated film

breather film vacuum bag

to vacuum pump

base sealant

Fig. 1.17 Typical illustration of vacuum bagging setup

mould

1.7 Different Fabrication Methods of Fibre Reinforced …

19

bubbles, resin and in achieving uniform thickness. Complex shapes and mediumsized parts could be fabricated. This process provides better control over the thickness of the laminate than the wet lay-up method. During fabrication, wet hand lay-up placed inside a vacuum bag over which a perforated sheet placed and above which a breather. The perforated sheet acts as a passage for excessive resin flow and avoids sticking the breather with wet lay-up. The breather acts as a medium of absorbing material.

1.7.3 Spray Lay-Up Technique This process is suitable for manufacturing chopped strand FRP laminates. In this process, a continuous spray of chopped strands of fibre and resin sprayed onto a mould which is of required shape through different nozzles (Fig. 1.18). Spray speed and movement of the nozzle defines the thickness of the laminate. The main concerning factor about this is that there is no control over the formation of void content. In the fabrication of chopped FRP laminates, this is a cost-effective method.

1.7.4 Filament Winding Method This method is used for fabricating hollow or convex-based hollow cylindrical structures such as pipes or tanks. Here, a spool of long fibre is passed through a resin bath and then wound onto a mandrel. The mandrel is removed after achieving the required resin bath resin + chopped fibre sprayer valve to control flow resin chopped fibre fibre spool

Mould

Fig. 1.18 Schematic illustration of spray lay-up procedure

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1 Introduction to Fibre Reinforced Polymer …

Fig. 1.19 Schematic representation of filament winding process

mandrel

resin impregnated fibres resin bath

direction of fibre movement

fibre spool

product (Fig. 1.19). Mandrel speed defines the thickness of the laminate, and the final product outer surface is not aesthetically pleasing. For the production of cylindrical FRP composite, this method is one of costs of the productive process. However, for low viscous resin, this process becomes tedious, and fibres cannot be wounded along the tube’s longitudinal axis. However, the winding pattern and winding angle decide the strength of the tube, which are controlled by adjusting the mandrel axis.

1.7.5 Pultrusion Method In this method, the long fibre is drawn from a creel and passed through a resin bath (Fig. 1.20). Then the wetted fibres are moved into a compressing die. The opening of the die decides the thickness of the laminate. After this, the compressed laminate chopper or cutter fibre positioning unit resin bath

fibre spool

Fig. 1.20 Schematic illustration of pultrusion procedure

forming die

heater

1.7 Different Fabrication Methods of Fibre Reinforced …

21

moved through an oven. Finally, the cured laminates are cut by a cutter depending on the required length of the component. This method is a continuous process and mainly used for long solid rectangular or cylindrical type of components.

1.7.6 Resin Transfer Moulding Process Here, the fibre layers laid inside a mould which is of the required shape and size (Fig. 1.21). Then, the upper portion of the mould is placed onto it. After that, the upper portion of the mould is positioned according to the required thickness. Now, from one side, the resin is injected into the mould under a defined pressure depending upon the viscosity of the resin used. This method is used to fabricate small-to-medium-sized aircraft structures.

1.7.7 Autoclave In this type of fabrication method, the prepregs of fibres are placed inside a closed chamber where temperature and pressure raised (Fig. 1.22). By this method, large parts such as aircraft wings, ship hulls can be made. Also, complicated shapes are manufactured by this method. mould up and movement pressurised resin in

vacuum assistance & resin out

laminate

Fig. 1.21 Schematic illustration of resin transfer moulding procedure

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1 Introduction to Fibre Reinforced Polymer …

closed chamber

to vacuum

sealant mould

laminate

base

Fig. 1.22 Schematic illustration of autoclave procedure

1.8 Micro- and Macromechanics of Fibre Reinforced Composite Laminates Since the use of fibre reinforced polymer (FRP) composites in aircraft during World War II, demand and importance for these materials increased exponentially. Need for these materials increased in the sectors where lightweight and high-strength components are required. Thus, the understanding the mechanical properties of these composite materials via mathematical formulation is important. Hence, the mathematic formulations were segregated into two sections: (i) micromechanics and (ii) macromechanics. The micromechanics of the laminate is based on the rule of mixture, while the macromechanics of the laminate is based on the classical laminate plate theory. The predicted basic constituent and strength values from these methods are not accurate; however, it provides near accurate results which can be considered reasonably while designing the structure. There are various factors which aid for not predicting accurate results; among them are anisotropic behaviour and heterogeneity.

1.8.1 Micromechanics of Lamina Micromechanics deals with the behaviour of composite constituent materials at the microscopic level. Microscopic properties such as lamina density, fibre–matrix interface behaviour, void content, strength in longitudinal and transverse strength.

1.8 Micro- and Macromechanics of Fibre Reinforced …

23

Predicted properties greatly influenced by fibre and matrix properties. Properties predicted are exponential and upper bound of the lamina. Assumptions made during lamina property estimation are as follows: • • • • • •

Lamina behaves linearly elastic Lamina is free from defects Lamina is homogeneous and isotropic Fibre and matrix are perfectly bonded No sliding exist between fibre and matrix under load Fibre spacing is regular and parallel to each other.

The volume of the composite is the sum of the volume of fibre and matrix (Eqs. 1.1–1.4). υc = υ f + υm

(1.1)

Vf =

vf vc

(1.2)

Vm =

vm vc

(1.3)

Vm + V f = 1

(1.4)

where υ c —Volume of composite material in mm3 υ f —Volume of fibre in mm3 υ m —Volume of matrix in mm3 V m —Volume fraction of matrix V f —Volume fraction of fibre. Weight of the composite is defined as the sum of the weight of the fibre and matrix (Eqs. 1.5–1.8). wC = w f + wm

(1.5)

Wf =

wf wC

(1.6)

Wm =

wm wC

(1.7)

W f + Wm = 1

(1.8)

24

1 Introduction to Fibre Reinforced Polymer …

where W m —Matrix weight fraction W f —Fibre weight fraction wf —Fibre weight in grams wm —Matrix weight in grams wc —Composite weight in grams. The density of the composite material is determined by its weight and volume (Eq. 1.9). 1 = ρC = ρ f V f + ρm Vm or ρc



Wf ρf



 +

Wm ρm

 (1.9)

where ρ c —Composite density in g/cc ρ f —Fibre density in g/cc ρ m —Matrix density in g/cc V f —Fibre volume fraction V m —Matrix volume fraction W f —Fibre weight fraction W m —Matrix weight fraction. Void contents are majorly introduced during manufacturing of the composite. These act as stress-concentrating sites and reduce the strength of the composite. It is predicted by using Eq. 1.10. υVOID =

(ρt − ρe ) 100 ρt

(1.10)

where ρ t —Theoretical density of composite ρ e —Experimental density of composite Modulus of the composite along the fibre direction is known as the longitudinal modulus (Fig. 1.23). It can be predicted from Voigt model and is developed based on the rule of mixture technique (Eq. 1.11). Few assumptions made are as follows: • Fibre and matrix experience the same strain • Fibre geometry is uniform

1.8 Micro- and Macromechanics of Fibre Reinforced … Fig. 1.23 Schematic representation of tensile load along the fibre direction

25 z

• Fibres are aligned correctly and are bonded perfectly with the matrix • Distribution of the fibre is uniform • Voids are not present The equation for longitudinal modulus is as follows: E C = E f V f + E m Vm

(1.11)

where E C —Longitudinal modulus of composite in GPa E f —Longitudinal modulus of fibre in GPa E m —Modulus of matrix in GPa V f —Fibre volume fraction V m —Matrix volume fraction. Modulus perpendicular to the fibre length is called transverse modulus and is also known as Reuss model (Eq. 1.12) (Fig. 1.24). This prediction is based on the rule of mixture, and it is always smaller than the longitudinal modulus. Assumptions made for this technique: Fig. 1.24 Schematic representation of transverse tensile loading along the Z-direction

z tensile load

tensile load

26

• • • • •

1 Introduction to Fibre Reinforced Polymer …

Geometry and properties of the fibre remain uniform Matrix and fibres bonded perfectly No void content is present Fibres alignment and distribution is uniform Fibre and matrix carry an equal amount of load Vf 1 Vm = + Ec Ef Em

(1.12)

where E C —Longitudinal modulus of composite in GPa E f —Longitudinal modulus of fibre in GPa E m —Modulus of matrix in GPa V f —Fibre volume fraction V m —Matrix volume fraction.

1.8.2 Macromechanics of Laminate Laminate macromechanics deals with the structural behaviour or properties of a laminate. A laminate is a collection of several laminates which are all adhered by an adhesive material usually known as resin. Laminate properties are predicted by several methods, but the classic laminate theory is the most widely used technique. This theory is an extension of the isotropic plate theory developed by Kirchhoff, and the assumptions made are known as Kirchhoff’s hypothesis for plate and Kirchhoff– Love hypothesis for thin and shells. These assumptions may vary when applied to beams, bars and rods. Classic laminate plate theory (CLPT) makes few assumptions for thin plates and are as follows: • The thickness of the laminate is much less when compared to its width. Further, the midplane deflection is small compared to the thickness of the plate • Shear strains are assumed to zero • Before and after application of the load, the cross section of the laminate remains the same • Bonding between the fibre and matrix is perfect, and no slipping takes place. Further, under bending the normal out-of-plane strains are assumed to be zero • Stresses along through-thickness direction (σ z = 0) are equal to zero • Laminate behaviour is linearly elastic under load • Laminate is homogeneous • Laminate is defect-free.

1.8 Micro- and Macromechanics of Fibre Reinforced …

27

The relation between force, strain, moment and curvature is given by Eq. 1.13 ⎡

⎤

⎡

A11 NX ⎥ ⎢ ⎢ A21 ⎢ NY ⎥ ⎢ ⎥ ⎢ ⎢ ⎢ ⎢ N X Y ⎥ = ⎢ A16 ⎥ ⎢ B11 ⎢ ⎥ ⎢ ⎣ MY ⎦ ⎢ ⎣ B12 MXY B16

A12 A22 A26 B12 B22 B26

A16 A26 A66 B16 B26 B66

B11 B12 B16 D11 D12 D16

B12 B22 B26 D12 D12 D26

⎡ ⎤ ⎤ ε0X B16 ⎢ 0 ⎥ ⎢ε ⎥ B26 ⎥ ⎥⎢ Y ⎥ ⎥⎢ ε0 ⎥ B66 ⎥⎢ X X ⎥ ⎥ ⎥⎢ D16 ⎥⎢ γ X0 ⎥ ⎥ ⎥⎢ D26 ⎦⎢ γ 0 ⎥ ⎣ Y ⎦ D66 γ X0 Y

(1.13)

where N X —Normal force per unit length along X-direction in N N Y —Normal force per unit length along Y-direction in N N XY —Shear force per unit length in XY plane in N M X —Bending moment per unit length along X-direction in N-mm M Y —Bending moment per unit length along Y-direction in N-mm M XY —Twisting moment per unit length in XY plane in N-mm ε0X —Mid plane strain in X direction εY0 —Mid plane strain in Y direction γ X0 Y —Mid plane strain in XY plane k X —Mid plane curvature in X direction kY —Mid plane curvature in Y direction k X Y —Mid plane curvature in XY plane ⎡

⎤ n A11 A12 A16

 ⎣ A12 A22 A26 ⎦ = Ai j = (Q i j ) k (h k − h k−1 ), i, j = 1, 2, 3 . . . . k=1 A16 A26 A66 Aij —Extensional stiffness matrix ⎤ n B11 B12 B16

  ⎣ B12 B22 B26 ⎦ = Bi j = 1 (Q i j ) k h 2k − h 2k−1 , i, j = 1, 2, 3 . . . 2 k=1 B16 B26 B66 ⎡

Bij —Bending coupling stiffness matrix

28

1 Introduction to Fibre Reinforced Polymer …

⎡

⎤ n D11 D12 D16

  3  ⎣ D12 D22 D26 ⎦ = Di j = 1 Q k h k − h 3k−1 , i, j = 1, 2, 3 . . . 3 k=1 D16 D26 D66 Dij —Bending stiffness matrix A11 , A22 and A66 —Extensional coupling stiffness along X, Y and Z directions, respectively A16 , A26 —Shear extension coupling D16 , D26 —Bend twist coupling Q—Transformed reduced stiffness matrix which can be obtained from Q = [T ]−1 [Q][R][T ][R]−1 ⎡

⎤ 100 [R] = Reuter matrix = ⎣ 0 1 0 ⎦ 001 ⎤ ⎡ sin2 θ −2 sin θ cos θ cos2 θ [T ] = Transformation matrix = ⎣ sin2 θ cos2 θ 2 sin θ cos θ ⎦ sin θ cos θ − sin θ cos θ cos2 θ − sin2 θ [Q] = Reduced stiffness martix = [S]−1 ⎤ S11 S12 S16 [S] = Compliance matrix = ⎣ S12 S22 S26 ⎦ S16 S26 S66 ⎡

where S11 =

1 E1

S12 = − S22 = S66 =

ν12 E1

1 E2

1 = E6. G 12

1.9 Fibre Reinforced Polymer Nanocomposite Laminates

29

1.9 Fibre Reinforced Polymer Nanocomposite Laminates There are various methods to improve the properties of FRP laminates, ply hybridization and nanomaterial reinforcement are among these methods. For hybridization, generally, two methods are used, i.e. interply hybridization (where two different plies used in a laminate) and intraply hybridization (where two different fibres used within the ply). In recent years, hybridization is also be done using two different types of resins. In hybridization, selection of fibres plays an important role, and it should be in such a way that the difference between the coefficient of thermal expansion of these fibres should be minimum. However, in recent years use of secondary reinforcement in composites is gaining much primary focus to enhance the properties of composites which involves mixing of nanosized particles or fibres into the matrix material. Doping of these nanomaterial improves the fibre wettability, load transfer capacity from matrix to fibres. Further, the nanomaterial forms bridging between the matrix and fibres, which help in cracks arrest and inhibit the crack growth in composite materials. The primary problem associated with secondary or nanoreinforcement is achieving the uniform dispersion of the nanomaterials in the matrix phase. Nanocomposites A composite is said to be a nanocomposite if one of the constituent material dimensions is at the nanoscale (10–9 ) (Fig. 1.25). Nowadays, many nanoscale materials are available, and this is due to their dramatic changes in the material properties at that small scale. Few examples of nanomaterials are carbon nanotubes, nanoclay and graphene. Carbon Nanotubes Sumio Ijima discovered the carbon nanotubes (CNTs) at the beginning of the 1990s [1, 2]. These materials have excellent mechanical, electrical and thermal properties. agglomeration of nanomaterial macro fibre reinforcement nanomaterial reinforcement

Fig. 1.25 Schematic representation of advanced nanocomposite

30

1 Introduction to Fibre Reinforced Polymer …

Further, these materials have low density and high aspect ratio. Thus, these nanomaterials used as secondary reinforcing materials in composites. CNTs are classified based on the number of concentric cylindrical carbon walls such as single-walled carbon nanotubes (SWCNTs) which consist of a single graphene sheet rolled in a cylindrical tube-like structure. Multi-walled carbon nanotubes (MWCNTs) nothing but graphene sheets rolled concentrically (Fig. 1.25a, b). If the number of walls is two, then it is called as double-walled carbon nanotubes. If the number of walls is more than two, then it is called as multi-walled carbon nanotubes. Two types of models exist for MWCNTs and are parchment model and Russian doll model. In parchment model, a single layer of graphene sheet rolled into several concentric layers. In the case of the Russian doll model, several sheets of graphene layers rolled concentrically in a cylindrical tube structure. MWCNTs have better mechanical properties than the SWCNTs. Because MWCNTs have the telescopic action of multiple walls under loading, and there exist van der Waals forces between the layers.

1.10 Advantages of Fibre Reinforced Polymer Composite Laminates Various advantages and limitations are associated with these composite materials, and few of them mentioned below. Advantages of Composites • High strength-to-weight ratio results in reduced weight of the structure which is more useful in saving the fuel in case of aircraft and automobiles • Flexibility in tailoring the composite material properties in the required direction • Fewer assembly parts are required • Highly corrosion and chemical resistance which means less maintenance cost • Low manufacturing and material cost • Low coefficient of thermal expansion of composite materials results in high thermal and dimensional stability • Better performance under impact and fatigue loading. Limitations of Composite Along with various advantages, few limitations also associated with these composite materials which are as follows • Reparability of these composite materials is not easy • Damage inspection is difficult and complicated • Due to its anisotropic behaviour and heterogeneity, the mathematical damage formulation and prediction becomes a difficult task • Sudden and catastrophic failure without any yielding.

1.11 Application of Fibre Reinforced Polymer Composite …

31

1.11 Application of Fibre Reinforced Polymer Composite Laminates FRP composites have been used in the past by several civilization. The Egyptians and Mesopotamian civilization reinforced the bricks with straws. The Mongolians used the composite bow for hunting. These bows were generally made of bamboo and antler along with wood or leather. During the world wars, the FRP laminated composites found increased in applications. The wood laminate composites were used in the bomber of British Royal Air Force fighter planes. Since then the FRP composites have found widespread industrial applications in several sectors due to their high strength, lightweight, corrosion resistance, chemical attack resistance, fatigue resistance, good resilience and better reproductivity. More importantly, the properties of FRP composite laminates can be tailored in the required direction. These sectors are broadly classified as automobile, aircraft, space, turbines, marine, defence, civil construction, medical and sports equipment. Here, few practical applications are mentioned for each sections. Automotive The FRP composites are finding increased applications in the body parts of road transport vehicles due to design freedom, high strength-to-weight ratio and low cost. Automobile body parts are heater housing, bumpers, truck carriages, frames, hood/roof, shafts, door panels, etc. Few examples of automobiles which use FRP composites are BMW, Mercedes Benz E-Class and front end grill of the Ford Montagetrager. Further, the Massachusetts Institute of Technology (MIT) and automobile manufacturer Lamborghini teamed up to build the futuristic Lamborghini Terzo Millennio a futuristic sports car which has the capability of ‘self-healing’, whose body panels are mostly made of FRP composite laminates (Fig. 1.26a). Koenigsegg car manufacturer also uses the FRP composites. For example, almost all body parts of the Koenigsegg Agera R are made of carbon fibre such as intake panel of the engine, pressurized turbo pipes, valve pipes, body panels, door panels and wheels (Fig. 1.26b). Another car manufacturing company McLaren also uses the CFRP composites to build monoque for car body for model 720s (Fig. 1.26c). Similarly, Porsche Carrera GT monoque is also made from CFRP composite laminates. Further, side frame of the BMW Z22 is also made from FRP composite. Furthermore, several body parts of Volkswagen × 11 are made from CFRP composite laminates. Aircraft Fuselage, engine cowlings, floor panels, ceilings, seats landing gears doors, pilots cabin door panels, overhead bins, tail, rotor blades, wing trusses, rudder, radomes, fairing, helicopter blades, canopies of helicopter and antenna are few parts which extensively use FRP composite materials. The major reason for the use of FRP composites in aircrafts is the FRP composite laminates’ high-strength-to-weight ratio property along with good fatigue resistance and damage tolerance. The high strengthto-weight ratio of the FRP composite laminate property reduces the aircraft weight

32

1 Introduction to Fibre Reinforced Polymer …

(b) (a) (c)

Fig. 1.26 Automobiles Made from FRP composite laminates a Lamborghini Terzo Millennio (Photo Festival automobile international 2018), b Koenigsegg Agera RS (Photo 2015 Geneva motor show), c McLaren 720S (Photo Salon Prive Concours d-Elegance Classic Car Motor Show 2019)

which in turn increases the fuel efficiency and also reduces the emission of global warming gases to the atmosphere. The reduction in the weight of the weight of the aircraft is found be to be around 20–30% compared to parts made from conventional metals. Use of FRP laminated composites in aircraft increased since the world wars. During the World War II, wood laminated composites were used in the British Royal Air Force fighter planes. Boeing and Airbus aircrafts are most commonly known examples who adopted FRP composites progressively (Fig. 1.27a, b). Grumman E2A rudder was made of FRP composite (Fig. 1.27c). Airbus A300 made in 1972 used the FRP composite parts such as fins and fairing panels. In 1979, the parts of Airbus A300 further upgraded with FRP composite panels such as spoilers, rudders, brakes, landing gear doors, etc. In the following years 1980 to 1987, the use of FRP composites is extended to the Airbus 310 and Airbus A320, followed by Airbus A321, Airbus A330 and Airbus A340. Replaced aircraft parts include radome, fins, cabin, cabin floor, fairing, fuselage, vertical stabilizer, lading gears, trailing edge, engine cowls, spoilers, wing panels, landing gear doors, ailerons, flaps, etc. Similarly, Boeing company has also adopted FRP composite in their aircrafts progressively over last few decades. Boeing 707 is the first aircraft from Boeing to use FRP composites. Followed by Boeing 757, 737-300, 727 and 737-200 used the FRP composites in various parts. Parts that are replaced by FRP composites in different Boeing aircrafts are spoilers, fore-flaps, stabilizers, aileron, rudder, nacelles, deck floor, fairings, landing gear doors, floor beams, rudders, etc. The Avions de transportation Regional also uses FRP composite laminates in the model ATR42. In this aircraft, the wind box are made of FRP composite. Apart from civilian aircrafts, the defence or army aircrafts also uses the FRP composites. The fuselage, wing box, stabilizers, rudders, doors, horizontal tail,

1.11 Application of Fibre Reinforced Polymer Composite …

33

(a)

(b)

(c)

Fig. 1.27 Use of FRP materials in aircrafts a Boeing 787 Dreamliner [http://www.boeing.com/boe ing/companyoffices/gallery/images/commercial/787/index1.page], b Airbus A380 [Airbus A-380 Photo Gallery], c Grumman F-14 Tomcat Fighter Aircraft U. S. Navy [NASA F-14 Tomcat Photo Gallery]

fairing, brakes, flaps, aileron and wing box of aircraft are also made from FRP composites. Few examples of aircrafts used in military applications that use FRP composite laminates are AV-8B, F-14, F-15, F-16, F-18 and B-1. Similar to civilian and military aircrafts, the FRP composite laminates are also used in helicopters. First use of FRP materials as optimum pitch blade was in the XCH-47 twin rotor helicopter of Vertol Aircraft Corporation. Helicopter blades made of FRP composites show greater service life than the blades made of conventional metals. Further, the FRP composite blades also show no bend twist during rotation and require no frequent inspection. Other helicopters which use FRP composites are McDonnell Douglas MD 520N/MD 900, Eurocopter EC 135, MBB BK 117, Bell 206L, Bell 402, Windecker Eagle, etc. Apart from rotor blades, the FRP composites are also used in fuselage, stabilizers, cockpit, tail rotor blade, canopy frame and rotor hub. Marine In recent years, the ship hulls and decks are majorly made of FRP composites. The USS Zumwalt is a guided missile destroyer ship in which the deck portion of the ship is made of FRP laminate composite (Fig. 1.28a). The British Royal Navy used the HMS Wilton ship (minesweeper) made from glass fibre composite laminate manufactured by Vosper Thornycroft Company (Fig. 1.28b). The British Royal Navy further adopted the HMS Brecon ship (Hunt Class Mine Counter Measures Vessels) along with HMS Sandown ship as Single Role Mine Hunter. The Navy Visby class ship hull is made of PVC core sandwiched by carbon/vinyl ester skins (Fig. 1.28c).

34

1 Introduction to Fibre Reinforced Polymer …

Fig. 1.28 a USS Zumwalt ship (U.S. Navy photo by Petty Officer 3rd Class Anthony N. Hilkowski), b HMS Wilton ship (Photographed at Portsmouth), c Navy Visby (Photographed while anchored in the vicinity of Gotska Sandön, Sweden)

Nigel Irens designed the ocean petrol vessel named Ocean Eagle 43 is made from glass/carbon/epoxy hybrid composite laminate. For underwater application, impregnated carbon fibre composite laminates can be used (Fig. 1.29a). The glass and carbon fibre reinforced polymer composite tubes are

Fig. 1.29 a Carbon fibre cylinder tube for underwater application [1], b protective casing used for oceanic equipment protection [1]

1.11 Application of Fibre Reinforced Polymer Composite …

35

Fig. 1.30 Nautile, Ifremer Submersible Vehicle [1] (Photo copyright Ifremer/Olivier Dugornay)

used as protective casings (Fig. 1.29b). For deep sea explorations, the submersible vehicles are made of FRP composite laminate (Fig. 1.30). Civil construction Fibre reinforced polymer composites are finding more and more applications in civil construction industries in the form of rebars and flat panels. Rebars are used as reinforcement with concrete, while flat panels are used for bridge decks and girders. In recent developments, the FRP composites are also used as jackets/wraps for concrete columns, repairing the damaged concrete structures, and seismic retrofitting along with concrete grids and underground tunnels, tunnel supports, storage containers and power plant chimneys. Few examples of FRP usage in civil construction are highway Appia near Rome, Bridge A10062 at St. Louis was impacted-damaged by an overweight truck and repaired by using CFRP composites, Bridge A5657 near south of Dixon, Missouri, Bridge G270 of Missouri, Horsetail falls bridge of Oregon, Bridge column replacement and/or repair in various places of USA (Illinois I-57, Huntsville I-565 (Fig. 1.31), Nevada I-580, Salt Lake City I-80,), trail bridge in Santa Fe National Forest, FRP deck and truss bridge in Olympic National Park, and Fibre-line bridge Kolding, Denmark. Few examples of FRP rebars used in existing bridges are Verdasio bridge, Switzerland and Aberfeldy foot bridge, UK. Medical FRP composites’ weight is low and can be used as an endoskeleton artificial limbs. Epoxy, polyester and acrylic are the most commonly used polymer resins for orthoses.

36

1 Introduction to Fibre Reinforced Polymer …

Initial cracks found on bridge girder

Bridge girder after repair with CFRP laminate

Fig. 1.31 Use of fibre reinforced polymer composite laminates in I-565 Highway Bridge Girder at Huntsville, Alabama, USA [2]

Further, glass fibre, carbon fibre, nylon fibre and Kevlar fibres are the most commonly used fibre materials for medical applications. The ankle foot orthoses are made from carbon fibre impregnated composite laminates (Fig. 1.32a, b). The upper limb is protected by FRP composite laminate. The fibre glass can be used to avoid the ulnar collateral ligament injuries. The carbon fibre composite laminates are also used in extended rigid shanks. The fibre glass counters are used as the stabilizer for shoes.

Fig. 1.32 Ankle foot orthoses made from carbon fibre impregnated composite laminates a ankle foot orthoses made by Ottobock [Image source http://www.ottobock.co.uk/], b ankle foot orthoses made by Allard [Image source Allardusa.com], c knee brace made from carbon fibre composite laminate (From Goldberg and Hsu [3])

1.11 Application of Fibre Reinforced Polymer Composite …

37

Fibres are used in pelvic support orthoses. Recently, knee braces are made from carbon fibre composite laminates (Fig. 1.32c).

1.12 General Failure Mechanism of Fibre Reinforced Polymer Composite Laminates The damage mechanism in FRP laminates is a critical and complicated issue because it involves matrix and fibre failure at micro as well as at macroscale. The complexity further increases as the FRP composite laminates are anisotropic and heterogeneous which attributes to the unpredictable and uncertain failure mechanism. However, understanding these mechanisms is essential as damage induced in the laminate reduces the strength of FRP. This reduction in strength may be gradual or may be sudden. However, both result in unexpected and catastrophic failure without any warning. The damage mechanism is classified as barely visible and visible damage. The barely visible damage involves delamination and subsurface microcracks which can also be known as microlevel damage mechanism while the visible damage includes other than subsurface microcracks which are visible to naked human eyes, also known as macro-level damage mechanism. The damage mechanism or the failure modes associated with matrix and fibre materials are discussed in the subsequent section.

1.12.1 Fibre Damage Mechanism Failure of fibres can take place in various modes depending upon the applied load. Fibre Fracture This kind of failure takes place when fibre breaks into two parts. Usually, this mode of failure occurs when a laminate subjected to axial loading, i.e. along the fibre length or fibre axis (Fig. 1.33). Fibre breakage takes place whenever the applied load reaches fibres yielding or maximum strength. Since the fibre breakage is brittle, no yielding and flat broken surface observed. Fibre Buckling This mode of failure takes place when a laminate is subjected to compressive forces (Fig. 1.34). This failure mechanism starts in those areas where fibre misalignment is present or when the applied load reaches the maximum compressive strength of the fibre. It is also called fibre kinking. Moreover, this is the dominant failure mechanism involved in a laminate under compressive load.

38

1 Introduction to Fibre Reinforced Polymer … Loading along the fibre direction

Fibre

Fibre breakage

Matrix

Fig. 1.33 Schematic illustration of fibre failure in FRP composite under tensile loading

compressive load

fibre

fibre buckling

matrix

Fig. 1.34 Schematic representation of fibre buckling in FRP composites under compressive loading

Fibre Bending Here, fibre failure takes place under flexural loading that when a load acts perpendicular to fibre length (Fig. 1.35). The extent of bending of the laminate depends on the

1.12 General Failure Mechanism of Fibre Reinforced … Fig. 1.35 Schematic illustration of fibre bending in FRP composites under out-of-plane flexural loading

39

loading direction fibre bending

loading roller

fixed supporting rollers

stiffness of the laminate or stiffness of the fibre. Laminate bending or flexure property of the laminate depends upon the through-thickness property of the laminate. Fibre Splitting This type of failure mode observed when the laminate subjected to transverse or circumferential loading (Fig. 1.36). Whenever the load applied is perpendicular to that of the radius of the fibre, this failure type of mechanisms can be observed.

matrix region fibre

Fibre Splitting

Fig. 1.36 Schematic illustration of fibre splitting in FRP laminate under circumferential loading

40

1 Introduction to Fibre Reinforced Polymer …

Fig. 1.37 Schematic illustration of fibre pull-out in FRP composite under tensile loading

tensile load

fibrepull-out

tensile load

Fibre Pull-out Fibre pull-out failure mode takes place when fibre slips from the matrix (Fig. 1.37). This mode of damage is majorly due to the weak interfacial bonding between fibre and matrix under tensile loading.

1.12.2 Matrix Damage Mechanism Matrix cracking plays a vital role in deciding the crack growth direction. This crack growth may be in the transverse direction or longitudinal direction. If the matrix material involves transverse direction failure, then most probably the breakage of the fibres will be the primary failure mode associated. Whereas, if the crack growth in the matrix materials is longitudinal in direction, then the interfacial failure is the dominant failure mechanism observed. Matrix Cracking Matrix failure takes place when stress applied exceeds the strength of the matrix. Usually, there are two kinds of cracks developed, i.e. perpendicular to the fibre or parallel to the fibre (Fig. 1.38). Parallel stresses are mainly responsible for fibre– matrix interface failure while perpendicular cracks mainly cause fibre breakage.

1.12 General Failure Mechanism of Fibre Reinforced …

41

transverse matrix cracking

longitudinal matrix cracking

crack growth along through thickness of the fibre

Fig. 1.38 Schematic illustration of different types of matrix cracks in FRP materials

From either of the failure modes, parallel damage mechanism causes a severe reduction in strength of the laminate than the perpendicular cracks. Meantime, perpendicular cracks sometimes can go undetected, which is not a desirable property of any materials from the safety point of view. Interfacial Cracking This failure starts from the weakest point between the matrix and fibre. Here, crack from matrix region grows parallel to the fibre length and results in debonding of the fibre–matrix interface leading to interfacial failure, which will be decided by the chemical and mechanical bonding between the matrix and fibre (Fig. 1.39). Moreover, this type of failure modes is avoided to a certain extent by fabricating less defected laminate.

Fibre matrix Interface Failure

Fig. 1.39 Schematic representation of fibre–matrix interfacial failure in FRP composite

42

1 Introduction to Fibre Reinforced Polymer …

1.13 Summary The FRP composite laminates used in wide range of applications due to their lightweight and high strength. The FRP composite materials are basically made of two constituent materials: one is the fibre material the primary load-bearing member and second is the polymer matrix which transfers the applied to the fibre. There are different fibre and polymers materials which are used based on the applications. Along with type of fibre and polymer matrix, the FRP composite laminate material properties are also affected by laminate stacking design and fibre weaving architecture. Further, the FRP composite laminate properties can be improved by using nanomaterials. These FRP composite laminates can be fabricated from different fabrication processes. However, the choice of fabrication process strongly depends on the geometry and end application of the FRP composite laminates. The rule of mixture and classical laminate plate theory methods are used to predicted the lamina and laminate properties of these fabricated composite laminates. It is now evident that the under different loading condition, the FRP composite laminate fails differently. However, the matrix cracking is the first failure mode in the composite laminate.

1.14 Further Reading More detailed explanation regarding the fibre reinforced polymer composites and mechanics of fibre reinforced polymer composites could be found in the following books. 1. 2. 3. 4. 5.

An introduction to composite materials by D. Hull and T. W. Clyne Analysis and performance of fibre composites by B. D. Agarwal and L. J. Broutman Mechanics of composite materials by R. Jones and R. M. Jones Mechanics of composite materials by Autar Kaw Fibre-reinforced composites: materials, manufacturing, and design by P. K. Mallick.

References 1. Davies P (2016) Behavior of marine composite materials under deep submergence. In: GrahamJones J, Summerscales JBT-MA of AF-RC (eds) Woodhead Publishing series in composites science and engineering. Woodhead Publishing, pp 125–145. https://doi.org/10.1016/B978-178242-250-1.00006-5

References

43

2. Uddin N, Cauthen S, Ramos L, Vaidya UK (2013) Vacuum assisted resin transfer molding (VARTM) for external strengthening of structures. In: Uddin NBT-D in F-RP (FRP) C for CE (ed) Woodhead Publishing series in civil and structural engineering. Woodhead Publishing, pp 77–114. https://doi.org/10.1533/9780857098955.1.77 3. Goldberg B, Hsu JD (1997) Atlas of orthoses and assistive devices, 3 edn. Mosby, St. Louis

Chapter 2

Mechanical Testing of Fibre Reinforced Polymer Composite Laminates

Abstract FRP composite materials are promising materials for lightweight structures. However, careful evaluation of strength is an important task. However, testing of FRP composite laminate is difficult along with achieving good repeatability. Because, the FRP laminates are highly anisotropic and heterogeneous. Thus, each experimental test results can deviate from each other, which may not provide repeatability and reliability. Moreover, the mechanical properties of these materials are more susceptible to microcracks and environmental conditions. More importantly, the material properties significantly influenced by the manufacturing method. On the whole, careful evaluation of these materials is a tedious and challenging task. Thus, in this section, a few common testing procedures are explained briefly which are conducted according to the ASTM standards so that basic understanding of FRP composite laminates mechanical testing can be developed. Further, few basic failure modes observed under different loading condition are also covered briefly. Since, specimen preparation for testing is also an important aspect of mechanical testing. Thus, this chapter includes a section where preparation of FRP composites for different mechanical testing is explained. However, there are different testing standards for FRP laminates. Here, only ASTM standards are referred and explained. For more details, authors of this book request the reader to go through the corresponding ASTM standards for more details. Keywords FRP composites · Mechanical testing · ASTM Standards

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite Laminates According to ASTM Standards 2.1.1 Measurement of Constituent Content in FRP Composite Laminates The FRP composite laminate constituents such as fibre, matrix and void are measured according to ASTM D3171 standard [1]. There are two methods to determine the © The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_2

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constituent content of FRP laminate composites. In the first method, the polymer matrix content is completely removed by digestion or ignition or carbonization, while the second method uses weight and volume. However, from second method it is not possible to calculate the void content in the FRP composites. There are several approaches for conducting first method where the matrix content is removed either by dissolution or combustion. For dissolution, a liquid medium such as concentrated nitric acid (matrix made of epoxy), aqueous mixture of sulphuric acid and hydrogen peroxide (matrix materials made of epoxy or phenolic or polyamide or thermoplastic), mixture of ethylene glycol and potassium hydroxide (epoxy) is used, while for combustion a furnace is used. All the chemical digestion methods are preferred for thermoplastic polymer composite laminates, while the combustion process is preferred for thermoset polymer composite laminates. For first procedure, the specimen mass selected for testing should be between 0.5– 1.0 g. Test samples weight before and after the testing should be noted or measured. For second procedure, the specimen for testing should be 625 mm2 . Before carrying the test, the weight of the testing specimen should be noted. Procedure for specimen digestion in nitric acid The test specimen should be placed in a flask with concentrated nitric acid. Then, the flask or beaker with concentrated nitric acid and test specimen are heated constantly not exceeding 80 °C for not less than six hours or till all the matrix material disappears into the nitric acid. After that the remaining fibre reinforcement is filtered out carefully and washed several times with water. Further, final washing of the remaining reinforcement should be done by acetone. Then, the washed left out reinforcement dried in an oven at 100 °C for one hour or till the reinforcement dries out completely. Now, measure the reinforcement weight using crucibles and calculate the constituents according the formulae as shown below. Procedure for specimen digestion in the mixture of sulphuric acid and hydrogen peroxide Add the specimen to a beaker and add a minimum quantity of 20 mL sulphuric acid. Then heat the beaker till fumes are seen. Once the solution attains constant colour, then add hydrogen peroxide into the beaker along the side walls of the beaker. Once all the matrix is digested by the acid mixture, stop heating the solution and allow it cool to room temperature. Now, filter and wash the left-out reinforcement several times with water and use the acetone for final washing. Heat the reinforcement in an oven for one hour at 100 °C. Once the fibres are completely dried use the crucibles to measure the weight of the reinforcement and use the formulae given below to calculate the constituent contents. Procedure for specimen digestion in the mixture of ethylene glycol and potassium hydroxide Prepare the ethylene glycol and potassium hydroxide solution in a beaker and place the specimen into the mixture. Now, connect the beaker to reflux condenser and apply the heat to boil the solution till all the matrix is digested. Filter the reinforcement

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

47

and wash two times with dimethylformamide then several times with water. Final washing should be done with acetone. Dry the reinforcement by placing in an oven at 100 °C for one hour. Then weigh the reinforcement in using a crucible. Use the measured values for calculating the constituent contents according to the equation discussed in the subsequent sections. Matrix burn-off procedure Measure the geometrical dimensions and weight of the testing specimen. Place the specimen in a crucible and heat the specimen using a muffle furnace. The temperature range should be between 500 and 600 °C or depending on the matrix decomposition temperature. Wait till all the matrix burns out leaving behind the traces of reinforcement. Take the crucible from the furnace and measure the weight of the reinforcement. Now, use the formulae below for further calculations. Calculation of Reinforcement Content in Weight Percent  Wr =

Mf Mi

 × 100

where W r = Weight percent of reinforcement; M i and M f = Initial and final (after digestion or combustion) mass of the test specimens respectively in grams. Calculations related to first method. Calculation of Reinforcement Content in Volume Percent  Vr =

Mf Mi



 ×

ρc ρr

 × 100

where V r = Volume percent of reinforcement; ρc and ρr = Density of composite and reinforcement, respectively in g/cm3 . Calculation of Matrix Content in Weight Percent  Wm = where W m = Weight percent of matrix.

Mi − M f Mi

 × 100

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2 Mechanical Testing of Fibre Reinforced Polymer …

Calculation of Matrix Content in Volume Percent  Vm =

Mi − M f Mi



ρc ρm

 × 100

where V m = Volume percent of matrix; ρm = Density of matrix in g/cm3 . Calculation of Void Volume Vv = 100 − (Vr + Vm ) where V v = Void volume in mm3 . Calculation of Specimen Density in (g/cm3 ) ρc =

Mi specimen volume

Calculations related to second method. Calculation of Reinforcement Content in Weight Percent Wr =

Ar × N p × 0.1 ρc × h

where Ar = Mass of one sheet of reinforcement/unit area in g/m2 ; N p = Number of plies in the test sample; h = Test specimen thickness in mm. Calculation of Reinforcement Content in Volume Percent Vr =

Ar × N p × 0.1 ρr × h

Calculation of Matrix Content in Weight Percent  Wm = 100 −

Ar × N × 0.1 ρr × h



2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

49

Calculation of Matrix Content in Volume Percent Vm = Wm ×

ρc ρm

Other ASTM standards used for determining the constituent content in FRP composite laminates are D3529 and D3530 [2, 3]. The ASTM D3529 is used for FRP composite prepreg, where the constituent content is determined in two methods. In first method, the FRP composite prepregs (mostly thermoset FRP composite prepregs) are dissolved in an organic solvent, where the matrix is soluble and reinforcement is insoluble. The second method involves ignition loss of FRP prepreg where the reinforcement mass should not change upon exposure to heat. If the FRP composite laminate contains any volatile content, then ASTM D3530 standard is used. According to the ASTM D3530 standard, the volatile content in the FRP prepreg composite is measured after the cure or consolidation temperature. However, the temperature limitation for this standard is 300 °C.

2.1.2 Tensile Test The tensile test is conducted to determine the in-plane tensile strength property of a given FRP composite specimen according to ASTM D3039 [4]. Here, a constant monotonic load is applied to the sample using a universal testing machine until the specimen breaks or fails. The ultimate tensile strength of the FRP composite laminate is equal to that is responsible for fibre breakage. Along with various other parameters, the specimen tabbing plays a crucial role in getting the accurate tensile properties and failure modes of the FRP composite laminates. The tabbing thickness should be less than that of the specimen thickness. The sample geometry used in this standard is a flat rectangular coupon with overall specimen length of 250 mm and width of 25 mm along with 50 mm end tab length (Fig. 2.1). In this testing procedure, the geometry is simple because the FRP composites consist of plies where fibres’ orientation may range from 0° to 90°. If the specimens are cut according to a dog bone shape, then it introduces unnecessary defects and stress concentration to the specimen which may

sample width

tab thickness

specimen thickness tab length

span length

Fig. 2.1 Schematic representation of FRP tensile test specimen with end tabs

tab wedge angle

adhesive thickness

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2 Mechanical Testing of Fibre Reinforced Polymer …

not yield valid data points. Thus, samples are flat coupon geometry with end tab. Detailed explanation and importance of the end tabs for testing FRP composites are given in the subsequent sections. The ultimate tensile strength of the tested specimen is calculated according to the Eq. (2.1). UTS =

Pmax A

(2.1)

where UTS = Ultimate tensile strength in MPa; Pmax = Maximum force before failure in N; A = Average cross-sectional area of the test coupon in mm2 . There are other ASTM standards used to conduct the tensile test of FRP composite laminates which include D638, D5083 and D5450/D5450M [5–7]. ASTM D638 is used for conducting tensile test for dumbbell-shape specimen geometry (Fig. 2.2a). The specimen thickness is limited to 14 mm only. The stress concentration at the radii region of the specimen geometry is highly unfavourable in which the specimen consists of highly oriented plies. For same specimen dimensions, the tensile creep test is conducted according to ASTM D2990 standard. Further, the specimen used in this standard according to D638 is not suitable for continuous FRP composite laminates. In that case, the specimen geometry should be prepared according to the ASTM D3039 standard. (a)

grip

grip span length

(c)

anchor for gripping

span length

(b)

(d)

test specimen bar

span length

grip

(e)

pin

pin bearing strength

metal block adhesive layer test specimen test specimen

bearing strength

Fig. 2.2 Schematic representation of FRP composite laminate test specimen according to various ASTM standards a FRP bar specimen according ASTM D693, b cylindrical tube specimen according to ASTM D5450/D5450M, c anchors used for bar specimen used in civil construction according to ASTM D7205/D7205M, d flat FRP specimen used for tensile testing according ASTM D7291/D7291M e comparison of bearing strength test specimens according to ASTM D953 and ASTM D5961/D5961M

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

51

Similarly, ASTM D5083 standard is used to conduct the tensile testing of thermosetting reinforced polymer composites without using tab. This standard is limited mostly to glass fibre reinforced thermoset polymer composites with limited specimen thickness of 14 mm fabricated by injection moulding or compression moulding process. The width of the test specimen should be 25 mm, total specimen length should be 250 mm, and thickness of the specimen may vary between 2 and 14 mm. For conducting tensile test for cylindrical geometry, ASTM D5450/D5450M is used (Fig. 2.2b). This test standard is limited to hoop wound FRP composite cylinders where only transverse tensile properties are determined. Here, test specimens are glued to the grip or fixture to conduct the tensile test. Further, the glue material tensile strength property should be more than the test specimen. ASTM D7565/D7565M is followed to conduct the tensile test of flat FRP composite laminates used in strengthening of civil structures [8]. Most of the testing conditions used in this testing standard are similar to that of ASTM D3039 standard. ASTM D7205/D7205M standard is used to test the tensile strength of FRP composite bars or rebars used in concrete [9]. Similar to the tabbing used for flat coupons, here also anchors are used to prevent the damage induced by grip (Fig. 2.2c). The standard bar specimen cross-sectional area should be similar to the steel bar used in concrete according to the ASTM A615/615M standard. The total bar length should be free length and two times the anchor length. The free length of the bar should not be less than 380 mm. The ultimate tensile strength of FRP bar is calculated according to the Eq. 2.2. ASTM D7291/D7291M is used to determine throughthickness tensile strength of FRP composite flat specimen (Fig. 2.2d). In this test standard, the specimens are glued to thick metal end table. UTS =

Pmax A

(2.2)

where UTS = Ultimate tensile strength in MPa; Pmax = Maximum force prior to specimen failure in N; A = Cross-sectional area of the specimen in mm2 . The tensile test of a specimen with a hole at the centre of the span length is conducted according to ASTM D5766 which is similar to that of ASTM D3039/D3039M standard [10]. Now, the tensile and compression of a test specimen with filled hole are conducted according to the ASTM D6742/6742M standard [11]. The test can be extended to test the close tolerance fastener or pin installed in the hole. Further, for conducting the static tensile pin-bearing test, the ASTM D953 is used [12]. In this test standard, one fastener and double shear pin bearing are used. Furthermore, ASTM D5961/5961M standard is used to conduct the static bearing tensile test for FRP composite laminate [13]. In this test, one and two fastener double with one shear test specimen is loaded in tensile loading. The difference between pin bearing and bearing tensile test is shown in Fig. 2.2e. For conducting the tensile creep rupture test for FRP bar specimen, the ASTM D7337/7337M standard is used [14]. The specimen geometry for this test procedure is similar to that of the ASTM D7205/D7205M standard.

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2.1.3 Compression Test The compression test of FRP composite laminates provides the in-plane compressive strength of a laminate when subjected to in-plane compressive loading. There are different ASTM standards to conduct the compression test where the major difference lies with the FRP specimen grip attachment. Conducting compression test is comparatively more difficult than the tensile test due to various reasons. The compression tests of FRP composite laminates are conducted on a universal testing machine by applying a monotonic load. The compression test can be conducted according three ASTM standards D695 or D3410/3410M or D6641/D6641M [15–17] (Fig. 2.3). The ASTM D695 standard is used for conducting compression test on reinforced rigid plastics at low strain rate. Here, a supporting jig is used for thin specimens to avoid the buckling in specimen. Specimen geometry is dog bone (Fig. 2.4a). The applied load is transferred to the test specimen by end loading. If the specimen thickness is between 3.2 and 6.4 mm, then the specimen dimensions should be 12.7 mm × 12.7 mm × 12.7 mm. If the specimen shows buckling during testing, then the height of the specimen should be reduced. Further, if the preferred specimen geometry is cylindrical, then the specimen dimensions become 12.7 mm × 12.7 mm × 25.4 mm for prism and 12.7 mm (diameter) × 25.5 mm (height) for cylinder. To avoid the specimen failure by buckling phenomenon, the slenderness of the specimen should be in the range of 11 to 16:1. Then according to this criteria, the specimen dimension changes to 12.7 mm × 12.7 mm × 50.8 mm for prism and 12.7 mm (diameter) × 50.8 mm (height) for cylinder specimen. For rod and tubes, the specimen length should be twice and 25.4 mm, respectively. This testing standard is not suitable if the

Fixed End

Front View

Side View

(b)

gauge length

FRP sample

h

tab length

loading direction

Width = 6-25mm

specimen length = 140-155mm

(a)

Flat FRP sample

Fig. 2.3 Schematic illustration of FRP compression testing sample a block specimen under loading and b rectangular coupon

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

53

(a) grip

(b)

grip

span length

ASTM D3410/D3410M

ASTM D6641/D6641M

Fig. 2.4 Compression test specimen geometry and fixtures a Schematic representation of compression test specimen according to ASTM D695, b difference in the fixture used in conducting the compression test according to ASTM D3410/D3410M and D6641/D6641M

laminate consists of high ply orientations. Similarly, ASTM D3410/3410M is used to conduct the compression test on FRP composite laminates. Here, the specimen geometry is flat rectangular coupon. The applied compressive load is transferred to the test specimen by shear at the interfaces of wedge grip using a specially designed fixture. Specimens with and without end tabs can be tested. Further, no supports are used in the gage section during the testing to avoid the buckling. Another and preferred ASTM standard for compression testing is the ASTM D6641/6641M. In this testing standard, the applied load is transferred to the test specimen by combined shear and end loading using a specially designed fixture (Fig. 2.4b). The surface of the fixture grip assembly should be smooth, polished, lubricated and nick-free. The wedge grip is used for flat rectangular tape like geometry, and end loading is used for small cylindrical-shaped specimens. The testing flat rectangular coupons can be tested either with or without end tabs. However, flatness of the tab along alignment of the tab with the test specimen influences the end test results data. Hence, care should be taken with respect to the tab alignment with the test specimen and flatness of the tab. Here, the test specimen thickness and gauge lengths are not fixed and selected in such a way that the selected gauge length must be free from Euler Buckling, and applied stress should be transferred axially to the specimen when the load is applied. Generally, the test specimen span length is around 10 mm. The ultimate compressive strength is calculated using Eq. 2.3. For testing transverse compressive strength of FRP composite cylinders, ASTM D5449/D5449M standard is used [18]. This test standard is limited to hoop wound (90°) cylinders. Further, during testing the cylindrical test specimens must be glued to the grips or fixtures.

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2 Mechanical Testing of Fibre Reinforced Polymer …

SC =

Pmax A

(2.3)

where S C = Compressive strength in MPa; Pmax = Maximum force before failure in N; A = Cross-sectional area of test specimen in mm2 . For conducting open-hole compression test, the ASTM D6484/D6484M standard is used [19]. For filled open-hole compression test, the ASTM D6742/D6742M is used. For testing pin bearing and bearing joints of FRP composites under static compression loading, ASTM D953 and D5961 standards are used.

2.1.4 Flexural Test This test is conducted either on a three-point bend or on a four-point bend test according to ASTM standard D7264/7264M [20] (Fig. 2.5). The testing provides the deflection behaviour of the specimen under the applied load. In three-point bend test, the specimen is subjected to centre point load on a simply supported beam, while in four-point bend test, the specimen is subjected to two-point load on a simply supported beam. Further, in three-point load the loading point is precisely at the midsection of the span or specimen length and located at equidistance from end supports. Similarly, in four-point bend test the two-point loads are located at equidistance from mid-section of the test specimen and also located at equidistance from the support end. Another major difference between three and four-point bend test is the formation of specimen length

width = 13mm

thickness

loading roller Over hanging length 10% of the specimen length

loading rollers

specimen

support roller

Fig. 2.5 Schematic illustration of flexure test under three- and four-point loading

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

55

three point bending

four point bending

shear force diagram

shear force diagram

bending diagram

bending diagram

Fig. 2.6 Schematic representation of shear force and bending diagrams of a test specimen subjected to three- and four-point bending

maximum bending moment and maximum flexural stress. In three-point bend test, the bending moment and flexural stresses are concentrated around the loading nose in the specimen, whereas in four-point bend test, the flexural stresses and bending moment are distributed uniformly throughout the specimen (Fig. 2.6). For flexural testing, both flat and curved rectangular coupons can be used. However, the test geometry should be such a way that no out-of-plane shear deformation and associated failure modes occur in the test specimen. The major standard span-to-thickness ratio is 32:1, but it can be varied according to 16:1, 20:1, 40:1 and 60:1. Further, the selected span-to-thickness ratio should produce failure at the outer surface of the specimens due to bending moment. Total specimen length is the sum of 20% of support span and specimen span length. The width of the specimen is 13 mm irrespective of the span-to-thickness ratio. The flexural strength of the tested specimen is calculated using Eqs. (2.4) and (2.4) for three- and four-point bend test, respectively. Another standard used for three-point flexural testing is ASTM D790 [21]. Moreover, same test fixture and specimen are used to conduct the flexure creep test by following ASTM D2990 standard. σ =

3P L 2bh 2

(2.4)

σ =

3P L 4bh 2

(2.5)

where σ = Flexural strength in MPa; P = maximum applied load in N; b = width of the test specimen in mm; h = thickness of the test specimen in mm.

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2 Mechanical Testing of Fibre Reinforced Polymer …

loading end

Curved test specimen

fixed end Fig. 2.7 Schematic representation of four-point bend test of a curved test specimen

ASTM D6272 is another standard used for conducting four-point bend test for reinforced insulating polymer composites [22]. The curved FRP composite specimen strength is determined using ASTM D6415/D6415M [23]. In this test standard, the curved FRP composite laminate test specimen is loaded into a four-point bend test fixture and applied a compressive load (Fig. 2.7). The angle between the two arms of the curved test specimen should be 90°. The scatteredness in the test data is high.

2.1.5 Short Beam Shear Test This test is conducted according to the ASTM D2344/D2344M standard [24]. Here, the test specimen is placed on a two-point simply support and loaded by single point. In this test, the specimen geometry can be flat or curved (Fig. 2.8). From this test, interfacial and shear failure in a laminate is evaluated. The test specimen geometries are limited, and the test specimen at least should have a minimum span length-tospecimen thickness ratio of 4 with minimum specimen thickness of 2 mm. The total specimen length is equal to the span length and 10% of the span length. Span length and width of the test specimen are calculated according Eqs. (2.6) and (2.7). Further, for curved specimens, the arc should be within 30°. The shear strength of the tested specimen is calculated according to Eq. (2.8). Specimen span length = thickness × 6

(2.6)

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

57

loading direction curved specimen

nose with 6 mm diameter flat specimen support roller with 3mm Diameter 12mm span length

specimen length

Fig. 2.8 Schematic illustration of three-point bend test method

Specimen width = thickness × 2 SBS =

0.75 × Pm b×h

(2.7) (2.8)

where SBS = Short beam shear strength in MPa; Pm = Maximum load at failure in N; b = width of the test specimen in mm; h = thickness of the specimen in n = mm.

2.1.6 Impact Test ASTM D7136/D7136M standard is used to determine the impact or damage resistance of a laminate under out-of-plane loading [25]. This ASTM standard is the most commonly and widely used for determining the impact damage resistance of FRP composite laminates subjected to low velocity impact or inducing the damage into the FRP composite laminate for further subsequent testing. Moreover, it provides reallife impact situations and helps in understanding the wave propagation in a laminate under impact loading. In this test method, the laminate is clamped with a specified boundary condition (Fig. 2.9). Then according to the required impact energy, the impactor mass and height is adjusted. Now, the impactor is dropped onto the specimen from guided rail (Fig. 2.9). The force, displacement and time data are collected for analysis. Further, non-destructive inspection is carried to determine the internal damage and extent of failure. Both boundary condition and impactor shapes may

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2 Mechanical Testing of Fibre Reinforced Polymer …

guiding column hemispherical tip impactor circular boundary condiƟon

specimen dimension

150 mm

rigid clamp corner clamped boundary condiƟon

100 mm laminate

Fig. 2.9 Schematic illustration of drop weight impact machine and general boundary condition used during the testing

vary. The test specimen geometry is square with length 100 mm and width 100 mm. The nose geometry is hemispherical with 18 mm diameter. Calculation of impact velocity vi =

  W12 t1 + t2 + g ti − t2 − t1 2

where vi = Impact velocity in m/s; W 12 = Length between the lower and upper flag prongs in m; g = acceleration due to gravity t 1 = time measure by the lower flag prong in seconds; t 2 = time measured by the upper flag prong in seconds; t i = time at impactor contact with test specimen in seconds. Calculation of impact energy Ei =

mvi2 2

where E i = measured impact energy in J; m = mass of the impactor in kg. Calculation of velocity versus time t

v(t) = vi + g.t − ∫ 0

F(t) dt m

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

59

where. v(t) = Impactor velocity at time ‘t’ in m/s; t = time during the test in seconds; F(t) = Force at time in N. Calculation of displacement versus time   t t F(t) g.t 2 d(t) = di + vi .t + −∫ ∫ dt dt 2 0 0 m where d(t) = impactor displacement at time ‘t’ in mm; d i = impactor displacement from reference location when time t = 0; Calculation of absorbed energy versus time   m vi2 − v(t)2 E a (t) = + m.g.d(t) 2 Izod Method ASTM D256 is used to conduct the test where specimen impacted by the swinging arm of a pendulum and the positioning of the specimen in the fixture resembles the cantilever beam (Fig. 2.10) [26]. There are various methods to conduct this test. In the first method, the test specimen is placed similar to a vertical beam. The notch in the specimen is facing towards the impactor and does not consider toss energy of the impacted sample. While, the second method is similar to the first method; however, the second method considers the toss energy. The third method estimates the notch sensitivity measured on the impact strength of FRP specimen by varying the radius of the notch. The fourth method is similar to the first method; however, the notch

10.16±0.05 mm specimen

notch angle = 24.50±0.0 mm

pendulum arm striking direction 31.8±1.0 mm

fixed jaw

moveable jaw 63.5±2.0 mm

Fig. 2.10 Schematic representation of a Izod impact testing b test specimen geometry

12.7±0.2 mm

(b)

(a)

60

2 Mechanical Testing of Fibre Reinforced Polymer …

face is placed away from the pendulum arm. Creation of notch in the test specimen increases the probability of brittle failure of the specimen and reduced the plastic deformation. ASTM D4812 is used to measure the impact strength of the un-notched plastics [27]. Here, the specimens are kept similar to cantilever beam as that of the ASTM D256 standard. This testing standard is useful for reinforced specimens because the test specimens are un-notched. Notching may mask the influence of orientation thus creating the discrepancy in the test results. The depth, length and width of the test specimen are 12.7 mm, 63.5 mm and 3.17 mm, respectively. Further, the width may vary, but it should be equal to the striker’s face or width. Charpy Method This testing method is another way of testing the impact resistance of a material. In this method, ASTM D6110-17 followed in conducting the experiment [28]. The only difference between Izod and Charpy test is the position of the specimen during testing (Fig. 2.11). In Izod, the sample is placed as a cantilever beam while in the Charpy test, the sample is positioned as a simply supported beam. In this test, a notch is created onto the specimen using milling and engine lathe. The notch either can be perpendicular to the impactor or parallel to the impactor. Other ASTM Standards for impact testing ASTM D5420 standard is used to conduct the impact test on rigid plastic using a falling weight method [29]. In this method, the impactor with required or specified weight is dropped from a specified or required height onto the specimen. The impactor movement is guided through a rail or tube. The applied impact energy either should generate crack or break the flat rigid plastic. If the cracks are not visible in the test specimen after impact test, then crack developing dye agents are applied. The specimen geometry is either square or circular plate with at least 25 mm dimension or at least greater than the hole diameter of the fixture. The impactor geometry is cylindrical with round nose. Another ASTM standard used to determine the impact resistance of flat and rigid plastic specimen under free fall dart impact is the D5628 [30]. In this testing standard, the dart is dropped onto the specimen without any (b)

22.50

10.16±0.05 mm

12.7±0.15 mm

(a)

test specimen

loading direction 63.5 mm 127 mm

Fig. 2.11 Schematic illustration of a Charpy impact testing and b test specimen geometry

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

61

guided rail or support. The impactor nose shape is either conical or hemispherical. Both diameter of the impactor and specimen dimensions varied. ASTM D3763 is used to test the high-speed puncture resistance properties of rigid plastics [31]. The test specimen length and width should be greater than the inside diameter of the circular clamp fixture. In this standard load and displacement sensors, a circular clamp with hole of diameter 76 mm is used to clamp the test specimen. ASTM D8101/D8101M standard is used to measure the penetration resistance of flat FRP composite panels with blunt impact under free falling weight condition [32]. This standard specially adopted if the thickness of the FRP composite test specimen is small compared to width and length such that the span-to-thickness ratio is greater than 50. The impact velocity used in this standard ranges between 100 and 500. The clamping fixture has a circular geometry with a centre hole. The testing specimens are clamped using 28 through bolts. A gas gun is used to generate the impact required impact velocity. Unlike guided impacting in this test standard, the impactor flight not guided through any rail or support. The thickness of the specimen ranges between 2 and 4 mm.

2.1.7 Compression After Impact Test This test investigates the residual compressive strength of FRP composite laminates after subjecting to impact. The test is conducted according to ASTM D7137/D7137M and on a uniaxial compression test machine [33]. However, before testing, the samples are subjected to impact loading to introduce residual stresses into the laminate which can be done either by quasi-static indentation or drop weight impact test according to their respective ASTM standards D6264/D6264M and D7136/D7136M, respectively. The specimen geometry is flat rectangular plate. This testing method only provides the residual compressive strength of impact damaged specimens. Further, to determine the compressive strength of undamaged specimens, the ASTM standards such as D6641/D6641M should be used. For conducting compressive after impact testing a specially designed fixture is used (Fig. 2.12). The compressive strength of the specimen after the impact is measured according to the Eq. (2.9). SCAI =

Pmax A

(2.9)

where S CAI = Compressive strength after impact in MPa; Pmax = Maximum load before failure in N; A = Cross-sectional area of the test specimen in mm2 .

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2 Mechanical Testing of Fibre Reinforced Polymer … undamaged specimen

laminate

impactor Compression loading direction

impact damage induced laminate

damage

fixture

Fig. 2.12 Schematic representation of steps involved in compression after impact test

2.1.8 Double Cantilever Beam Test The double cantilever beam test method is used to determine the Mode I interlaminar fracture toughness for FRP composites laminates. The ASTM D5528 standard is used for determining the interlaminar fracture toughness (Fig. 2.13) [34]. In this method, an artificial crack is inserted into the FRP laminate as a crack initiator. Generally, the thin films (thickness = 13 µm) such as polytetrafluoroethylene (PTFE) for lowtemperature curing matrix and polyimide for high-temperature curing matrix are used as inserter. Then, either piano hinges or loading blocks are glued onto the FRP composite laminates towards the crack inserted side. Now, the prepared specimen is loaded into the universal testing machine and applied the tensile loading. The crack Sample length125 mm thickness 20 – 25 mm

crack initiator applied load

crack propagation applied load

Fig. 2.13 Schematic illustration of Mode I testing

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

63

initiator must be thin, free from tear fold and should not dis-bond from the laminate. Otherwise, the abrupt jumps observed in the obtained results. The initial load can be applied either by a piano hinge or through loading blocks. During testing, two types of failures are observed: one being cohesive-type failure and the other being adhesive-type failure. The specimen dimensions of the test specimen should be at least 125 mm in length and 20–25 mm in width. The interlaminar fracture toughness is calculated by three different method such as modified beam theory, compliance calibration and modified compliance correction. It is still not clear which method is more justifiable to calculate the interlaminar fracture toughness. However, the modified beam theory provide more conservative values, thus, preferred to calculate the interlaminar fracture toughness of FRP composite laminate. For investigating the Mode I, fatigue delamination initiation under cyclic or fatigue loading ASTM D6115 standard is followed [35]. Modified beam theory method GI =

3Pδ 2ba

where. GI = Interlaminar fracture toughness in kJ/m2 ; P = Applied load in N; δ = load point displacement in mm; b = Width of the test specimen in mm; a = delamination length in mm. Since the above formula overestimates the interlaminar fracture toughness value, a correction factor ‘’ is added. The ‘’ is calculated by generating least square curves using compliance (C) along y-axis and delamination length (a) along x-axis. After adding the compliance correction the modified beam theory formulae changes according the Eq. (2.10). GI =

3Pδ 2b(a + ||)

(2.10)

Compliance calibration method

  Here, least square curves are plotted with log Pδii along y-axis and log(ai ) along x-axis. Using the slope of the graph to calculate ‘n’. Now, the interlaminar fracture toughness can be calculated according to Eq. (2.11). GI =

n Pδ 2ba

(2.11)

Modified compliance calibration method In this method, the least square curve is plotted with compliance correction on x-axis and delamination length normalized by thickness (a/h) along y-axis. The slope of

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2 Mechanical Testing of Fibre Reinforced Polymer …

this curve ‘A1 ’ is used to calculate the interlaminar fracture toughness according to Eq. (2.12). 2

GI =

3P 2 C 3 2 A1 bh

(2.12)

2.1.9 End Notch Flexural Test The end notch flexural test is used to investigate the Mode II interlaminar fracture toughness of materials. ASTM D7905/790M is used for determining the Mode II interlaminar fracture toughness of FRP composite laminate [36]. The test is conducted according to three-point bend method on a universal testing machine, where the specimen is inserted with a non-adhesive crack initiator (Fig. 2.14). The specimen geometry is flat rectangular coupon. Unlike in double cantilever beam testing method, here no piano hinges or loading blocks are required. Further, the FRP laminate should have a symmetric lay-up. The test specimen length should be minimum 160 mm, while test specimen width should be in the range from 19 to 26 mm. The crack insert criteria is similar to that used in Mode I or double cantilever beam testing procedure. The test specimen edge should be coated with white or silver paint or any other paint or coating where the delamination or crack propagation should be visible to the naked eyes. A light marking should be done on the coating within the crack insertion at a distance of 20 mm, 30 mm and 40 mm. Now the loading and unloading should be done onto the test specimen at mark 20 mm, then at 40 mm and finally at 30 mm. Using these loading conditions, load vs displacement curve should be plotted and join the each points using a linear regression line. Since, the calculation related to the Mode II fracture roughness is complex; hence, authors sincerely request the readers to refer ASTM D7905/7905M standard for more details. Calculation of Compliance Using intercept and slope from the linear curve fitting obtained by joining compliance calibration, calculate the compliance according to Eq. (2.13).

crack initiator

hemispherical nose loading roller

sample

thickness supporting rollers

Fig. 2.14 Schematic illustration of end notch flexure testing of FRP composite laminate

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

C = A + ma 3

65

(2.13)

where C = Compliance; A and m = Intercept and slope of the linear regression curve respectively obtained from compliance calibration. a = crack length cubed. Calculation of Mode II fracture toughness GQ =

2 a02 3m Pmax 2B

where GQ = Test specimen Mode II interlaminar fracture toughness; m = slope of the linear fit of compliance versus crack length cubed; Pmax = Maximum value of force obtained in force versus displacement curve during loading; a0 = delamination length inserted into the test specimen; B = Width of the test specimen.

2.1.10 Mixed Mode Interlaminar Fracture Test ASTM D6671/D6671M standard is used to conduct the mixed mode interlaminar fracture toughness of FRP composite laminate [37]. In this standard, the test specimen with crack initiator is loaded under bending load (Fig. 2.15). The test standard is limited to unidirectional FRP composite such as carbon fibre. The specimen geometry loading end

loading roller

test specimen

crack initiator piano hinge

supporting roller

fixed end

Fig. 2.15 Schematic representation of mixed mode fracture toughness test

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2 Mechanical Testing of Fibre Reinforced Polymer …

is flat which consists of crack inserter. The crack inserter or initiator must be thin nonadhesive film. The test specimen one end where the crack initiator is present is glued with piano hinge and applied tensile loading, whereas at the centre of the test specimen, a concentric load is applied similar to three-point bend test. For better results, proper positioning and insertion of the crack initiator is necessary. The specimen length should be at least 137 mm, while the thickness should be 25 mm. Highlight the edges of the test specimen with white paint or highlighter (water soluble type writer fluid) then mark at every 1 mm distance inside the insert section and at every 5 mm beyond the insert region. Now, fix the test specimen into the fixture and apply the load. Observe the delamination growth as it moves from marked points on the edges using a magnifying glass or microscope. Following equations are used for the calculation of Mode I, Mode II and GC . Calculation of C cal Ccal =

2L(c + L)2 E cal bcal t 3

where C cal = Calibrated compliance in mm/N; bcal = Calibrated width of the test specimen in mm; E cal = Calibrated modulus value in MPa; t = Test specimen thickness calibration in mm. Calculation of C sys Csys =

1 − Ccal m cal

where mcal = Slope of calibration curve in N/mm; C sys = System compliance in mm/N. Calculation of E 1f E1 f

 8(a0 + xh)3 (3c + L)2 6(a0 + 0.42hx)3 + 4L 3 (3c + L)2   = 16L 2 bh 3 m1 − Csys

where E 1f = Modulus of elasticity of the test specimen along the fibre direction measured in flexure in MPa; aO = Initial delamination length in mm; m = slope from the load versus displacement curve in N/mm.

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

67

Calculation of GI GI =

12P 2 (3c − L)2 (a + xh)2 16b2 h 3 L 2 E 1 f

Calculation of GII GII =

9P 2 (c + L)2 (a + 0.42xh)2 16b2 h 3 L 2 E 1 f

Calculation of G G = G I + G II Calculation GII /G G II G II = G G I + G II where GI = Mode I fracture toughness in kJ/m2 ; GII = Mode II fracture toughness in kJ/m2 ; G = Total strain energy release rate in mixed mode in kJ/m2 ; a = delamination length in mm; b = test specimen width in mm; c = length of the mixed mode bending fixture in mm; h = half thickness of the test specimen in mm; L = half span length of the mixed mode bending fixture in mm; P = applied load in N; x = crack length correction parameter. Calculation of ‘x’

 2    E 11  x = 3−2 11G 13 1+ where  = Transverse modulus correction parameter. Calculation of ‘’ √  = 1.18

E 11 E 22 G 13

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2 Mechanical Testing of Fibre Reinforced Polymer …

Fig. 2.16 Schematic illustration of translaminar fracture toughness test

loading end and direcƟon test specimen

notch

fixed end

where E 11 = Longitudinal modulus of elasticity of the test specimen measured in tension in MPa; E 22 = Transverse modulus of elasticity of the test specimen measured in tension in MPa; G13 = Shear modulus of out-of-plane of test specimen in MPa.

2.1.11 Translaminar Fracture Toughness Test The translaminar fracture toughness (K TL ) test of FRP composite laminate is conducted according to the ASTM E1922 standard [38]. The test specimen is flat which consists of a notch with two drilled holes at the end of the test specimen (Fig. 2.16). These drilled holes in the test specimens are used to grip the test specimen into the fixture. The test specimen is loaded into the machine and applied tensile loading. The damage zone at the crack tip length should be less than the total length of the crack or notch.

2.1.12 Fatigue Test Fatigue is a cyclic loading which can be applied to the specimen either in tension–tension or compression–compression or tension–compression mode. ASTM D3479/D3479M is followed in conducting the tension–tension fatigue test for FRP composite laminates [39]. This test gives the behaviour of FRP laminate under tensile cyclic loading condition. In this testing process, specimens are not notched and the amplitude of the uniaxial cyclic loading is constant. Constant amplitude is applied to

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69

the specimen in two ways, one by controlling the stress and the second by controlling strain along the loading direction. In stress-controlled process, the test specimen is cycled between minimum and maximum in-plane strength or stress with a required frequency. Similarly, in strain controlled process the test specimen is cycled between minimum and maximum in-plane axial strain at a desired frequency. The maximum and minimum stress and strain values are determined from tensile testing of test specimen according to the ASTM D3039/D3039M standard. Similar to tensile testing, the crushing of the test specimen by the machine grip is avoided by tabbing the test specimen ends. Care should be taken to avoid fatigue failure in the tab region, and this unwanted failure may be due to the improper selection of tabbing material. The sample geometry and dimension are the same as that of the tensile testing sample. For conducting tension–tension fatigue test for the specimen with open-hole at the centre of the span length, ASTM D7615/D7615M standard is followed [40]. The complete testing procedure is similar to the ASTM D3479/3479M standard, while the specimen preparation is similar to the ASTM D5766/D5766M and D6468/D6468M standards. The test specimen used in ASTM D5961/D5961M is used to conduct the bearing fatigue or cycling loading test according to the ASTM D6873/D6873M [41].

2.1.13 Lap Shear Test The tensile test of adhesively bonded lap joint FRP composite laminates is conducted according to the ASTM D5868/D5868M standard. In this standard, two FRP laminates are bonded using an adhesive [42]. The adhesive area should be 25.4 × 25.4 mm2 . The dimensions of the FRP composite used in lap joint should be 25.4 mm in wide and 101.6 mm in length. Further, the adhesive thickness should be 0.76 mm. If tabs are used, then they should be glued only one side of the each towards the adhesive side (Fig. 2.17).

2.1.14 Quasi-Static Indentation Test The damage resistance of FRP composite laminates can also be determined by applying a concentrated indentation force under quasi-static loading condition. This process is another way of investigating the damage resistance of the composite materials, and ASTM D6264 is used for experimentation [43]. Here, the specimen is clamped to a fixture, and an out-of-plane quasi-static load is applied to the test specimen (Fig. 2.18). The indenter may be hemispherical, conical and ogival. However, generally blunt or hemispherical nose with 13 mm tip diameter is preferred. The test specimen should be a flat rectangular plate with length 150 mm and width 150 mm. Fixing of the sample can be done either by edge fixing or by rigid fixing by a circular plate with hole. The boundary condition can be circular or rectangular. Unlike in the case of drop weight impact testing, the wave propagation in the specimen cannot

70

2 Mechanical Testing of Fibre Reinforced Polymer … 25.4 mm Fixed end

101.6 mm FRP laminate

Adhesive (0.76 mm thick) lap joint

25.4 mm

tensile loading

Fig. 2.17 Schematic diagram of adhesive lap joint of FRP composite laminate for tensile loading

hemispherical indenter

Specimen (size = 150 x 150) fixture circular boundary condition with diameter = 125±3

Fig. 2.18 Schematic illustration of quasi-static indentation test

be measured. The test is also used for inducing damage into the test specimen for further analysis such as compression after impact. Further, ASTM D2583 is used to conduct the indentation hardness test on FRP composite laminate [44]. The indenter tip used in this test standard is flat.

2.1 Mechanical Testing of Fibre Reinforced Polymer Composite …

71

Calculation of Energy Following equation provides the energy value at any indenter displacement. δ

E(δ) = ∫ F(δ)dδ δ0

where E(δ) = Energy at indenter displacement value in N-m; δ = Indenter displacement in mm; δ 0 = Indenter displacement at initial contact with the test specimen in mm; F(δ) = Measure contact force at any indenter displacement in N.

2.1.15 In-Plane Shear Test The in-plane shear tests on FRP composite laminates are conducted according to ASTM D3518/D3518M, D4255/4255M, D5379/5379 M, D5448/D5448M and D7078/D7078M [45–49]. ASTM D3518/3518M standard is used to measure the in-plane shear properties of FRP composite laminate consisting of only ±45° plies by applying tensile loading. Similar to ASTM D3039, here also a uniaxial tensile loading is applied onto to a FRP composite laminate made of ±45° plies. Shear strength of the tested specimen is calculated according to Eq. (2.14). The testing procedure of this standard is simple and cost-effective, thus widely used. However, the test specimen does not provide accurate in-plane shear strength. m = τ12

Pm 2A

(2.14)

where m τ12 = Maximum in-plane shear strength in MPa, Pm = Maximum force at or below 5% engineering shear stress in N, A = Test specimen cross section in mm2 . Another ASTM standard to measure the in-plane shear strength of FRP composite laminate is D4255/D4255M. This standard uses specially designed fixture to grip the specimen. The testing procedure is difficult and has poor reproducibility because huge stress concentrations are generated at the grip section of the fixture during testing. ASTM D5379/D5379M standard can also be used to determine the shear strength of FRP composite laminates by V-notched beam method. In this test standard, the FRP composite specimens are cut with v-notches on both sides symmetrically at the centre of the geometry (Fig. 2.19). Then the specimen is loaded in compressive loading condition using a specially designed fixture. Strain gages are used to measure the strain development in the test specimen during loading. The strain gages should be attached 45° to the loading axis. End tabs are recommended when the thickness

72

2 Mechanical Testing of Fibre Reinforced Polymer …

(a)

(b)

loading end

notch angle 90 0

3.8 mm

38 mm

19 mm

76 mm

fixed end thickness

Fig. 2.19 Schematic diagram of in-plane shear test specimen a geometry of test specimen, b general illustration of force application on the double v-notched specimen used in in-plane shear test

of the test specimen is 2.5 mm. This test standard provides better and reliable test results. The shear strength of the test specimen is calculated according to Eq. (2.15). Another standard used to determine the in-plane shear strength of FRP composite is ASTM D7078. Similar to ASTM D5379/D5379M, in this test standard also the specimen has v-notches at the centre and towards the test specimen edges. However, the specimen dimension varies from that of the ASTM D5379/D5379M standard. The v-notch test specimens are subjected to tensile loading using a specially designed fixture. SS =

Pm A

(2.15)

where SS = Shear strength in MPa; Pm = Maximum load before test specimen failure in N; A = Test specimen cross-sectional area in mm2 . The shear strength of hoop wounded cylindrical FRP composite is measured using ASTM D5448/5448M standard. In this method, the FRP specimen is applied a torsional load. This test method is only applicable to hoop wound FRP composite cylinders made of only 90° plies. For testing, the test specimen must be glued to the grips or fixture. Similarly, ASTM D7617/D7617M standard is used to determine the transverse shear strength of FRP bar which is used as reinforcement in concrete structures [50]. In this test method, the test specimen is cut using a fixture where the fixture applied a double shear cutting force onto the test specimen. The out-of-plane shear strength of FRP composite laminate is determined by ASTM D3846 [51]. The test specimens are applied compressive load which consists of v-notches. The test specimens are loaded with supports to avoid buckling failure.

2.2 Failure Modes in Fibre Reinforced Polymer Composite …

73

2.2 Failure Modes in Fibre Reinforced Polymer Composite Laminates 2.2.1 General Failure Modes Observed Under Tensile Loading There are different ASTM standards to conduct the tensile test FRP composite materials. However, ASTM D3039/D3039M standard is most widely used to conduct the tensile test on rectangular flat coupons. Further, when FRP composite laminates subjected to tensile loading according to ASTM D3039/D3039M standard, then few general failure modes are observed when the test specimen fails (Fig. 2.20). These failure modes can be categorized as acceptable (Fig. 2.20a–c) and unacceptable (Fig. 2.20d–g) failure modes. The unacceptable failure modes of FRP composite laminates under tensile loading are breakage of FRP composite laminate under tab region (Fig. 2.20a), tab slipping or debonding of tab from the FRP composite laminate (Fig. 2.20b), and breakage of test specimen near the tab region or near the edge of the tab (Fig. 2.20c). The acceptable failure modes in FRP composite laminate when subjected to tensile loading are failure of the test specimen through ply delamination (Fig. 2.20d), flat fracture of the test specimen along the width (Fig. 2.20e), laminate splitting along the span length (Fig. 2.20f) and breakage of the test specimen into two halves at an angle along the test specimen width (Fig. 2.20g). Loading end Loading end

Loading end

(a)

(b)

(c)

Fixed end

Fixed end

Fixed end

Fixed end

Loading end

(f)

(e)

(d)

Loading end

Loading end

Loading end

(g)

Fixed end

Fixed end

Fixed end

Fig. 2.20 General failure modes observed in FRP specimen subjected to tensile loading (a–c) Unacceptable failure modes and (d–g) acceptable failure modes

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2 Mechanical Testing of Fibre Reinforced Polymer …

Loading end Loading end

(b)

Loading end

Loading end

Loading end

Loading end

Loading end

(c)

(d)

(e)

(f)

(g)

(h)

Span length

(a)

Loading end

Fixed end

Fixed end

Fixed end

Fixed end

Fixed end

Acceptable failure modes FRP composite laminate

Fixed end

Fixed end

Fixed end

Unacceptable failure modes Tab

Fig. 2.21 Schematic illustration of general failure modes observed in FRP composite laminate under compression loading (a–d) Acceptable failure modes and (e–h) unacceptable failure modes

2.2.2 General Failure Modes Observed Under Compression Loading When a FRP composite laminate is tested under ASTM standards related to compression test, then the tested specimen should fail in certain manner which are called acceptable failure modes, while the tested specimen does not fail by these accepted failure modes then they are called as unacceptable failure modes (Fig. 2.21). The acceptable failure modes are transverse shear failure (Fig. 2.21a), brooming failure at the middle of the gage or span section (Fig. 2.21b), through-thickness failure near the tab region within the span length (Fig. 2.21c) and several long splitting in the gage or span section (Fig. 2.21d). During testing, the test specimen may fail by certain unacceptable failure such as debonding of tab from the test specimen (Fig. 2.20e), through-thickness failure in the tab section (Fig. 2.21f) and end crushing of the test specimen in the tab section (Fig. 2.21g–h).

2.2.3 General Failure Modes Observed Under Low Velocity Impact or Quasi-Static Indentation The FRP composite laminate when subjected to drop weight impact (low velocity impact) or quasi-static indentation according to ASTM standard then few general failure modes can be observed. These failure modes can be broadly categorized into two sections: (i) visible damages and (ii) subsurface or non-visible damages

2.2 Failure Modes in Fibre Reinforced Polymer Composite …

75

(a)

(b)

(e)

(c)

(d)

(f)

Visible damage

Subsurface/Non-visible damage

Fig. 2.22 Schematic diagram of common failure modes observed in FRP composite laminates subjected to low velocity impact

(Fig. 2.22). The visible damages include formation of dent at impact side or fibre splitting and cracks on the opposite side of the impact site or combination of both (Fig. 2.22a–d), whereas the subsurface or non-visible damages include delamination or subsurface microcracks or combination of both (Fig. 2.22e, f). The subsurface damage modes are also called as barely visible impact damages. All these failure modes’ occurrence in FRP composite laminate depends on the impact velocity applied. As the impact velocity increases, the damage modes transform from barely visible impact damage to visible damage.

2.2.4 General Failure Modes Observed Under Compression After Impact Loading When tested impacted specimen under compression loading according to ASTM D7137/D7137M standard, then the specimen fails by certain acceptable failure modes. These failure modes include damage propagation from the impact site through the middle of the specimen (Fig. 2.23a), damage propagation slightly away from the damage cite (Fig. 2.23b), long splitting of the test specimen around the impact site or away from the impact site (Fig. 2.23c), delamination propagation from the impact site towards edges of the test specimen either along the width or length (Fig. 2.23d, e).

76

2 Mechanical Testing of Fibre Reinforced Polymer … (b)

(a)

(d)

(c)

(e)

Fig. 2.23 Schematic illustration of common failure modes observed in FRP composite laminates subjected to compression loading after subjecting to impact loading

2.2.5 General Failure Modes Observed in Specimen Under Out-Of-Plane Shear Loading The FRP composite laminates when subjected to shear loading according to ASTM standard (D2344—short beam shear strength) following general modes can observed in the tested specimen. General failure modes which could be observed in the test specimen when subjected to short beam shear loading are (i) interlaminar failure where the crack is generated between the two adjacent plies which are oriented in different direction (Fig. 2.24a), (ii) interlaminar failure at the edges of the laminate (Fig. 2.24b), (iii) compression failure at or near the loading roller in the form of fibre bending or buckling or kinking (Fig. 2.24c), and tensile failure at the opposite side of the load roller in the form of crack formation of ply splitting or fibre breakage (Fig. 2.24d).

2.2.6 General Failure Modes Observed Under Shear Loading Few general failure modes are observed in FRP composite laminate when subjected to in-plane shear loading according to ASTM D5379/D5379M standard. These failure modes can broadly categorized into two types: (i) unacceptable failure modes

2.2 Failure Modes in Fibre Reinforced Polymer Composite … (a)

77

(c)

loading direcƟon loading roller laminate failure

SupporƟng roller (d)

(b)

`

Fig. 2.24 Schematic illustration of general failure modes observed in FRP composite laminate subjected to short beam shear test

(Fig. 2.25a–c) and (ii) acceptable failure modes (Fig. 2.25d–g). The unacceptable failure modes are crack formation near the notch region or adjacent to the notch region (Fig. 2.25a), vertical crack formation around the notch region (Fig. 2.25b) and long angled crack formation across the middle region of the test specimen (Fig. 2.25c). The acceptable failure modes include any crack formation in the gage section (between top and bottom) of the test specimen (Fig. 2.25d–g). (a)

(d)

(e)

(f)

(g)

(b)

(c) Acceptable failure modes

Unacceptable failure modes

Fig. 2.25 Common failure modes observed in FRP composite laminate under in-plane shear loading (a–c) Unacceptable failure modes and (d–g) acceptable failure modes

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2 Mechanical Testing of Fibre Reinforced Polymer … (a)

(b)

(c)

(d)

(g)

(h)

(e)

(f)

Fig. 2.26 Schematic representation of common failure modes observed in FRP composite laminates adhesive lap joint when subjected to tensile loading

2.2.7 General Failure Modes Observed in Adhesive FRP Composite Lap Joints The adhesive lap joint of FRP composite laminates under tensile loading is conducted according to the ASTM D5868/D5868M standard. The failure modes associated are shown in Fig. 2.26. General failure modes observed according to this standard are adhesive failure, where the adhesive completely adheres to one part of the test specimen (Fig. 2.26a), cohesive failure where the adhesive is observed on both part of the test specimen (Fig. 2.26b, c), fibre tear failure with adhesive where the fibres from one part of the test specimen is peeled apart (Fig. 2.26d, e), breakage of one of the laminate of the test specimen near the adhesive region (Fig. 2.26f) and adhesive promoter failure (Fig. 2.26g, h).

2.2.8 General Failure Modes Observed in Filled or Non-Filled Open-Hole Specimens Subjected Tensile or Compression The filled open-hole tensile or compression test of FRP composite laminates show few general failure modes when subjected to tensile or compression loading according to corresponding ASTM standard. The acceptable failure modes are those which takes placed near the filled hole (Fig. 2.27). General failure modes observed in the failed test specimen after subjecting to tensile or compression loading are

2.2 Failure Modes in Fibre Reinforced Polymer Composite …

Fixed end

Fixed end

79

Fixed end

Fig. 2.27 Schematic diagram of failure modes observed in open-hole FRP composite laminate specimen

failure of the test specimen when the crack propagates across the filled hole and breaks the specimen into two halves. Further, delamination and cracks may be present extensively or moderately around the filled hole region.

2.3 Summary Testing of FRP composite laminates is a complex and time-consuming procedure. Because, the FRP composites are heterogeneous and anisotropic which significantly affects the results and makes difficulty in achieving good repeatability and reliability in experimental data. Further, the test specimen preparation is the one of the key factor in getting better and reliable results. Furthermore, the specimen preparation also influences both acceptable and unacceptable failure modes. End tabbing of test specimen is necessary where the specimen has to be held by or between the grips to avoid crushing. There are more than one ASTM standards for few testing. In such scenarios, careful study is necessary to select the most appropriate testing standard.

2.4 Further Reading Corresponding ASTM standards.

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2 Mechanical Testing of Fibre Reinforced Polymer …

References 1. ASTM D3171-15 (2015) Standard test methods for constituent content of composite materials. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D3171-15 2. ASTM D3529-16 (2016) Standard test methods for constituent content of composite prepreg. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D3529-16 3. ASTM D3530-20 (2020) Standard test method for volatiles content of composite material prepreg. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D3530-20 4. ASTM D3039/D3039M-17 (2017) Standard test method for tensile properties of polymer matrix composite materials. ASTM International, West Conshohocken, PA. https://doi.org/10. 1520/D3039_D3039M-17 5. ASTM D638-14 (2014) Standard test method for tensile properties of plastics. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D0638-14 6. ASTM D5083-17 (2017) Standard test method for tensile properties of reinforced thermosetting plastics using straight-sided specimens. ASTM International, West Conshohocken, PA. https:// doi.org/10.1520/D5083-17 7. ASTM D5450/D5450M-16 (2016) Standard test method for transverse tensile properties of Hoop wound polymer matrix composite cylinders. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5450_D5450M-16 8. ASTM D7565/D7565M-10 (2017) Standard test method for determining tensile properties of fiber reinforced polymer matrix composites used for strengthening of civil structures. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7565_D7565M-10R17 9. ASTM D7205/D7205M-21 (2021) Standard test method for tensile properties of fiber reinforced polymer matrix composite bars. ASTM International, West Conshohocken, PA. https:// doi.org/10.1520/D7205_D7205M-21 10. ASTM D5766/D5766M-11 (2018) Standard test method for open-hole tensile strength of polymer matrix composite laminates. ASTM International, West Conshohocken, PA. https:// doi.org/10.1520/D5766_D5766M-11R18 11. ASTM D6742/D6742M-17 (2017) Standard practice for filled-hole tension and compression testing of polymer matrix composite laminates. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6742_D6742M-17 12. ASTM D953-19 (2019) Standard test method for pin-bearing strength of plastics. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D0953-19 13. ASTM D5961/D5961M-17 (2017) Standard test method for bearing response of polymer matrix composite laminates. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/ D5961_D5961M-17 14. ASTM D7337/D7337M-12 (2019) Standard test method for tensile creep rupture of fiber reinforced polymer matrix composite bars. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7337_D7337M-12R19 15. ASTM D695-15 (2015) Standard test method for compressive properties of rigid plastics. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D0695-15 16. ASTM D3410/D3410M-16e1 (2016) Standard test method for compressive properties of polymer matrix composite materials with unsupported gage section by shear loading. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D3410_D3410M-16E01 17. ASTM D6641/D6641M-16e2 (2016) Standard test method for compressive properties of polymer matrix composite materials using a combined loading compression (CLC) test fixture. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6641_D66 41M-16E02 18. ASTM D5449/D5449M-16 (2016) Standard test method for transverse compressive properties of Hoop wound polymer matrix composite cylinders. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5449_D5449M-16 19. ASTM D6484/D6484M-20 (2020) Standard test method for open-hole compressive strength of polymer matrix composite laminates. ASTM International, West Conshohocken, PA. https:// doi.org/10.1520/D6484_D6484M-20

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20. ASTM D7264/D7264M-21 (2021) Standard test method for flexural properties of polymer matrix composite materials. ASTM International, West Conshohocken, PA. https://doi.org/10. 1520/D7264_D7264M-21 21. ASTM D790-17 (2017) Standard test methods for flexural properties of unreinforced and reinforced plastics and electrical insulating materials. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D0790-17 22. ASTM D6272-17e1 (2017) Standard test method for flexural properties of unreinforced and reinforced plastics and electrical insulating materials by four-point bending. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6272-17E01 23. ASTM D6415/D6415M-06a (2013) Standard test method for measuring the curved beam strength of a fiber-reinforced polymer-matrix composite. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6415_D6415M-06AR13 24. ASTM D2344/D2344M-16 (2016) Standard test method for short-beam strength of polymer matrix composite materials and their laminates. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D2344_D2344M-16 25. ASTM D7136/D7136M-20 (2020) Standard test method for measuring the damage resistance of a fiber-reinforced polymer matrix composite to a drop-weight impact event. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7136_D7136M-20 26. ASTM D256-10 (2018) Standard test methods for determining the Izod pendulum impact resistance of plastics, ASTM International, West Conshohocken, PA. https://doi.org/10.1520/ D0256-10R18 27. ASTM D4812-19 (2019) Standard test method for unnotched cantilever beam impact resistance of plastics. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D4812-19 28. ASTM D6110-18 (2018) Standard test method for determining the Charpy impact resistance of notched specimens of plastics. ASTM International, West Conshohocken, PA. https://doi. org/10.1520/D6110-18 29. ASTM D5420-21 (2021) Standard test method for impact resistance of flat, rigid plastic specimen by means of a striker impacted by a falling weight (Gardner impact). ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5420-21 30. ASTM D5628-18 (2018) Standard test method for impact resistance of flat, rigid plastic specimens by means of a falling dart (tup or falling mass). ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5628-18 31. ASTM D3763-18 (2018) Standard test method for high speed puncture properties of plastics using load and displacement sensors. ASTM International, West Conshohocken, PA. https:// doi.org/10.1520/D3763-18 32. ASTM D8101/D8101M-18 (2018) Standard test method for measuring the penetration resistance of composite materials to impact by a blunt projectile. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D8101_D8101M-18 33. ASTM D7137/D7137M-17 (2017) Standard test method for compressive residual strength properties of damaged polymer matrix composite plates. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7137_D7137M-17 34. ASTM D5528-13 (2013) Standard test method for mode I interlaminar fracture toughness of unidirectional fiber-reinforced polymer matrix composites. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5528-13 35. ASTM D6115-97 (2019) Standard test method for mode I fatigue delamination growth onset of unidirectional fiber-reinforced polymer matrix composites. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6115-97R19 36. ASTM D7905/D7905M-19e1 (2019) Standard test method for determination of the mode II interlaminar fracture toughness of unidirectional fiber-reinforced polymer matrix composites. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7905_D7905M19E01 37. ASTM D6671/D6671M-19 (2019) Standard test method for mixed mode I-Mode II interlaminar fracture toughness of unidirectional fiber reinforced polymer matrix composites. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6671_D6671M-19

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38. ASTM E1922-04 (2015) Standard test method for translaminar fracture toughness of laminated and pultruded polymer matrix composite materials. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/E1922-04R15 39. ASTM D3479/D3479M-19 (2019) Standard test method for tension-tension fatigue of polymer matrix composite materials. ASTM International, West Conshohocken, PA. https://doi.org/10. 1520/D3479_D3479M-19 40. ASTM D7615/D7615M-19 (2019) Standard practice for open-hole fatigue response of polymer matrix composite laminates. ASTM International, West Conshohocken, PA. https://doi.org/10. 1520/D7615_D7615M-19 41. ASTM D6873/D6873M-19 (2019) Standard practice for bearing fatigue response of polymer matrix composite laminates. ASTM International, West Conshohocken, PA. https://doi.org/10. 1520/D6873_D6873M-19 42. ASTM D5868-01 (2014) Standard test method for lap shear adhesion for fiber reinforced plastic (FRP) bonding. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/ D5868-01R14 43. ASTM D6264/D6264M-17 (2017) Standard test method for measuring the damage resistance of a fiber-reinforced polymer-matrix composite to a concentrated quasi-static indentation force. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D6264_D6264M-17 44. ASTM D2583-13a (2013) Standard test method for indentation hardness of rigid plastics by means of a barcol impressor. ASTM International, West Conshohocken, PA. https://doi.org/ 10.1520/D2583-13A 45. ASTM D3518/D3518M-18 (2018) Standard test method for in-plane shear response of polymer matrix composite materials by tensile test of a ±45° laminate. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D3518_D3518M-18 46. ASTM D4255/D4255M-20 (2020) Standard test method for in-plane shear properties of polymer matrix composite materials by the rail shear method. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D4255_D4255M-20 47. ASTM D5379/D5379M-19e1 (2019) Standard test method for shear properties of composite materials by the V-notched beam method. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5379_D5379M-19E01 48. ASTM D5448/D5448M-16 (2016) Standard test method for inplane shear properties of hoop wound polymer matrix composite cylinders. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D5448_D5448M-16 49. ASTM D7078/D7078M-20e1 (2020) Standard test method for shear properties of composite materials by V-notched rail shear method. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7078_D7078M-20E01 50. ASTM D7617/D7617M-11 (2017) Standard test method for transverse shear strength of fiberreinforced polymer matrix composite bars. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D7617_D7617M-11R17 51. ASTM D3846-08 (2015) Standard test method for in-plane shear strength of reinforced plastics. ASTM International, West Conshohocken, PA. https://doi.org/10.1520/D3846-08R15

Chapter 3

Low Velocity Impact on Fibre Reinforced Polymer Composite Laminates

Abstract A structural component may experience several mechanical loading conditions such as dynamic, cyclic, and static under various environmental conditions. Examples for each loading conditions are as follows: dynamic loading such as impact, cyclic loading such as fatigue, static loadings such as creep and chemical corrosion for environmental condition. These may degrade the material properties with time or may lead to an instantaneous reduction in the strength of the laminates. In both the scenarios, the composite structural material fails catastrophically without any yielding. Any material fails prematurely within the service life of that particular structure than it is a loss of money and time or it may be a matter of life and death of a passenger travelling in an aircraft or an automobile. In order to tackle these problems, it is necessary to understand the FRP laminates behaviour under these situations and environmental conditions. Among various loading conditions, impact is one of the critical loading conditions for FRP composites where the entire impact phenomenon occurs within a short time period. The fracture and damage mechanism involved with the FRP composite materials under impact loading is complicated because of its anisotropic behaviour and heterogeneity. Further, the FRP materials have weak through-thickness properties; thus, these materials become highly vulnerable and susceptible to impact loading, particularly if the impact velocity is low. Because the low velocity impact (LVI) creates barely visible impact damage (BVID) in the structure. This BVID is not visible to the naked eyes and sometimes may go unnoticed during non-destructive inspection. Moreover, the BVID severely reduces the residual strength of the structure and the damage formed in the laminate grows throughout the structure with time leading to sudden catastrophic failure. Thus, this chapter covers various aspects involved with low impact loading condition of FRP composites. Further, different damage detection methods are discussed briefly. The finite element analysis is a power full tool which can be used to understand the complex damage mechanism or failure events associated with FRP composite materials subjected to LVI. These numerical analysis methods use various mathematical models to define the material damage or failure criteria. Moreover, different contact formulations are also used to define the contact between impactor and composite laminate along with the plies interfaces. This chapter provides brief information regarding the common

© The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_3

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material models and contact formulation used to conduct the finite analysis of FRP composite laminates under LVI. Keywords FRP composites · Low velocity impact · Damage mechanism · Finite element analysis · Material models · Contact definition

3.1 Basics of Impact Loading ‘Sudden, instantaneous load or force acting on a limited area of on an object or structure is termed as impact’. FRP composite laminates are vulnerable to LVI, and it is dangerous to the structure because LVI induces subsurface cracks which are barely visible to naked eyes and may also introduce visible cracks depending upon the velocity of the impact. In case, if the LVI induces the barely visible impact cracks into the laminate, then these cracks grow throughout the laminate over a period of time and eventually reduces the residual strength of the structure. Which leads to catastrophic failure. Thus, it is much necessary to study the LVI damage mechanism of the FRP composites along with their behaviour under LVI. The damage formation due to the applied impact force differ from high velocity to low velocity impact along with the impactor mass. For example, the low mass with high velocity impact which represents the debris impact on aircraft during take-off and landing is well imitated by the gas gun experimental setup (Fig. 3.1). During in this type of impact testing, different waves such as compressive, shear and surface are propagated into the FRP composite laminate and away from the impact site in all three direction. The damage in the FRP composite under high velocity impact is generated when the reflected compressive waves creates sufficient tensile stresses at back of the 2 10 8

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Fig. 3.1 Basic schematic illustration of gas gun experimental setup

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composite plate. Similarly, the large mass with low velocity impact which represents the tool drop on a FRP composite in an assembly line or a service station is recreated using drop weight impact machine. During the low velocity impact, the stress waves generated are propagated many times back and forth in the laminate before final descent in all material direction. The damage in the FRP composite laminate under LVI is induced when the appropriate bending moment is created. As mentioned in the earlier sections, the vulnerability of the FRP composite laminate increases under LVI. Because the LVI creates the delamination in the FRP composite laminate which significantly reduces the residual strength of the FRP composite plate. The FRP laminates are highly anisotropic, and the damage mechanism is highly complex involving various mechanisms from subsurface microcracks to delamination or matrix failure to fibre failure. The impact energy absorption by the FRP composite broadly takes place by barely visible impact damage (delamination, matrix and fibre microcracks) and permanent damage involving visible fibre breakages (indentation, penetration and perforation). Another way of classifying the impact damage is interlaminar damage (delamination) and intralaminar damage (matrix cracks, fibre fracture and fibre-matrix debonding). However, a generalized failure mechanism of FRP composite laminates under LVI is given below. Elastic deformation of the laminate FRP composites are brittle and yield small deformation. Thus, a small amount of energy is absorbed out of the total amount of energy supplied during the impact loading. Plastic deformation of the laminate This stage absorbs a significant part of the total energy supplied during the impact as it involves various irreversible failure modes such as matrix cracking, delamination and fibre breakage. Other stages It involves various other parameters such as the generation of sound produced during the impact, generation of heat between the impactor and laminate during the impact. This stage absorbs an insignificant amount of energy provided during the impact loading. Failure events associated with the FRP laminate under LVI are as follows and represented in Fig. 3.2. Fibre breakage: It takes place when the fibre reaches its fracture strain. Matrix deformation and cracking: When the applied energy reaches the elastic limit of the matrix. Matrix cracks are off two types: (i) tensile cracks and (ii) shear cracks. Fibre and matrix debonding: In this event, separation of matrix and fibre takes place due to the cracks generated during loading, which is running parallel to the fibres. Here, both chemical and secondary bonds between the matrix and fibre interface are entirely separated.

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Fig. 3.2 Schematic representation of different failure events associated with FRP materials after subjecting to low velocity impact

Fibre pull out: When fibres reinforced in a tough matrix, this phenomenon will take place; here, fibre breaks at the weak cross section.

3.2 Classification of Impact Loading In impact, there is no distinct classification of impact loading conditions. However, generally, it is classified based on the velocity of the impact such as low velocity impact (LVI), medium velocity impact (MVI), high velocity impact (HVI), ballistic impact and hyper-velocity impact. Here, there is no confined boundary for the range of impact velocity. However, it is widely considered that if the impact velocity is between 0 and 10 m/s, then it is called low velocity impact (Example: Dropping of tools on a FRP composite structure in a production line or assembly line or service station). If the impact velocity ranges from 10 to 100 m/s, then it is considered as medium velocity impact (Example: Dropping of hails onto a FRP composite structure). If the impact velocity ranges between 100 and 200 km/s is known as high velocity impact (Example: Runway debris hitting an aircraft during take-off and landing). If the impact velocity is between 200 and 500 km/s, then it comes under ballistic impact (Bullet penetration or grenade/TNT explosions), and if the impact velocity is more than 500 km/s, then it is considered as hyper-velocity impact (Example: Space debris collision with the spacecraft).

3.3 Parameters Influencing the Impact Performance of Fibre Reinforced Polymer Composite Laminates There are various parameters which affects the impact resistance of FRP composite laminates (Fig. 3.3). Understanding these parameters influence over the impact resistance properties of FRP composite is essential. For example, the damage

Cross-ply, quasiisotropic, etc.

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created by the high mass impactor with low velocity may be different from the damage created by the low mass with high impact velocity even though the impact energy in both scenario is same. Further, if the impact velocity is low, then the damage becomes global phenomenon where the complete FRP structure provides the response. Whereas, if the impact velocity is high, then the damage becomes localized and that particular region of the FRP structure responds to the applied load. The damage propagation in the FRP composite laminate during LVI is strongly influenced by the laminate thickness. For thick laminates the damage starts from top play to the bottom ply whereas for thin laminates the damage starts from bottom to top ply. Similarly, the impactor shape also affects the damage modes in the FRP composite laminate. For example, if the impactor nose shape is flat, then the laminate shows more global damage failure, whereas the sharp nose impactor shows more localized failure involving penetration and fibre breakage. Figure 3.4 provides the illustration of different impactor nose shapes. Thus, the fundamental understanding helps in developing mathematical models which intern helps in predicting the FRP composite laminate behaviour and developing enhanced material system.

3.4 Damage Detection Methods Determining the defect free or damage free or material quality is long been with humans. For example, tapping method is used to determine the quality of the wood

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or detection of fatigue cracks in railroad in early days. The quantitative determination or estimation of the damage induced in the FRP composite laminate is important particularly when the damage induced is from LVI. The quantitative measurement of induced damage can be done by two techniques: (i) destructive techniques and (ii) non-destructive techniques. The destructive methods are: 1. Cross-sectional fractography where the FRP composite laminate is cut using a diamond saw and the damage extent is observed using optical microscope. 2. Deply technique where the damaged specimen is soaked in gold chloride solution with an isopropyl carrier then dried and pyrolysed in an oven around polymer decomposition temperature. The non-destructive technique include ultrasonic, X-ray radiography, infrared thermography, vibrothermography, tapping technique, radiography, etc. Selection of the damage detection technique depends on the FRP composite laminate. For example, the high intensity damage detection (non-destructive) method is used for translucent materials such as GF/epoxy or Kevlar/epoxy where the impact induced damage is quantified using high intensity background light (Fig. 3.5). However, this method only gives the projected damage area. Technological advancement has given different methods for damage detection. The FRP composite laminates might contain fabrication defects such as voids, misalignment of fibre, etc. These defects grow into damage in the FRP composite with time during their service. Further, the external loads such as impact further induces additional damage into the FRP laminates. Thus, regular inspection or damage detection of structures made of FRP composite laminate at regular interval time is necessary to check-in the quality and criticality of the FRP composite laminates. Because both quality and damage analysis affects the performance of FRP composite laminates. Further, the FRP composite laminates are heterogeneous and anisotropic, thus the damage detection in FRP composite laminate becomes challenging. Moreover, it is much necessary to understand which non-destructive testing method suits which

2 4 1 3 1 – high intensity background light source; 2 – translucent FRP composite laminate; 3 – impact induced damage; 4 – observer Fig. 3.5 Schematic illustration of high intensity background light setup used for observing the impact damage in translucent FRP composite laminate

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type of damage. Because any one non-destructive method will not detect all types of damage or defects. Damage detections such as interlaminar disbands, matrix cracks, fibre cracks, fibre-matrix interface failure, and damage area by the naked eyes is impossible. Thus, non-destructive techniques are used to check the quality of the fabricated FRP laminates or defects induced during the fabrication process such as voids, contaminations, air bubbles at resin pocket regions, or scratches. Here, few of the non-destructive methods are explained briefly. Ultrasonic The basic working principle of ultrasonic method is that an emitting transducer emits high frequency sound through a test specimen and on the other side the receiver transducer receives these frequencies (Fig. 3.6). The results obtained from the inspection are compared with the data corresponding to the undamaged specimen. Any variation in the frequency data from that of the standard specimen indicates the presence of damage or defect or flaws. There are different scan methods to conduct the ultrasonic inspection such as A, B and C scans. The ultrasonic C-scan method is extensively used than other scanning modes because this scanning method is good at porosity detection and density measurement along with delamination (when specimen in normal to the wave) detection. Further, there are two types of ultrasonic inspection method. One is immersion ultrasonic inspection method where the test specimen is immersed in a liquid medium for inspection. Here, the flat test coupon is immersed in a water tank such that the test specimen is kept parallel to the base. Now, the high frequency ultrasonic pulses are passed through the test specimen. The receiver transducer receives the signal on the other side of the test specimen. Now, the variation in Transduced and receiver on the same side

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Fig. 3.6 Schematic illustration of ultrasonic non-destructive damage detection testing

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frequency of tested specimen and reference material is checked. In another method, the specimen in not immersed in any liquid medium; instead, the inspection of the specimen is carried by direct contact approach. The ultrasonic inspection technique is used in aerospace industry to check the defects present in aircraft panels. Advantages of this non-destructive method are portability, good accuracy and high sensitivity. Primary limitation of ultrasonic non-destructive inspection method is the inability to detect the damage when the wave and damage are parallel. Other challenges in ultrasonic damage detection method are misalignment of specimen, emitter and receiver, unwanted geometrical discontinuities such as thickness, in accurate gradient measurement, etc. Thermography This technique analyses the surface temperature distribution over a region of interest (Fig. 3.7). There are different approaches for this inspection process which can be broadly categorized into contact and non-contact type thermography methods with direct and indirect heating methods. Similar to acoustic emission, this method can be used during specimen testing. Unlike, acoustic emission where the damage position cannot be predicted, here the position of the damage position can be predicted. There are two methods chemical and electronic by which the thermal images can be produced. In electronic approach, the infrared radiation emitted from the test specimen is detected using an infrared camera then converted into visual thermal image. In chemical approach, liquid crystals are used where these liquid crystals behave as liquid as well as optical crystals. The liquid crystal used for chemical thermography is the cholestric.

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Fig. 3.7 Schematic illustration of basic working principle of thermography damage detection method

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Infrared Thermography This method detects the damage or defects based on the thermal diffusion mechanism. In this method, the test specimen with region of interest is excited with the heat. Then the results are recorded in terms of surface temperature distribution. The major advantages of this process include scanning of large area with short time, the testing is non-contacting, no receiver is required on the other end and involves no harmful radiation. Further, following are few limitations of this method: (i) complete clearance of the testing surface is required, (ii) uniform heat application is required, (iii) the testing surface should have good emissivity. Vibrothermography In this method, a low amplitude mechanical vibrations are induced to generate the localized heating in FRP composite laminate, and corresponding change in heat patterns is recorded (Fig. 3.8). Then, the recorded data are compared with the reference data. Based on the difference in results between the undamaged and damage specimens, the damage extent is estimated. Acoustic emission In this testing procedure, the transducer is coupled to the testing specimen using adhesive and then detected for the transient elastic stress waves which are produced as results of applied stresses (Fig. 3.9). This procedure is useful to detect the damage in pipes, tanks and pressure vessels. Real-time or in situ damage detection is possible using this technique. However, the test specimen has to be loaded with mechanical loads or temperature or pressure or any other means to stimulate the stress waves. Further, other limitations are the data prediction requirement of intense data filtering; from this method, particular failure in the test specimen cannot be detected.

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Acoustoultrasonics In this non-destructive inspection method, an acoustic or stress waves are generated at one of the test specimen. Then on the other end of the test specimens, the signals are detected. The results are compared to with standard results and any deviation with the standard results indicate the presence of damage or defect in the specimen. Scanning acoustic microscopy Similar to ultrasonic testing method, the basic working principle of scanning acoustic microscopy is same. Here also, a high frequency is pointed and attenuated. This method can give a resolution of 20 µm damage or defect. Major limitation is the region of scanning area limited or small. Thus, the larger area of interest is covered though several scans. Further, this inspection method is also limited by damage depth in the test specimen. Computed radiography In this process, an electric charge beam is generated using a high voltage current on a cathode which are accelerated through a vacuum tube towards anode. This anode is the target. When the generated electron beam strikes the target, a rapid deceleration is observed in the electron beam. During this deceleration of electron beam, the Xrays are generated which are then directed to the test specimen. Then a 2D image is collected on a digital screen and used to detect both defects and damage in the laminate. The major advantage of this process is that this technique inspects the entire volume of the test specimen rather surface. Further, advantages include: (i) the technique does not use any pre chemical processing of the specimen and (ii) good sensitivity towards material density and thickness. The limitations of this technique include: (i) not sensitive to surface cracks, (ii) radiations concern, (iii) the damage detection sensitivity reduces as the specimen thickness increases, (iv) high initial

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cost, (vi) difficult to achieve good contrast if the atomic number of the composite substructures are between low to medium. Radioscopy The only working principle difference between the radiography and radioscopy is that the radioscopy technique provides near real-time non-film damage detection. Further, the major advantage of this process is that the process allows the operator to work from a radiation free distance. The limitations of this process are the cost associated with the machine is high. Further, the sensitivity and resolution of the images obtained from this technique are not as good as the images obtained from radiography technique. Holographic interferometry This non-destructive testing process is good at detecting the matrix cracks, delamination between plies and fracture in fibres with high resolution. In this method, a 3D dimension image of the test specimen is constructed using a 2D photographic emulsion. Procedure for conducting holographic inspection of specimen as follows. First, the test specimen is placed in a room till the temperature of the specimen reaches the room temperature. Then the test specimen is clamped and maintained an angle of 65 °C between the object and reference beam where both the beams intersect. Now, by maintaining the reference to object beam intensity ratio of 0.5, the real-time hologram is recorded. Here, both thermal pulsing and real-time hologram used to determine the optimum thermal cycle. After every exposure, the test specimen is allowed to cool till the test specimen temperature reaches to the room temperature. X-Ray radiography In this inspection method, photons or X-rays are used to detect the damage. Here, the damage in the test specimen should be parallel to the X-ray path. High density penetrant used to cracks and/or delaminations present in the test specimen. Here, the X-rays are directed onto the test specimen. Then the images are collected on film. The collected images are in black and white, where the light region indicates the presence of crack or damage or defect. Several penetrants are used to enhance the X-ray radiography. However, the voltage required for penetrant enhanced X-ray radiography is higher than the non-penetrant enhanced X-ray radiography. There are several pentrants such as zinc iodide. One among them is the tetrabromoethane is used to enhance the quality and damage in the FRP composite. Toxicity and resin solvent are the limitations of tetrabromoethane usage. X-Ray Tomography The X-ray radiography with computed tomography known as X-ray tomography. This non-destructive testing technique extends the damage detections in the test specimens to all three dimension or viewpoints. Further, in X-rays, the defects or damage which are parallel to the X-ray path goes undetected; however, the X-ray tomography detects the damages or defects in all three directions irrespective of

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damage or defect position in the test with respect to the X-ray path. Factors that define the damage detection quality and capacity are the ability to absorb the X-rays by the test specimen, X-ray density and the distance between test specimen and X-ray source. The basic working principle of X-ray tomography is similar to that of X-Ray radiography. However, only difference between X-ray tomography and X-ray radiography is that the unlike the X-ray radiography, the X-ray tomography uses multiple radiographs to collect the image along 360°. After collecting the multiple images, they are reconstructed using mathematical algorithms to construct 3D image. Computed tomography In this technique, X-rays are passed through the test sample along different direction and a 3D cross section is reconstructed. This method provides clear 3D crosssectional view which helps in better understanding of the damage. The images obtained from this method are clearer for interpretation. Major limitation with this technique is the size of the test specimen. Further, the large area inspection is also not possible and reconstruction of the images takes time. Furthermore, it is difficult to obtain the contrast for composite materials with low atomic number substrates. Leaking testing In this damage detection method, fluid or gas is passed through the test specimen at a pressure or concentration gradient. The damage or leak location is determined either by tracer probe or by detector probe. The tracer probe method is used when the tracer gas is externally used after complete evacuation. The detector prove is used when the test specimen is pressurized with the tracer gas itself. It is used for composite tubes. This method gives more accurate results than the liquid penetration method. Shearography This technique detects the damage in a specimen when the test specimen experiences a gradient load. The basic working principle of shearography is the interferometric method. In this technique, the test specimen is subjected to a certain preload or proof load to detect the damage. Once the excitation is done, the images are captured. Then the before loading images and after loading images are superimposed which detects the damage in the test specimen. The major advantages of this process are (i) cost effective, (ii) the damage detection can be done to a large area, (iii) influence of testing environment on the testing results are minimum, (iv) minimum skills are required to operate the machine. The major limitation of this detection method is that the testing specimen surface condition may interfere the results.

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3 Low Velocity Impact on Fibre Reinforced Polymer …

3.5 Numerical Simulation of Fibre Reinforced Polymer Composite Laminates Subjected to Low Velocity Impact The critical question is that why do we need FEA tools for simulation..? There are several reasons for that, one being to reduce the cost and time associated with experimental work. More importantly, FRP laminates are highly anisotropic and involve more complicated damage mechanisms/subsurface microcracks. This becomes a more cumbersome and tedious process; hence for better understanding and to predict the damage mechanism of an FRP laminate under impact loading, FEA tools are must and are a more efficient way to do the analysis. John Hallquist invented the new DYNA FEA simulation tool for crash testing. This tool provides more real-life situations to simulate, especially impact related situations. In these FEA tools, there are various ways or methods to model the FRP composite to predict the damage and its related mechanism. Such as MAT_022, 054, 055, 058–059 are explained in the following section. In the subsequent section, only material models and contact definitions are explained briefly which are extensively used for the analysis LVI behaviour of composite materials.

3.5.1 Material Models MAT_022 material model in LS-DYNA uses Chang-Chang composite failure criteria. According to this failure criteria, a composite material fails when the following condition is satisfied (Eq. 3.1). Fmatrix > 1 and Ffibre > 1

(3.1)

where  σ2 2 Fmatrix = +τ S2  2 σ1 Ffibre = +τ S1       σ2 σ2 2 C2 −1 +τ = + 2S12 2S12 C2 

FComposite

(3.2)

(3.3)

(3.4)

3.5 Numerical Simulation of Fibre Reinforced Polymer …

 Fibre Matrix Shearing = τ = 

97

τ12 2G 12

2 S12 2G 12

2



+

+

3 4 ατ12 4



 3  4  αS12 4

(3.5)

S 1 –Longitudinal tensile strength, S 2 —Transverse tensile strength, S 12 —Shear strength, C 2 —Transverse compressive strength, alpha—Nonlinear shear stress parameters calculated from in-plane stress–strain equations (Eqs. 3.6–3.8).  1 (σ1 − ν1 σ2 ) E1   1 ε2 = (σ2 − ν2 σ1 ) E2   1 3 τ12 + ατ12 2ε12 = G 12 

ε1 =

(3.6) (3.7) (3.8)

In the case of MAT_054 and MAT_055 material models, MAT_054 uses Chang matrix failure criterion while MAT_055 uses Tsay-Wu matrix failure criterion. These two failure criteria’s segregate the failure due to fibre and matrix under compressive and tensile loading. Chang-Chang criteria involve tensile fibre failure mode, compressive fibre failure mode, tensile matrix failure mode compressive matrix failure mode and given below (Eqs. 3.9–3.13). Tensile fibre failure mode, σ aa > 0 then  e2f

=

σaa Xt

2



σab +β Sc



 ≥ 0 Failed −1 ≤ 0 Elastic

(3.9)

E a = E b = Gab = ν ab = ν ba = 0 Compressive fibre failure mode, σ aa < 0 then  eC2 =

σaa XC

2

 ≥ 0 Failed −1 ≤ 0 Elastic

(3.10)

E a = ν ab = ν ba = 0 Tensile matrix failure mode, σ bb > 0 then  em2 =

σbb Yt

2

 +

σab Sc



 ≥ 0 Failed −1 ≤ 0 Elastic

E b = ν ba = 0 as Gab to zero Compressive matrix failure mode, σ bb < 0 then (Eq. 3.12)

(3.11)

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3 Low Velocity Impact on Fibre Reinforced Polymer …

 ed2

=

σbb 2SC



2 +

   2   YC σbb σab 2 ≥ 0 Failed −1 −1 + ≤ 0 Elastic 2SC YC Sc

(3.12)

E b = ν ba = 0 as Gab to zero Tsai-Wu failure criteria consider the same fibre failure criteria in tensile and compressive as in case of Chang-Chang failure criteria, and the only difference is matrix failure under tensile and compressive loading and is given by Eq. 3.13.  2 emd

=



σbb YC Yt

 +

σab SC

2

 +

(YC − Yt )σab −1 Yt YC



≥ 0 Failed < 0 Elastic

(3.13)

MAT_058 model uses strain failure criteria where if a particular element reaches the strain, then that element is removed. Similarly, MAT_059 uses Hashin failure criteria which is interactive type failure criteria and predict the failure of the fibre and matrix separately. According to this theory, failure will take place when Eqs. 3.14– 3.19 criteria satisfied. σ11 ≥ 1 Failure along the fibre direction 

σ11 Xt

2

 +

2 2 + σ13 σ12 2 S12



 =

≥ 1 Failure < 1 No failure

(3.14)

σ11 < 0 Compressive fibre failure 

σ11 XC



2 =

≥ 1 Failure < 1 No failure

(3.15)

σ22 + σ33 > 0 Tensile Matrix Failure 

σ22 + σ33 Yt

2

 +

2 − σ22 σ33 σ23 2 S23



 +

2 2 σ12 + σ13 2 S12



 =

≥ 1 Failure < 1 No failure

(3.16)

σ22 + σ33 < 0 Compressive Matrix Failure 

YC 2S23

+

2

 −1

σ22 + σ33 YC



2 − σ22 σ33 σ23 σ2 + σ2 + 12 2 13 2 S23 S12



 (σ22 + σ33 )2 2 4S23  ≥ 1 Failure = < 1 Nofailure

+

σ33 > 0 Interlaminar Tensile Failure

(3.17)

3.5 Numerical Simulation of Fibre Reinforced Polymer …



σ33 ZT

2

 =

≥ 1 Failure < 1 No failure

99

(3.18)

σ33 < 0 Interlaminar compression failure 

σ33 ZC

2

 =

≥ 1 Failure < 1 No failure

(3.19)

3.5.2 Contact Definition The contact definition defines the interaction between two separate entities. The contact definition also enables the unmerged Lagrangian elements to interact among themselves. Basically, to conduct the finite element analysis of LVI on FRP composites, two separate contacts have to be defined. First contact is between the impactor and the laminate. Second is between among the plies. The LS-DYNA uses three methods to handle the contact definitions: (i) kinematic constrained-based method, (ii) penalty-based method, and (iii) distributed parameter method. To define the contact definition, there are different models which are to be selected considering various parameters. Few of the contact definition names commonly used for FRP composite laminates under LVI are given below. AUTOMATIC_GENERAL AUTOMATIC_NODES_TO_SURFACE AUTOMATIC_NODES_TO_SURFACE_SMOOTH AUTOMATIC_ONE_WAY_SURFACE_TO_SURFACE AUTOMATIC_ONE_WAY_SURFACE_TO_SURFACE_TIEBREAK AUTOMATIC_ONE_WAY_SURFACE_TO_SURFACE_SMOOTH AUTOMATIC_SINGLE_SURFACE AUTOMATIC_SINGLE_SURFACE_SMOOTH AUTOMATIC_SINGLE_SURFACE_TIED AUTOMATIC_SURFACE_TO_SURFACE AUTOMATIC_SURFACE_TO_SURFACE_TIEBREAK AUTOMATIC_SURFACE_TO_SURFACE_SMOOTH CONSTRAINT_NODES_TO_SURFACE CONSTRAINT_SURFACE_TO_SURFACE ERODING_NODES_TO_SURFACE ERODING_SINGLE_SURFACE ERODING_SURFACE_TO_SURFACE NODES_TO_SURFACE NODES_TO_SURFACE_INTERFERENCE NODES_TO_SURFACE_SMOOTH ONE_WAY_SURFACE_TO_SURFACE

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3 Low Velocity Impact on Fibre Reinforced Polymer …

ONE_WAY_SURFACE_TO_SURFACE_INTERFERENCE ONE_WAY_SURFACE_TO_SURFACE_SMOOTH SURFACE_TO_SURFACE SURFACE_TO_SURFACE_INTERFERENCE SURFACE_TO_SURFACE_SMOOTH TIEBREAK_NODES_TO_SURFACE TIEBREAK_NODES_ONLY TIEBREAK_SURFACE_TO_SURFACE TIED_NODES_TO_SURFACE TIED_SHELL_EDGE_TO_SURFACE TIED_SHELL_EDGE_TO_SOLID TIED_SURFACE_TO_SURFACE TIED_SURFACE_TO_SURFACE_FAILURE To define the contact, there must be two bodies or objects. Among these two, one is considered as slave and another one is considered as master. Correspondingly, the nodes which present in slave are called as slave nodes. Similarly, the nodes which are present in the master are called as master nodes. In penalty-based method distinction, between slave and master is irrelevant. However, it is much necessary to distinguish between master and slave nodes if the kinematic constrain and distributed parametric methods are used. Because in both the methods, the slave nodes slide over the master nodes during the impact. After the impact, the slave node must remain over the mater node till a tensile force develops between them. Kinematic Constraint Method In this method, the nodal displacement components corresponding to the slave nodes along with the contact interfaces are transformed to the global equations. During the transformation process, the normal degree of freedom of the nodes is eliminated. The global degree of freedom of master nodes is coupled efficiently by considering the lumped mass and explicit time integration. The uncertainty in the kinematic constraint method arises when the slave surface zoning is not finer than the master surface zoning because the few nodes in the master surface may penetrate the slave nodes surface creating kink lines (Fig. 3.10).

Fig. 3.10 Illustration of surface nodes from master surface to slave surface (indicated by x) [LSDYNA Theory manual r:9320 2017—Livermore Software Technology Corporation]

3.5 Numerical Simulation of Fibre Reinforced Polymer …

101

Penalty Method In penalty-based contact algorithms, the discrete spring elements are placed between the penetrating nodes and contact surface where the discrete spring elements transfers the force and moment resultants from slave nodes to master nodes (Fig. 3.11). This method creates comparatively less noise than the nodal constraint approach. In this method, the momentum is conserved. There are three different penalty formulations: (i) standard penalty formulation, (ii) soft constraint penalty formulation and (iii) segment-based formulation. In the standard penalty formulation, the stiffness of the interface is approximately equal to the stiffness of the interface element normal to the interface. When the slave node penetrates the master surface, then the standard penalty formulation activates, where the interface forces are applied to the slave nodes and their contact points. The soft constraint penalty formulation method is used to define the contact between soft body and a rigid body, for example, the contact between foam and steel. During this penalty formulation as one surface is soft which may result excessive penetration. This limitation is tackled by using scale factor. This type of formulation does not work with NODES_TO_SURFACE type contacts. In the segment-based formulation, the penetration by one segment into another segment is detected and then the penalty forces are applied to both the segments. Further, initial penetration is ignored. This penalty formulation is majorly used for solid and shell elements. The segment formulation is more suitable if the surface of the contact is rough or if the contact region has very sharp edges or corners. The segment-based penalty formulation is also suitable for airbag type contact. In this formulation, segment masses are considered instead of nodal masses. For shell segments, the segment mass is equal to the element mass, whereas the segment mass is equal to half the element mass for solid segments. The stiffness of the contact in penalty method is defined according to Eqs. 3.20 and 3.21 for solid and shell elements α K A2 . V

(3.20)

αK A Max(shell diagonal)

(3.21)

k= k= where k = contact stiffness K = material bulk modulus α = penalty scale factor A = segment area V = element volume.

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3 Low Velocity Impact on Fibre Reinforced Polymer …

Distributed Parameter Method In this method, the slave element half mass of each element in contact is dispersed over entire master surface area. This method shows less mesh instabilities during the simulation because the constraints are imposed to the salve nodes acceleration and velocity. Further, here the slave nodes are not allowed to enter into the master nodes. One-Way Contact Types In this contact approach, the coarser mesh part should be considered as master. Further, the rigid or impacting is preferred as master, while the nodes are considered as slave. This method is cost effective than the twoway type contact definitions. Following are the few examples which falls in the category of one-way contact types such as NODES_SURFACE, ONE_SURFACE_TO_SYRFACE, AUTMOATIC_NODES_TO_SURFACE and ONE_WAY_AUTOMATIC_SURFACE_TO_SURFACE. Single Surface Contacts These types of contact are used to define the contact between parts. Most commonly used contact type for impact analysis. Further, this contact type is useful to calculate the Eigenvalues and Eigenvectors of the interface if TIED option is used. This type of contact always considers the thickness offset and the consideration of orientation of segment is not necessary. This method has robust bucket sorting. This type of contact is generally suitable for laminated composites. The contacting surface fails if the defined shear value between the interfaces is reached. In this type only, the slave side is defined and also considers the thickness offsets. It is similar to NODES_TO_SURFACE type contact. In this contact type, the slave part is considered as segment instead of set of nodes. The contact types such as SINGLE_SURFACE, AUTOMATIC_SINGLE_SURFACE, ERODING_SINGLE_SURFACE and AUTOMATIC_GENERAL are few examples. SMOOTH Contact Any contact definition with SMOOTH used to indicate the master segment. For this type of contact condition, a smooth curve fitter surface is required to assign as master segment. Use of contact definition with SMOOTH helps in representing the actual surface. Further, the contact definition with SMOOTH helps in reducing the contact noises generated during the analysis. Further, use of contact definition with SMOOTH also helps in getting better results with coarser mesh. For contact algorithms SURFACE_TO_SURFACE and SINGLE_SURFACE with SMOOTH, the slave and master segments are smoothed every cycle. GENERAL Contact Both AUTOMATIC_GENERAL and AUTOMATIC_SINGLE_SURFACE are single surface contact type, which are beam-to-beam contact where the penetration is

3.5 Numerical Simulation of Fibre Reinforced Polymer …

103

checked along the beam length. Generally suitable to define the contact for shell edgeto-edge or beam-to-beam. The AUTOMATIC_GENERAL creates null beam automatically at exterior edges of shell elements; if the INTERIOR option is used, then the AUTOMATIC_GENERAL contact definition applied the null beam to all the shell elements along with unshared edges as well as the shared edges. This contact type is not efficient as well as expensive than the AUTOMATIC_SINGLE_SURFACE. TIED Contact TIED type contact definition is more suitable for slid-to-solid type elements, where there is no rotational degree of freedom. Further, it is used to define the contact between two parts that are co-planar with dissimilar mesh. Use of this contact definition to the beam elements or plate results in desirable results as the handling of rotational degree of freedom is not efficient. This contact definition formulation is based on the constraint based method. This type of contact definition is used for the solid mesh transition where the two meshed parts are co-planar or adjacent. This contact definition provides a smooth interpolation of displacement and stress for adjacent surface with dissimilar mesh. For example, the TIED_NODES_TO_SURFACE and TIED_SURFACE_TO_SURFACE contact definitions can be used for the solid elements with no rotational degree of freedom. The TIED type contact is also used to define the shell-to-shell and beam-to-shell contacts, where nodes have rotations degree of freedom. The contact definitions include TIED_SHELL_EDGE_TO_SURFACE. The contact definition TIED_SHELL_EDGE_TO_SOID is used for defining the contact between shell edges to solid nodes as well as for defining the contact between beam ends to solid where only slave nodes have rotational degree of freedom. In TIED contact algorithms, sudden transition of zoning is allowed solid meshed elements (Fig. 3.11). In other words, this contact type is good if the contact surfaces of two bodies are meshed with two different sizes. There is no need to match the nodes across the two objects whose interfaces are merged if TIED contact definition is used. ERODING_SURFACE In this type of contact algorithm, the elements are deleted present on the free surface as they reach the material failure criteria defined for the materials. The deleted nodes or elements belongs to the slave part. Further, it is based on the AUTOMATIC_SINGLE_SURFACE where the segment orientation is not important. This type of contact definition is more suited for high velocity impact analysis. The nodes sets associated should be considered completely. The element deletion takes place after every bucket sorting at each time step. Example for this type of contact is the contact between the bird strike and blade. TIEBREAK_SURFACE Here, the failure is based on the tensile and shear stresses. Once the failure occurs, then reverted to non-automatic. Failure of the elements or nodes is based on the

104

3 Low Velocity Impact on Fibre Reinforced Polymer …

Fig. 3.11 Illustration of TIED contact definition used for mesh transition at tied interface [LSDYNA Theory manual r:9320 2017—Livermore Software Technology Corporation]

tensile and shear force. Once the failure occurs, then reverted to non-automatic NODES_TO_SURFACE contact. SURFACE_TO_SURFACE. Both tensile and shear stress failure criteria (Eq. 3.22) are used to during the formulation of this type of contact. Failure criteria 

|σn | NFLS

2

 +

|σ S | SFLS

2 ≥1

(3.22)

where NFLS and SFLS are tensile failure stress and shear failure stress, respectively. TIEBREAK_NODES The failure criteria used in this type of contact is given below (Eq. 3.23) Failure criteria 

| fn | NFLF

NEN

 +

| fn | SFLF

MES ≥1

(3.23)

where NEN and MES are exponent for normal and shear force, respectively. Similarly, the NFLF and SFLF are normal and shear failure force. The f n and f n are normal and shear interface forces, respectively. SURFACE + INTERFERENCE This type of contact is useful in defining the contact between the parts which are already shrink or prestressed in the initial stage. This contact

3.5 Numerical Simulation of Fibre Reinforced Polymer …

105

type considers the shell thickness offset for deformable bodies are meshed with shell elements. Material cards which are based on this type of contact definition are CONTACT_NODES_TO_SURFACE_INTERFERENCE, CONTACT_ONE_WAY_SURFACE_TO_SURFACE_INTERFERENCE and CONTACT_SURFACE_TO_SURFACE_INTERFERENCE.

3.6 Summary Impact loading is an eminent threat to the FRP composite structures. The impact loading type is classified based on the impact velocity. Further, the impact performance of FRP composite laminates affected by various parameters which are broadly categorized as material parameters, impact testing parameters and testing environmental parameters. Furthermore, there are several destructive non-destructive techniques to detect the impact induced damage for FRP composite laminate. However, the choice of testing method depends on the material, geometry, accuracy and type of damage detection. Use of finite element analysis can enable for better understanding of LVI damage behaviour of FRP composite laminate. There are several material models and contact formulations to carry the finite element analysis of LVI on FRP composite laminate.

3.7 Further Readings 1. 2. 3.

Impact on composite structures by S. Abrate Impact behaviour of fibre reinforced composite materials and structures by S. R. Reid and G Zhou LS-DYNA Theory manual 2015 Livermore Software Technology Corporation (LSTC).

Chapter 4

Low Velocity Impact on Carbon Fibre Reinforced Polymer Composite Laminates

Abstract Carbon fibre reinforced polymer composites show high strength, high stiffness and low weight. Thus, these composite laminates are used in many industries such as aerospace, defence and automobile. Thus, in this chapter, behaviour of carbon fibre reinforced polymer composite laminates under LVI is presented. The literature presented here covers different aspects such as laminate thickness, impact velocity, impact energy, stacking sequence which affect the impact resistance properties of the carbon fibre reinforced polymer composite laminates. Further, effect of hybridization of the carbon fibre with other fibres such as glass, basalt on the impact properties is also covered. The final section of this chapter covers the effect of addition of nanomaterial on the impact performance of carbon fibre reinforced polymer composite laminate. Keywords Carbon fibre reinforced polymer composites · Low velocity impact · Hybridization · Nanomaterial

4.1 Low Velocity Impact Test on Controlled Carbon Fibre Reinforced Polymer Composite Laminates Cantwell et al. [1] conducted LVI test on CF composite laminates for different stacking sequence. Experimental results showed that the CF composite laminates’ impact resistance properties were sensitive to stacking sequence. Addition of woven plies by replacing ±45° plies into the stacking sequences further enhanced the impact resistance properties of CF composite laminate. Dorey et al. [2] studied the influence of use of two different matrix materials on the damage resistance properties of carbon fibre. The two difference matrix materials used were PEEK and epoxy. Experimental results showed that the CF/PEEK specimens showed less damage than the CF/epoxy composite laminate. Further, the CF/PEEK composite showed less matrix cracks than the CF/epoxy composite laminates. Bishop et al. [3] fabricated the CF/PEEK composite laminate, then compared the impact resistance properties with the CF/epoxy composite laminate. The impact testing results showed that the CF/PEEK is more notch sensitive and has better impact resistance than the CF/epoxy © The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_4

107

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4 Low Velocity Impact on Carbon Fibre Reinforced Polymer …

laminate. Further, the damage inspection in CF/PEEK by non-destructive testing was more advantageous than the CF/epoxy laminate. Cantwell et al. [4] used high-strain carbon fibres and subjected to low velocity impact. Experimental results showed that use of high-strain carbon fibres improved the impact resistance. Further, the tensile and compressive strength after impact increased by 100% and 30%, respectively. Curtis [5] examined the impact strength of seven types of CF when reinforced with five different matrix materials where carbon fibre modulus and strain were varied from medium to high. Experimental results showed that the thermoplastic matrix materials showed good impact resistance properties. Liu et al. [6] investigated the matrix cracking and delamination interaction when subjected to sub-perforation impact loading and studied the effect of laminate thickness, stacking sequence, boundary conditions, type of loading and geometrics on the matrix crack and delamination interaction. Glass/epoxy plates were fabricated using autoclave method assisted by vacuum bagging. Laminates were cured at 340 °F under 40 psi for 40 min. A blunt-end impactor of mass 14.3 g was impacted onto a laminate using an air gun with 20–100 m/s impact velocity. Cracks in the delamination were not interrupted by the delamination on the second interface. Three different types of clamping, i.e. boundary conditions were used, such as rectangular shape with completely clamped, rectangular shape with simply supported and circular shape with clamped completely. The rectangular boundary conditions with fixed and simply supported showed no significant difference in the matrix cracking pattern except near the boundary of the laminate. Similar behaviour was also observed in the circular boundary condition. The [30°5 /−30°5 /30°5 ] and [0°5 /90°5 /0°5 ] lay-ups showed similar cracking pattern except for the pattern angle, which was rotated by 30°. The [0°3 /90°3 /0°3 /90°3 /0°3 ] and [0°/90°]10S did not show the same delamination area. In the case of [0°5 /90°5 /0°5 ] and [0°3 /90°3 /0°3 ] lay-ups, the matrix cracking pattern was the same except the crack density. Thick plates showed dense matrix cracks. It was clear from bright light and edge replication that the tensile stresses were the reason for matrix cracks. Higher the delamination area, the significant matrix cracking was the sign of proof that both delamination and cracking depend on each other. Cantwell and Morton [7] studied the effect of specimen geometry (thickness, length, width and shape) on the impact resistance properties of CF composite laminates. Experimental results showed that the damage in short and thick specimen was initiated at the impactor edge, whereas in long and thin specimen, the damage was initiated at lower plies by fibre splitting. Cantwell and Morton [8] investigated the effect of impactor mass on the damage formation in CF composite laminate under low velocity impact loading condition. Keeping the total impact energy same, the low mass with relatively high velocity created more concentrated damage around the impactor where most of the energy is absorbed by fibre shear failure. Cantwell and Morton [9] compared the impact response of CFRP composite laminate when subjected to low and high velocity impact loading. In low velocity impact, the materials’ elastic energy absorption capacity is important, and the target geometry defined the impact response, whereas in high velocity impact, the geometry

4.1 Low Velocity Impact Test on Controlled Carbon …

109

of the target showed no significant effects. However, high velocity impact created greater level of damage to the structure than the low velocity impact. Morton and Godwin [10] investigated the use of two matrix system on the impact resistance properties of carbon fibre. Two matrix used were PEEK and epoxy. Experimental results showed that the CF/PEEK laminates showed better impact resistance properties than the CF/Epoxy laminate. Further, the laminate design with 45° plies at the periphery showed better impact resistance properties. Ghasemi Nejhad and Parvizi-Majidi [11] studied the effect of two different matrix materials (PEEK and PPS) on the energy absorption capacity of carbon fibre. The CF/PPS composite laminates showed resistance to perforation by forming large delamination, whereas the CF/PEEK composite laminates showed localized damage area with increased impact resistance properties. Further, the perforation energy for CF/PPS was more compared to CF/PEEK composite laminates. Prichard and Hogg [12] used PEEK and epoxy matrix materials to fabricate the CF composite laminates. The fabricated CF/PEEK and CF/epoxy laminates were subjected to LVI. Experimental results showed that the CF/PEEK composite laminate performed better than the CF/epoxy composite laminate. Cantwell and Morton [13] studied the damage mechanism of CF composite laminate when subjected to low and high velocity impact. Conical damage was observed in the tested specimen, and the conical damage was not affected by the laminate thickness and stacking sequence. Further, for low velocity impact, the threshold impact limit was dependent on the areal dimensions where for high velocity impact, the threshold impact limit was independent on the areal dimensions of the test specimen. Dost et al. [14] investigated the effect of stacking sequence on the damage resistance property of CF composite laminates. Stacking sequence strongly influenced the delamination area and impact resistance property of CF composite laminate. The laminate with uniform stacking sequence and increment stacking sequence along with repeatable pattern in the stacking sequence affected the delamination size. Further, for any stacking sequence, the sub-laminate-type delamination was equal to (2n + 1) where ‘2n’ is the repeating unit number. Effect of moisture on the mechanical properties of carbon fibre (CF) reinforced with different matrix materials was investigated by Selzer et al. [15]. Matrix materials used were unmodified epoxy (EP), toughened epoxy and polyether-ether ketone (PEEK). Ply designs were [0°]16 , [90°]16 for unidirectional, [0°, 90°]4S for bidirectional and [0°/±45°/90°]2S for multi-directional fibres. Introduction of moisture was done by a desiccator at 23, 70 and 100 °C. Various mechanical tests such as tensile, compression, fatigue, compression after impact, Mode I and Mode II tests were investigated and conducted according to DIN 29971, 65350, 10100, 65561 and EGF standards, respectively. There was no significant effect of moisture on the tensile strength of all laminates fabricated at 0 °C. However, a sharp decrease in tensile strength is observed in CF/EP and CF/modified Ep laminates fabricated at 90 °C. CF/PEEK showed no significant effect of moisture content on compression strength at 0 °C, but a clear decrease in compression strength was observed for CF/modified EP and CF/toughened EP laminates along with an increase in moisture content. Similar behaviour was also observed at 90 °C. As the impact energy was increased,

110

4 Low Velocity Impact on Carbon Fibre Reinforced Polymer …

Fig. 4.1 Different stacking sequences considered for the impact testing [16]

the damaged area also increased for all types of laminates. A similar trend was also observed for GIC and GIIC properties. From the analysis, it was found that the moisture content greatly influenced the interface failure. The water temperature did not influence the laminate properties. Finally, it was concluded that the CF/PEEK performed better than the other laminates because this laminate water uptake was much less compared to other fibre laminates. In a study, Morioka et al. [16] investigated the influence of stacking sequence on the impact strength and failure modes for CF/epoxy composite laminate (Fig. 4.1). The impact tests were performed using Charpy impact machine. Different stacking sequence designs were unidirectional, orthotropic and quasi-isotropic. The unidirectional composite laminate showed good impact strength compared to other stacking designs. From the fracture analysis, it was found that the fracture in 90° ply occurs in advance than the ply oriented along 0°. Further, the fracture mode was zigzag for orthotropic laminate design. Scarponi et al. [17] examined the LVI damage-detection efficiency of ultrasonic non-destructive testing technique against the CF/epoxy, GF/polyester and KF/polyester composite laminates. Initially, the ultrasonic method was validated against the aerospace materials made of thin CF/epoxy and CF/PEEK composites. Then, in second phase, the LVI damage was detected in GF/polyester and KF/polyester materials. The ultrasonic detection technique efficiently detected the delamination and its location. Further, it was concluded that the damage resolution and accuracy were dependent on choice of probe frequency along with the skills of the operator. In an experimental investigation, Sohn et al. [18] studied the LVI performance of CF/epoxy composite laminate. Here, the CF/epoxy composite laminate was reinforced with other interlayer fibres. The interlayer fibre material used was Kevlar, Zylon, poly(ethylene-co-acrylic acid) film and polyamide web. The CF/epoxy composite laminate with short Zylon fibre showed minimized residual strength and enhanced the interlaminar fracture resistance, whereas the CF/epoxy composite laminate with poly(ethylene-co-acrylic acid) films showed reduced damage area. It was also observed that the coarse distribution of Kevlar fibres was not effective to improve the impact performance of CF/epoxy composite laminate. The CF/epoxy composite laminates showed different failure modes such as matrix cracking, delamination,

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intraply cracking and fibre failure. However, the failure modes were dependent on the interlayer material. Olsson [19] examined the effect of mass on FRP laminate under impact loading. It was given that the impact response by the laminate depends upon the impactor versus plate mass ration but not on the impact velocity. Small mass impacts lead to higher delamination than the large mass impact because in small impact, high impact loading was observed. Long impacts were influenced by plate size and boundary conditions, whereas short impacts were un-affected by these conditions. For small mass impact response, no need to include boundary conditions or surrounding substructure in the model based on this a numerical was developed and compared with experimental results available in the literature and found a good correlation between each other. Will et al. [20] examined the impact strength of cylindrical tube made of two different stacking sequences under LVI (Fig. 4.2). The stacking sequences [−35°/+35°/90°3 /−35°/+35°/90°3 /−35°/+35°] have comparatively low energy dissipation than the [90°6 /(−35°/+35°)3 ]. Further, the stacking sequence was also influenced the delamination area formation. Lee et al. [21] studied the crack introduction behaviour of non-carbon tissue (NWCT) incorporated with CFRP laminate under GIC loading when the artificial cracks were introduced at different positions (upper, middle and lower). Design of the laminate was [0°24 ] for CFRP and [0°12 /T/0°12 ] for NWCT incorporated CFRP laminates. GIC test was conducted using an end-notch flexural method according to ASTM D5045-9 standards with a crosshead speed of 0.5 mm/min. Fracture toughness damage mechanism and crack growth were investigated by optical and scanning electron microscopies. From the experimental results, it was observed that the mean GIIC of the NWCT specimen with a crack at the upper, middle and lower positions was 26%, 260% and 205% higher than that of the corresponding CFRP, respectively. From optical and scanning electron microscopies, it was concluded that the NWCT experiences the short bridging effect due to the random in-plane short carbon fibre and, hence, results in better interlaminar fracture toughness.

Fig. 4.2 Clamping of CF/epoxy tube [20]

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Pirvu et al. [22] studied the delamination behaviour of carbon/vinyl ester and glass/vinyl ester laminates reinforced with wood under different environments. Vacuum assisted resin transfer moulding/SCRIMP method was used to fabricate the laminates. Un-diluted creosote, copper naphthenate, diesel and mineral spirits were used as different environmental conditions. ASTM D3039, D2344, D4065, D2559 and D905 standards were followed to test the tensile, interlaminar, shear strength, dynamic mechanical analysis, resistance to composite delamination and resistance to shear by compression loading, respectively. The experimental results showed 1855 and 825 MPa tensile strength for carbon and glass fibre laminates. No degradation in interlaminar shear strength was observed in composite laminates. Glass transition temperature in the longitudinal direction for carbon and glass fibre is observed to be 139.7 °C and 145.3 °C, respectively. In dry test condition, 78% of the wood failure was occurred when treated with copper naphthenate, and 70% was observed in mineral spirits in wet conditions. E-glass failure takes place at 39% and 32% (highest) with diesel environment in both dry and wet conditions, whereas carbon fibre failure takes place at 43% in mineral spirits condition in dry test and 46% in copper naphthenate in wet condition. Finally, it was concluded that vinyl ester forms better interface bonding with carbon fibre than glass fibre, and VARTM/SCRIMP method was used as an efficient laminate fabrication process for wood structures. Mitrevski et al. [23] investigated the effect of impactor nose shape (hemispherical, ogival and conical) on the impact performance of thin CF/epoxy composite laminate. When the laminate was impacted by a conical nose impactor, then the CF/epoxy laminates absorbed the impact energy as a local penetration. The CF/epoxy composite plate showed the highest peak and minimum contact duration when impacted by hemispherical nose impactor and also produced barely visible impact damage. Further, it was also observed that the damage threshold limit was the highest when the CF/epoxy composite laminate was impacted by hemispherical impactor. Morais et al. [24] studied the effect of laminate thickness under repeated low velocity impact loading on glass, carbon, Kevlar laminates. Carbon laminate consists of three, four and six plies, while glass and Kevlar fibres were made of six, eight and ten laminates. Laminates were prepared by using vacuum bagging assisted by autoclave method and curing time for all the laminates were varied based on the fibre type used. A semi-spherical nose tip impactor of 12.7 mm diameter and 765 g of mass was impacted repeatedly onto the laminates from different heights (1 and 0.5 m) using a drop weight equipment. Macroscopic and microscopic failures were studied by visual and radiography techniques, respectively. From the experimental analysis, it was observed that the glass fibre laminates performed better than other laminates and concluded that below certain energy level, it does not matter the type of fibres used. As the impact energy increased with the number of repeated impacts, the laminate thickness also becomes a crucial parameter among other parameters such as fibre properties. Jiang et al. [25] investigated carbon/carbon fibre sandwiched structure analysis under low velocity impact. The sandwich structure consists of two outer skins made of CF/epoxy with thickness and plies orientation of 1.25 mm and [0°/45°/9°/−45°/0°]S , respectively. The core material is made of the same material with ply thickness

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and orientation of 1 mm and [0°/90°/90°/0°]S , respectively, and the structure was impacted with a hemispherical nose impactor of diameter 6.35 mm. From results, it was found that internal CF/epoxy with [0°/90°/90°/0°]S sheets helped in uniform distribution of energy throughout the laminate by reducing local displacement of the core panel. The static and axial collapse of CFRP square tubes were studied by Mamalis et al. [26] for varying length. LS-DYNA was used to simulate the damage mechanism using MAT_55 material card. Mode I fracture toughness was modelled by eroding single surface material card. Tie break surface to surface was used to model the bonding between the bundles of the plies tube wall. The variation in experimental and numerical results was 4.5%, 96.5–103.5% and 33% for the dynamic test, Mode I and Mode III, respectively. Finally, it was concluded that the existence of error between numerical and experimental works was because of the LS-DYNA material card MAT_55 as it does not consider the strain rate effect of the composite material. Puente et al. [27] developed a new analytical model for carbon/epoxy composite under high velocity impact. The model was based on energy balance for a spherical projectile, which includes laminate crushing, linear momentum transfer and tensile fibre failure. Laminate consists of ten layers of carbon fibre prepregs. The projectile was made of tempered steel with 7.5 mm in diameter and 1.73 g of mass. A singlestage gas gun was used to fire the projectile with velocity up to 500 m/s. The developed model was in good agreement with the obtained experimental results (Fig. 4.3). Tita et al. [28] examined the effect of stacking sequence and impact energy on the impact damage behaviour of thin CF/epoxy composite plates under LVI. The Hills failure criteria was used for the numerical analysis. The stacking sequences used were [0]10 , [0/90/0/90/0]S and [+45/−45/+45/0/90]S . The experimental results showed that if the absorbed energy was within 35%, then no failure was initiated. As the absorbed impact energy falls between 35 and 75%, then matrix cracks and delamination were observed. Finally, if the laminate absorbed more than 75%, then fibre rupture, delamination and matrix cracks were observed. It was concluded that the indentation test can be used to represent the LVI for laminate which shows quasi-elastic behaviour.

Fig. 4.3 Bottom surface of the impacted laminate at 112 m/s a Experimental simulation and b Numerical simulation [27]

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In this investigation, fibre type, lay-up, fibre architecture and arrangement were studied by Hufenbach et al. [29] under impact loading. Unidirectional T700, T300 plain-weaved and T800 four harness satin-weaved carbon fibres were used to fabricate the different laminates using resin transfer moulding technique. Impact test was conducted using Charpy method with impactor mass, velocity, arm length and energy were being 2 kg, 3.85 m/s, 390 mm and 15 J, respectively. DYNA-3D was used for numerical analysis with material card MAT-162, which allows modelling the different reinforcement types. From experimental results, it was observed that T300 unidirectional—[0°]16 core—and T300 bidirectional—[±45°]—as the outer skin and outer skin absorbed the highest energy (at 0° test angle) but yielded sudden failure. T300 bidirectional [0°]8 and metal as lower protective layer showed maximum deflection (at 0° test angle). Finally, it was concluded that tough fabric layers as outer skin could improve the impact resistance, and shear nonlinearities were not accurate to simulate the impact testing. Lopes et al. [30] tried to improve the impact resistance of CF/epoxy composite laminate via dispersed stacking sequence method. Here, a baseline stacking design [±45/90/0/45/04 /−45/02 ]S was optimized using genetic algorithm. Based on the optimization following dispersed stacking sequences [±45/0/70/ −70/0/15/10/−10/−15/15/−15]S and[±45/80/5/20/−20/10/−80/−10/−5/15/−15]S , it was observed that the optimized stacking sequences showed no improvement in the impact resistance properties compared to the baseline stacking sequences. In this experimental investigation, Heimbs et al. [31] analysed the low velocity impact behaviour of CFRP laminates. Three samples with 24 plies lay-up and [−45°/0°/45°/90°]3S were made of non-crimp fabric, and [−45°/0°/45°/90°] and [0°/45°/0°/−45°/020 /45°/0°/−45°/9020 /0°]S were made of prepregs tested using drop tower at 40 J of impact energy (Fig. 4.4). From LS-DYNA and experimental results, it was found that non-crimp fabric absorbs more impact energy. Delamination depends on the number of ply orientation. Finally, it was concluded that less delamination action was between the layers of similar fabric orientation. Hadavinia et al. [32] investigated the effect of delamination crack growth on the energy absorption capacity of hybrid unidirectional (U) and twill (T) carbon fibre laminates. Mode I and Mode II tests were performed using an end-load split and double cantilever beam method. LS-DYNA was used for numerical analysis where

Fig. 4.4 Compression of energy absorption at 40 J a Un-loaded and b Loaded [31]

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material model 54 was used to study the damaged modes, which involves ChangChang failure criterion. Out of [T0 /U0 ]7 , [T0 /U90 ]7 , [T90 /U0 ]7 , [T90 /U90 ]7 , [T90]10 and [U90 ]10 laminate lay-ups were used. The [T0 /U0 ]7 and [T90 /U0 ]7 absorbed the maximum energy compared to other hybrid composite boxes. Further, [T0 /U90 ]7 and [T90 /U90 ]7 performed almost the same as that of [U90 ]10 laminate. From numerical and experimental analyses, it was concluded that the crack generation for hybrid laminates in GIC and GIIC material properties plays a crucial role than its propagation. Saito et al. [33] studied the damage morphology of CFRP laminates as a function of ply thickness made of 24 and 96 plies with ply orientation of [45°/0°/−45°/90°]. Drop weight impactor was used with an impactor weight of 0.998 kg and impacted with the energy of 2.7 J. Compression after impact test was conducted according to the ASTM standard D7137. Ultrasonic C-scan and sectional fractography were used to evaluate the impact damage extent and damage behaviour. From the experimental investigation, it was observed that the compression after impact for thin laminates was 23% higher than the thick laminates. From ultrasonic and fractography analyses, circular and fan blade-shaped damages were observed for thick and thin laminates, respectively. Finally, it was concluded that the thin laminates showed fewer and localized transverse cracks, and delamination was propagated mainly in the mid-plane. In contrast, the thick laminates underwent buckling and delamination in through-thickness direction. The effect of carbon fibre laminate thickness on the dissipation of energy and deceleration effect of the impact between the plies was analysed experimentally, numerically and analytically by Obradovic et al. [34]. From analytical results, it was found that in cylindrical tubes, total external is equal to the sum of work required to bend the fibre layers, straining of fibres, matrix cracking and to overcome the frictional forces. From experimental and numerical results (Fig. 4.5), it was found that as the wall thickness increased, the corresponding displacement of laminates also increased. Finally, it was concluded that the force peak was much sensitive to the wall thickness.

Fig. 4.5 Compression test a Experimental and b LS-DYNA simulation [34]

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Double-ply, mirror-symmetric, balanced and quasi-static laminate designs effect on damage mechanism and fibre failure under low velocity impact were investigated by Hongkarnjanakul et al. [35]. Unidirectional carbon fibre laminates were fabricated according to the mentioned laminate designs. An experimental test was carried out on a drop tower machine using 16 mm diameter and 2 kg mass impactor according to Airbus Industries Test Method (AITM 1-0010). ABAQUS and VUMAT were used for numerical simulation, which include fibre failure, intraply failure, matrix cracking in intraply and delamination in interply. Square-shaped mesh style was used for 0° and 90° plies, and parallelogram-shaped mesh style was used for 45° and − 45° laminates. From experimental, numerical model and ultrasonic c-scan results, it was evident that this new compressive law used was more effective and reduced the calculation time, material parameters required, and no coupling parameter was required for failure modes. Xu et al. [36] conducted both experimental and numerical analysis to study the damage behaviour of carbon/epoxy laminates under low velocity impact. Two laminate designs [0°4 /90°4 ]S and [0°2 / ± 45°/90°2 ]S were impacted with a spherical impactor (mass = 2.48 kg, diameter = 25 mm) at 1.17 m/s impact velocity. Finite element analysis was conducted on ABAQUS/Explicit and VUMAT where C3D8R and COH3D8 were used for plies and cohesive plies, respectively. Delamination area obtained from experimental and numerical analyses was 392 mm2 and 458.64 mm2 , respectively, for [0°4 /90°4 ]S laminate. There were 17% and 10.2% delamination area was existed between numerical and experimental results for [0°4 /90°4 ]S and [0°2 / ± 45°/90°2 ]S laminates, respectively. In this research investigation, He et al. [37] used the eddy current pulsed thermography method (Fig. 4.6) to study the damage mechanism of carbon fibre laminates subjected to impact with 4 J to 12 J energy levels. A CFRP laminate consisting of 12 layers was prepared using polyphenylene sulphide matrix. Prepared laminates were subjected to the following energy levels in sequence 2, 4, 6, 8, 10 and 12 J. From experimental results, it was observed that lower conductivity of carbon fibre laminate could be detected directly in the heating phase. As the energy level of impact increased, corresponding hot area around it was also increased. Finally, it

Fig. 4.6 Set-up of a Reflection mode and b Transmission mode [37]

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was concluded that the eddy current method could be used as an effective technique to study the damage mechanism of the CFRP laminate. Agarwal et al. [38] carried out a comparative study on chopped and bidirectional plain-woven carbon fibre by subjecting dynamic and mechanical testings’ for varying fibre content. Hand lay-up method was used to fabricate the bidirectional FRP laminates, whereas short carbon fibre was mixed with resin (epoxy + hardener mixture) and poured into a mould and followed by curing process at room temperature for 24 h and post-curing at 50 °C for 15 min. Tensile, flexural, interlaminar and hardness were carried out to compare the mechanical properties of the composites. Meanwhile, storage modulus, loss modulus and tanδ are used for damping analysis to investigate the dynamic characteristics of the laminates. ASTM D3039-76, D2344-84 and D256 standards were followed and conducted the tensile, ILSS and impact tests on the universal testing machine, three-point bend test and Charpy impact testing machine, respectively. From experimental results, the hardness (diamond indenter) of chopped CFRP laminates was better than bidirectional woven composites. However, in all mechanical and dynamic testings’, bidirectional CFRP yielded the better properties than chopped CFRP laminates (except damping properties at 10 wt.% of chopped CFRP). Finally, it was concluded that as the fibre loading increased (10, 20, 30, 40 and 50%), the corresponding properties were also increased, and if this trend was observed, then the laminate might contain void content in it. The bidirectional CFRP properties were higher than chopped CFRP because these exhibited poor interfacial bonding. Elanchezhian et al. [39] investigated the mechanical properties of CFRP and GFRP composites. Tensile, flexural and impact tests were conducted to study the behaviour of both the composites. Scanning electron microscope images were used to analyse the damage and fracture behaviour of the laminates. Tensile test was conducted on UTM according to ASTM D638 (Fig. 4.7) at varying strain rates (1.5, 2.5 mm/min) and temperatures (350 °C and 750 °C). The flexural test was conducted using three-point bend fixture according to ASTM D7264. Samples were prepared

Fig. 4.7 Tensile specimens, according to ASTM D638 standard [39]

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using hand lay-up technique using epoxy. From experimental results, it was seen that maximum tensile strength for CFRP laminate was 913.861 N/mm2 with elongation of 4.76% at 350 °C with a strain rate of 2.5 mm/min. Meanwhile, GFRP showed less tensile strength 114.149 N/mm2 with 5.1176% elongation at a strain rate of 1.5 mm/min at 35 °C temperature. Flexural strength of CFRP was 31.578 N/mm2 with the displacement of 6.1 mm at a strain rate of 2.1 mm/min, and for GFRP, flexural strength was 9 N/mm2 with a displacement of 5.7 mm at a strain rate of 1.5 mm/min. The energy absorption capacity of CFRP and GFRP was 11 and 6 J, respectively. It was concluded that GFRPs are less prone to catastrophic failure than CFRP because GFRP has a high percentage of elongation when compared to CFRP. CFRP laminates were more impact resistance than GFRP as CFRP has high stiffness compared to GFRP. Intralaminar and interlaminar crack initiations and progressions under low velocity impact were studied by Shi et al. [40]. The authors proposed a new numerical model based on stress and fracture criterion and considered the interface cohesive element model, delamination/interlaminar cracking, ply splitting and transverse crack for developing an interlaminar crack. Damage formulation and damage laws were developed based on Hashin’s failure criterion and quadratic laws. Drop weight impact method and ABAQUS/Explicit FEM tool were used for experimental and numerical investigations, respectively. Carbon/epoxy [0°/90°]2S laminate was fabricated by autoclave method assisted by vacuum bagging. Laminate was impacted with an impactor of mass 1, 1.5 and 2 kg to obtain impact energy of 7.35, 11.03 and 14.7 J by dropping from a height of 0.75 m. The obtained experimental and numerical results were compared and found that the interlaminar cracking reduced the variation of results from 14% to less than 5.2%. The X-ray radiography analysis was also in close agreement with numerical analysis. Kim et al. [41] analysed the high velocity impact behaviour of carbon fibres in epoxy resin with [45°/0°/−45°/90°]ns quasi-symmetric and [0/90]ns ply orientation (where n = 4, 6 and 8). Steel core aluminium and armour piercing impactors with conical shapes were used, where impactor core and skin were made of steel and copper, respectively, and impacted with approximately 600 m/s and 800 m/s impact velocity. Damage shapes in the laminate were analysed by C-scan, and impact velocity of the impactor was measured using high-speed camera. LSDYNA was used for numerical analysis. ERODING_SURFACE_TO_SURFACE material card was used for simulating composite plate and projectile. MAT_ COMPOSITE_FAILURE_SOLID (MAT59) material card was used for damage analysis of composite. Material card MAT_PLASTIC_KINEMATIC (MAT3) was used to model the impactor. Along with ply thickness, the impact velocity also played a major role in the failure of the laminates. There was 2.60% of error between experimental and numerical analyses for steel core ammunition projectile subjected to 616.8 J of energy for laminate with [45°/0°/-45°/90°]24s orientation. Finally, it was concluded that the experimental and numerical results were agreed well. Caputo et al. [42] investigated the low velocity impact on CFRP panel with 6, 10, 13 J impact energy levels. Experimental and numerical analyses were carried out using drop weight impactor and ABAQUS FEA tool. An eight-layered laminate

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with [45°/90°/45°/−45°]S ply orientation was impacted with a hemispherical tip impactor of diameter 19 mm and mass of 3.5 kg. The test was carried out according to ASTM D7136/D7136-05 standards. Variation in numerical analysis results from experimental results was 0.76% concerning the delamination area for 10 J impact energy. Finally, it was concluded that the numerical and experimental results were in good correlation. Artero-Guerrero et al. [43] explored the damage behaviour of CFRP/epoxy as a function of impactor mass under low velocity impact. Other parameters, such as laminate thickness and impact energy, were also investigated. In-plane and throughthickness damage area extents were examined using traditional ultrasonic C-scan, and phased array ultrasonic methods were adopted. CFRP laminate with 20 and 30 plies was fabricated by autoclave method and impacted on a drop weight tower with a hemispherical nose impactor of diameter 20 mm according to the ASTM D7136 conditions. Three impactor mass 3.8, 7.8 and 15.8 kg were impacted with 10 and 110 J of impact energy. From C-scan and phased array method, it was un-ambiguous that different impactor mass played no significant role in damage behaviour (i.e. extension and topology). However, the impact energy and laminate thickness parameters were influential in deciding the damaged area and its extent. Time-dependent behaviour of thermomechanical properties of carbon fibre reinforced epoxy laminates were studied by Dong et al. [44]. Vacuum assisted resin transfer moulding method was used to fabricate the laminates with three curing stages. During the first stage, the laminates were manufactured at 90 °C for two hours followed by post-cured at 110 °C for one hour, and then, in third stage, laminate was cured at 130 °C for four hours. Then, the laminates were allowed to cool to room temperature. In-plane and out-of-plane coefficients of thermal expansion of carbon fibre were determined by dilametero. Finite element model (ABAQUS) was used for numerical analysis where CATIA and HyperMesh were used to create a 2D model and meshing followed by numerical analysis in ABAQUS. Warp yarns experienced fist thermal expansion and shrinkage in the in-plane direction along with slight transverse expansion. In out-of-plane direction, thermal stresses were primarily generated, which were influenced by carbon yarns and glass transition temperature. It was evident that both in-plane and out-of-plane coefficients of thermal expansion were in good agreement with the numerical model. Schwab et al. [45] proposed a new approach to simulate the impact mechanism of CFRP laminate using ABAQUS. An eight-layered laminate with [0°/45°/0°/45°]S plies orientation was fabricated by autoclave at 177 °C for 2.5 h. Drop tower impact testing was used with a cylindrical impactor with a hemispherical nose striker of diameter 20 mm. Impact velocity and energy during testing used were 20 m/s and 400 J, respectively. Orthotropic energy-based continuum damage mechanics approach was used to simulate the damage and failure behaviour of the laminate under the impact. Intradamage and interdamage mechanisms and failure behaviour of CFRP laminate results were in close agreement with experimental results (Fig. 4.8). The durability of the CFRP rods was examined by Tanks et al. [46] when exposed to alkaline and saline solutions under the Charpy impact test. Three unidirectional CFRP laminates were prepared: two samples were carbon fibre with vinyl ester, and one was carbon fibre with epoxy. A pendulum impact tester was used, and the impact

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Fig. 4.8 Damage of the composite impacted at 140 J a Simulated model and b Experimental sample [45]

Fig. 4.9 Charpy impact test result of notched and un-notched specimens [46]

test was conducted according to ASTM D6110-10 standards. Impact velocity and energy were 3 m/s and 330 J, respectively. During Charpy impact testing, notched and un-notched samples were tested. Samples were treated with alkaline solution at 20 °C and 60 °C for three months. From experimental results, notched CF/epoxy performed better than other samples and lowest being CF/VE exposed to solution environment. From the result, it was concluded that samples made of CF/epoxy (notched) performed better than un-notched, CF/VE and environmentally treated samples (Fig. 4.9). Arachchige et al. [47] used the first-order deformation theory to predict the contact force history for a cured composite under low velocity impact and proposed a new numerical method, i.e. spring mass model (Fig. 4.10). Carbon fibre and steel impactor were used, and the model was used for curved laminates with [(±45°/02 )2 /±45°/0°/90°]2S lay-up. Lateral stiffness was used as a variable to

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Fig. 4.10 Spring mass system to model dynamic force [47]

develop the mathematical model, which predicted the impact response of laminates. Initially, this model was compared with the literature results to validate the results and then implemented for four geometry designs. Finally, it was concluded that the used mathematical model could reduce the complicated time significantly. Ahmad et al. [48] carried out a research investigation on the effect of moisture content in the unidirectional CFRP composites when subjected to low velocity impact. The laminate was made of unidirectional carbon fibre reinforced with epoxy matrix using vacuum bagging technique. Laminates were preconditioned in a drying oven at 50 °C until no further change in the weight was observed, and then laminates were immersed in water at 80 °C for varying time cycles. A hemispherical impactor with 3.44 kg of mass and nose radius of 5.5 mm used with impact energy and velocity was approximately 0.7 m, 23.62 J and 3.71 m/s, respectively. The energy absorption for the sample immersed in water for 12,611.29 h was approximately 10.5 J when compared to the laminate immersed in water for zero hours was 12.5 J. The displacement for non-immersed laminate was approximately 27 mm and that of composite immersed in water for 12,611.29 h was 32 mm (Fig. 4.11). Then, it was concluded that the moisture content in CFRP laminates induced an adverse effect and reduced the properties such as impact peak force, threshold delamination load and energy initiation. Meanwhile, the flexural rigidity and damage area were increased. Han et al. [49] studied four different composite samples, both experimentally and numerically. Low velocity impact test and compression after impact tests were used for experimentation, while ABAQUS/Explicit was used for numerical analysis. Four different samples CCF800/epoxy, CCF300/epoxy, CCF800/bismaleimide and CCF300/bismaleimide were prepared with stacking sequence of [45°/0°/−45°/9°]4s and [45°/0°/−45°/90°]3s with a layer thickness of 0.125 mm and 0.19 mm, respectively. Two energy levels 4.43 J/mm and 6.67 J/mm were used for impact testing. ASTM D7136 and ASTM D7137 standards were used for conducting impact and compression after impact testing, respectively. Finally, it was concluded that the carbon fibre 300 with epoxy was better damage tolerant than carbon fibre 800 with

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Fig. 4.11 Displacement of laminates with time [48]

bismaleimide. In numerical analysis, interlaminar damage and delamination were simulated by using multi-scale failure criterion and cohesive element, respectively, which were a new numerical method used in the analysis, and it yielded a good correlation with experimental results. Qu et al. [50] developed a new numerical model to examine the damage mechanism and accurately predict the damage extent of FRP laminates under low velocity impact. Unidirectional carbon fibre and epoxy laminates with [45°/0°/−45°/90°]4S , [450/00/−450/90°]3S , [45°/0°/−45°/90°]2S , [0°/90°]4S and [45-45]4S ply orientation were fabricated. Drop weight impact test was conducted according to ASTM D7136 standards and laminates impacted at the centre of the laminate using a hemispherical nose impactor with 16 mm diameter with 6.67 J/mm impact energy. The numerical model developed consisted of induced and tangential delamination methods. Further, Hashin failure criteria considered to evaluate the damage failure modes. From testing, it was evident that [45°/0°/−45°/90°]4S absorbed more energy than other laminate designs. Based on ultrasonic C-scan, [45°/0°/−45°/90°]4S showed maximum damage area. Finally, it was concluded that the new model based on continuum damage mechanics and cohesive zone model was in good agreement with experimental results. Singh et al. [51] developed a new analytical model to assess the contact, bending, shear and membrane stiffness of the laminate. CFRP laminates were analysed and compared with experimental results available from the literature. ABAQUS/Explicit was adopted for numerical simulation. The laminate design used was [0°2 /90°2 /0°2 /90°2 /0°2 ]S and modelled using C3D8R elements. Spherical impactor with 6.5 kg was modelled and impacted with 5 J and 10 J of impact energy. The numerical simulation results were in good agreement with the experimental

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results, and prediction accuracy was mainly due to the incorporation of interlaminar and intralaminar damage effects in the developed analytical model. A new multi-scale analysis model was developed by Lou et al. [52]. It combined the micromechanics of failure and cohesive damage modes to predict both interlaminar and intralaminar damages as well as failure modes. CFRP laminate with [−45°/90°/45°/0°]2S lay-up was fabricated using a hot pressing method and impacted with a hemispherical impactor of mass 2.11 kg and nose diameter of 20 mm at impact energies of 3.40 J, 5.36 J and 7.35 J. ABAQUS/Explicit and VUMAT are used for numerical analysis. Python and FORTRAN languages are used to develop the model, respectively. Computational efficiency is improved by using two different meshing systems, i.e. C3D8 at the centre of the model and C3D8R at the periphery of the models were used. Numerical analysis showed that matrix cracking and delamination were the primary failure mechanism, whereas fibre breakage was limited, and these numerical results were supported by C-scan analysis. Front and back layers undergo failure due to compression and tensile stresses, respectively. From the analysis, it was found that the new numerical model developed was in good agreement with experimental results obtained from drop tower and C-scan. Li et al. [53] used computed tomography and numerical analysis method to study the shear impact damage modes of 3D braided composites under punch shear loading. Further, the influence of braiding parameters, impact strain and specimen thickness on the damage modes were also investigated. The 3D braided CF/epoxy laminates were fabricated using vacuum assisted resin transfer moulding, and testing was carried out on split Hopkinson compression pressure bar. Experimental results showed the braided angle 25° with 8 mm thickness at 2000 S−1 strain rate. Further, at 2000 S−1 strain and 25 °C braided angle, the 8 mm thick laminate absorbed more energy than that of 5 mm thick laminate. Mechanical properties of the carbon fibre/epoxy laminates by incorporating functionalized electrospun polyacrylonitrile nanofibre (PANF) were investigated by Neisiany et al. [54] under various mechanical loading conditions. Laminates were made of carbon fibre, and five layers of nanofibres were sandwiched between the carbon fibre plies. Laminates were fabricated by using hand lay-up followed by vacuum assisted resin transfer moulding technique to ensure proper wetting between fibres and matrix. Laminates were cured at 27 mm/Hg for ten hours and then at 50 °C for 15 h. Impact energy absorbed by control laminate, PAN and PAN-g laminates was found to be 39.13 ± 2.73 J, 43.85 ± 2.66 J and 47.40 ± 2.06 J, respectively. It was clear that the functionalized nanofibre incorporation was the critical factor in better performance of the laminate. Finally, it was concluded that this enhancement was due to better interfacial bonding between fibre and matrix, which was developed by PAN nanofibre which resulted in improved out-of-plane properties thus improving the mechanical properties of the laminates. Stanciu et al. [55] developed a finite element approach to investigate the damage propagation in CFRP laminates. Laminate was made of 12 layers of 45° ply orientation (each layer) used. Two materials with different material mode were used to impact the laminate. The simulation was carried out at 6 J of impact energy with spherical impactor of diameter 20 mm and mass 2.045 kg for stiffer and ductile composites.

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The displacement from the numerical analysis was 0.835 and 1.786 mm, respectively. From numerical simulation, it was concluded that elastic material absorbed more shock than the stiffer composite, and hence, the delamination occurrence was reduced. However, in the case of stiffer composites, shock propagation leads to delamination even though the material showed high strength. Vo et al. [56] established a numerical model to evaluate the flexural behaviour of laminates and sandwiched composite by using quasi-3D theory. Further, layup combinations, boundary conditions and span-to-height ratios were examined. Predicted results were then compared with that of the existing models available in the literature and found a good correlation between each other. Finally, it was concluded that the laminated deformation was more pronounced in the case of clamped–clamped boundary condition with asymmetric lay-up than others, and laminates composite was also greatly influenced by the lay-up sequence and span-to-height ratio. Topal et al. [57] enhanced a numerical method called teaching–learning-basedoptimization method (TLBO) to determine the simply supported beam frequency for an asymmetric FRP laminate. The model was developed using the first-order shear deformation theory and compared with artificial bee colony (ABC) algorithm. All numerical models were developed in MATLAB platform. From numerical simulation results, it was evident that the new method was better, accurate and took less computational time than the ABC method. Soufeiani et al. [58] optimized the carbon fibre ply orientation to study the dynamic response under human walking loading. Ply orientation used were 0°, ±45° and 90°, and dynamic evaluation parameters were natural frequency, displacement and peak acceleration. Analysis and optimization were carried out on ANSYS and MATLAB, respectively. From the analysis, it was found that natural frequency depends on stiffness matrices which were greatly influenced by 0° and ±45° ply orientation. Out of 256 cases, 54 optimal cases were obtained from the optimization process and were in acceptable range according to AIS and ACI serviceability standards. Single and multiple impact behaviours of CFRP cables under pretension condition using drop weight impactor were examined by Xiang et al. [59]. Samples consist of seven CFRP cables and two reactive powder concretes. Impact tests were conducted using a hemispherical nose impactor of diameter 20 mm and mass of 180 kg. Samples were subjected to single impact under 40 kN pretension showed 16 kN transverse impact resistance. The energy absorption capacity for a single impact was 727 J, and it reduced to 367 J and 458 J after two successive impacts. Transverse resistance, cable tension, transverse stiffness and energy dissipation capacity were reduced by 37%, 30%, 39% and 20%, respectively. Finally, it was concluded that there were three basic damage patterns observed in CFRP cables, i.e. complete, partial fracture and indentation. In this work, Alam et al. [60] investigated the effect of parameters such as FRP type, FRP wrapping direction, CFRP wrapping layers, wrapping length and impact velocity on concrete-filled steel tubular (CFST) with carbon fibre (CFRP) and glass fibre (GFRP) laminates under lateral impact loading conditions. Drop hammer was used to carry out the lateral impact test at mid-span with a flat-headed hammer of 592 kg mass and 100 mm diameter from a height of 550 and 1274 mm to achieve an

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impact velocity of 3.28 and 5 m/s, respectively. From experimental results, it was clear that concrete-filled steel tube wrapped by two CFRP layers in longitudinal direction yielded minimum global deformation energy, i.e. 2.9 kJ and found 18.7% reduction in the displacement of CFST strengthened by FRP. However, severe debonding was absorbed in CFRP strengthened CFST samples. Thus, to minimize the effect, GFRP and CFRP wrapped can be effectively used. CFRP laminates were used to minimize the damage and failure of CFST under the lateral impact. Experimental and numerical transverse impact behaviour investigations were carried out on braided I-beam composites by Zhou et al. [61]. Laminates were fabricated using vacuum assisted resin transfer moulding using carbon fibre and epoxy resin, followed by three-stage curing process: one at 90 °C for two hours, 110 °C for one hour and 130 °C for four hours. Transverse impact test was conducted on split Hopkinson impact test at 0.2, 0.4 and 0.6 MPa pressure, which generates impact velocity of 7, 12 and 17 m/s. Eight-noded C3D8R reinforcement (carbon) and fournoded C3D4 elements were used to model the laminate, and the damage modes (initiation, propagation) and failure modes were conducted according to J2 isotropic plastic theory. From the experimental and numerical results, it was clear that the energy absorption capacity of the laminate increased with increase in impact energy. For 0.2, 0.4 and 0.6 MPa of pressure, the energy absorption capacity was 0.2, 48.12 and 99.68 J, respectively. From the analysis, it was evident that the stress progression in laminate greatly influenced by braiding structure. Numerical and experimental results were in good correlation. Effect of matrix crack density on carbon nanotubes reinforced plates under nonlinear low velocity impact loading was examined by Fan et al. [62] where the plate rested on visco-pasternak foundation. Two stacking sequences [0°C /90°F /0°C /90°F /0°C ] and [90°F /0°C /0°C /0°C /90°F ] were used. Refined selfconsistent model, modified Hertzian model, shear deformation and von Karman models were used to evaluate the laminate stiffness degradation of the laminate by matrix cracking, impactor-plate contact and motions equations of the plate, respectively. From the analysis, the following denouncements were made, such as matrix cracks play the least significant role in contact force. Deflection at the centre of the plate can be reduced effectively by increasing the foundation stiffness, and functionally graded carbon nanotubes composites showed a significant effect on laminates under low velocity impact. Zhao et al. [63] investigated the impact localization of CFRP composite under low velocity impact loading. Genetic algorithm and hybrid particle swarm optimization method were used to predict the optimum impact localization, which was used for triangulation localization technique and wavelet transform signals received by strain gauge (Fig. 4.12). Here, CFRP panels with [0°/90°]4 s and [0°/45°/−90°45°/0°/90°/−45°/0°] orientations were used. The experimental results were in better correlation with the optimization method used. Choi et al. [64] investigated the behaviour of T300 and T800 carbon fibre composites. Eight plies of unidirectional laminates were prepared with [0°/90°]4s orientation with epoxy resin. Impact test was conducted using a drop tower impact machine with hemispherical impactor of nose radius 6.5 mm and 3.44 kg of

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Fig. 4.12 Signal time frequency diagram [63]

Fig. 4.13 Comparison of energy absorption of T800 and T300 carbon fibres [64]

mass. The drop height was 0.7 m, impact velocity was 3.71 m/s and 23.67 J of impact energy. Numerical analysis was carried out using LS-DYNA with materials card MAT_COMPOSITE_FAILURE_SOLID (MAT59). By analysing the results, it was concluded that less fibre breakage observed in T800 than T300 carbon fibres (Fig. 4.13) and T300 showed wider delamination area than T800, and T300 carbon fibre absorbed more energy than T800. Both numerical and experimental results were in good correlation. Kamar et al. [65] tried to improve the mechanical properties of carbon fibre reinforced with styrene-butadiene methyl methacrylate (SBM) and also studied the effect of fibre sizing on the laminate properties. Fibre sizing was done by UV-ozone (UVO) treatment process at varying parts per hundred resin (phr). Scanning electron microscope (SEM) was used to determine the void content and fracture mechanism of the samples. Flexural strength, density and Mode I tests were performed according to

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the ASTM D790, D792 and D5528, respectively. The dynamic mechanical analysis was carried out at 3 °C/min from room temperature to 250 °C. From experimental results, it was clear that UVO-treated 10 phr and 15 phr carbon fibre reinforced with SBM performed better than others. There was no effect of sizing on glass transition temperature. Mode I fracture toughness of CFRP was enhanced by 290% at 10 phr. From SEM analysis, it was concluded that the enhancement was due to the nanoscale cavitation, void growth and concomitant matrix shear yield. Alzeanidi et al. [66] investigated the impact behaviour of CFRP laminates with 4.125 mm and 2.625 mm thick under selected impact velocity (100– 500 m/s) and developed a finite element model which gave the guidelines to repair the laminates. Composite panels were made of 7 and 11 layers of woven fibre with [(0°/90°)/(±45°)/(0°/90°)/(±45°)/(0°/90°)/ (±45°)/(0°/90°)/ (±45°)/(0°/90°)/(±45°)] and [(±45°)/(0°/90°)/(0°/90°)/(±45°)/(0°/90°)/(0°/90°)/ (±45°)] lay-up and cured for 100 min at 180 °C under 100 psi pressure. The air gun was used for high velocity impact testing with an impactor of 11.97 ± 0.01 mm diameter was made of stainless steel projected with an impact energy of 141 J. In numerical modelling, MAT_ELASTIC_PLASTIC_HYDRO and MAT_COMPOSITE_FAILURE_SOLID_ MODELs were used to model the projectile and laminate, respectively. Along with high velocity impact, the residual strength of the repaired laminates was also evaluated by compression after impact. From the experimental and numerical analyses, it was concluded that the numerical and experimental results were in close agreement with kinetic energy and shape of the damaged area. Liu et al. [67] investigated the impact damage resistance of hybrid carbon fibre laminate and validated it by numerical simulation. The hybrid laminate was made by unidirectional carbon fibre sandwiched between five harness satinweavedcarbonfibre(5HS). Theplydesigns were[5HS/−45°/+45°/90°/0°/−45°/+45°/ 90°/+45°/−45°/90°/+45°/−45°/0°/90°/+45°/−45°/5HS] {denoted as A} and [5HS/ 0°/0°/+45°/−45°/0°/0°/0°/−45°/+45°/0°/0°/0°/+45°/−45°/ 0°/0°/5HS] {denoted as B}. Resin infusion under flexible tooling method was used to fabricate the hybrid laminates and impact test was carried out on a drop weight tower machine according to ASTM D7136 standards with a hemispherical nose impactor of diameter 12.7 mm and mass 6.4 kg made of steel. Computational damage model was developed using the interlaminar damage model, which accounts unidirectional and woven fibre. This developed model was executed using ABAQUS/Explicit numerical analysis tool. In Lay-up A, the indentation depth was maximum, i.e. 0.218 ± 10.1 mm at 25 J and minimum, i.e. 0.075 ± 6.7 mm at 10 J. Similarly, for Lay-up B, maximum indentation depth was maximum, i.e. 0.185 ± 18.9 mm at 25 J and minimum, i.e. 0.133 ± 17.3% at 10 J of impact energy. Damage area was maximum for Lay-up A (1904.7 ± 2.73%) at 25 J, which was higher than Lay-up B at the same impact energy. As the impact energy increased, the corresponding damage area was also increased. The lay-up designs were also played a crucial role in damage resistance. Results obtained from the developed damage model for hybrid laminates made of the same material but with different architecture were in good agreement with the experimental results. Finally,

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woven plies as skin faces can reduce the damage extent in the laminate under impact loading. Uniform, linear and nonlinear distributions of moisture and temperature effect on CNT reinforced composites plates were studied by Ebrahimi et al. [68] under eccentric low velocity impact loading. FEM analysis was carried out with the help of higher order shear deformation and Karman geometrical nonlinearity along with Halpin Tsai model to calculate the material properties. Simply supported boundary condition was applied to the plate. From results, it was observed that the CNTs addition to the polymer increased its impact resistance, but doping percentage was limited to 1–2 wt.%. The central deflection was increased as the temperature and moisture contraction were increased.

4.2 Low Velocity Impact Test on Hybrid Carbon Fibre Reinforced Polymer Composite Laminates Harris and Bunsell [69] conducted Charpy impact on carbon and glass fibre hybrid composite laminates. Fracture energy of the hybrid composite was dependent on carbon and glass fibre content. Further, the notch dimensions directly influenced the failure modes of carbon and glass fibre hybrid composite laminate. Naik et al. [70] studied the impact and post-impact behaviour of glasscarbon/epoxy hybrid composites Fig. 4.14 for different stacking sequencing by using drop weight impact test and NASA 1142 post-impact compressive test fixture. Plainweaved E-glass and twill-weaved carbon fibres were used. Five FRP samples were made, including neat glass and carbon fibre. GFRP and CFRP were made of 19 and

Fig. 4.14 Damage shape and size for different carbon glass/epoxy laminates [70]

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13 layers, and hybrid laminate sequencing were as follows: [C4/G4]S , [G4/C4]S and [G/C]4S . From the analysis, it was concluded that hybrid composites have less notch sensitivity compared to GFRP and CFRP. In this research, Mamalis et al. [71] worked on thick-walled CFRP components made of hybrid sandwich composite. The crash behaviour and energy absorption of the component under axial compressive loading were studied using LSDYNA. Numerical results were in good agreement with the experimental results and concluded that LS-DYNA could be used as a useful tool to simulate and study the impact damage mechanism of a composite with less time and limited resources. Hosseinzadeh et al. [72] examined the impact behaviour of CF/epoxy, GF/epoxy and CF/GF/epoxy hybrid composite laminates. The experimental results were validated with numerical analysis. The Chang-Chang failure criteria was used to capture the composite damage. The impact energy applied was varied from 30 to 100 J. Further, two different impactor mass 2.5 kg and 5.5 kg were considered for impact energy 50 J. The CF/epoxy laminate showed no damage for impact energy 30 J. However, the CF/epoxy plate was collapsed suddenly at impact energy 50 J. Further, upon addition of GF plies to the CF, the damage resistance improvement was observed. Finally, the numerical analysis predicted damage threshold with good accuracy. However, the predicted failure modes were not in agreement with the experimental observation. The impact behaviour of woven aramid-carbon fibre folded (zigzag pattern) laminates was studied by Heimbs et al. [73] under low and high velocity impact loadings. Drop weight tower method was used to test the energy absorption capacity of the laminates with a hemispherical steel impactor of diameter 25.4 mm and mass 1.56 kg (Fig. 4.15). During testing, the impact energy was varied from 5 J (at velocity 2.5 m/s) to 60 J (at velocity 9 m/s). From experimental results, it was observed that woven aramid and carbon fold cores showed ductile and brittle behaviours, respectively. Zigzag structure inhibited the delamination growth in the laminate during impact testing. The LS-DYNA emerged as an efficient numerical analysis tool for analysing the impact damage mechanism of composites due to its flexibility and accuracy. In this research investigation, Badie et al. [74] studied the effect of natural frequency, buckling strength, fatigue strength and torsional strength on a composite drive shaft made of carbon, glass and carbon/glass hybrid laminates. In this study, six different laminates [±45°]4 glass, [±45°]4 carbon, [90°]4 glass, [90°]4 carbon,

Fig. 4.15 a LS-DYNA numerical simulation and b Experimental samples after the impact [73]

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Fig. 4.16 Flax Z-pinning arrangement in hybrid composite [75]

[(45°)2 glass/ (90°)2 carbon] and [(45°)2 carbon/ (90°)2 glass] were prepared using hand lay-up technique and cured at room temperature. Finite elemental analysis was done using LUSAS tool in which Q158 shell elements made of quadratic and quadrilateral elements were used. From results, it was observed that the natural frequency, buckling torque, fatigue life and torsional strength were increased as fibre orientation was increased to 90° and ±45°, respectively. Finite element results showed a deviation of 0.78% and 9.8% with respect to natural frequency and buckling torque when compared to the analytical solution. An effect of stacking sequence is seen in the case of buckling strength and fatigue life of laminate, but the same was not true for natural frequency. Hybrid laminates behaviour made of natural flax Z-pinning (Fig. 4.16), carbon and glass fibre under low velocity impact using Charpy impact tester was investigated by Ghasemnejad et al. [75]. Two hybrid composites with [C0°/G90°]3 and [C90°/G0°]3 laminates were fabricated. The LS-DYNA was used for numerical analysis where the laminate was modelled by using quadrilateral shell element. MAT_54 material card was used to model the laminate. Both experimental and numerical results were in good correlation and concluded that the [C90°/G0°]3 ply orientation yielded the maximum impact resistance. Numerical and experimental studies of shape memory alloys embedded in composite were carried out by Meo et al. [76] using drop weight impact testing and LS-DYNA. Two specimens PPS/CF and PPS/CF/SMA were fabricated by prepreg technique assisted by hydraulic compression. Drop weight impact machine with a hemispherical tip impactor of mass 6.84 kg was used during testing. PPS/CF was modelled using Chang-Chang failure criterion with MAT30 (MAT_SHAPE_MEMORY) material card. (Fig. 4.17). From results, it was observed that due to the increase in elasticity, the hybrid laminates absorbed more energy and

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Fig. 4.17 LS-DYNA and experimental results comparison of PPS/CF at 24.56 J [76]

reduced ply failure. It was concluded that high strain to failure property of SMA and its ability to recover strain stress were the key factor behind the enhancement in hybrid laminate properties. Pandya et al. [77] carried out ballistic impact test on satin-weaved carbon fibre and plain-weaved glass fibre hybrid composites. Total six laminates were fabricated out of which two were [glass/epoxy] and [carbon/epoxy] followed by four hybrid laminates [G3 /C2 ]S , [C2 /G3 ]S , [C2 /G2 ]S and [G2 /C2 ]S . A single-stage gas gun with hardened steel spherical projectile was impacted onto the fabricated laminate. Glass/epoxy and carbon/epoxy laminates were absorbed 30.83 J and 21.06 J of impact energy, respectively. Meanwhile, the hybrid laminate [C2 /G2 ]S and [G2 /C2 ]S absorbed 22.65 and 24.30 J of impact energy. From results, it was evident that hybrid laminate in which glass/fibre as exterior and carbon fibre as the interior layer performed better than other hybrid and neat laminates of the thickness. Dynamic analysis of C/Sic composite was examined by Li et al. [78] under split Hopkinson pressure bar (SHPB) and gas gun. Failure mechanism, void contents and interfaces were studied by using several microstructure characterizations such as X-ray chromatography, dynamic X-ray phase contrast imaging (XPCI), scanning electron microscopy and optical velocimetry. XPCI was used to examine the microstructure of C/Sic composite under dynamic compression experiment at 102 to 103 s−1 with the help of SHPB. Gas gun experiment was carried out at a higher strain rate (104 to 105 s−1 ) from this void collapse and spallation was studied. Fracture mechanism of C/SiC laminates under compression and tension was examined by scanning electron microscope. From experimental results, it was concluded that outof-plane loading, compression induces fracture via void collapse and shear damage

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Fig. 4.18 Ultrasonic view of damage area of CFRP and Al/CFRP after low velocity impact [79]

bending. Nevertheless, in the case of in-plane, delamination dominates the fracture in the laminates. In these findings, Bienia´s et al. [79] conducted a comparative experimental study between Al-CFRP hybrid composite and traditional CFRP laminates under low velocity impact. Unidirectional carbon fibre and thin aluminium sheets were reinforced with epoxy resin. Laminates were prepared using autoclave method with a curing cycle of 2 °C/min up to 135 °C, and the laminates were maintained at 135 °C for 2 h. The [02 /902 ] fabric orientation was sandwiched between aluminium sheets on both sides as outer layers. The impact test was conducted using a drop weight impact tester, according to ASTM D7136. Hemispherical impactor with nose diameter 38.1 mm and mass of 1.4 kg was impacted with slightly varying velocities ranging from 1.10 m/s to 2.10 m/s with impact energy levels 1.5 J, 2.5 J and 5 J (Fig. 4.18). From experimental results, it was observed that control CFRP energy absorption mechanism was associated with elastic response and damage of the laminates, whereas the hybrid laminates energy absorption mechanism included an additional parameter that was the plastic deformation of the metal layers and thus resulted in more energy absorption than control CFRP. Moon et al. [80] investigated a new hybrid composite design and subjected to high velocity impact. The design consists of three parts: initial bumper, intermediate fabric layer and a rear plate. CFRP, aluminium, polyethylene, polycarbonate and polymethyl methacrylate were used as a rear plate to fabricate five different laminates. A laminate frame was used to un-strain and provide a gap for intermediate fabric layers so as to allow the fibre pull out mechanism. Finally, it was concluded that carbon/Kevlar/carbon (Fig. 4.19) sequence samples as a bumper, intermediate fabric and rear plate, respectively, showed lowest energy absorption, as the rear plate is stiff and not allowing fabric material to deform completely. The fabric pull out and energy absorption were maximum in samples made of carbon/Kevlar/polycarbonate as a bumper, intermediate fabric and rear plate, respectively. Because the rear plate polycarbonate was weak, it was allowing maximum fabric pull out. Rathore et al. [81] investigated the flexural strength at varying temperature (room temperature, 70, 90, 110 °C) and creep behaviour for carbon fibre (CF), glass fibre

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Fig. 4.19 Carbon/Kevlar/carbon a Back surface, b Inside view, c Penetration hole and d Penetration size [80]

(GF) and carbon/glass hybrid laminates. Samples were made of seven plies with different CF volume fraction. Different lay-up designs were G7 , C7 , G5 C2 , C2 G5 and C2 G3 C2 out of these C2 G3 C2 performed better among other combinations immediately after C7 . Strength and modulus were increased by 83% and 112%, respectively, when compared with G7 at 90 °C. However, this trend was changed at 110 °C because at this temperature, matrix enters the rubbery phase where the polymer chain started to move more freely. However, the carbon content restricted this movement, and hence, improvement was observed in C2 G3 C2 combination. Flexural failure at the bottom layer propagated towards the middle surface. Observed failure modes were debonding, ply splitting and delamination. C2 G3 C2 sample does not fail catastrophically as that of C7, and at a higher temperature, the failure trend was similar for all laminates, and G5 C2 showed lower strength than C2 G5 . At 70 °C, better strength in the laminate was observed due to the free movement of polymer chains. It was observed that the coefficient of thermal expansion (CTE) generates thermal stresses which was dependent on CTE difference and temperature gradient; i.e. fibre expands at a slower rate than the matrix material which leads to debonding and delamination which resulted into microcracking of the matrix. The hybrid laminate performed better than others where tensile and compression were observed in outer layers and shear stresses in mid-plies.

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4.3 Low Velocity Impact Test on Nanomaterial-Doped Carbon Fibre Reinforced Polymer Composite Laminates A comparative study was conducted by Xu et al. [82] between neat epoxy and MWCNTs reinforced thin epoxy films. Elastic modulus of the thin films was tested by using a shaft-loaded blister test. These experimental results were compared with the linear and nonlinear elasticity models. From the results, it was observed that a 20% increase was in the elastic modulus when reinforced with 0.1 wt.% of MWCNTs. From molecular mechanics simulation and micromechanics analysis, it was found that the presence of substantial shear stress was between MWCNTs and epoxy matrix. Enhancement in mechanical properties of two (carbon nanofibre + resin) and three (carbon nanofibre + carbon fibre + resin) phase composites due to the addition of cup stacked nanofibre was investigated by Iwahori et al. [83]. Further, the effect of aspect ratio (10, 50) and weight percentage (5%, 10%) of cup stacked nanofibre was also investigated. Laminates were fabricated using vacuum bagging method with curing and post-curing temperature of 100 °C and 175 °C for 75 min and four hours, respectively. Tensile, compressive, flexural properties for two-phase (without carbon fibre) CSNF were 3.628 GPa, 2.933 GPa and 2.707 GPa for aspect ratio 10 with weight 10 wt.%, whereas for three-phased CSNF (with CFRP), they were 53.25 GPa, 50.17 GPa and 53.5 GPa, respectively. Similarly, for aspect ratio 50 and 10wt.%, tensile, compressive and flexural properties for two and three-phase laminates were 3.602 GPa, 3.975 GPa, 3.277 GPa and 56.24 GPa, 48.74 GPa, 19.7 GPa, respectively. From experimental results, it was evident that the aspect ratio 50 with 10 wt.% percentage yielded better results than other laminates two-phase composite and decrease in properties were observed in the case of three-phase composites. Enhancement was observed in the case of three-phase composites for aspect ratio 50 and 10 wt.%, but it was not significant. Finally, it was concluded that the tension, compression and flexural enhancement were observed in two-phase CSNF composites. Yaping et al. [84] investigated the flexural and impact strength of CFRP laminates doped with amine group functionalized and pristine MWCNTs. Doping weight percentage ranges from 0.2 to 1.0 wt.% in 0.2 incremental steps. ASTM D256 and ASTM D790 standards were followed for impact and flexural strength testing, respectively. The optimum mechanical values were obtained for 0.6% doping weight percentage. There was 100% improvement in bending strength, 58% improvement in bend modulus, and impact strength was improved by almost two times when compared to the CFRP laminates with pMWCNTs. Impact strength of nanocomposites was 8.5 kJ/m2 and 15.5 kJ/m2 for pMWCNTs and NH2 -MWCNTs at 0.6 wt.%, respectively. Finally, it was concluded that enhancement in MWCNTs-NH2 mechanical properties was superior to pMWCNTs due to improved interfacial bonding (Fig. 4.20). Bekyarova et al. [85] studied the effect of MWCNTs doping by electrophoresis technique on mechanical properties of CF/epoxy composites (Fig. 4.21). Laminates were reinforced with MWCNTs fabricated by vacuum assisted resin transfer

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Fig. 4.20 Effect of functionalization of MWCNTs on impact strength [84]

Fig. 4.21 Electrophoresis technique for depositing CNTs on carbon fibre [85]

moulding. Short beam method was adopted to measure the interlaminar shear strength according to ASTM D2344 standards. From experimental results, 30% improvement in interlaminar shear strength was observed in comparison with the neat CF/epoxy laminates. This enhancement is mainly due to the preservation of in-plane mechanical properties. Meanwhile, enhancement in out-of-plane electrical conductivity was also observed. In this research work, Qian et al. [86] investigated the effect of MWCNTs on the mechanical properties of polystyrene. Ultrasonic assisted solution evaporation method was used to achieve homogeneous dispersion of MWCNTs in polystyrene matrix, and 36–42% enhancement in tensile property of polystyrene was achieved by the addition of 1 wt.% nanotube and 25% increase in elastic modulus as well as break stress when compared with neat polystyrene. It was concluded that the enhancement in mechanical properties was mainly due to the large aspect ratio of MWCNTs. Investigation of CNTs reinforced with epoxy laminate under impact loading was carried out by Visco et al. [87] according to ASTM D256 standards. From experimental results, it was seen that the curing rate was improved in the presence of CNTs when compared to pristine resin. It was concluded that up to 0.5% of doping of CNTs resulted in increased kinetic reaction properties of the resin, but beyond this weight percentage, mechanical performance and reduced molecular chain mobility.

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Fig. 4.22 Compression after impact versus impact energy for both non-doped and doped samples [88]

Kostopoulos [88] studied the effect of MWCNTs doped with quasi-isotropic carbon fibre reinforced with epoxy laminate when subjected to impact and after impact. Dispersion of MWCNTs in the matrix was carried out by shear mixing process, and CFRP laminates were fabricated by using autoclave process assisted by vacuum bag technique at 10 bar pressure. Two curing cycle phases were used, the first phase was for 60 min at 80 °C, and the second phase was for 240 min at 140 °C. The laminate contains unidirectional carbon fibre with quasi-isotropic orientation [0°/+45°/90°/−45°]2S with a thickness of 2.4 ± 0.1 mm. Drop weight tower impactor was used with hemispherical aluminium impactor of diameter 20 mm with 3.01 kg of mass, and test was carried out according to the ASTM D5628-07 standard. Five impact energy levels 2, 8, 12, 16 and 20 J were used during impact testing. The ultrasonic method was used to examine the extent of damage in the laminates after impact. Experimental data suggest that MWCNTs had a more pronounced effect and can be used as reinforcing material at higher impact energy levels. MWCNTs improved the strength and modulus of composites after impact compared with neat resin. Finally, it was concluded that MWCNTs would affect only when laminates were subjected to higher energy levels and significantly improved the compression after impact and compression-compression fatigue after impact performance of CFRP laminates (Fig. 4.22). Davis et al. [89] used the functionalized and non-functionalized single-walled carbon nanotubes (SWCNTs) to enhance the mechanical properties of CFRP laminates. Composites were fabricated using heated vacuum assisted resin transfer moulding and cured at 122 °C for two hours, followed by post-curing at 177 °C for two more hours. SWCNTs content was varied for 0.2 to 0.5 wt.%, and tensile strength was improved by 10 and 19% for strength and stiffness for a-SWCNTs over

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controlled CFRP laminate, whereas 0.2 wt.% SWCNTs showed strength and stiffness improvement by 7% and 13% over neat laminates, respectively. Better tensiontension fatigue life improvement was observed for a-SWCNTs over neat CFRP. From scanning electron microscopy and optical microscopy analyses, it was concluded that the enhancement in properties was due to the enhanced interfacial bonding between fibre/matrix due to the presence of a-SWCNTs. In this research experiment, Ashrafi et al. [90] used a commercially available unidirectional carbon fibre prepreg with a modified epoxy system which contains 0.1 wt.% and 0.2 wt.% of surface modified SWCNT. Impact, compression after impact, Mode I and Mode II tests were conducted according to ASTM D7136, ASTM D7137, ASTM D5045 and D5528, respectively. Drop weight impact tester with hardened steel hemispherical impactor of 15.9 mm diameter with 6.3 kg of impactor weight was used. From experimental results, it was found that for 0.1 wt.% SWCNT reinforcement 5% reduction in damage area, CAI strength was improved by 3.5%, Mode I toughness was increased by 13%, and 28% increase in Mode II fracture toughness was found when compared with the laminates without SWCNTs. In this research experiment, Soliman et al. [91] studied on- and off-axis tensile strength of CFRP laminates when doped with pristine and functionalized MWCNTs. Varying MWCNTs weight percentage from 0.1%, 0.5%, 1.0% and 1.5% of pristine as well as –COOH was doped into the laminates and fabricated using hand lay-up technique consisting of six layers of fabric. ASTM D5687, D3039, D3518 and D6272-10 standards were followed to carry out laminate fabrication, in-plane shear testing, offaxis tension testing and flexural testing, respectively. From the experimental results, it was found that failure strain rate, ultimate strength and toughness of off-axis tension strength are increased for samples doped with functionalized MWCNTs (1.5 wt.%). From TGA analysis, it was concluded that the functionalized MWCNTs form better and strong interfacial bonding between matrix and reinforcement, which results in significant improvement in properties of CFRP composite. In this experimental work, Joshi et al. [92] studied the interlaminar fracture toughness of MWCNTs reinforced with CFRP laminates by conducting Mode I and Mode II tests. Mode I and Mode II experiments were carried out by double cantilever beam and end-notch flexural test, respectively. Six layers of woven CFRP composite with [0°/90°] orientation were fabricated by using hand lay-up method. From experimental data, at 1.32 g/m2 of MWCNTs, optimum fracture toughness was observed. It was concluded that MWCNTs enhanced the interlaminar fracture toughness only at the optimum weight percentage of CNTs per unit area of prepreg, and the improvement was due to the interlocking and formation of bridging between matrix and carbon fibre by MWCNTs. Soliman et al. [93] used the multi-walled carbon nanotubes (MWCNTs) to investigate the behaviour of laminate under impact loading. Here, five energy levels: 15, 24, 30, 60, 120 J and three volume fractions of MWCNTs: 0.5%, 1.0% and 1.5% were analysed. CFRP laminates were prepared by using hand lay-up technique assisted by vacuum bagging technique according to the ASTM D5687 standards. Each laminate contained ten layers of carbon fibres with an orientation of (0°–90°). Drop tower impact tester was used for impact testing with a hemispherical striker of diameter

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Fig. 4.23 a Penetration energy versus CNTs weight % in composite b Damage height versus CNTs weight % in composite [93]

Fig. 4.24 X-ray radiographs of impacted samples a CFRP and b MWCNTs [94]

12.7 mm and 0.63 kg of mass. From experimentation, it was observed that below penetration limit, 50% improvement in penetration energy was seen for 1.5% COOHMWCNTs by weight of epoxy. The reinforcement showed no improvement above the penetration limit regardless of applied energy, and penetration limit decreased as the applied energy increased. From the study, it was concluded that MWCNTs content in CFRP significantly reduced the damage size and is effective only below the penetration limit under low velocity impact (Fig. 4.23). In this research findings, Tehrani et al. [94] studied the impact resistance and damage progression of MWCNTs infused with plain-woven carbon fabric and epoxy laminates. Five-layered laminates were fabricated by using the hand lay-up method assisted by vacuum bagging technique. ASTM D4065 and D790 standards were followed to prepare the laminate samples, and a single-stage gas gun is used with speed up to 500 ms−1 with an impactor of mass 17 g. MWCNT-infused CFRP laminates showed multiple-stage delamination. However, the neat CFRP laminates absorbed less energy for penetration with abrupt delamination. CFRP energy absorption capability is improved by 21%. X-ray radiography was used to study the impact damage mechanism, and from analysis, it was concluded that MWCNTs have greater capability to attenuate impact shocks generated during impact (Fig. 4.24). MWCNTs improved interlaminar and intralaminar performances, which resulted in improved

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Fig. 4.25 Optical microscopy analysis for impact damage of a CF/epoxy, b CF/epoxy/CNTs, c CF/epoxy/CNT aged and d CF/epoxy/CNT functionalized [95]

impact resistance. From experimental result analysis, it was concluded that intermediate dispersion of MWCNTs in the matrix material could yield better improvement in the mechanical properties of CFRP laminates with negligible increase in weight of the composites. Siegfried et al. [95] investigated the out-of-plane low velocity impact test for three different types of CFRP laminates doped with non-functionalized, aged and functionalized CNTs. Functionalization of CNTs was done with an amino group, and twill 2/2-woven carbon fibre and epoxy resin were used. Resin transfer moulding process is used to fabricate CFRP laminate with fabric orientation of [(+45°/−45°), (0°/90°)]4S . Precuring was done at 70 °C for 60 min, and post-curing was done at 150 °C for 60 min. Drop weight impactor is used for impact testing, according to ASTM D7136. A hardened hemispherical steel impactor with a tip of diameter 16 mm and mass of 5.536 kg was impacted from a height of 0.49 m to achieve the impact energy of 26.8 J. C-scan was used to detect the delamination in laminates after impact. From experimental results, it was observed that functionalization of CNTs helps in better dispersion. Impact properties of CNT-doped laminates were improved due to the bridging mechanism of cracks and CNTs pull out (Fig. 4.25). It was concluded that the dispersion of CNTs in resin plays a crucial role in enhancing the mechanical properties of laminates, and aged CNTs performed better than others. Islam et al. [96] carried out experimental work on dynamic mechanical analysis and low velocity impact test at 2 wt% nanoclay, and 0.3 wt% MWCNTs was doped separately with carbon fibre reinforced with epoxy. Laminates were fabricated using hand lay-up technique and cured for 4 h in a compression mould and again cured at 120 °C with the pressure of 120 kN. Drop weight impact tester was used for low velocity impact testing of laminates at varying energy levels 30, 40 and 50 J. From experimental data, it was found that all samples infused with nanomaterials performed better than the laminates which were not doped. The energy absorption capacity of nanoclay and MWCNTs at 30 J was 29.11 and 31.15, and for 40 J, it was

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Fig. 4.26 Thermoscopic damage analysis of a, b Control laminate, c, d Nanoclay-modified composite e, f MWCNTs-modified composites [96]

41.05 and 41.87, respectively. There was no much difference in the energy absorption capacity of MWCNT and nanoclay-doped laminates, and still, MWCNT performed better than nanoclay (Fig. 4.26). Low velocity impact resistance of CFRP laminates doped with MWCNTs was investigated by Singh et al. [97]. Doping wt.% ranges from 0 wt.% to 5 wt.%. A quasi-isotropic laminate [(0°/90°)/(+45°/−45°)/(+45°/−45°)/(0°/90°)//(+45°/−45°) /(0°/90°)/(0°/90°)/(+45°/−45°)] was fabricated by hand lay-up method assisted by vacuum bagging. Dispersion of MWCNTs in epoxy was carried out by probing ultrasonication method for 1 h. The test was conducted on drop weight impactor according to ASTM D7136 standards with a hemispherical nose impactor of diameter 20 mm, impact velocity of 6 m/s and impact energy of 94.14 J, respectively. From experimental results, it was observed that 2 wt.% MWCNT-doped CFRP absorbed more energy (Fig. 4.27) than other laminates. Pyramidal height and damage for 2 wt.% MWCNTs-doped CFRP were 5 mm and 19.6 mm, respectively, which were the lowest compared to neat and other doped CFRP composites. Finally, it was concluded that the enhancement in CFRP impact resistance for 2 wt.% doping was due to the formation of bridging and inhibition of crack growth by MWCNTs. However, enhancement was not observed in the case of 5 wt.% doping; the impact resistance property of the CFRP was reduced due to agglomeration of MWCNTs which inhibited the formation of bridging and introduces the void content in laminates. In this research work, Rawat et al. [98] investigated the optimal doping percentage of MWCNTs in CF/epoxy composites to enhance the damage tolerance. Doping of MWCNTs was ranged from 0.25 wt.% to 1.0 wt.% with an increment of 0.25 wt.%. Hand lay-up method assisted by vacuum bagging was

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Fig. 4.27 Energy absorption of asymmetric CFRP laminates a Neat laminate b 2 wt.% of MWCNTs c 5 wt.% of MWCNTs [97]

used to fabricate eight-layered CFRP with symmetric orientation of [(0°/90°)/ (+45°/−45°)/(+45°/−45°)/(0°/90°)//(0°/90°)/(+45°/−45°)/(+45°/−45°)/(0°/90°)]. Dispersion of MWCNTs in the epoxy was done by using sonication for 1 h. Drop weight tester with hemispherical nose impactor of diameter 12.7 mm and mass of 10 kg was used during impact testing. Testing was carried out according to ASTM D7136 standard with impact energy and velocity of 94.14 J and 6 m/sec, respectively. From experimental results, it was observed that 0.25 wt.% MWCNTs-doped laminate absorbed the highest impact energy (Fig. 4.28) than any other laminates. Damage in weft and warp for the same doping wt. % was 17.12 and 17.12 mm, respectively, which was the lowest than other CFRP-doped composites with pyramidal damage height of 9 mm. Experimental results showed that adding MWCNTs in CFRP laminates improved the damage resistance property. There was an optimal nanomaterial doping value into the laminate beyond which no enhancement is observed.

4.4 Summary The carbon fibre reinforced polymer composite laminates impact performance majorly depends on the stacking sequence, impact energy, matrix material and fabric architecture. Further, the impact performance of CFRP composite laminates can be improved by fibre hybridization with glass, basalt, Kevlar, etc. Both interply and

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Fig. 4.28 Energy versus time graph for comparing energy absorption of various laminate doped with different weight fraction of MWCNTs [98]

intraply hybridizations increase the impact performance of CFRP composite laminates. Furthermore, use of nanomaterials such as CNTs, and nanoclay also enhances the impact strength of the CFRP composite laminates.

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66. Alzeanidi N, Ghasemnejad H (2017) Analysis of the effect of impact damage on the repairability of composite panels. Compos Struct 176:304–312. https://doi.org/10.1016/j.compstruct.2017. 05.054 67. Liu H, Falzon BG, Tan W (2018) Experimental and numerical studies on the impact response of damage-tolerant hybrid unidirectional/woven carbon-fibre reinforced composite laminates. Compos Part B Eng 136:101–118. https://doi.org/10.1016/j.compositesb.2017.10.016 68. Ebrahimi F, Habibi S (2018) Nonlinear eccentric low-velocity impact response of a polymercarbon nanotube-fiber multiscale nanocomposite plate resting on elastic foundations in hygrothermal environments. Mech Adv Mater Struct 25:425–438. https://doi.org/10.1080/153 76494.2017.1285453 69. Harris B, Bunsell AR (1975) Impact properties of glass fibre/carbon fibre hybrid composites. Composites 6:197–201. https://doi.org/10.1016/0010-4361(75)90413-9 70. Naik NK, Ramasimha R, Arya H et al (2001) Impact response and damage tolerance characteristics of glass-carbon/epoxy hybrid composite plates. Compos Part B Eng 32:565–574. https:// doi.org/10.1016/S1359-8368(01)00036-1 71. Mamalis AG, Manolakos DE, Ioannidis MB, Kostazos PK (2003) Crushing of hybrid square sandwich composite vehicle hollow bodyshells with reinforced core subjected to axial loading: numerical simulation. Compos Struct 61:175–186. https://doi.org/10.1016/S0263-822 3(03)00069-2 72. Hosseinzadeh R, Shokrieh MM, Lessard L (2006) Damage behavior of fiber reinforced composite plates subjected to drop weight impacts. Compos Sci Technol 66:61–68. https:// doi.org/10.1016/J.COMPSCITECH.2005.05.025 73. Heimbs S, Cichosz J, Klaus M et al (2010) Sandwich structures with textile-reinforced composite foldcores under impact loads. Compos Struct 92:1485–1497. https://doi.org/10. 1016/j.compstruct.2009.11.001 74. Badie MA, Mahdi E, Hamouda AMS (2011) An investigation into hybrid carbon/glass fiber reinforced epoxy composite automotive drive shaft. Mater Des 32:1485–1500. https://doi.org/ 10.1016/j.matdes.2010.08.042 75. Ghasemnejad H, Soroush VR, Mason PJ, Weager B (2012) To improve impact damage response of single and multi-delaminated FRP composites using natural Flax yarn. Mater Des 36:865– 873. https://doi.org/10.1016/j.matdes.2011.05.018 76. Meo M, Marulo F, Guida M, Russo S (2013) Shape memory alloy hybrid composites for improved impact properties for aeronautical applications. Compos Struct 95:756–766. https:// doi.org/10.1016/j.compstruct.2012.08.011 77. Pandya KS, Pothnis JR, Ravikumar G, Naik N (2013) Ballistic impact behavior of hybrid composites Kedar. Dyn Deform Damage Fract Compos Mater Struct 44:128135. https://doi. org/10.1016/B978-0-08-100080-9.00015-4 78. Li T, Fan D, Lu L et al (2015) Dynamic fracture of C/SiC composites under high strain-rate loading: microstructures and mechanisms. Carbon N Y 91:468–478. https://doi.org/10.1016/j. carbon.2015.05.015 79. Bienia´s J, Jakubczak P, Surowska B, Dragan K (2015) Low-energy impact behaviour and damage characterization of carbon fibre reinforced polymer and aluminium hybrid laminates. Arch Civ Mech Eng 15:925–932. https://doi.org/10.1016/j.acme.2014.09.007 80. Moon J-B, Yoon SH, Kim C-G (2016) High velocity impact test of a hybrid sandwich composite shield with unrestrained boundary fabric. Compos Struct 153:60–68. https://doi.org/10.1016/ j.compstruct.2016.05.088 81. Rathore DK, Prusty RK, Mohanty SC et al (2017) In-situ elevated temperature flexural and creep response of inter-ply glass/carbon hybrid FRP composites. Mech Mater 105:99–111. https://doi.org/10.1016/j.mechmat.2016.11.013 82. Xu X, Thwe MM, Shearwood C, Liao K (2002) Mechanical properties and interfacial characteristics of carbon-nanotube-reinforced epoxy thin films. Appl Phys Lett 81:2833–2835. https:// doi.org/10.1063/1.1511532 83. Iwahori Y, Ishiwata S, Sumizawa T, Ishikawa T (2005) Mechanical properties improvements in two-phase and three-phase composites using carbon nano-fiber dispersed resin. Compos Part A Appl Sci Manuf 36:1430–1439. https://doi.org/10.1016/j.compositesa.2004.11.017

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Chapter 5

Low Velocity Impact Test on Glass Fibre Reinforced Polymer Composite Laminates

Abstract The glass fibre reinforced composite laminates are widely used due to their reasonable strength properties at low cost. In this chapter, various aspects of the GFRP composite under LVI are covered. The first section of this chapter covers the literature works that deal with the pure GFRP composite laminates where different parameters that affected the impact properties and failure mechanism of GFRP composites are covered. Parameters such as stacking sequence, impactor geometry, impactor diameter, GFRP plate thickness and matrix material. The second section of this chapter deals with the effect of hybridization of GFRP composite materials with aramid, flax, hemp, jute and carbon fibres under LVI. The final section of this chapter briefly accounts the doping of nanomaterials to the GFRP composite laminates and their influence on the impact properties and damage mechanism. Keywords GFRP · LVI · Stacking sequence · Laminate thickness · Impactor shape · Impactor diameter

5.1 Low Velocity Impact Test on Controlled Glass Fibre-Reinforced Polymer Composite Laminates Morris et al. [1] studied the impact resistance properties of GF/polyester and bornfibre/aluminium composite laminates used in turbine blades. It was found that the internal damages, delamination and debonding between fibre-matrix were the major failure modes observed at low impact energies which significantly reduced the bend strength of the composite laminates. Husman et al. [2] studied the after impact tensile strength of the GF/epoxy composite laminate. The experimental results were compared with the tensile strength of the GF/epoxy composite laminate containing known and artificially inserted flaw. The prediction of residual tensile strength was accurate for both impact-induced and flaw-induced GF/epoxy composite laminates impacted well within the penetration limit. Further, the GF/epoxy composite laminate containing flaw with size same as impact-induced damage showed corroborative residual tensile strength. Lifshitz et al. [3] studied the angle plied GF/epoxy composite laminates impact performance. The angle plied stacking sequences considered for © The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_5

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investigations were ±30°, ±35°, ±41°, ±45°, ±49°, ±55° and ±60°. The experimental results showed that the failure mechanisms of the angle plies GF/epoxy composite laminates changed with the fibre orientation. Further, the theoretical prediction of failure for specimens with fibre orientation above ±45° agreed well with the experimental failure mechanism. Zhou et al. [4] examined the impact strength of thick GF/polyester composite laminates. Two thicknesses 10 and 25 mm were used for the experimentation. The impactor nose shape used was flat end. It was observed that the prediction of impact damage threshold was independent of damage inspection. However, the ultimate impact threshold was dependent on the in-plane geometry. Furthermore, the force and energy characteristics were found to be an effective approach to monitor the damage propagation. Mittal et al. [5] developed a theoretical model to explain the impact resistance properties of GF/epoxy composite laminates. The model considered the fibre volume fraction, laminate thickness, impactor mass and the laminate thickness. Using these parameters, maximum force, contact time and deflections were predicted. Morri et al. [6] examined the influence of resin matrix hybridization on the impact resistance performance of GF composite laminates. Two epoxy resin systems, one was normal epoxy and other was epoxy with glycol acrylate, were used in the experimental investigation. The experimental investigation was conducted by incorporating normal GF/epoxy plies and GF/modified epoxy sheets at different positions in the laminate (Fig. 5.1). The impact energy absorption capacity of the modified epoxy was increased. The experimental results showed that the modified epoxy at impact site showed better impact resistance properties than the normal epoxy at impact site. It was found that the impact energy was consumed only in damage development and propagation. Further, the modified epoxy absorbed the impact energy in damage development, propagation and recoverable deformation. Analysis of repeated impact test on E-glass/epoxy laminates was examined by Kishore et al. [7]. The number of impacts required to cause the tup to penetrate the laminates for varying thickness was investigated. The following points were concluded from the testing: yield strength of the laminate was lower at the impacted zone and as the thickness of the laminate increased the number of impacts required to cause the tup to penetrate inside the laminate also increased. Okoli et al. [8] developed the numerical model to predict the impact, flexural properties of randomly oriented FRP laminates. In this investigation, randomly oriented glass/epoxy laminates were used and tested under low (1.7, 8.3, 17.0, 83.0 mm/s) to high (10, 100, 1000, 2000 mm/s) strain rates on a three-point bend test according to ASTM 3039 standard. Stiffness of the laminate at 4 mm/s was 33.503 GPa and 35.805 GPa for unfiltered and filtered data type for FEA model in the meantime, and it was 43.373 GPa and 39.173 GPa for experimental and numerical testing, respectively. The flexural peak load of unfiltered and filtered data from the FEA model were 1.124 kN and 0.685 kN when compared to experimental results, i.e. 1.152 kN and 1.137 kN, respectively. Finally, it was concluded that a little variation in the results of FEA and experimental was may be due to the Chang–Chang model, which do not account the strain rate. However, a correction function was used to overcome the underestimation of the FEA model prediction, which can improve the accuracy of the model.

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Fig. 5.1 Different configurations of glass fibre reinforced with normal and modified epoxy (‘C’ in unshaded region represents the normal epoxy; ‘F’ in shaded region represents the modified epoxy) [6]

Mechanical properties of acrylic-based ductile polymer reinforced with glass fibre laminates were investigated by Kanie et al. [9]. Different fibres such as unsilaned, silaned glass fibre laminates of varying laminate thickness (1, 2, 3 and 4 mm) were sandwiched between denture powder matrix. Furthermore, laminates were pressed under 2 MPa force and cured at 70 °C for one hour, followed by post-cured at 100 °C for one hour. Flexural and impact tests were carried out on a three-point bend test and a flywheel-type impact testing machine, respectively. Crosshead speed for flexural and impact tests was 2 mm/min and 543 mm/min, respectively. The unsilaned-based D-type laminates with 2 mm and 1 mm thickness showed maximum and minimum flexural strength, respectively, whereas D- and N-based laminate with 4 mm and 2 mm thickness showed maximum and minimum impact strength, i.e. 2142.5 and 1125.5 J/m2 , respectively. In the case of silanized-based laminates with N-types (1 mm thickness) and S-type (4 mm thickness) showed good and poor flexural properties. However, the impact strength of the controlled laminate with 4 mm thickness was better compared to other laminates. Finally, there was an enhancement in flexural and impact strength of L-, W- and T-type laminates by 11, 17, 14% and 31, 58 49%, respectively.

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Scarponi et al. [10] studied the delamination detection in CFRP, GFRP and KFRP composite laminates subjected to LVI using ultrasonic technique. The resin transfer moulding process was used to fabricate all the FRP composite laminates. The ultrasonic non-destructive damage detection process was able to detect the delamination in all the fabricated composite laminates. Further, the resolution of the obtained images was high. Naik et al. [11] tried to predict the matrix cracking and delamination of a FRP composite laminates subjected to LVI for different stacking sequence designs. The stacking sequences included both woven and unidirectional GF/epoxy plies. The in-plane and interlaminar failures were determined by using quadratic failure criteria. Further, the stress state in the FRP composite plate was determined using 3D transient finite element analysis. Use of both unidirectional and woven plies in the laminate showed better impact resistance. It was concluded that the 3D orthogonal woven GF/epoxy composite laminates showed highest impact resistance. Okoli et al. [12] examined the effect of strain rate as a function of fibre volume fraction for randomly oriented glass fibre/epoxy laminates. Two kinds of samples were tested where a 3 mm thick randomly oriented GFRP laminate was subjected to varying strain rate (1.7, 8.3, 17.0, 83.0 mm/s) in which tensile strength of the laminate was found to increase with increase in crosshead speed and at 83.0 m/s tensile strength was maximum, i.e. 318 MPa. The energy required for the failure of GFRP laminate was increased with an increase in fibre content of the laminate. At 26.9%, volume fraction energy required was 7.438 J, whereas above and below this fibre volume fraction energy required to fail the laminate was low. Finally, it was concluded that poor performance of laminate at higher fibre volume fraction might be due to the poor flowability of the matrix in the laminate. It resulted in poor wettability of fibre by the matrix, which led to weak interfacial bonding between matrix and fibre. In this research work, dissipation of kinetic energy of glass fibre during impact loading on unidirectional and woven layers with different fabric orientations was examined by Belingardi et al. [13]. [0/90]s, [0/+60/−60]s and [0/+45/−45]s lay-ups were used for testing. Drop dart testing machine with a hemispherical impactor of 10 mm nose radius was used, and the tests were conducted according to ASTM D3029 standards. The impact energy of 400 J was generated using an impactor of mass 20 kg with an impact velocity of 6.28 m/s. From observation, two forces were categorized one that caused no damage and second which caused damage to the laminate. Both woven and unidirectional laminates [0°/90°]S showed better saturation energy than the other laminate orientation. Saturation energies for woven and unidirectional laminates were 48.14 J and 53.08 J, respectively, for [0°/90°] laminate orientation. Fan et al. [14] studied the circular GFRP tubes, concrete-filled GFRP and steel under flexural loading as a function of concrete filling, central hole, tube-in-tube with concrete between them and different laminate designs. Four-point bend method was used to examine the flexural behaviour of the samples. From flexural testing, it was observed that local buckling was avoided by filling concrete. After the crack initiated, the stiffness of the structure was mainly due to the laminate design and diameter to thickness ratio. Among all other laminates, the concrete-filled pultruded

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GFRP tubes resulted in higher stiffness. Higher strength-to-weight ratio for flexural properties was achieved by the central hole in the core. Finally, it was concluded that the flexural properties of the laminates were a function of stiffness diameterto-weight ratio and less dependent on concrete strength. Fibres were the main load carrying members, whereas concrete improved the ductility of the composites. Mirmiran et al. [15] studied the E-glass fibre tube filled with and without concrete under compressive impact (shock loading) loading and evaluated by using stress wave propagation method. E-glass/epoxy laminates with [0°/0°/+45°/−45°]10 ply orientation was used, which contains 51% of fibre material. The diesel impact hammer was used during field testing. The hammer was 12.77 kN in weight, 3.85 m long with 1.2–2.9 m of stroke length, which generated 60–38 blows per minutes. Further, the energy generated during the blow varied from 15.6 to 37.0 k-Nm. From field test and stress wave propagation analysis, no damage was observed in FRP tubes filled with concrete both at the upper face and at the interface. Further, the empty FRP tubes underwent buckling failure with steel mandeer or in soft soil or at shallow depth. Shyr et al. [16] studied the impact resistance of GF/polyester composite laminates made of non-crimp, woven and non-woven mat fabrics. The impactor geometry was hemispherical with 1.27 cm nose diameter and 47.53 kg of mass. Simple hand lay-up method was used to fabricate the GF/polyester composite laminates. It was observed that as the ply numbers in the laminate decreased the dominant damage mode shifts from fibre failure to delamination. It was concluded that the non-crimp fabric architecture showed better impact energy absorption capacity than other fabric architecture. Aslan et al. [17] examined the effect of impactor mass and in-plane dimension on the impact resistance property of GF/epoxy composite laminates. The stacking sequence of the GF/epoxy composite laminate was [0/90/090]S . The impactor mass significantly influenced the force history curve. The impact resistance of GF/epoxy composite laminate was also dependent on the in-plane dimensions. Further, as the in-plane dimensions of the GF/epoxy composite laminate was reduced, the contact duration increased. Furthermore, the delamination area formed in the laminate was also affected by the in-plane laminate dimensions. Mitrevski et al. [18] studied the combined effect of preloading and impactor geometry on the impact resistance property of GF/polyester composite laminates. Figure 5.2 shows experimental setup for preloading the composite laminates. Impactor geometry used were hemispherical, conical, ogival and flat. The biaxial tensile load applied was 500 and 1000 micro-strain. The fabricated GF/epoxy composite laminates ends were tabbed with aluminium to avoid crushing of material at grip section during biaxial tensile preload. The composite laminates were fabricated by hand lay-up process. The stacking sequence of the GF/polyester was [0/90, ±45, 0/90]S . The experimental results showed that flat nose impactor showed highest peak force and the sharp nose impactor showed highest contact time. The preloading had insignificant influence on the impact resistance property of GF/polyester composite laminates. Hosseinzadeh et al. [19] examined the low velocity impact-induced damage behaviour in thick GF/epoxy, thin GF/epoxy, CF/epoxy and CF/GF/epoxy composite

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Fig. 5.2 Experimental setup used to preload the GF/polyester composite laminate under biaxial tensile loading [18]

laminates. The impactor diameter was 19 mm. The impact energy was varied from 30 to 100 J. At impact energy 30 J, both thick and thin GF/epoxy composite laminates showed damage, whereas the CF/epoxy laminate showed no damage. Further, as the impact energy increased, the damage also increased significantly in thick and thin GF/epoxy composite laminates. Furthermore, at impact energy 50 J, the CF/epoxy composite laminate showed sudden collapse. Unlike the thick GF/epoxy, thin GF/epoxy and CF/epoxy composite laminates the hybrid CF/GF/epoxy composite laminate showed no damage at low impact energy or collapse at high impact energy. Sevkat et al. [20] conducted both numerical and experimental study on S2 -glass fibre/epoxy laminates subjected to ballistic impact. The laminate consisted of 24 unidirectional layers of S2 glass/epoxy laminate with different stacking sequences. Drop weight and gas gun testing equipment were used for low and high velocity impact, respectively. For drop weight impact testing, a hemispherical impactor of mass 5.1 kg was used, and for ballistic impact, a 22-calibre bullet with copper coating was used. LS-DYNA was used to simulate the impact mechanism under both the impact conditions. MAT 002 material model was used to simulate the orthotropic elastic impact behaviour of the laminates under low velocity impact testing. For ballistic impact MAT_03 material card was used to model the 22-calibre impactor and Chang–Chang model was employed to study the damage mechanism. Based on the study, few conclusions were made and are as follows: cross-ply fabric

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orientations yielded the highest impact resistance when compared to other configurations. Delamination was predominant in low velocity impact, whereas delamination, matrix cracking and fibre breakage were predominant in ballistic impact. Both the experimental and the numerical results were in good correlation with each other. The effect of cross section, fibre orientation and condition deformation of Eglass/epoxy were investigated by Shokrieh et al. [21]. Numerical simulation was carried out by using LS-DYNA. It was observed that as the amount of fibre in the axial direction decreased the energy absorption also decreased, resulting in fibre breakage. This phenomenon was mainly took place between fibre orientations 30°– 60°, but from 75° to 90° orientation angle, matrix cracking took place as the fibres absorb more impact energy. Crushing failure load of the rectangular tube was less in comparison with circular tubes. From results, it was observed that both experimental and numerical results were in better agreement with each other. Impact-induced damage in an asymmetric GFRP laminate with circular precrack in it was investigated by Singh et al. [22] under low velocity impact loading. Symmetric and asymmetric laminate designs were [0°4 /90°4 /90°4 /0°4 ]S and [0°4 /90°4 /90°4 /0°4 //90°4 /0°4 /0°4 /90°4 ], respectively. Laminates were fabricated using hot press and cured at 150 °C under 1 MPa. However, the fabrication of the laminate was carried out in two phases. In the first phase, the bottom and upper halve about the mid-plane were fabricated. Then, in second, both the halves were joined using epibond and 7.5% of XNBR along with the introduction of circular pre-crack at the mid-plane using a Teflon film (15 μm thick). Airgun was used to induce the impact with a mild steel impactor with an impact energy of 12 J at 50 m/s of impact velocity. Damage extent and failure modes were investigated by optical microscopy. Average damage area (DAE) for symmetric and asymmetric laminates without epibond and XNBR were 0.76 ± 0.11cm2 and 1.53 ± 0.13cm2 , respectively. For samples with and without XNBR, it was 1.04 ± 0.17 and 0.42 ± 0.15cm2 , respectively. Finally, it concluded that XNBR reduced the damaged area in asymmetric laminate than with and without epibond laminate. Singh et al. [23] investigated the effect of adhesively bonded static toughness and impact-induced damage behaviour of symmetric and asymmetric GFRP laminates joined by epibond 1590 A/B. Further, XNBR (high molecular carboxylated acrylonitrile butadiene) content was varied (0, 5, 7.5, 10) in epibond to find out the optimum doping quantity. Ply design for asymmetric laminate was [[0°2 /45°2 /−45°2 /90°2 ]S /AI/[0°2 /45°2 /−45°2 /90°2 ]S ], and for symmetric laminate, the ply design was [0°16 /AI/0°16 ]. These designs were fabricated in two phases; in the first phase, the laminate bottom and upper half from the mid-plane were fabricated using hot press and cured at 180°C under 1 MPa; then, both halves were abraded using emery paper and cleaned with acetone and then dried. Then, both the halved were adhered using epibond in the first set of samples, and the second set, the samples adhered with epibond mixed with XNBR. Mode-I and mode-II tests were conducted using double cantilever and end notch flexural methods, respectively. Impact test was conducted by dropping a mild steel impactor with 12 J of impact energy. From experimental and optical micrographs, it was evident that symmetric laminated bonded with epibond adhesive with 7.5% XNBR performed better than

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other laminate designs. Thus, it was concluded that 7.5% of XNBR as the optimum value to get better mechanical properties. Icten et al. [24] evaluated the impact damage behaviour of GF/epoxy composite laminates impacted at 20 °C, −40 °C and −60 °C. The environmental-controlled chamber was integrated to the impact machine to conduct LVI tests at low temperatures. The quasi-isotropic stacking sequences used for the experimentation was [0/90/45/−45]S . The damage tolerance of GF/epoxy was same for all the testing temperatures below impact energy 20 J. Beyond the 20 J impact energy, the testing temperature significantly influenced the impact performance of GF/epoxy composite laminates. It was also observed that the damage threshold of GF/epoxy composite laminate was increased with decrease in testing temperature. Meola et al. [25] adopted the infrared thermography technique to analyse the impact behaviour of GF/epoxy composite laminates. The stacking sequence used for the analysis was [02 /902 ]S . The experimentations were carried on a Charpy impact machine using ogival and hemispherical impactor shape. The nose diameter of the impactor used was 18 and 24 mm. The infrared thermography results showed that during impact the surface of the GF/epoxy composite laminates were first cooled down due to thermoelastic properties of the laminate. Then, the laminate surface gets heated up due to energy dissipation. Using infrared thermography temperature maps, it was possible to differentiate between damage initiation and propagation zones during the LVI. LeBlanc et al. [26] investigated the numerical and experimental analysis of the underwater ballistic impact on E-glass laminates. 0° and 90° ply orientations were used. The prepared laminate thickness was 3.33 mm and 4.82 mm for 13 and 19 plies laminates, respectively. MAT_059 material card was used for numerical simulation in LS-DYNA. The MAT_COMPOSITE_FAILURE_SOLID_MODEL material card was used to study the laminate damage mechanism. The quality of the experimental and numerical correlation was quantified by Russel compressive error measurement technique. Mathivanan et al. [27] investigated the low velocity impact behaviour of woven glass fibre epoxy laminates under different velocities. Varying laminate thicknesses 2, 4 and 6 mm were fabricated by hand lay-up method and cured for 12 h at 0.2 MPa pressure and post-curing was carried out at 120 °C for four hours. Low velocity impact tests were conducted using the drop weight test method and carried out according to ASTM D3029 standards. A hemispherical nose impactor with a mass of 10 kg and a radius of 10 mm was used and impacted with 150 J of impact energy and 6.25 m/s of impact velocity. Impact energy was classified into two categories: one was saturation impact energy at which the laminate can take maximum energy without perforation, and the second one was damage degree, i.e. the amount of energy transmitted and dissipated to the laminate. Damage progression of glass/epoxy cross-ply laminates was studied by Azouaoui et al. [28] under low velocity impact fatigue loading (Fig. 5.3). Laminate consisting of nine prepregs was fabricated using autoclave method and impacted at a regular time interval with a hemispherical nose impactor. The process was continued until the complete penetration of the laminate took place. Crushing and deformation of

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Fig. 5.3 Impact fatigue apparatus [28]

underneath fibres resulted in the formation of a hemispherical crated shape. From experimental results, it was observed that damage took place in several phases; in the first phase, the plate stiffness decreased moderately. During second phase, the delamination reached the saturation state. In the third stage, stopping of delamination progression occurred. It took 36,860 and 67,527 impacts to create damage of 12.6% and 18% in the laminate at 5 J and 4 J impact energy levels, respectively. Eighty percentage of the damage was induced into the laminate after 448,229 and 572,423 impacts corresponding to 5 J and 4 J impact energy and finally concluded that the delamination and crater formation were the two factors that deteriorated the material properties. The experimental and numerical study was conducted by Karakuzu et al. [29] on glass/epoxy laminate with [0°/30°/60°/90°]S ply orientation. Impact energy of 40 J, impactor mass of 5 kg and impact velocity of 2 m/s were adopted. From experimental and numerical results, it was found that the energy absorption capability of the laminate at equal velocity was more than the equal mass–energy absorption. Higher the mass, more is the corresponding delamination and contact time. Finally, it was concluded that for every FRP composite, there exists an optimum delamination value which corresponds to maximum laminate failure. In this research analysis, the contact force history, deflection, in-plane shear strains and stress waves of an aluminium plate sandwiched between glass/polyester laminates were studied by Payeganeh et al. [30] under low velocity impact. The analytical investigation was carried out using the first-order shear deformation theory and Fourier series method. [0°/90°/0°/90°/0°]S lay-up sequence was impacted with a steel impactor of diameter 12.7 mm, a mass of 2 kg and an impact velocity of 1.0 m/s. Two analytical methods were used to improve the accuracy of theoretical prediction

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of impact resistance properties. Finally, it was concluded that the placing of Al sheet near the impact zone improved the fibre metal laminate properties, whereas farther away, the Al sheet from the impact zone reduced the properties significantly. Arun et al. [31] studied the stress intensity factor (SIF), interlaminar shear strength and impact toughness as a function of normal and seawater environment conditions with varying seawater treatment timings. Hybrid composite made of plain weave E-glass, fibre textile fabric and epoxy resin was fabricated using hand lay-up technique. The fibre volume fraction of glass fibre was varied from 15 to 45% weight percentage (wt.%), and the corresponding textile fabric was also varied by keeping constant epoxy wt.%, i.e. 40 wt.%. Samples were fabricated according to ASTM E1922, D2344 and D256 for end notched (ENT), short beam bend strength (SBB) and impact test, respectively. ENT, SBB and impact tests were carried out on a universal testing machine, three-point bend test and pendulum-type impact testing machine, respectively. From scanning electron microscopy (SEM), it was found that the type of loading and nature of environment decided the nature of fracture. Further, the experimental results showed that among other seawater-treated hybrid composite, the laminate with 45 wt.% of glass and 15 wt.% of textile fabric yielded better results. Glass/epoxy laminates were studied by Menna et al. [32] under low velocity impact. GF/epoxy laminates were fabricated by hand lay-up technique assisted by press cure. Curing of laminates was carried out at 120 °C under 0.1 MPa of pressure for 2 h. [(0°, 90°)n /(+45°,−45°)n ] ply orientation was used, and 0.96 mm of laminate thickness was achieved for n = 1 and 1.92 mm for n = 2. Hemispherical impactor with nose diameter 16 mm and a mass of 3.6 kg was used for testing. During numerical analysis in LS-DYNA, MAT_59 (COMPOSITE_ FAILURE _SOLI D _MODEL), material card was used to define the composite laminate. From both experimental and numerical analyses, it was found that the damage in laminate was increased linearly with increase in impact energy. Aktas et al. [33] studied influence of thermal exposure on the impact resistance property in GF/epoxy composite laminates. The impact tests on GF/epoxy composite laminates were carried at 40, 60, 80 and 100 °C. The two stacking sequences used in the experiment were [0/90/0/90]S and [0/90/45/−45]S . The impactor mass and impactor nose diameter were 5.02 kg and 12.7 mm, respectively. Experimental results showed that the impact resistance properties of GF/epoxy composites were decreased with increase in testing temperature. However, the damage area was decreased with increase in temperature. Caprino et al. [34] analysed the influence of impact energy, laminate thickness and impactor nose diameter on the impact resistance behaviour of GF/epoxy and basalt fibre/epoxy laminate. The stacking sequence used for the experimentation was [(0, 90)n /(+45, −45)n /(+45, −45)n /(0, 90)n ] where n = 1, 2, 3 and 4. Both basalt fibre and GF composite laminates showed different failure modes. Evci et al. [35] studied the LVI damage behaviour in unidirectional GF/epoxy, woven GF/epoxy and woven aramid/epoxy composite laminate. It was concluded that the unidirectional composite laminates performed better under quasi-static loading compared to woven composite laminates. However, under dynamic loading, the woven composite laminates showed better LVI resistance than unidirectional composite laminate. Further,

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the damage formed was confined in a smaller area in woven composite laminates compared to unidirectional composite laminates. Among all composite laminates, the aramid/epoxy composite laminate showed best impact resistance and performance compared to unidirectional GF/epoxy and woven GF/epoxy composite laminates. Icten et al. [36] examined the effect of impactor diameter on the impact damage behaviour of GF/epoxy composite laminates under LVI. Four different impactor diameters 12.7, 20, 25.4 and 31.8 mm were selected for testing. Experimental results revealed that as the impactor diameter increased correspondingly the contact force was also increased. Further, the penetration and perforation thresholds were also increased as the impactor diameter increased. Amaro et al. [37] studied the LVI effect on the delamination damage behaviour of GF/epoxy composite laminates containing open holes. The GF/epoxy composite laminates had one or two open holes (Fig. 5.4). It was observed that the presence of hole did not affect significantly on the impact resistance property of GF/epoxy composite laminate. However, an insignificant influence was observed on the damage area formed. A modified analytical model was proposed by Mohan et al. [38] to predict the ballistic impact behaviour of unidirectional glass/vinyl ester laminate with [0°/90°] ply orientation. The developed analytical model was based on the energy balance method. It included cone formation at the back surface of the composite laminate, deformation of the secondary yarns, primary yarn failure under tension, matrix cracking, fibre/matrix delamination and projectile laminate friction during the penetration. For experimental testing, the glass/vinylester laminates were fabricated using hand lay-up method. Furthermore, post-curing of the laminate was carried out at 180 °C for four hours and cooled to room temperature naturally. The ply design [0°/90°/0°/90°]2S was impacted with conical-, truncated conical- and hemisphericalshaped impactors. The experimental ballistic limit for conical, truncated conical and hemispherical nose impactor was 73.2, 76.7 and 80.8 m/s, whereas the predicted ballistic limit were 69.8, 74.8 and 75.6 m/s, respectively. Predicted and experimental energy absorption were 170.5, 195.8, 200 and 187.5, 206.1, 228.5 J, respectively. It was observed that as the impactor cone angle decreased the ballistic limit and energy absorption capacity of the laminate was increased. Predicted and experimental energy absorption and damage area result above the ballistic limit 159.3 J, 172 J, 187.2 J; 173.5 J, 183.3 J, 197.4 J and 11,642 J, 14,298 J, 16,529 J; 12,560 J, 16,404 J, 19,874 J, respectively. Even above the ballistic limit, both experimental and predicted damage area increased as the cone angle of the impactor decreased. Finally, it was concluded that the new modified analytical model predicted accurate results for unidirectional composite plates. Aktas et al. [39] studied the impact and post-impact behaviour of glass/epoxy laminates for different stacking sequencing. Here, rib, Milano knitted and 2D woven fabrics were used to fabricate eight-layered laminates using hand lay-up technique assisted by hot pressing. Curing was carried out under 8 MPa constant pressure at 110 °C for 100 min. A hemispherical nose diameter of 12.7 mm and mass of 5.07 kg was used as an impactor and struck onto the laminate at an impact energy levels of 10, 20, 30, 40, 42.5 and 45 J. Compression after impact test was carried out according to the ASTM D7137. A 3DIMPACT code was used to carry out the finite element analysis.

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Fig. 5.4 GF/epoxy composite laminates containing open hole subjected to LVI a GF/epoxy composite laminate with no open hole. b GF/epoxy composite laminate with single open hole. c GF/epoxy composite laminate with two open hole [37]

From experimental results, it was observed that the highest impact and compression after impact properties were showed by Rib-knit sandwiched between two woven fabric and knit/knit hybrid composite sandwiched between Milano fabrics. Both numerical and experimental results were in good correlation. Perillo et al. [40] studied the damage development in stitch bonded GF/epoxy composite laminates subjected to LVI. Cross-ply stacking sequence designs [0, 0, 90, 90]S , [0, 90, 90, 0]S and [0, 90, 0, 90]S were adopted for the LVI test. The stitch GF/epoxy composite laminates were fabricated by vacuum-assisted resin infusion process. The position of 0° and 90° plies in the laminate significantly affected the delamination area in stitch GF/epoxy composite laminates. However, the position of 0° and 90° plies in the laminate showed no significant effect on the force–displacement curve. An experimental and numerical study was conducted by Hassan et al.

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[41] on GFRP laminate under low velocity impact for the laminate thickness of 2, 3 and 4 mm at impact energy ranging from 9.8 to 29.4 J. Hand lay-up technique was used to fabricate the laminate and cured for 24 h at room temperature. Free-fall drop dart machine with hemispherical nose impactor with 12.7 mm in diameters made of steel was used during impact testing, and ASTM D5628 standard was followed for testing. MSC.MAR® MENTAT® software was used to carry out the numerical analysis. It was concluded that the impact resistance of the GFRP laminates was much sensitive to laminate thickness and impact energy level. Raja et al. [42] studied the effect of varying weight percentage (wt.%) of fly ash on mechanical properties of E-glass fibre reinforced with polyester resin. The laminates were fabricated using hand lay-up technique assisted by a mould box. Mechanical properties such as tensile, flexural, impact, hardness and compressive tests were performed according to the ASTM D638, D7264, D256, D2583 and D695, respectively. Flyash, glass fibre and polyester interface interactions were studied by using scanning electron microscopy (SEM). From the experimental analysis, it was observed that the controlled sample with 70% polyester and 30% chopped E-glass yielded better impact, tensile, compressive and hardness of the material and found that at 10wt.% of fly ash, 20wt.% fibre and 70% of resin produced good results compared to other samples due to its better adhesion which was supported by SEM analysis. An experimental and numerical investigation on glass fibre/epoxy laminates under low velocity impact was carried out by Chandekar et al. [43]. Experimental and numerical testings were conducted by using drop tower and LS-DYNA, respectively. Fabrication of composite was done by using heated vacuum-assisted resin transfer moulding. From experimental results, it was found that due to the tensile stresses produced during the impact at the bottom surface leads to fibre breakage; thus, the damage was seen more at the bottom surface of the laminate. Results were overpredicted by the numerical analysis than the experimental results. Evci et al. [44] developed a new numerical model to determine the laminate thickness variation effect on energy dissipation and damage mechanics of woven/unidirectional GFRP laminates under low velocity impact loading. Both unidirectional and woven laminates with [0°/90°] lay-ups of varying thickness from 2 to 8 mm were fabricated using hand lay-up technique and then pressed at 140 °C under 1.96 MPa pressure. Impact test was conducted with a spherical tup impactor of diameter 10 mm according to EN ISO 6603-2 standards at 4, 6, 8 J/layer of impact energy. The numerical model developed uses the Hertzian failure, and the main failure was forced to determine the damage threshold of a laminate, whereas penetration and perforation energies were used to examine energy threshold. Numerical and experimental results showed that as the thickness of the laminate was increased the energy absorption capacity and damage resistance capacity also increased correspondingly. The woven laminates yielded better load bearing capacity than the unidirectional laminates. It was concluded that the damage initiation, damage propagation and failure modes were mainly depended on the impact energy and laminates thickness as well.

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Fig. 5.5 Effect of volume fraction of energy absorption during impact loading [45]

Reghunath et al. [45] studied the optimal glass fibre volume fraction under impact loading. A bidirectional woven glass fibre laminates were fabricated by vacuum bagging process. Impact testing was conducted for varying velocity (2–4 m/s) according to ASTM D3029 standard. Resin burn-off method was used to find out the fibre volume fraction according to the ASTM D7584 standard, and scanning electron microscope was used to study the damage mechanism. From experimental analysis, it was concluded that the glass fibre volume fraction of 43–48% offered maximum impact resistance (Fig. 5.5). In this experimental work, Singh et al. [46] studied the energy absorption capacity of the symmetric and asymmetric GFRP laminates under low velocity impact. Bidirectional woven glass fibre was reinforced with epoxy laminates were fabricated by using the hand lay-up method assisted by vacuum bagging under 0.96 atmospheric pressure and cured at room temperature for 24 h. The symmetric and asymmetric laminate designs were [(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)]S and [(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)//(+45°, −45°)/(0°, 90°)/(0°, 90°)/(+45°, −45°)], respectively. Impact testing was conducted on drop weight tower according to ASTM D7136 using a hemispherical impactor of mass 5.23 kg at an impact velocity of 3 m/s, 4 m/s and 6 m/s with corresponding impact energies of 23.54 J, 41.84 J and 94.14 J, respectively. The experimental results were validated by using LS-DYNA with material card 20 to model the impactor and MAT059 to model the laminate. From experimental results, it was observed that at 3 m/s impact velocity energy absorption of symmetric and asymmetric laminates was 8.86 and 10.7 J, respectively. Similarly, at 4 m/s it was 32.63 and 35.96 J, respectively. However, at 6 m/s, the fibre damage took place with energy absorption of 59.70 and 45.98 J for symmetric and asymmetric laminates, respectively. The pyramidal damage height from the experimental testing for symmetric and asymmetric laminate was 4.25 and 5.67 mm, respectively. Similarly, pyramidal height for numerical simulation was 5 mm, 6 mm for symmetric and asymmetric laminates, respectively. Finally, it was concluded that the symmetric laminate absorbed more energy than asymmetric laminate. Ma et al. [47] examined the damage behaviour of GF/epoxy composite laminates under LVI. The GF/epoxy specimens were fabricated from vacuum infusion process and stored at room temperature (295 K), dry ice (199 K) and liquid nitrogen (100 K). The cryogenic temperature-treated GF/epoxy composite showed

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lower apparent damage area compared to GF/epoxy specimens impacted at room temperature or non-cryogenic temperature-treated specimens. However, the energy absorption capacity of cryogenic temperature-treated specimens were less compared to the specimens tested at room temperature. Further, cryogenic treatment increased the GF/epoxy composite stiffness and caused embrittlement of the composite specimens. Furthermore, the interfacial bonding between fibre and matrix was improved as the GF/epoxy composites were treated at cryogenic temperature. Berk et al. [48] conducted a comparative analysis on LVI resistance between GF/epoxy and aramid/epoxy composite laminates. Both the composite laminates were impacted at 20 J and 30 J of impact energy. The experimental results showed that the GF/epoxy composite laminate absorbed more energy than the aramid/epoxy composite laminate. Further, at impact energy 30 J, the impactor was rebounded from GF/epoxy composite laminate, while at the same impact energy, the impactor perforated the aramid/epoxy composite laminate. Amaro et al. [49] worked on examining the multi-impact effect on the damage behaviour of GF/epoxy composite laminates tested at room temperature, 60 and 90 °C. The specimens were impacted at 1 and 3 J of impact energies. It was concluded that at higher impact energies the temperature had no significant effect on the impact properties of GF/epoxy composite laminates. However, at low impact energies, the temperature had significant influence on the impact resistance of GF/epoxy composite laminate. The number of impacts was also strongly dependent on the LVI testing temperature. It was also found that the testing temperature strongly affected the bending stiffness of the GF/epoxy composite laminate which in turn affected the impact strength. In this research, Ansari et al. [50] conducted an experimental and numerical analysis on GFRP laminates under ballistic impact using drop tower and ANSYS/AUTODYN analysis tool, respectively. Laminates were fabricated according to the ASTM D3039 standard using hand lay-up method. GFRP laminate was used at room temperature for 24 h under 250 N pre-pressure, and then, laminates were post-cured at 80 °C in a hot air oven for 3 h. Using the Lagrangian process, hexahedron brick elements were used to model the ogival-shaped impactor. In experimental work, an ogival-shaped 52 g of mass and 19 mm diameter bullet made of steel was used and fired using a pneumatic gun. Damaged area for FE model was 63.85 mm in the x-direction and 65 mm in y-direction, whereas experimental results showed 58.94 mm in the x-direction and 61.4 mm in the y-direction. Thick composites made from glass fibre and epoxy used in wind/marine industry were studied by Perillo et al. [51] under impact loading. Fabrication of GF/epoxy laminate was done by vacuum-assisted resin transfer technique with a pre- and postcuring time of 24 h at room temperature and 15 h at 80 °C, respectively. Three ply orientations of L1 [0°/0°/90°/90°]S , L2 [0°/90°/90°/0°]S and L3 [0°/90°/0°/90°] were subjected to impact testing with the help of drop tower machine and were conducted according to ASTM D7136 standard using a hemispherical impactor of mass 5.02 kg and nose diameter of 20 mm. Impact energy levels were varied from 45.5 J to 68.25 J. From the experimental and numerical data analysis, it was observed that the numerical and experimental results were in good agreement.

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Schwab et al. [52] presented the study on GF/epoxy laminates under high energy impact by large deformable bodies. Autoclave method was used to manufacture the E-glass/epoxy prepregs with 14 layers and [0°/45°/0°/45°/0°/45°/0°]S lay-up. Drop weight impactor was used for impact testing with an impact velocity of 20 m/s, impact energy of 400 J and drop weight of 2 kg. ABAQUS was used to simulate the impact testing. The use of shell element-based modelling helped in improving the computational parameters within the reasonable bounds. From the simulation, it was found that the energy dissipation over the large surface area was the critical factor in withstanding the impact. It was concluded that within reasonable resources and time shell elemental modelling approach was a favourable approach to simulate the significant composite components under high impact energy (Fig. 5.6). Ansari et al. [53] investigated the shock effect on GFRP composite due to impact under impact loading using numerical simulation (AUTODYN hydrocode). A blunt steel impactor with clamped and fixed boundary conditions were considered for numerical analysis. The numerical results were compared with experimental results from the literature. From the numerical simulation, it was concluded that the damage and penetration of a bullet were less under impact loading when the composite plate was in the fully restrained boundary condition. The thick composite plate showed high damage area value than the thin composite laminates, and the damage was localized near the impact point. As the span of the laminate increased, the corresponding ballistic velocity limit decreased. The numerical results obtained from AUTODYN hydrocode were in good agreement with the experimental results available in the literature.

Fig. 5.6 Damage of inner casing of turbine blade a, b inner and outer surface of the casing, respectively, with 60 plies and c, d inner and outer surface of the casing, respectively, with 100 plies [52]

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Yang et al. [54] developed a new numerical model to study the microscopic damage mechanism of GFRP laminate under low velocity impact by introducing transition model where embedded cell model was linked to the macrolaminate model. The new model considered Druker–Prager yield criterion and cohesive zone models to model the matrix material and interplay interface of the laminate, respectively. The developed model was executed using ABAQUS tool where the impactor was modelled with a hemispherical nose. The numerical results were compared with the experimental results available in the literature. From the numerical analysis, it was observed that the bottom ply just under the impactor tip was more susceptible to damage. The damage follows matrix cracking at the bottom of the laminate, progression of crack, interface and delamination. Finally, the fibre pull-out and breakage were observed at the bottom of the laminate. On comparison of numerical and experimental results, it was convincing that both the results were in good agreement because the transition zone model developed considered both inter- and intraply effect. In this research work, Ansari et al. [55] studied both experimental and numerical analyses of GFRP laminates under the normal and oblique impact (Fig. 5.7). Hand lay-up method was used to prepare the unidirectional GF/epoxy laminates with [0°/90°/90°/0°] fabric orientation according to ASTM D3039 standard. Impact test conducted using a pneumatic gun, which can fire the projectile with an impact velocity ranging between 150 and 300 m/s. A blunt impactor was used consisting of a mass 52 g with an impactor diameter of 19 mm. Oblique angles 0°, 30°,45° and 60° were used during impact testing. From experimental results, it was observed that as the impact angle of the projectile increases the residual velocity of the projectile decreases. Finally, it was concluded that the amount of damage decreases as the impact angle increases. Delamination in laminates under oblique impact was due to interlaminar stresses, whereas the delamination in laminates subjected to normal impact was due to the failure of the matrix in tension. Li et al. [56] used the pultruded glass fibre/isophthalic resin composite to study the damage development and progression. Drop weight tower impactor with hemispherical nose striker of diameter 20 mm and 1.2 kg of mass was used to conduct the impact tests according to ASTM D7136 standard. Impact energy, impact velocity and impact height were varied from 16 to 67 J, 2.49 to 4.29 m/s and 311 to 1424 mm,

Fig. 5.7 Oblique impact and fibre stacking design [55]

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respectively. From results, it was clear that the damage in the laminate was increased linearly with increase in impact energy up to a limiting value, and a good correlation was observed between numerical and experimental results. In this study, Duarte et al. [57] compared the extended finite element model (XFEM) and Hashin damage criterion. A numerical study was conducted on glass/vinyl ester laminates with [0/908 /01/2 ]S , [908 /0/01/2 ]S , [90/08 /901/2 ]S and [08 /90/901/2 ]S ply orientations with different hole radii (0.25, 2.50, 5.0 mm) at the centre using ABAQUS. From the numerical analysis, it was found that as the radius of the hole increased the corresponding strength of the laminate decreased. Laminate with more 0° oriented plies performed better than others. Since XFEM disregarded the lower resistance of shear in 0° plies, however, the failure of fibre took place at a higher load than Hashin’s failure. Finally, it was concluded that there is no requirement of the initial defect to seed the matrix for possible cracking, and these methods provided identical numerical results. Reddy et al. [58] tested the tensile and water absorption capacity of glass/epoxy laminates by introducing voids and holes artificially. Void sizes were 0.5, 1.5 and 2 mm, whereas the drilled hole diameters were 1, 2, 2.5, 3 and 3.5 mm. The laminates were prepared by using the hand lay-up method. Water-soaked gel balls were introduced to create defects. Both tensile and water absorption tests were carried out according to D638 and D570, respectively. From the tensile test, it was observed that as the void and drill sizes increased, the corresponding reduction in strength was observed. For laminate, without defect and drill hole, the tensile strength was 81.84 MPa and 81.993 MPa, whereas with a defect and drilled hole (at 2 and 3.5 mm, respectively), the tensile strength was 42.335 and 30.577 MPa, respectively. Water absorption of laminates in both drilled and void defect with 2 and 0.5 mm size performed better compared to with and without defect and drilled holes at various dimensions in the laminates. Finally, it was concluded that the degradation of material properties would take place. Thus, care should be taken to avoid such kind of situations during the FRP laminates service life. Balaganeshan et al. [59] studied the effect of impact on GFRP laminates at 0°, 30° and 60° temperatures. GFRP laminates were fabricated using vacuum bagging method, and filler weight percentage used was 1, 2, 3, 4 and 5%, and these laminates were made of three layers of 0°/90° orientation. Impact testing was carried out on a air gun with hemispherical nose impactor of mass 7.6 kg and nose diameter of 9.5 mm. From experimental results, it was evident that a slight enhancement in energy absorption capacity was observed in the laminate doped at 5 wt.% nanoclay impacted at 30 °C, whereas vice versa was observed in the case of impact at 0 and 60 °C. The delamination improvement observed in all laminates impacted at 0, 30 and 60 °C. The improvement was only up to 3 wt.% beyond which delamination increased. The energy absorption capacity of the CNTs was less compared to nanoclay but followed the same trend as that of the nanoclay. Overall the 3 wt.% nanoclay and 2 wt.% CNTs showed better damage resistance properties, beyond and below which reduction in properties can be observed and finally concluded that better performance of clay than CNTs was due to the better dispersion of clay in the matrix material.

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Post-impact reduction in compression and flexural properties of glass/epoxy laminate under impact loading was studied by Hart et al. [60]. 2D and 3D braided laminates were fabricated by vacuum-assisted resin transfer moulding method at 70 °C and cured at 121 °C (3 °C/min) for eight hours and cooled to room temperature with a cooling rate of 1 °C/min. Beam testing was done on drop weight tower with cylindrical impactor of diameter 50.8 mm and impacted energy of 25 J at the centre of the beam samples. Plate samples were tested with similar experimental conditions except impact energy ranging to 100 J. Flexural after impact (FAI) was conducted according to D6272 standard with 2 mm/min crosshead speed on a four-point bending fixture. ASTM D7137 standard was used to conduct the compression after impact (CAI) with 1 mm/min crosshead speed. From results, it was observed that both flexural and compression strength of both 2D and 3D laminate decreased linearly with an increase in impact strength except 2D laminate. It showed a slight increase in flexural strength from ≈500 to ≈550 MPa. The same trend was also observed in compression strength, but the reduction was much less as the impact energy increased. Nevertheless, the 3D braided laminates showed less reduction in properties compared to 2D braided GFRP laminates. Finally, it was concluded that the Z-tows in 3D braided GFRP composites resists the buckling failure thus becomes a critical factor in increasing the performance of the laminate. Vengalrao et al. [61] examined the effect of resin transfer moulding (RTM) process parameters such as injection pressure and its influence on void formation in the laminate. Four-, five- and six-layered chopped glass mat and polyester laminates were fabricated using resin transfer moulding process at different pressure, i.e. 1.96, 2.45, 3.43 and 3.92 bar. The void presence calculated according to the ASTM D273494 standard, and it was clear that as the number of layers was increased the pressure required to reduce the void content also increased. There exist an optimal pressure value for all layered glass/polyester laminate beyond which the void content started to increase; i.e. for four-, five- and six-layered laminates, 2.45, 2.94 and 3.43 were the optimal pressure values to get the minimum void content, respectively. Izod impact test was carried out according to ASTM D256 and found that same optimum pressure values corresponding to minimum void contents yielded better impact strength. For four-, five- and six-layered glass/polyester laminates, the maximum impact strengths were 351.58, 435.94 and 468.19 J/m; these correspond to the pressure values 2.45, 2.94 and 3.3 bar, respectively. Scanning electron microscopy (SEM) was used to examine the damage mechanism and void formation in the laminate. Finally, it was concluded that the void formation during fabrication of glass/polyester laminate from RTM process, which depends on the pressure, greatly influenced the impact properties of the laminates. Extent of damage recovery of E-glass/poly(ε-caprolactonel) (PCL)/epoxy blend laminates under low velocity impact was investigated by Cohades et al. [62]. Laminates were fabricated using vacuum-assisted resin infusion (VARI) method with and without PCL. The PCL content in the laminate was 25 volume% with fibre lay-up [(+45°/−45°)/(0°/90°)2 /(+45°/−45°)]4 and cured at 180 °C for three hours. Impact test was conducted according to ASTM D7136 standards with 5.5 kg impactor mass at three impact energy levels, i.e. 8.5, 17 and 34 J. After the impact, the laminates

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were undergone thermal mending at 150 °C for 30 min and recovered the damaged area from 40 to 90%. The healing efficiency for 34, 17 and 8.5 J was 39.9% ± 0.6%, 55.4% ± 2.7% and 94.8% ± 2.7%, respectively. In some case, the recovery failed in which the fibres were damaged, and large cracks were present. At 8 J of impact energy, full healing of the laminates was observed, and it was concluded that PCLepoxy laminates could be used to reduce or eliminate the formation of subsurface cracks when subjected to low velocity impact loading. Dynamic behaviour of GFRP laminates under ballistic impact loading with different boundary conditions and impactor nose angles (conical) was studied by Ansari et al. [63]. The pneumatic gun was used for conducting impact testing with a conical nose impactor of mass 52 g and diameter of 19 mm for different boundary conditions CCCC (fully clamped from all edges), CFCF (edges are clamped or free) and SSSS (simply supported at all edges). ANSYS/AUTODYN was used for numerical modelling to validate the experimental results. From experimental results, it was found that the composite plates offered high penetration resistance for blunt impactor and simply supported boundary conditions. More damage was observed at back face rather than at front face of the laminate due to the pressure wave variation. Kim et al. [64] developed a new progressive 3D failure model to investigate the nonlinear mechanical response of GFRP laminate under impact loading. LS-DYNA was used for numerical simulation. Hashin failure criterion was used to develop the material property degradation model. The simulation results were compared with the experimental results available in the literature. The impact simulation test was conducted using a hemispherical nose impactor of mass 5.5 kg and diameter of 16 mm. The impact energy and impact velocity considered for numerical simulation were 20.1 J and 4.3 m/s, respectively. The damage prediction between adjacent plies and the contact interfaces was modelled using cohesive zone mode and hexahedral elements. The simulation results for delamination and damage area were in good agreement with the experimental results present in the literature. It was concluded that this model can be used for predicting damage and impact behaviour but also the interlaminar delamination can be established. Tang et al. [65] used an energy balanced method to develop an analytical model to predict the damage for ceramic faced light armours when subjected to impact loading with a flat ended face. The composite was made of ceramic AD85 face tile and S2glass/phenolic FRP laminate as back plates. A spherical ball of US M33 0.5 in calibre with 46.8 g of core and boat tail was used for impact testing. From the numerical model, it was found that deformation and erosion were the two main mechanisms that absorbed the maximum amount of impact energy. Followed by compression and fragmentation, shear failure and perforation of back plates absorbed the applied impact energy and played a crucial role in absorbing impact energy and damage mechanism of the composite. Influence of steel ball impactor mass on damage and wave propagation in GFRP laminate was investigated by Nassr et al. [66]. Two lay-up designs [0°]9S and [45°/0°]4S were fabricated. Three steel balls of mass 255, 535 and 1115 g were impacted onto a laminate using air cannon device at a speed of 9 m/s with a 550 kPa pressure. Successive impacts were carried out on [0°]9S laminate with 18, 19

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and 20 layers on which multiple impacts were made with impact velocity 28, 50, 60, 69 and 70 m/s. Out of all combinations and testings, GFRP laminate with [0°]11 impacted by a steel ball of a diameter of 65 mm and mass of 115 g with velocity 19 m/s yielded less damage area of 1000 mm2 . Meanwhile, GFRP laminate with [0°]20 yielded maximum damage area of 88,700 mm2 when impacted by a steel ball of diameter 39.7 mm and mass of 255 g with an impact velocity 91 m/s. From C-scan and experimental observation results, it was found that strain and delamination were large when impacted with small mass than the large mass. [0°] lay-up design offered greater impact resistance than [45°/0°]. The successive impact created an increased cumulative damage/delamination areas. It was concluded that flexural waves were the predominant system which propagated with different speeds in different directions. Nassir et al. [67] examined the off-centre impact loading influence on the damage initiation in GF/epoxy composite laminates. The parameters considered during the testing were geometrical and GF/epoxy plate thickness. The experimental results showed that the peak load was high when impacted at centre of the laminate. The damage initiation load was increased with increase in thickness. It was concluded that the off-centre impacts were more serious threat to the composite laminate than the central impact.

5.2 Low Velocity Impact Test on Hybrid Glass Fibre Reinforced Polymer Composite Laminates Mechanical properties of oil palm/glass fibre reinforced with phenol–formaldehyde resin were investigated by Sreekala et al. [68] as a function of varying volume fraction of glass and oil palm fibres. The glass fibre content was varied from 10, 23, 27, 40 and 45 wt.%, whereas in hybrid laminates, the oil palm fibre content was varied from 0, 0.3, 0.5, 0.7, 0.9, 0.92, 0.96 and 1.0 wt.%. Tensile, flexural and impact tests were conducted according to the ASTM D638-76, D790 and D256, respectively. Scanning electron microscopy and optical microscopy were used to study the void content fracture mechanics of the laminates. Fibre weight (wt.%) percentage at 45% showed better tensile strength (≈92 MPa). Maximum flexural strength for control and hybrid laminates were approximately 100 MPa and 60 MPa for 40wt.% of glass and 0.6 volume fraction of oil palm fibre, respectively. At 40 wt.% of glass fibre the impact strength was maximum for control glass fibre laminates (≈225 kJ/m), whereas it was approximately 270 kJ/m for 0.74 volume fraction of oil palm fibres. From this experimental analysis, it was concluded that the oil palm fibre incorporation reduced the void content. Thus glass/oil palm hybrid laminates resulted in high performing lightweight structures with cost-effectiveness. The behaviour of glass, carbon and aramid fibre laminate under repeated velocity impact was investigated by Morais et al. [69]. Laminates were fabricated by vacuum bagging method assisted by autoclave technique. Repeated impact tests

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were conducted on a drop weight impact machine from 0.5 m and 1 m height. Hemispherical impactor with 12.7 mm nose diameter was impacted with 3.13 m/s and 4.42 m/s impact velocities with a corresponding impact height of 0.5 m and 1 m, respectively. Optical and scanning electron microscopy characterization were used to study the void content and fibre distribution in the laminate. Visual inspection and X-ray radiography were used to examine the macro- and microdamage mechanism after the repeated impact. It took 7, 58 and 1500 more impacts from 0.5 m impact height to cause the failure in aramid, glass and carbon laminates. Further, from 10 m impact height 2, 4 and 12 repeated impacts were sufficient to cause the failure in aramid, glass and carbon fibre laminates. Finally, it was concluded that the high elastic strain energy was the reason behind the better performance of carbon fibre which delayed the damage initiation and propagation. Anisotropic behaviour of aramid fibre was the reason for poor performance which resulted in sudden delamination and collapse of the laminate. Mamalis et al. [70] studied the behaviour of foam core/sandwiched by glass fibre laminates under crushing loading. Five layers of glass fibres were used as face sheets while, PMF foam, two different grades of PVC and polyurethane (PUR) were used as core materials. In-plane compression, the load was applied along the edgewise direction according to ASTM C364. Experimental results showed different failure modes such as unstable buckling with mode-I, unstable buckling with mode-II, the disintegration of sandwiched composites by delamination and buckling, stable progressive and crushing. The mode of failure depends on the core foam material properties. Most common failure mode was mode-II, i.e. disintegration, and occurred at a frequency of 25% of occurrence provided the buckling of the face sheets was avoided. Finally, it was concluded that other than the progressive-end crushing failure mode, all other modes were catastrophic. Theoretical, experimental and numerical comparison of a GFRP girder and GFRP girder structure with CFRP flanges were investigated by Hejll et al. [71]. GFRP decks were fabricated and left in a clamp for five days for curing process, whereas CFRP flange GFRP decks were manufactured using vacuum infusion method which provided rough surface finishing. Four-point bend test with 0.2 mm/min crosshead was adopted to carry out the experimental test, whereas MARC mentat 2003® was used for numerical simulation. Shell element 75 bilinear thick shell elements were used to model the bridge deck. Theoretical, experimental and numerical results were in good correlation. Finally, it concluded that the CFRP flange with GFRP beams performed better than GFRP beams without CFRP flanges. Ahmed et al. [72] examined the impact performance of GF-jute/isothalic-polyester hybrid composite laminates. The hybrid laminates were fabricates using hand layup process. Experimental results showed that the impact strength of jute/polyester composite laminates was less than the impact strength of GF/polyester composites. Further, the impact strength hybrid composite laminate with 17.2% glass fibre showed good impact strength which was comparable to pure GF/polyester composite laminate. Effect of incorporation of hollow glass fibre into the glass/epoxy and carbon/epoxy laminates under impact and flexural loading conditions was studied by Trask et al.

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[73]. A 16-ply laminate with [0°/+45°/90°/−45°]2S was fabricated using autoclave method and cured at 120 °C for one hour. The hollow glass fibres were placed between four 0°/45° ply in E-glass. In the case of CFRP with ply orientation [−45°/90°/45°/0°/−45°/90°/45°/0°]2S laminate, the hollow glass fibres introduced at 0°/−45° ply with 70 and 200 μm spacing. Impacting of laminates was done by hardened steel with a hemispherical nose of diameter 4.63 mm. Flexural properties of undamaged specimens were investigated on a four-point bend test according to ASTM D6272-02 with 5 mm/min crosshead speed. The damaged specimen flexural properties were studied using a three-point bend test. The spacing of hollow glass fibres and failure modes studied from optical microscopy. From experimental results, it was evident that the damaged laminates with hollow glass fibres showed minimal degradation of flexural properties and incorporation of hollow glass fibres does not create any site weakness in the laminate. Finally, it was concluded that if the crack generated was more than 300 μm, then the capillary forces are insignificant to repair the damage in the laminate. Sudarisman et al. [74] investigated the effect of flexural loading on hybrid laminates at different span to depth ration (S/d). Six ply laminates were prepared using E-glass (E) and S-glass (S) with the laminate design of [E6 ], [S6 ], [E4 S2 ] and [E2 S4 ]. These laminates were fabricated using autoclave method and cured at 120 °C for 16 min. Post-cured at 0.3 MPa pressure under 1.0 MPa compressive pressure held at 120 °C for 30 min. The flexural properties were calculated according to ASTM D79003 standard with three-point bend test, and failure modes were examined under optical microscopy. From experimental and optical microscopy results, it was concluded that 23% enhancement in the flexural property was achieved when 33% of E-glass fibre was replaced by S2 -glass fibre. Finally, fibre kinks and buckling were observed for S/d = 32 and S/d = 64, whereas shear cracking was seen for S/d = 16. Rosa et al. [75] studied the post-impact effect on the mechanical properties of Eglass/jute hybrid composites. Resin transfer moulding technique was used to fabricate the sandwiched and intercalated hybrid laminates consisting of 14 layers of glass and four layers of jute fibre. In the case of sandwiched laminate design, the jute fibre laminate was used as core material. In both the designs, the fibre volume fraction maintained was 50 wt.%. Impact testing was conducted on a drop weight tower with a hemispherical impactor of varying mass with nose radius of 12.7 mm and impact velocity of 2.5 m/s. Impact energy was varied from 5 and 15 J in five different steps, and ASTM D3763 was used to conduct the testing, and post-impact analysis was done by conducting the flexural test. ASTM D790 standard was followed with 5 mm/min crosshead speed. Scanning electron microscopy was used to examine the fabrication homogeneity of the hybrid laminate. Scanning acoustic emission and pulsed thermography were used to examine the impact and post-impact failure modes. From characterization and experimental results, it was observed that the sandwiched laminate design performed better at 10.5 J, whereas intercalated lay-up design was at 12.5 and 15 J of impact energy. However, the sandwiched structure yielded better residual mechanical properties and sudden failure, whereas no such phenomenon was seen in intercalated laminate lay-up design.

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Sevkat et al. [76] studied the impact behaviour of hybrid plain weaved glass–graphite/epoxy composite at different velocities. Hybrid laminates were fabricated using vacuum-assisted resin transfer moulding technique. Composite with four different stacking sequences was analysed: [GL37], [GL9/GR18/GL9], [GR8/GL16/GR8] and [GR28]. Tests were carried out using a steel impactor with a hemispherical nose of 16 mm diameter and 6.15 kg of mass. Numerical analysis was carried out in LS-DYNA. The following material model MAT_USER_DEFINED_MATERIAL_MODELS (MAT_43) was used to analyse the damage in laminates, CONTACT_ AUTOMATIC_ SURFACE_TO_ SURFACE _TIEBREAK was used for glass–glass, graphite–graphite interface, MAT_ADD_EROSION was used for the erosion of the element, and ERODING_SURFACE_TO_SURFACE was used to model the contact between steel impactor and composite. From results, it was evident that GL and GR offered high and least resistance to the impact compared to the other laminates, respectively. Out of hybrid laminates, GL/GR/GL yielded better results than other hybrid laminates. Hybrid laminates were more vulnerable to delamination than pure laminates. Both experimental and numerical results were in good agreement. The maximum delaminated area was observed for GL/GR/GL approximately 61 cm2 at 6.3 m/s. Highest energy absorption was observed for GR laminate approximately 120 J at 6.3 m/s (Fig. 5.8). Analysis of multi-site high velocity impact (Fig. 5.9) on S2 -glass/SC/epoxy composite was investigated by Deka et al. [77]. Vacuum-assisted resin transfer moulding method was used to fabricate the laminate with [0°/90°]S lay-up. For

Fig. 5.8 Energy absorption of composites with different velocities [76]

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Fig. 5.9 Multisite impact testing a Experimental, b numerical [77]

impact testing single-stage gas gun and a steel ball projectile of diameter, 7.95 mm and mass of 2.03 g were used. Material model 162 was used to model the specimen, and it used the progressive failure mode of laminates during transverse impact. CONTACT_ERODING_SINGLE_SURFACE material card used to model the contact between laminate and impactor. Two projectiles were impacted on to the laminates below the ballistic limit with velocity 201.2 and 201.9 m/s. The energy absorbed per projectile was 41.3 J and 41.6 J from experimental results and 38.76 J and 35.6 J from the numerical simulation. From the experimental and numerical analysis, energy absorbed and the surface area created were 76 J and 106.5 cm2 , respectively. From results, it was pointed out that the sequential impact resulted in increased damage over simultaneous impact. Balsa wood and PVA foam as core materials in a sandwich structure behaviour under impact loading were studied by Atas et al. [78] when sandwiched between glass fibre with [±45° Glass/core/±45°] stacking sequence. A drop weight impact machine was used with a hemispherical nose impactor of mass 5 kg and nose diameter of 12.7 mm. The test was carried out for varying impact energy levels, i.e. 5, 15, 30, 45 and 75 J. External damage was studied by visual inspection method, whereas internal damage and damage mechanism analysed by destructive method, i.e. water jet machining. Load deflection curves, energy diagrams and damage mechanism were analysed to study the impact behaviour of the balsa wood and PVA sandwiched laminates. From the analysis, it was found that balsa wood was stiffer than PVA. Thus, less deflection was observed, and hence, delamination was seen in PVA. Due to the weak interface bonding between balsa and glass, fibre debonding was the primary failure in that laminate. Due to the stiffness of balsa wood, it absorbed less energy than PVA sandwiched laminate. The behaviour of single- and multi-delaminated glass/epoxy and carbon/epoxy composites under impact loading conditions was studied by Ghasemnejad et al. [79]. Unidirectional glass fibre (GF) and carbon fibre (CF) were used to fabricate the [G0 ]4 , [G90 ]4 , [C0 ]6 , [C90 ]6 , [G0 /C0 ]3 , [G0 /C90 ]3 , [G90 /C0 ]3 and [G90 /C90 ]3 . To study the delamination behaviour, a Teflon film of 13 μm thickness was placed at three different positions, i.e. H = 0.5, 0.25 and 0.125 individually. Charpy impact machine was used for impact testing with a pendulum of mass 0.5 kg, 200 mm of arm length

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and 2.83 m/s of impact velocity. Experimental results were validated by finite element modelling of hybrid composites using LS-DYNA. Belytschko-Lin-Tsay beam model with quadrilateral shell elements and material model 54 was used for analysis. From the experimental and numerical analysis, it was found that [G0 /C90 ]3 absorbed the highest amount of energy, when delamination was at H/t = 0.25 and H/t = 0.125. Finally, it was concluded that the hybrid effect was more prominent and effective only when delamination was closer to the impacted zone. Fiore et al. [80] investigated the effect of stacking sequence and number of basalt fibres on the mechanical strength of glass/epoxy hybrid laminates. GFRP composites were fabricated using a vacuum bagging method and cured at room temperature for 24 h followed by post-cured at 60 °C for eight hours. Flexural and tensile strength were evaluated according to ASTM D790-03 and D3039, respectively. The obtained experimental results were validated by finite element analysis using ANSYS. For numerical analysis SHELL 99, eight nodded with six degrees of freedom elements were used. From experimental results, it was concluded that the presence of two unidirectional external basalt fibre with E-glass fibre mat as inner core performed better when compared to GFRP laminates. The numerical results were in good correlation with the experimental results. Unidirectional GFRP and SiC/Al composites behaviour under different strain rates were examined by Huang et al. [81]. For GFRP 300, 620 and 1050 S−1 and SiC/Al with 0.001, 200, 700 and 1200 S−1 , strain rates were used using split-Hopkinson tension bar, respectively. Single and bimodal Weibull statistical constitute mode were used to find out the effectiveness of rate dependency of stress–strain of unidirectional composite materials. From experimental analysis, it was concluded that the composites were rate sensitive. Strength and ductility of the composites were directly proportional to the strain rate, and results were in close agreement with the constituted model. Ahmadi et al. [82] investigated the behaviour of the aluminium-glass/epoxy sandwich laminates under high velocity impact for varying thickness of the composite laminate. A 2/1 laminate consisting of aluminium as outer skin and glass as the inner core was fabricated by hand lay-up method and wet lay-up was pressed under 8 bar pressure for 8 h. For experimentation, the gas gun was used consisting of a blunt cylindrical projectile with a diameter of 8.7 mm. In LS-DYNA, MAT_RIGID material card was used to model the projectile, MAT_PLASTIC_ KINEMATIC material card was used to model the aluminium panel and MAT_COMPOSITE_ DAMAGE material card was used to model the laminate. Each ply contact was modelled by using CONTACT_ AUTOMATIC_SURFACE_TO_SURFACE and CONTACT _ERODING_SURFACE_TO_ SURFACE. It was concluded that as the laminate thickness increased, the deformation in the laminate got reduced. Effect of fibre volume fraction on flexural strength of glass/carbon hybrid laminate was investigated by Dong et al. [83]. Optimization process was carried out to find the optimal fibre volume fraction. Glass fibre and carbon fibre volume fraction were varied, i.e. 30, 50 and 70 wt%. Three-point bend test was carried out according to ASTM D790-07, and finite element analysis was carried out using ANSYS. Various fibre combinations of laminates were [0°8C ], [0°G /0°7C ], [0°2G /0°6C ], [0°3G /0°5C ],

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Fig. 5.10 Impact time response a epoxy resin, b hybrid epoxy resin, c hybrid epoxy resin + 0.1 wt.% of MWCNTs, d hybrid epoxy resin + 0.5 wt.% of MWCNTs and e hybrid epoxy resin + 1.0 wt.% of MWCNTs [84]

[0°4G /0°4C ], [0°5G /0°3C ], [0°6G /0°2C ], [0°7G /0°C ] and [0°8G ]. From experimental and numerical analysis, it was found that the maximum flexural strength observed when carbon and glass fibre content was 50 wt.% and an increase of 43.46% and 83.57% were found when compared to neat carbon or glass laminates. Venkatanarayanan et al. [84] studied the medium velocity bullet impact response and vibration damping characteristic of GFRP. Characteristics of six-layered GFRP laminates were made of different types of matrix material. Different matrix materials were epoxy, hybrid (Epoxy: Polyester::3:2) and hybrid resin doped with MWCNTs. Samples were fabricated using hand lay-up technique assisted by hot pressing at 120 °C under 30 bar pressure for 4 h. From experimental results, it was found that damage mechanics consists of matrix cracking, debonding, delamination and fibre breakage. Damping efficiency of the laminates improved significantly, for 0.1 wt.% of MWCNT hybrid panel 175.51% increment was achieved and at higher weight percentage of MWCNTs yielded lesser response. It may be due to MWCNTs entanglement and agglomeration. The damaged area before and after the bullet impact was studied using an ultrasonic pulse–echo method. It was observed that epoxy showed higher damaged area than hybrid resin nanocomposite reinforced with MWCNTs. It was concluded that hybrid resin reinforced with MWCNTs showed improved impact resistance and damping characteristics as MWCNTs absorbs most of the shock (Fig. 5.10). Zhang et al. [85] studied the effect of stacking sequence on the strength of glass/carbon hybrid laminates by conducting three-point bend, tensile and compression tests according to the ASTM D790, D3039-76 and NASA’s short block compression fixture, respectively. Plain glass and twill carbon weave fibres were used during the wet lay-up fabrication process. [C8], [C2 /G2 ]S , [CG3 ]S , [CGCG]S and [G8 ] laminate designs were used for experimentation. In all testing, [C8 ] performed better than other laminate design, and among other hybrid combinations, [C/G/C/G]S showed better results. From optical micrographs, it was evident that the fibre pullout was significant in hybrid laminates except [G8 ] and [C/G/C/G], while the [C8 ] laminate showed sudden catastrophic failure when compared to other laminate layups. Finally, it was concluded that 50:50 fibre content of glass and carbon was the optimal fibre content with carbon fibre sequencing either in alternative or in external formation.

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Numerical and experimental study of the flexural property of the glass/carbon hybrid laminate was studied by Dong et al. [86] using a three-point bend test according to ASTM D790-07. Laminates were fabricated by using hand lay-up technique and cured at room temperature for 24 h. The stacking sequences were [C6 ], [G1 /C4 ], [G2 /C3 ] and [G5 ]. Numerical analysis was carried out using ANSYS with eight nodded PLANE 183 elements. Failure modes were studied using optical micrographs. From experimental results, it was evident that there was an increase in flexural strength by 8% and 3.2% when glass content was 24% in comparison with neat carbon and glass fibre laminates. Flexural, impact and post-impact flexural properties of ternary hybrid composites made of flax (F). hemp (H), basalt (B) and glass (G) fibre were investigated by Petrucci et al. [87]. Vacuum infusion method was used to fabricate the hybrid laminates and cured at room temperature and followed by a two-stage curing process, i.e. 60 °C for three hours and 80 °C for four hours. Three pure laminates, i.e. flax, hemp and basalt fibres, with [0°/90°]2S lay-ups were fabricated, and the hybrid laminate design includes GFBBFG, GHBBHG and FHBBHF. Drop weight impact tower was used for impact testing with an impactor of mass 1.25 kg and diameter of 12.7 mm. ASTM standard D7136 was followed with an impact height and energy being 3 m and 36.75 J, respectively. D790 standard was followed for flexural testing with a crosshead speed of 1.7 mm/min. Flexural strength before impact was 339.01 ± 30.15 MPa for basalt laminate, which was highest, whereas flax laminate performed poor among neat as well as hybrid laminates. In the case of hybrid laminate, GFBBFG and GHBBHG yielded maximum (271.71 ± 12.34 MPa) and minimum (149.92 ± 12.34 MPa), respectively. Impact penetration energy required was maximum for basalt (167 ± 1 J) and FHBBHF (25.1 ± 2.4 J) laminate among neat and hybrid laminates, whereas H (8.2 ± 0.9 J) and GHBBHG (14.3 ± 1.7 J) yielded minimum penetration energy. Post-impact flexural properties were maximum for B (246.20 ± 21.89 MPa) and GFBBFG (218.55 ± 22.55 MPa) and minimum for F (6.08 ± 0.94 MPa) and GHBBHG (109.71 ± 9.81 MPa) for both neat and hybrid laminates, respectively. From results, it was concluded that glass incorporation in hybrid laminates offered better results if and only if in the presence of flax fibres rather than hemp fibres. Tensile, flexural, shear and impact strength performance of abaca–jute–glass fibre hybrid laminates were investigated by Ramnath et al. [88]. [Glass/Abaca/Jute]S , [Glass/Abaca/Abaca]S and [Glass/Jute/Jute]S specimens were fabricated by using hand lay-up method with a curing period of 8–12 h. ASTM D638, D790, D5379 and D256 were followed to study the tensile, flexural, double shear and impact strength, respectively. Tensile strength of glass/abaca, glass/jute and glass/abaca/jute laminates was 44.5, 46.5 and 57 MPa, respectively. Flexural strength of glass/abaca, glass/jute and glass/abaca/jute FRP composite was 12.5, 11.9 and 12.1 MPa, respectively. Shear strength results for glass/abaca, glass/jute and glass/abaca/jute were 73.5, 71.1 and 77.6 MPa. Glass/abaca, glass/jute and glass/abaca/jute absorbed impact energy of 16, 15 and 12 J, respectively. Based on scanning electron microscopy characterization and experimental data, it was concluded that hybrid composite, i.e. glass/abaca/jute, was the best option in the case of cost-effectiveness when compared to other neat FRP laminates.

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Gonzalez et al. [89] conducted the LVI on interply hybrid composite laminates. The interply hybrid composite was made of woven glass fabric, woven carbon fabric and unidirectional carbon fabric. The intraply hybrid composites were fabricated by resin transfer moulding technique. Experimental results showed that non-hybrid GF/epoxy composite laminates showed better impact strength and damage resistance to LVI compared to non-hybrid woven CF/epoxy and non-hybrid unidirectional CF/epoxy composite laminates. Among hybrid laminates, it was observed that the hybrid laminate with GF/epoxy at mid-plane or core showed good damage resistance to the applied LVI. It was concluded that the use of GF and CF hybrid composite laminates can reduce the material cost and can improve the in-plane properties without compromising the impact properties. Petrucci et al. [90] studied the hybrid GF/epoxy composite laminates behaviour under LVI. The hybridization of GF was carried by using hemp, basalt and flax fibres. The stacking sequences of hybrid composite laminates were GHB and GFB where G, H and B represent the glass, hemp and basalt fibres, respectively. The experimental results showed that among GHB and GFB, the GFB hybrid composite laminate showed better impact resistance property than the GHB hybrid composite laminate. Further, the GHB hybrid composite laminate showed fibrillation due to the presence of hemp fibre. In this experimental and numerical analysis, Kim et al. [91] tried to replace the glass mat thermoplastic bumper with the glass/carbon mat thermoplastic to reduce the weight of the bumper. The micro-genetic algorithm was used for the optimization of laminate design (i.e. carbon and glass). The properties were calculated theoretically, using classical laminate plate theory, and LS-DYNA was used to simulate the impact loading condition onto the beam bumper. CATIA was used to model the design, and ABAQUS was used to mesh the designed model. From theoretical and numerical design optimization, it was observed that as the carbon content increased the weight of the bumper reduced and at 28% of unidirectional carbon fibre and 35% of woven glass fibre volume fraction the bumper beam yielded a reduction of 33% weight. Both theoretical and simulated results were in good agreement. Bienias et al. [92] studied the performance of aluminium–glass/epoxy composite when subjected to low velocity impact. Further, they investigated the nature and area of the damage created during the impact. Laminates lay-ups were 2/1, 3/2, 4/3 with a corresponding thickness of 1.5, 2.5 and 3.5 mm. Ply orientation of the laminate was [0°/90°], and laminate was fabricated at a heating rate of 2 °C/min to 135 °C for 2 h at 0.5 MPa pressure. ASTM D7136 standard was followed in conducting the impact test which performed on a drop weight impactor using a hemispherical striker of nose diameter 12.7 mm and impactor mass of 2 kg. Velocity and impact energy during impact testing were 3.16, 4.98 m/s and 10, 25 J, respectively. From results, it was found that the surface damage area increased with impact energy (Fig. 5.11). It was observed that damage in the composite mainly occurred at the interface of the fibre–fibre, metal fibre and along with matrix cracking. Finally, it was concluded that results obtained from VUMAT code used for analysing the damaged structure were in good correlation with the experimental results.

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Fig. 5.11 Damage in laminates after impact a 2/1 composite, b 3/2 composite and c 4/3 composite [92]

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Theoretical and experimental tensile strength results of glass/epoxy, carbon/epoxy and glass-carbon/epoxy were compared by Naresh et al. [93] using two parameters Weibull distribution under quasi-static to high strain loading conditions. Under quasistatic condition (8.3 S−1 strain rate), CFRP laminates performed better and GFRP laminates showed minimum tensile strength, i.e. 520.2 MPa and 338 ± 11.9 MPa, respectively. GFRP laminates yielded better results at 542S−1 strain rate (655 ± 41.3 MPa) and weak at 221 S−1 (405 ± 18.2 MPa). Similarly, CFRP and hybrid laminates yielded better results at 542 S−1 , i.e. 565.7 ± 17 and 578.29 ± 29, respectively. Fibre/matrix interface microcracking was observed in scanning electron microscopy (SEM) under quasi-static loading. Fibre/matrix debonding, matrix cracking and matrix failure were observed at high strain rates. From experimental results, it was clear that strain rate played a significant role in GFRP and hybrid laminates while the less significant effect on CFRP composites. Tensile strength increased with the corresponding strain rate increment. At high strain rate, glass fibre laminates provided better tensile strength than other composites. Naik et al. [94] developed a numerical model to predict the ballistic limit and conducted an experimental investigation on plain weave glass fibre and twill weave carbon fibre laminates under ballistic impact loading. A flat-headed impactor of mass 2.8 kg and diameter of 5 mm was impacted onto a six-layered glass/fibre and five-layered twill weave carbon fibre laminates. The predicted and experimental ballistic limits were 159 and 150 m/s for plain weave glass/fibre laminate, respectively whereas 99 and 105 m/s for twill weave carbon fibre laminate (for impactor mass = 1.8 g), respectively. The damage size was 9.6 and 10 mm in radius for predicted and experimental results for plain weave glass fibre, respectively. Finally, it was concluded that the ballistic limit for E-glass fibre was better than twill weave carbon fibre laminate because in glass fibre composite the energy absorption took place by deformation of secondary yarn and tensile failure of primary yarn, whereas only secondary yarn deformation and shear plugging were involved in case of twill weave carbon fibre laminates. Ansari et al. [95] studied the progressive damage analysis of sandwiched glassKevlar fibre/ epoxy composite plates. Various plate thicknesses, i.e. 3.12 mm, 6.24 mm and 9.36 mm, were impacted under ballistic impact with various shapes of projectiles (conical and blunt). The plates were prepared by hand lay-up method. The diameter and mass of both conical- and flat-headed impactor were 19 mm and 52 g, respectively. Numerical simulation were carried out by using AUTODYN hydrocode. Performance of the composite under ballistic impact depends on the shape of the impactor. As the laminate thickness increased, the ballistic limit also increased irrespective of projectile shape. From experimental and numerical results, Kevlar/epoxy laminates as the outer layer and GF/epoxy as inner core performed better than epoxy laminate as inner and GF/epoxy as the outer layer. The energy absorption capacity of composite tubes with metallic structures under axial impact loading was studied by Dlugosch et al. [96] and followed by numerical simulation using the adhesive interface model between the two discrete phases. Steel and unidirectional fibre (glass and carbon) was used to fabricate the specimens. Fabrication lay-ups considered were [0°/90°2 /0°]2S and [−45°/0°/45°/0°]2S with a

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steel plate thickness of 1.5 and 1.0 mm. The crash test conducted with 8.3 m/s velocity and sledge mass of 715 kg for [0°/90°2 /0°]2S and 795 kg for [−45°/0°/45°/0°]2S. ABAQUS tool with Hashin’s failure criteria were used during crash testing. From the experimental and numerical analysis, it was concluded that the cohesive model was complex but yet yielded accurate results than the tie modelling used for the hybrid interface modelling. Further, the hybrid GFRP steel tube composites were efficient design models in terms of cost-effectiveness. Bodur et al. [97] examined the hybrid glass fibre/waste cotton laminates. The glass content was varied from 2.5, 5 and 10 wt.%., whereas the waste cotton used was 12.5 and 25 wt.%. Further, the effect of maleic anhydride treatment was also examined. Single screw extruder was used to fabricate the hybrid laminates. Tensile, flexural and impact strength were examined according to ASTM D638-08, D790 and EN ISO 179, respectively. Flexural and impact tests were conducted on flexural and Izod impact testing machines, respectively. From experimental results, it was evident that as the glass content in the laminate was increased the corresponding enhancement in mechanical properties of the laminates was seen and found that with 25 wt.% of used cotton reinforced hybrid laminate yielded better in flexural and tensile testing. However, the hybrid laminate with 12.5 wt.% content of used cotton showed better impact resistance properties than the rest of the hybrid laminates. The properties were reduced in all the cases when fibres treated with maleic acid. Finally, it was concluded that glass fibre dominated the rheological properties when compared to cotton fibres. Three-point bend test and drop weight tower methods were used to evaluate the bend and impact strength of dual core sandwich composite structure used in a hydroelectric turbine by Ouadday et al. [98]. The sandwich composite consists of dual core made of extruded polystyrene foam (innermost) and aluminium trihydrate (ATH) filled with epoxy which sandwiched between glass/epoxy face sheets. Vacuumassisted resin infusion method was used to fabricate the glass/epoxy face sheets with [90°/0°/90°] ply orientation. ATH/epoxy core was prepared by mixing process for 15 min at room temperature then cured for 24 h at room temperature. Threepoint bend test was conducted according to ASTM D790 standards with 2 mm/min crosshead speed. The laminate impacted by using a 22.7 kg of mass and 25.4 mm of diameter with hemispherical nose striker using a drop weight tower machine at 8, 14, 18 and 24 J of impact energy. Exothermic reaction behaviour was studied according to ASTM D5470-06 using thermocouple by measuring thermal conductivity. Results showed a reduction of 74 °C and 117 °C for 40 and 60 wt.% further, reducing the time, i.e. 18 min and 51 min, respectively. As the weight percentage of ATH increased, the sandwiched structure showed more brittle nature. For 0 wt.%, 4 wt.% and 63 wt.%, the flexural strength observed was 76, 63 and 46 MPa, respectively. LS-DYNA was used for numerical simulation of face sheets and modelled by using eight nodded linear solid elements and continuum damage failure model. For ATH and polystyrene core, plastic damage model and MAT_LOW_DENSITY_VISCOUS_ FOAM (MAT_73) were used during the modelling of sandwiched composite structures. From the analysis of results, it was evident that the dual core absorbed 50% of the impact energy. Face sheets and ATH/epoxy were vital in deciding the impact resistance.

5.3 Low Velocity Impact Test on Nanomaterial …

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5.3 Low Velocity Impact Test on Nanomaterial Reinforced Polymer Composite Laminates The effect of fabrication technique on the strength of GF/epoxy laminates doped with MWCNTs was studied by Fan et al. [99]. Vacuum-assisted resin transfer moulding and injection double vacuum-assisted resin transfer moulding (IDVARTM) methods were used for fabrication of the laminates, and IDVARTM yielded better dispersion of MWCNTs in the laminate. From ILSS, it found that the MWCNTs improved the strength of the laminate by 33% when compared to the control laminates. Experimental results showed that the short beam test for ILSS was not reliable. Meantime compression shear test gave a reliable result for ILSS. For 0.5 wt.%, 1 wt.% and 2 wt.% of O-MWCNTs, reinforced laminates yielded 9.7%, 20.5% and 33.1% and improved ILSS properties compared to neat laminates (Fig. 5.12). Finally, it was concluded that the dispersion of MWCNTs in the matrix material played a crucial role in improving the ILSS property of the laminate. Anbusagar et al. [100] used the nanoclay-doped GF/polyester composite laminates to sandwich the polystyrene foam. The fabricated sandwiched specimens were subjected to LVI. The nanoclay was modified before using it in GF/polyester composite laminate. The nanoclay content used was 2, 4 and 6 wt.%. The 4 wt.% nanoclay-doped sandwiched composite showed highest peak load. It was found that the addition of nanoclay reduced the damage degree. Agarwal et al. [101] studied the impact behaviour of silicon carbide-filled GF/epoxy composite laminates. Five different wt.% 0, 5, 10, 15 and 20 were added to the GF/epoxy composites. Charpy impact tests were performed on the fabricated specimens. Addition of silicon carbide reduced the void content in the GF/epoxy

Fig. 5.12 Interlaminar shear strength (ILSS) of hybrid O-MWCNT/glass/epoxy composites made by injection and double vacuum-assisted resin transfer moulding (IDVARTM) method [99]

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laminates. Further, the impact strength of GF/epoxy composite was increased upon addition of silicon carbide. Dolati et al. [102] used the nanoclay in GF/epoxy composite laminates to study the hail impact damage. The experimental investigation was conducted for both single and repeated impacts. The nanoclay content used was 0.5, 1.5 and 3 wt.% by epoxy weight. The nanoclay was mixed in epoxy using high shear homogenizer. Experimental results showed that the 1.5 wt.% of nanoclay yielded optimal results compared to other 0, 0.5 and 3 wt.% doping. It was concluded that the addition of nanoclay enhances the impact strength and influences the damage area formation. Impact behaviour of a glass bubble, nanoclay and hybrid (glass bubble + nano \clay)-doped GFRP laminates was investigated by Koricho et al. [103]. Hand lay-up technique was used to fabricate the laminate with 1 wt.% of nanoclay, glass fibre, glass bubble and hybrid fillers and cured at 60 °C for two hours followed by post-cured at 94 °C for four hours. The laminates consist of eight layers with [45°/−45°/0°/ 90°]S lay-up. In the case of hybrid doping, the glass bubble and nanoclay-doped laminates were placed outside and inside the laminate about the mid-plane. Reason for this design was that the glass bubbles have a better impact and scratch resistance, whereas the nanoclay has crack prevention properties. 1 wt.% nanoclay and glass bubble-doped GFRP laminates showed lower (2.78 ± 0.11 kN) and highest (3.28 ± 0.26) stiffness. Meantime 1 wt.% nanoclay-doped GFRP and neat GFRP showed highest (160.27 ± 7.52 J) and lowest (154.41 ± 4.63 J) energy absorption capacity. Hybrid-doped laminate showed intermediate properties. Thus, hybrid reinforcement provided better tailorability of laminates. Alemi-Ardakani et al. [104] examined the fibre weaving architecture on the impact resistance property in GF/epoxy composite laminates. Four different types of weaving architectures plain woven, twill woven, unbalanced twill woven and unidirectional were considered for the experimentation. The LVI tests results showed that both plain woven and unidirectional GF/epoxy composite laminates showed severe local damage, whereas the twill GF/epoxy composite laminate showed more uniformly distributed damage rather localized. The twill weave GF/epoxy composite laminate absorbed more energy than the plain woven GF/epoxy composite. It was found that among all weaving architectures the unidirectional composite laminate performed better in terms of absorbed energy, maximum reaction force and central deflection. Balali et al. [105] used the shear thickening fluid in GF/epoxy composite laminate to improve the low velocity impact resistance properties. Here, the nanoclay was added to the shear thickening fluid. The nanoclay content used was 1 and 3 wt.%, while the nanosilica content was 30 and 20%. Experimental results showed that the addition of nanosilica and nanoclay increased the peak load and energy absorption because addition of nanosilica and nanoclay mitigated the crack propagation, while use of shear thickening fluid increased the fibre pull-out load. Effect of manufacturing methods on impact response of the unidirectional glass/epoxy laminates doped with nanoclay under low velocity impact loading was examined by Meyobodi et al. [106]. Along with that effect of nanoclay types (Cloisite 30B and Cloisite 15A) on impact damage area was also studied. Six ply laminates

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with nanoclay content (0, 3, 5 and 7 wt.%) and [0°/90°/0°] were fabricated using hand lay-up and vacuum-assisted resin transfer moulding method. The quasi-static test was conducted with 2 mm/min crosshead speed, and low velocity impact test was performed on drop weight tower with an impactor of mass and diameter 3.27 kg and 10 mm, respectively. From the experimental results, it was observed that the nanoclay has less effect on peak load, which was primarily influenced by the fibre type used. The damaged area reduced as the doping weight percentage of the nanoclay increased from 3 to 7wt.%. Among Cloisite 15A and Cloisite 15B, nanoclay 15A absorbed more energy than 30B; meanwhile, laminates fabricated with vacuumassisted resin transfer moulding methods showed the less damaged area and more energy absorption than the hand lay-up methods. Rafiq et al. [107] studied the nanoclay-doped GF/epoxy composite behaviour under LVI. The nanoclay content 1 wt.% and 3 wt.% was mixed into the epoxy resin using high shear mixing. The fabricated GF/epoxy/nanoclay nanocomposites were impact at 10 and 50 J of impact energy. Experimental results showed 23 and 11% enhancement in peak load and stiffness for 1 wt.% doping whereas 3 wt.% nanoclaydoped GF/epoxy showed 14% improvement in peak load. The enhancement in impact performance after doping of nanoclay into GF/epoxy was mainly due to transition in failure mechanisms where the crack formed was mitigated and crack propagation path was tortuous.

5.4 Summary The damage mechanism of GFRP composite laminates under LVI is affected by the laminate thickness, stacking sequence, impactor diameter and impactor geometry. Further, the stacking sequence, matrix material and impact energy strongly influenced the impact strength of the GFRP composite laminate. In hybrid composite laminates, the presence of GFRP plies at core or as face sheets proved to be more effective. Incorporation of GF fabric increases the flexibility of the hybrid composite laminate. Addition of nanomaterials increased the impact performance of GFRP laminates. However, the extent of dispersion depends on the nanomaterial dispersion state in the GFRP composite laminates. Further, the damage mechanism of GFRP composite laminates changes upon addition of nanomaterials.

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77. Deka LJ, Bartus SD, Vaidya UK (2009) Multi-site impact response of S2-glass/epoxy composite laminates. Compos Sci Technol 69:725–735. https://doi.org/10.1016/j.compsc itech.2008.03.002 78. Atas C, Sevim C (2010) On the impact response of sandwich composites with cores of balsa wood and PVC foam. Compos Struct 93:40–48. https://doi.org/10.1016/j.compstruct.2010. 06.018 79. Ghasemnejad H, Furquan ASM, Mason PJ (2010) Charpy impact damage behaviour of single and multi-delaminated hybrid composite beam structures. Mater Des 31:3653–3660. https:// doi.org/10.1016/j.matdes.2010.02.045 80. Fiore V, Di Bella G, Valenza A (2011) Glass–basalt/epoxy hybrid composites for marine applications. Mater Des 32:2091–2099. https://doi.org/10.1016/j.matdes.2010.11.043 81. Huang W, Huang Z (2010) Dynamic behavior of unidirectional fiber-reinforced composites: experiments and modeling. J Reinf Plast Compos 30:53–59. https://doi.org/10.1177/073168 4410386636 82. Ahmadi H, Sabouri H, Liaghat G, Bidkhori E (2011) Experimental and numerical investigation on the high velocity impact response of GLARE with different thickness ratio. Procedia Eng 10:869–874. https://doi.org/10.1016/j.proeng.2011.04.143 83. Dong C, Davies IJ (2012) Optimal design for the flexural behaviour of glass and carbon fibre reinforced polymer hybrid composites. Mater Des 37:450–457. https://doi.org/10.1016/j.mat des.2012.01.021 84. Venkatanarayanan PS, Stanley AJ (2012) Intermediate velocity bullet impact response of laminated glass fiber reinforced hybrid (HEP) resin carbon nano composite. Aerosp Sci Technol 21:75–83. https://doi.org/10.1016/j.ast.2011.05.007 85. Zhang J, Chaisombat K, He S, Wang CH (2012) Hybrid composite laminates reinforced with glass/carbon woven fabrics for lightweight load bearing structures. Mater Des 36:75–80. https://doi.org/10.1016/j.matdes.2011.11.006 86. Dong C, Ranaweera-Jayawardena HA, Davies IJ (2012) Flexural properties of hybrid composites reinforced by S-2 glass and T700S carbon fibres. Compos Part B Eng 43:573–581. https:// doi.org/10.1016/j.compositesb.2011.09.001 87. Petrucci R, Santulli C, Puglia D et al (2013) Mechanical characterisation of hybrid composite laminates based on basalt fibres in combination with flax, hemp and glass fibres manufactured by vacuum infusion. Mater Des 49:728–735. https://doi.org/10.1016/j.matdes.2013.02.014 88. Vijaya Ramnath B, Junaid Kokan S, Niranjan Raja R et al (2013) Evaluation of mechanical properties of abaca–jute–glass fibre reinforced epoxy composite. Mater Des 51:357–366. https://doi.org/10.1016/j.matdes.2013.03.102 89. González EV, Maimí P, De Aja JRS et al (2014) Effects of interply hybridization on the damage resistance and tolerance of composite laminates. Compos Struct 108:319–331. https://doi.org/ 10.1016/j.compstruct.2013.09.037 90. Petrucci R, Santulli C, Puglia D et al (2015) Impact and post-impact damage characterisation of hybrid composite laminates based on basalt fibres in combination with flax, hemp and glass fibres manufactured by vacuum infusion. Compos Part B Eng 69:507–515. https://doi.org/10. 1016/j.compositesb.2014.10.031 91. Kim D-H, Kim H-G, Kim H-S (2015) Design optimization and manufacture of hybrid glass/carbon fiber reinforced composite bumper beam for automobile vehicle. Compos Struct 131:742–752. https://doi.org/10.1016/j.compstruct.2015.06.028 92. Bienias J, Jakubczak P, Dadej K (2016) Low-velocity impact resistance of aluminium glass laminates—experimental and numerical investigation. Compos Struct 152:339–348. https:// doi.org/10.1016/j.compstruct.2016.05.056 93. Naresh K, Shankar K, Velmurugan R, Gupta NK (2018) Statistical analysis of the tensile strength of GFRP, CFRP and hybrid composites. Thin-Walled Struct 126:150–161. https:// doi.org/10.1016/j.tws.2016.12.021 94. Naik NK, Doshi AV (2005) Ballistic impact behavior of thick composites: analytical formulation. AIAA J 43:1525–1536. https://doi.org/10.2514/1.11993

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Chapter 6

Low Velocity Impact Test on Other Fibre Reinforced Polymer Composite Laminates

Abstract There are various other fibres reinforced polymer composite laminates apart from carbon and glass fibre reinforced polymer composite laminates. This chapter covers low velocity impact behaviour of these various other fibre reinforced polymer composites such as graphite, Kevlar, aramid, basalt, kenaf, hemp, jute and flax. Further, the chapter also covers the effect of hybridization on the impact and damage behaviour of these other fibre reinforced polymer composite laminates. The final section of this chapter covers the use of nanomaterials, particularly carbon nanotubes, graphene and nanoclay to improve the impact strength of these other fibre reinforced polymer composites laminates. Keywords Graphite · Kevlar · Aramid · Basalt · Kenaf · Hemp · Jute · Flax · Hybrid · LVI · Nanomaterials · CNTs · Graphene · Nanoclay

6.1 Impact Test on Controlled Other Fibre Reinforced Polymer Composites Laminates In this research investigation, Joshi [1] studied the impactor dynamics and contact modelling. A graphite/epoxy laminate with [90°5 /0°5 /90°5 ] lay-up was impacted by a rigid cylindrical impactor. The crack nucleation near the stress filed played a crucial role in defining the fracture initiation in the laminate during the impact loading. Elastic waves were propagated at a different speed in different directions because of the anisotropic nature of the laminate. Finally, it was concluded that the crack nucleation to a certain extent could be effectively predictable by coupling target dynamics with the impactor dynamics using static indentation law which was effective if and only if the impact duration was higher than the time required for longitudinal waves to travel through the thickness. Williams et al. [2] studied the effect of 24 different epoxy resin systems on the impact damage resistance of graphite composite laminates. It was found that the tensile properties of the epoxy matrix material played a significant role on the impact performance of graphite laminates. Further, the graphite composite laminates with

© The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_6

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matrix content greater than 40% of total volume of the laminate showed good damage tolerance. Wardle et al. [3] examined the impact resistance of Kevlar-49, Kevlar-29, aramid Thornel-300 graphite and glass fibre composite laminates. The Kevlar-29 composite laminate absorbed more energy than Kevlar-49, aramid, Thornel-300 and glass fibre reinforced polymer composites. However, the glass fibre reinforced polymer composite laminate showed intermediate impact resistance between Kevlar-29 and Kevlar-49 composite laminates. It was concluded that the impact resistance property of FRP composite laminates consisting of high strength fibres depends on the trade-off with the stiffness. However, the choice of fibre depends on the requirement. Sjoblom et al. [4] investigated better parameter characteristic to assess the impact damage of graphite/epoxy and graphite/polyethylether ketone (PEEK) FRP laminates under low velocity impact. Laminates consist of 16 and 48 plies with [0°/45°/−45°/90°]nS lay-up design. A pendulum type of impact testing machine was used with an impactor of mass 1 kg dropped from a height of 0.1 m which generated an impact energy of 1 J. This newly developed impact testing model yielded the results with 0.1% of accuracy. From experimental testing, it was observed that the laminates made of epoxy showed two distinct failures, i.e. delamination and damaged back face of the laminate, which included matrix cracking and fibre failure. Meantime, laminates with PEEK resin did not show any delamination. Finally, it was concluded that impact force history was a better parameter to assess the impact damage of the laminate than the total energy. Further, the energy loss was the direct measure of the damage than the impact of energy. Masters [5] incorporated thin and ductile resin film as interleaving between the interfaces of lamina in graphite/epoxy and graphite/bismaleimide composite laminates. The low velocity impact experimental results showed that nearly 80% of the impact resistance property of the laminate was improved and delamination formation in the laminate was reduced. Bogdanovich and Yarve [6] investigated the impact resistance behaviour of graphite/epoxy composite laminates. The experimental results showed that the damage in the graphite/epoxy composite was sensitive to impact velocity and impactor mass with impact energy being constant. Lane et al. [7] discussed in detail regarding the use of eddy current method which is detecting the damage in graphite/epoxy composite laminates induced by impact loading. Experimental results confirmed that the eddy current inspection process does not detect the delamination, whereas it effectively detects the fibre breakage in the laminate. Thus, it was concluded that the eddy current detection method should be used as a differentiator rather than as a detector. Choi et al. [8] investigated the relationship between impactor mass, stacking sequence and laminate thickness when graphite/epoxy composite laminates were subjected to low velocity impact. Experimental results showed that the different ply orientations distributed uniformly in the laminate have a good impact resistance. Further, both laminate sequence and impactor mass were found to have a greater influence on the impact resistance in graphite/epoxy laminate than the laminate thickness. Minguet et al. [9] investigated the influence of manufacturing process on the damage resistance property of graphite composite laminate under impact loading

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used in helicopter drive shafts. The graphite composite laminates were fabricated using tow-placement process, filament winding process and resin transfer moulding process. Further, PEEK and/or epoxy and/or toughened epoxy resins were used as matrix materials. The graphite/PEEK shafts were fabricated using tow-placement and filament winding process, while the graphite/epoxy and graphite/toughened epoxy shafts were fabricated using resin transfer moulding process. Experimental results showed that the delamination as common failure mode in all shafts. The graphite/PEEK shafts made from tow-placement and filament winding showed better properties than graphite/epoxy and graphite/toughened epoxy shafts fabricated using resin transfer moulding. Demuts [10] examined the effect of laminate thickness and ply orientation on the impact damage property of graphite/PEEK and graphite/toughened BMI composite laminates. Experimental results showed that the thin graphite/PEEK composite was more damage tolerant than the graphite/toughened BMI composite laminate. Karasek et al. [11] investigated the influence of moisture on the impact resistance property of graphite/epoxy composite laminate as a function of temperature. Experimental results showed that the moisture did insignificantly reduced the impact resistance property of graphite/epoxy composite laminate while at elevated temperature, the moisture significantly reduced the graphite/epoxy composite laminates impact resistance property. Sanadi et al. [12] studied the impact strength of the kenaf/polypropylene composite laminates. The kenaf/polypropylene laminates were fabricated using thermokinetic mixer followed by injection moulding. The interfacial strength between kenaf fibre and polypropylene matrix was good because the used polypropylene matrix was maleated. Experimental results showed that the impact strength of kenaf/polypropylene composite laminate was less compared to glass/polypropylene. However, it was concluded that the notched impact strength of kenaf/polypropylene can be improved by using maleated copolymers. Devi et al. [13] conducted impact test on pineapple leaf fibre reinforced polyester composite laminates. Fibre surface treatment along with fibre content and fibre length was considered as testing parameters. The surface of the fibres was treated with NaOH and silane coupling treatment. Addition of pineapple fibre to the polyester matrix resulted in improvement of ductility of the matrix. The pineapple leaf fibre with length 10–30 mm showed good impact strength. Further, the optimal fibre weight fraction was found to be 30%. Furthermore, significant improvement in impact strength was observed after pineapple fibre surface treatment. Karnani et al. [14] studied the effect of matrix functionalisation and fibre surface modification on the impact resistance property of polypropylene-kenaf fibre reinforced composites. The polypropylene matrix was modified with maleic anhydride. Kenaf fibres were surface modified with siloxane chains. The Izod impact tests were conducted to investigate the impact resistance property of prepared composite laminates. From experimental results, it was observed that the matrix modified polypropylene-kenaf composite laminates showed better impact resistance properties than the unmodified polypropylene-kenaf composite laminates. It was found that the enhancement in the interfacial bond strength between polypropylene matrix and

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kanf fibre after matrix modification was the major reason for the improvement in the impact resistance property. Hill et al. [15] used different chemical treatment methods to improve the impact strength of oil palm and coir fibre reinforced polyester matrix. Different chemical treatments include acetylation or silane coupling or titanate coupling. The impact tests were performed according to BS2782 standards on a pendulum type impact machine. The acetylation chemical treatment increased the impact strength. However, no significant improvement in impact strength was observed for silane or titanate coupling chemical treatment. Wu et al. [16] examined the fibre treatment effect on the interfacial and impact strength in wood fibre reinforced polypropylene composite laminates. The wood fibres were pretreated with acid-silane coupling agents then subjected to Charpy impact test. The fibre pretreatment significantly improved the interfacial strength between wood fibre and polypropylene matrix which intern increased the impact strength of wood fibre reinforced polypropylene composite laminates. It was concluded that the matrix toughness was improved after fibre pretreatment which also avoided fibre premature brittle failure. Schoeppner et al. [17] fabricated the graphite composite laminates using different matrix materials such as epoxy, PEEK and BMI. Then all the fabricated graphite composite laminates were subjected to low velocity impact to investigate the matrix influence on the delamination threshold limit. Among epoxy, PEEK and BMI matrix materials, the graphite/BMI composite laminate showed highest damage threshold limit. Ray et al. [18] investigated the dynamic mechanical and thermal properties of jute fibre reinforced with vinylester as a function of fibre volume fraction, which was treated with alkali solution (5% NaOH). Samples were fabricated using hand lay-up method in a circular cylindrical shape and cured at room temperature for 24 h, followed by post-curing at 80 °C for four hours. The untreated fibre volume fraction varied from 23, 30 and 35%. Whereas the treated fibre volume fraction was 35% with varying treatment conditions, i.e. four and eight hours in 5% NaOH. Flexural testing carried out according to ASTM D790M-81 with 2 mm/min crosshead speed. Dynamic thermal properties were tested at a frequency of 1.0 Hz with a heating rate of 5 °C/min in the temperature range of 30–210 °C. From experimental results, in all combination, the 35% fibre volume fraction treated with 5% NaOH solution for four hours showed better properties than other laminates with a flexural modulus of approximately 15 GPa and loss modulus of 1158 MPa. Finally, it was concluded that the enhancement in that sample was due to improved interfacial bonding between fibre and matrix. However, in the case of eight hour treated fibres, the crystallinity was increased, leading to reduced mechanical and dynamic properties. Mohanty et al. [19] studied the tensile, impact, flexural, dynamic and thermal properties of MAPE treated jute/high density polyethylene (HDPE) laminate. MAPE treated jute/HDPE, untreated jute/HDPE and HDPE were fabricated at 120 °C with 25 rpm of rotor speed. In jute/HDPE, fibre volume fraction was varied from 10, 15, 30 and 45%, whereas in MAPE treated jute/HDPE, the MAPE was varied from 0.3, 0.5, 1 and 2%. Tensile, flexural and impact tests were conducted according to ASTM D638, D790 and D256 on the universal testing machine, three-point bend and Izod impact

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machine, respectively. The dynamic behaviour, interfacial and thermal properties were studied by using rheometrics, scanning electron microscopy, Fourier transformation infrared spectroscopy (FTIR) and thermogravimetry (TGA), respectively. From experimental results, it was observed that at 30wt.% of fibre content in untreated jute/HDPE yielded maximum result with tensile, flexural and impact strength of 27.24 MPa, 34.83 MPa and 51.25 J/m, respectively. Below and above this weight percentage, the properties of laminate was reduced significantly. MAPE concentration at 1% yielded better properties with tensile, flexural and impact strength of 41.14 MPa, 47.97 MPa and 65.59 J/m, respectively, and beyond this concentration value, the properties were reduced. From all the experimental and characterisation analysis, it was concluded that with optimum fibre and coupling agent, content in jute fibres reinforced with HDPE laminates could yield better results. In this study, Bledzki et al. [20] conducted a comparative analysis between abaca, jute and flax fibre reinforced polypropylene composite laminates in terms of impact resistance property. The fibre content of abaca fibre used was 20, 30, 35, 40, 45, and 50 wt.%, whereas the jute and flax fibre content used was 30wt.%. The impact strength of abaca fibre reinforced polypropylene composite laminate was increased with increase in abaca fibre weigh percentage upto 30 wt.%. The abaca fibre reinforced polypropylene composite showed good impact resistance properties compared to jute fibre reinforced polypropylene composite laminate. It was found that 40 wt.% of abaca fibre as the optimal fibre loading and use of maleated polypropylene improved the interfacial strength with the abaca fibre. Chakraborty et al. [21] developed a numerical model to study the delamination behaviour in unidirectional graphite/epoxy laminate under multiple impact loading. For numerical analysis, Newmark β method and Hertzian contact law were used. The cylindrical impactor used for multiple impact study with a time interval of 25, 50 and 105 μs. The new numerical model for multiple impact loading showed that along with the impactor mass and velocity, the impact time interval also decided the magnitude of contact force. Further, the two delamination coalescence into one large single delamination or remains two distinct delaminations. Gower et al. [22] studied the effect of conical and hemispherical shaped impactor on woven Kevlar/polyvinyl/butryl under ballistic impact. The objective of this study was to examine the back-surface signature for each impactor. From the experimental and numerical analysis, it was found that Kevlar-129 yielded less resistance to sharp (conical) impactor. Kevlar-29 exhibited the lower back-surface signature than Kevlar-129 under low velocity impact for both the impactor shape (Fig. 6.1). Kevlar/Polyvinyl laminates performance under ballistic limit was investigated by Soykasap et al. [23]. The tests were conducted using a 9 mm full metal jacket bullet of diameter 9 mm. The deflection of the plate under impact varied with temperature and maximum deflection was observed at −30 °C followed by 60, 30, −15 and 15 °C. Effect of intralayer and interlayer bridging of the carbon nanotubes reinforced composites under mode-I fracture, and tension bearing loading conditions investigated by Wicks et al. [24]. By chemical vapour deposition (CVD) method, carbon nanotubes (CNTs) were grown directly onto the aluminium fibre cloth (satin weave, Fuzzy fibre) and samples were fabricated by hand lay-up method assisted by

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Fig. 6.1 Maximum dynamic versus projectile velocity [22]

vacuum bagging. Further, the samples cured at room temperature for 24 h. Mode-I tests were carried out according to ASTM D5528 standard at 2 mm/min crosshead speed, and ASTM D5961 was followed in conducting the tension bearing testing. GIC fracture initiation for aluminium/epoxy was 21 ± 0.07 kJ/m2 , whereas for fuzzy fibre/epoxy, it was 2.02 ± 0.32 kJ/m2 and ultimate bearing strength for Al/epoxy was 236 ± 10 MPa; meantime, it was 248 ± 17 MPa for fuzzy-fibre/epoxy laminate. From scanning electron microscopy analysis, it was concluded that the enhancement was mainly because of CNTs bridging between the interlayer and intralayer of fibre and matrix. Srivastava et al. [25] used the shear thickening fluid to improve the impact resistance property of Kevlar fibre. The fabricated Kevlar composite laminates were subjected to impact loading using a drop weight impact machine. Experimental results showed that the Kevlar fibres treated by shear thickening fluid with appropriate padding pressure can improve the impact resistance property. Gopinath et al. [26] studied the role of interlaminar bonding between matrix and fibre when laminate was subjected to ballistic impact. Laminates were prepared by using Kevlar fibre and resin matrix. A 9 mm full metal jacket bullet was used as an impactor. LS-DYNA was used for numerical analysis. Damage mechanism was modelled using MAT_COMPOSITE_DAMAGE _MODEL, ERODING_SINGLE_SURFACE model was used to define the contact between laminate layers and MAT_PIECEWISE_LINEAR_PLASTIC was used to model the matrix material. From the analysis, it was observed that the stiffer matrix was less effective in slowing down the projectile speed than soft matrix but, the kinetic energy was effectively reduced by stiffer matrix. Majumadar et al. [27] subjected the shear thickening fluid treated Kevlar fibres to impact loading. Experimental results showed that the untreated Kevlar fibre absorbed 40% out of total applied energy at the time of failure, whereas the shear thickening

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fluid treated Kevlar fibre absorbed 80% of energy out of total energy applied. Feng et al. [28] proposed a new finite element model for precise prediction of throughthickness damage using progressive damage model based on continuum damage mechanics. Results from the developed model validated by drop weight and X-ray radiography experimentation. An [0°3 / + 45°/−45°]S lay-up of graphite/epoxy laminate was fabricated using vacuum bagging and cured between 85 °C – 125 °C temperature range. Further, the laminates were subjected to low velocity impact using drop weight tower and a hemispherical nose impactor of mass 2.34 kg with nose diameter of 12.5 mm. Impact energy was varied from 1 to 8.4 J. The numerical model was developed using ABAQUS/Explicit and VUMAT platform. The developed damage model consists of intralaminar and interlaminar failures. The developed model was not able to predict the through-thickness failure and damage but predicted the impact response resulted from the developed model was in good correlation with that of the drop weight and X-ray radiography experimental results. Thus, the authors suggested considering in-plane damage model to predict the damage response of the laminates under impact loading. Ansari et al. [29] studied the behaviour of graphite/epoxy laminates under low velocity impact considering various parameters such as laminate thickness, ply sequence, impactor size, impact velocity and boundary conditions. ABAQUS finite element tool was used to analyse these conditions and evaluated the damage mechanism behaviour using Hashin’s failure criterion. Ply sequence/orientations [45°/−45°/45°], [0°/90°/0°] and [30°/−30°/30°] were used. In numerical modelling, eight-noded continuum shell (QC8R) was used to model the laminate. The impactor used was cylindrical with a hemispherical nose of 5 and 10 mm diameter and 0.01475 kg of mass. From the numerical analysis, it was seen that as the plate thickness increased, the corresponding damaged area in the laminate was decreased. Damage initiation and propagation were less proactive in fully clamped boundary condition than other types of boundary conditions. Matrix cracking in tension or delamination was the main factor for the damage in laminates under low velocity impact. Out of three laminate the [45°/−45°/45°] stacking sequence showed better results in terms of stiffness and deflection than others. Boria et al. [30] studied the behaviour of thermoplastic matrix material under low velocity impact with varying mass and velocity. A polypropylene laminate of 53 0°–90° plies were prepared using hot press machine at 130 °C under 100 bar pressure for 2 h and impacted with a hemispherical dart impactor of 20 mm diameter with a varying mass of 7.3 and 67.3 kg. From experimental results, it was found that there was no perforation involved; only penetration and rebounding occurred due to the thickness of the laminate. It was also found that the laminates failed in ductile nature, mostly due to the yarn sliding mechanism. Graupner et al. [31] studied the mechanical properties of new bio-inspired petioles of red Rubarb hybrid laminates. The Ramie (R), Lyocell (L) fibres were reinforced with polyactide (PLA) and mixed well and cured at 105 °C for two hours, followed by hot pressing for five minutes at 105 °C. Six laminate configurations PLA, PLA/Ramie, PLA/Lyocell, PLA/L-R mixed, PLA/LR-L sandwiched and PLA/R-LR sandwiched were subjected tensile, flexural and impact testing according to ISO 257-2, 178 and 179, respectively. The failure and

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morphological studies were done by using a scanning electron microscope (SEM). Ramie/PLA performed better in tensile and flexural strength than others. However, Lyocell/PLA showed better energy absorption capability than others. From results, it proved that the impact resistance was increased by more than two factors. SEM provided the evidence of two stage energy absorption, i.e. straightening of fibres and pull-out of fibres. Thus, it was concluded that these bio-inspired fibres performed better when mixed with proper strengthening materials. Tensile and impact properties of cementious PVA and shape memory alloy (SMA) hybrid composites were tested by Ali et al. [32]. Effect of heat treatment on impact properties, i.e. prestressing was also studied, and the repeatability of the experiment was evaluated by two-parameter Weibull distribution method. The fabrication of samples were done by the gradual pouring of ingredients into a mould and left for 24 h curing followed by demoulding and post-cured for seven days. PVA volume was fraction kept constant (2%), and SMA fibre volume fraction (Vf ) was varied, i.e. 0.5, 1.0 and 1.5%. Impact test was carried using a drop weight impactor with 4.5 kg impactor at 20.167 J of impact energy. From experimental results, it was evident that 1.0% V f of SMA showed better characterisation than other volume percentages of SMA. However, the heat-treated samples showed even better impact resistance characteristics at the same fibre volume fraction. Finally, it was concluded that the enhancement was due to the bridging effect of SMA, but beyond 1.0% volume, fraction enhancement was not observed, which might be due to the fibre accumulation at one place leading to void formation and thus, reduced the impact resistance capacity of the laminate. In this study, Habibi et al. [33] addressed the damage tolerance of non-woven jute fibre reinforced epoxy composite laminates. Two impactors hemispherical and conical were used to impact the fabricated jute mat/epoxy composite laminates. Further, the fabricated laminates were impacted at 4, 6, 8, 10, 12 and 14 J of impact energy. The experimental results revealed that the conical impactor damage was significantly consists of matrix cracking and delamination along with smaller damage area compared to hemispherical impactor. It was concluded that the impactor shape and impact energy both significantly affected the damage tolerance of jute mat/epoxy composite laminates. Cuynet et al. [34] used the high speed imaging and stereo image correlation to characterise the LVI induced damage in flax/epoxy composite laminates. Eight plies of twill flax fabric were used to fabricate the flax/epoxy composite laminate. The fabricated specimens were subjected to LVI using drop tower impact machine. The specimens were subjected to 4–34 J of impact energy. The back face cracks were easily identified by high speed imaging. Puech et al. [35] captured the crack formation, propagation during LVI in short hemp fibre reinforced polypropylene composite laminates using high speed imaging. The hemp/polypropylene composite laminates absorbed high impact energy before failure compared to GF/polypropylene composite laminate. The high speed imaging method revealed that the crack propagation absorbed more energy in hemp/polypropylene composite than the glass/polypropylene composite laminates.

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Fig. 6.2 Schematic illustration of LVI experimental setup [36]

Sy et al. [36] examined the impact resistance behaviour of unidirectional and cross-ply flax/epoxy composite laminate. The stacking sequence of unidirectional flax/epoxy was [0]8S and cross-ply flax/epoxy composite laminate was [0/90]4S . The LVI tests were conducted using pendulum type impact machine (Fig. 6.2). Under LVI, the cross-ply flax/epoxy composite laminate showed 3 and 2.5 times greater penetration and impact toughness compared to unidirectional flax/epoxy composite laminate. Papa et al. [37] validated the indentation and penetration of semi-empirical models with the LVI experimental data obtained for glass fibre reinforced phenolic composite laminates and basalt fibre reinforced epoxy composite laminates. Two stacking sequences of GF/phenolic composite laminates used were quasi-isotropic [((0/90), (+45/−45))n ]S (n = 2–4) and cross-ply [(0/90)]n . The stacking sequence of basalt/epoxy composite laminate was [((0/90), (+45/−45))n ]S with n = 2– 4. Experimental results revealed no significant difference in impact performance between cross-ply and quasi-isotropic GF/phenolic composite laminates. Further, the basalt/epoxy composite laminates showed highest indentation resistance.

6.2 Low Velocity Impact Test on Other Hybrid Fibre Reinforced Polymer Composites Laminates Park et al. [38] fabricated the aramid and polyethylene intraply hybrid composite laminates using the open leaky mould process. The fabricated hybrid laminates were subjected to Izod impact. The experimental results showed that the impact strength of the aramid/polyethylene hybrid composites strongly and proportionally depends

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on the aramid fibre content. In an experimental, Park et al. [39] examined the influence of stacking sequence on the impact performance of hybrid glass/aramid/epoxy laminates. The stacking sequence of hybrid composite laminates was G, A, GG, AA, GA and AG, where ‘A’ is aramid fibre ply and ‘G’ is glass fibre ply. Addition of glass fibre with the aramid plies reduced the impact performance hybrid composite laminates. Position of aramid plies in the laminate strongly influenced the impact strength of hybrid glass/aramid/epoxy composite laminates. The hybrid composite laminate with aramid ply as face sheet showed highest impact strength. Pegoretti et al. [40] examined the impact property of intra and interply glass/poly(vinyl alcohol)/polyester hybrid composite laminates. Different stacking configuration of interply and intraply hybrid glass/poly(vinyl alcohol)/polyester composite laminates was used. The intraply hybrid composite laminate showed greater resistance to impact crack propagation than the interply hybrid composite laminates. Mueller et al. [41] evaluated the impact strength of flax, hemp and kenaf natural fibres based on fibre fineness, processing temperature and duration of preheating. Laminates were impacted by steel projectile by pressurised air through a barrel. Piezoelectric sensors were used to evaluate the growth of the damage during impact. From experimental analysis, it was concluded that the mid-range processing temperature laminates showed better results than other processing temperatures. Nevertheless, the glass fibre laminate was unaffected from processing temperature variations. Finally, it was concluded that fibre fineness, processing temperature and duration of processing played a significant role in deciding the impact strength of a laminate. Burgueno et al. [42] investigated the mechanical strength of hybrid biocomposites. Various combinations of fibres were involved: (1) Raw hemp fibre core was sandwiched between chopped glass strand mat. (2) Raw hemp fibre core was sandwiched between woven jute fibre. (3) Green hemp with woven jute hybrid laminates (4) Green hemp/glass strand mat hybrid material. Several material properties such as tensile, impact, coefficient of thermal expansion (CTE) and moisture absorption properties evaluated. Tensile and impact tests were conducted according to ASTM D790 and D256. CTE was measured from room temperature to 140 °C at 4 °C/min. Moisture absorption tests were carried out in an environment of 30 °C and 90% humidity. From experimental results, it was concluded that fibres used as core material sandwiched between glass mat face sheets performed better. An experimental investigation was conducted by Burgueno et al. [43] to test the load-bearing capacity of industrial hemp and flax fibre reinforced with unsaturated polyester. Natural fibres were dried in an oven at 80 °C under 100 kPa pressure for ten hours followed by hand mixing method technique and cured the hybrid biocomposites sample under 550 kPa for four hours. Further, the laminates were heat treated at 100 °C for two hours and at 150 °C for two more hours. Tensile, impact, coefficient of thermal expansion (CTE) and moisture absorption tests were conducted. Tensile strength of E-glass was found to be 2.3 times better than biocomposites. Impact strength of E-glass was 5.6 times better than natural fibre laminates, but impact strength for biocomposite found to be increased from 6.86 to 10.40 J/m as the fibre volume fraction of hemp fibre content increased from 13 to 25%. CTE of

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biocomposite material was 30% greater than E-glass laminates. Moisture absorption test conducted for 50 days under a controlled condition, and it absorbed four times moisture content than E-glass laminates. The analytical analysis which was carried out using Halpin Tsai model was in good agreement with the tested results, and it was concluded that the biocomposites could compete with the E-glass fibre in certain aspects. In an experimental investigation, Ahmed et al. [44] studied the impact strength of untreated jute and glass fibre reinforced isothalic polyester composite laminates. The hybrid jute/glass/polyester composite laminates were fabricated using hand layup process. The fabricated hybrid composite laminates were impacted by a hemispherical impactor of nose diameter 10 mm at 1.5 m/s impact velocity. The hybrid composite laminate with glass weight fraction 17.2% showed similar specific impact strength as that of glass/polyester composite laminates consisting of 60% glass fibre. In this experiment, Hosseinzadeh et al. [45] conducted experimental and numerical investigation to study the performance of body armour made of ceramic, FRP and rubber composite material under high velocity impact. LS-DYNA was used for numerical analysis with material card 7 to model rubber layer, which consists of hyper-elastic continuous rubber element. Isotropic elastic–plastic element model (material type 13) used for ceramic layer and material type 22 for the composite damage failure. A gas gun was used to carry out the experimental study under high velocity impact test. During the experiment, a spherical projectile with velocity 610 m/s was projected and observed that the projectile was trapped in a ceramic layer because it dissipated the energy quickly. The numerical model used for damage analysis of ceramic was based on a quasi failure strain criterion, which was a strength based model. Thus, a deviation in the numerical result from the experimental result was observed. Nevertheless, good agreement was observed between rubber interfaces, where delamination occurred mainly due to interlaminar shear stress rather than tensile stress. Wang et al. [46] fabricated the basalt/aramid hybrid interply and intraply FRP composite laminates using 3D stitching process. The fabricated composites were subjected to low velocity impact using a drop weight impact machine. The interply basalt/aramid hybrid composite laminates showed lower specific energy absorption than the interply basalt/aramid hybrid composites. Further, the interply hybrid composite laminates showed layer-by-layer failure and cracks propagation, while the intraply hybrid composite laminates showed brittle failure. Roy and Chakraborty [47] studied the impact resistance of hybrid Graphite/Kevlar/epoxy composite laminate under transverse impact consisting of holes. The stacking sequence of the hybrid composite laminates was [0/−45/45/90]2S . Under impact loading, the delaminations in the hybrid composite laminates were initiated from the free edges of the holes. From experimental results, it was found that the presence of Kevlar plies at the outer face of the hybrid composite increased the impact resistance. Wang et al. [48] studied the behaviour of long-run basalt/carbon fibre hybrid laminates when used in a cable bridge application. Various parameters such as static behaviour, dynamic behaviour, sagging effect, vibrational damping effect and aerodynamics stability of hybrid, CFRP and BFRP and steel cables were investigated.

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From the analysis, it was found that hybrid basalt/carbon hybrid laminate with 28% volume proportion of carbon fibre content exhibited better impact properties. These hybrid cables can possess much higher natural frequency and exhibited much better aerodynamics stability when compared to CFRP and steel cables. Finally, it was concluded that B/CFRP hybrid cables could be used in cable bridges instead of CFRP and steel cables in order to reduce the cost without compromising with any mechanical properties. In this study, Wang et al. [49] evaluated the strength and cost of the 1000–10000 m super long cable stayed bridges made of basalt hybrid FRP composites and compared with the steel, CFRP and BFRP laminates. Total five materials, i.e. steel, CFRP, BFRP, B/CFRP and B/SFRP were studied by considering safety factor, cable length and cost. From the analysis, it was found that B/SFRP composite can save up to 30% of the total material cost when compared to a steel cable and 50% of cost cutting can be done by using B/CFRP in comparison with the CFRP cables. Dekhordi et al. [50] studied the effect of nylon fibre content in basalt intraply hybrid laminate under impact loading. The nylon fibre content and impact energy were varied from 0–100% and 16–40 J, respectively. Hand lay-up method used to fabricate the four-layered intraply hybrid laminate with quasistatic orientation [(+45°, −45°)/(0°, 90°)]. Drop weight tower impact machine used to conduct the experimental work according to the ASTM D7136 standards. Hemispherical tip impactor with mass 1.22 ± 0.01 kg and diameter 12.7 mm was used. From the analysis, it was found that the fibre content of basalt and nylon decided the damaged area and energy absorption capacity of the hybrid laminates. From the visual analysis, the damaged shape found to be circular. The C-scan analysis gave the damaged area, and it was highest and least in nylon and basalt laminates, respectively. Meanwhile, 66% basalt and 35% nylon intrahybrid laminate performed better than other hybrid laminates in terms of damage resistance. Effect of varying relative fibre volume fraction of chopped and sisal fibre hybrid composites on mechanical properties were investigated by Idicula et al. [51]. Hybrid laminates sequence were [Banana/Sisal/Banana], [Sisal/Banana/Sisal] and [Sisal/Banana]. Mechanical properties such as tensile, flexural and impact properties were investigated on a universal and Izod impact testing machines by keeping constant fibre volume fraction (0.40). The hybrid composite prepared by hand layup method and cured at 30 °C for 24 h under 1 MPa of pressure. From experimental results, it was observed that the laminate with 47.5% banana fibre content performed better in the flexural property. In case of the sequencing of the bilayer (Banana/sisal), hybrid laminate yielded better flexural strength (73 MPa). Whereas, [Banana/Sisal/Banana] laminates showed better tensile strength (59 MPa). Both [Banana/Sisal] and [Banana/Sisal/Banana] laminates showed more or less same impact strength, i.e. 34 and 37 kJ/m2 , respectively, when subjected to 25 J of impact energy. Scanning electron microscopy used to study the interface and failure characteristics of hybrid FRP composites after tensile and impact testing. It was concluded that overall positive effect of hybridisation observed for tensile strength when the volume fraction of sisal fibre was 0.26, flexural strength when banana fibre volume

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fraction was 0.19 and impact strength increased when sisal fibre volume fraction increased. In this study, Wei et al. [52] investigated the bending and tensile strength of seawater solution treated basalt and glass fibre laminates for varying time scale, i.e. 10, 20, 30, 60 and 90 days. Gauge length and crosshead speed for both the materials were 150 mm and 2 mm/min. Nevertheless, bending strength for both the materials decreased continuously as the seawater treated time scale increased. To study the corrosion mechanism of both the fibres, the scanning electron microscopy (SEM) was used. It was found that the degradation in properties of both the laminates was due to diffusion and infiltration of seawater. Finally, it was concluded that basalt fibres were more stable than the glass fibre under seawater treatment. Lopresto et al. [53] studied the mechanical properties of E-glass and basalt FRP laminates. Laminates were prepared by hand lay-up method and cured at room temperature for 24 h. Ply orientation for tensile, compression, flexural, shear and impact testing of basalt laminates were [(0°/90°)]10 , [(0°/90°)]25 , [(0°/90°)]16 , [(0°/90°)]64 and [(0°/90°)]20 respectively. Similarly, ply orientation for glass fibre laminates were [(0°/90°)]8 , [(0°/90°)]10 , [(0°/90°)]10 , [(0°/90°)]36 and [(0°/90°)]10 used to conduct the tensile, compression, flexural, shear and impact respectively. For tensile, compression, flexural, shear and impact testing ASTM D3039/D3039M-00, D695-029, D793-03, D2344/D2344M-00 and EN-6038 standards were used respectively. Basalt fibre laminates showed all mechanical properties higher than that of the glass fibre laminates and were about 35–42%, 53–55%, 53–55%, 11–13% more with respect to tensile, flexural, compression and shear properties. Further, basalt fibre absorbed 10–12% more energy than the glass fibre laminate. Finally, it was concluded that the basalt fibre could compete with glass fibre in terms of cost and recyclability/biodegradability. Rahman et al. [54] used the polypropylene/glass/nanoclay hybrid laminate to study the dynamic, thermal and impact behaviour using thermal analysis instrument, differential scanning calorimeter and Charpy test, respectively. Laminates were fabricated by using a twin-screw extruder and injection moulding with varying nanoclay content 3, 6, 9 phr (parts per hundred parts of resin). Differential scanning calorimeter testing was carried out at a heating rate of 10 °C/min over a temperature range of 0 °C to 190 °C. ASTM standard E-23 was used to calculate the impact strength. Impact testing was conducted using an impactor mass of 6.448 kg, impact velocity of 0.9238 m/s and impactor energy of 2.512 J. Transmission electron microscopy (TEM) and scanning electron microscopy (SEM) was used to study the failure and fracture mechanism of the laminates under impact loading. From characterisation and experimental results, it was observed that as the nanoclay content increased, the melting point temperature of the laminate also increased. Meantime, there was no significant change in strength. Burger et al. [55] performed the numerical simulation of ceramic/fibre hybrid composite using ABAQUS/Explicit tool under ballistic impact. Here, two different grain sizes used with two different impact velocities, i.e. 765 and 665 m/s. In numerical simulation different failure models such as Johnson–Cook model to predict the behaviour of the structural material, JH- model to predict the ceramic material

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behaviour and 3D progressive failure model to predict the structural response of the composite and contact login to predict the debonding between the ceramic and fibre base were used. From the experimental and numerical analysis, it was evident that the adhesive model predicted the complete failure, whereas the adopted ceramic model did not show all failure stages involved in the damage mechanism. Further, it does not account for poor delamination and residual deformation. Dehkordi et al. [56] tried to improve the impact properties of basalt fibre by hybridising it with nylon fibre. The content of nylon fibre was 0, 25, 33.3, 50 and 100%. The hybridisation carried was intraply (Fig. 6.3). All these intraply hybrid basalt/nylon composite laminates were impacted at 16, 30 and 40 J of impact energy. The experimental results showed that the hybridisation of basalt fibre by nylon fibre did not improve the impact performance of basalt fibre composite laminate. Sarasini et al. [57] conducted low velocity impact test on aramid/basalt hybrid composite laminates. The resin transfer moulding process was used to fabricate the aramid/basalt hybrid composite laminates. There were two types of ply stacking sequence design (i) sandwiched (ii) intercalated. In sandwiched type design seven basalt fibre layers were sandwiched by three aramid fibre plies on each side whereas in intercalated design six basalt fibre plies and seven aramid fibre plies were placed

Fig. 6.3 Schematic illustration of intraply hybrid basalt/nylon/epoxy composite laminates [56]

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alternatively by keeping aramid fibre as skin or outer layer. The intercalated laminate design showed better energy absorption capability than the sandwiched laminate design. Jawaid et al. [58] investigated the mechanical and dynamic behaviour of jute and oil palm hybrid laminates as a function of fibre volume fraction of oil palm fibres. Bilayer hybrid composites were fabricated by hand lay-up method keeping total fibre weight fraction of 40% and cured at 105 °C for one hour and post-cured at 105 °C for 30 min. The tensile tests were conducted according to ASTM D3039 standard with a crosshead speed of 5 mm/min. The dynamic mechanical analysis carried out at 5 °C/min from −150 °C to 150 °C. From experimental results, it was concluded that oil palm and jute fibres with 1:4 fibre content performed better among other hybrid laminate combinations. In the meantime, a slight decrease in glass transition temperature observed. Scanning electron microscopy used to study the interface behaviour and found that a better interface adhesion observed at 1:4 oil palm and jute fibre combination. Finally, it was concluded that these hybrid laminates could compete with the synthetic FRP composites and can be used in the automotive and aerospace sectors. Sarasini et al. [59] studied the sandwiched and intercalated stacking sequence effect on the basalt/carbon hybrid laminates under low velocity impact loading. The laminates were fabricated using resin transfer moulding and cured at room temperature for 12 h and further, post-cured at 70 °C for four hours. In sandwiched type stacking sequence, seven carbon fibre plies sandwiched between three basalt fibre face sheets on either side. In the intercalated stacking sequence, seven layers of basalt fibres and six layers of carbon fibres stacked alternatively. Four-point bend tests under 2.5 mm/min crosshead speed were performed according to ASTM D6272. Falling dart impactor was used for impact testing with a hemispherical impactor of nose diameter 12.7 mm and mass of 6.929 kg. The laminates were struck at three impact energy levels 5, 12.5, 25 J. ASTM D2344 standard was followed to study the interface shear strength. Ultrasonic C-Scan and acoustic emission characterisation were used to study the damage progression. From experimental results, it was clear that the intercalated hybrid laminate performed better than sandwiched laminate because the sandwiched structures underwent interface failures. In an experimental work Dhakal et al. [60] investigated the influence of basalt fibre hybridisation with hemp fibre on the impact strength. All the fabricated composite laminates were impacted at 3, 6 and 9 J of impact energy. The compression moulding process assisted by hand lay-up process was used to fabricate the hybrid basalt/hemp/polyester composite laminate. The hemp fibre was in mat form, while the basalt fibre was in woven form (Fig. 6.4). Addition of basalt fibre to the hemp mat fabric improved the impact resistance and damage tolerance properties. However, it was observed that presence of hemp mat fibre at core resulted in sudden failure. Rajesh et al. [61] conducted Charpy impact tests on jute/banana/polyester intraply hybrid composite laminates according to ASTM D256 standard. Experimental results showed that the impact strength of jute/banana/polyester intraply hybrid composite laminate was higher compared to jute/polyester. Bozkurt et al. [62] studied the effect of position of basalt ply in basalt/aramid/epoxy hybrid composite laminates

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(a)

Hemp mat fabric

Basalt woven fabric

(b)

Fig. 6.4 Schematic representation, a hemp mat fabric and basalt woven fabric, b stacking sequence [60]

on impact strength. The hybrid composite laminates were fabricated by vacuum assisted resin transfer moulding process. Different stacking sequences of the hybrid basalt/aramid/epoxy were considered for the experimentation. Experimental results suggested that the presence of material at impact side influences both impact strength and damage of the hybrid basalt/aramid/epoxy composite laminate. The basalt plies restricted the deformation of aramid plies in the hybrid composite laminates. It was found that the hybrid composite laminate with basalt ply as outer layers in hybrid composite laminate exhibited the lower impact strength. Because the basalt ply at outer face restricted the deformation of aramid plies. The aramid fibre in hybrid composite laminates at outer layers showed highest impact strength. It was concluded that the impact performance of aramid fibre could be improved by the basalt fibre hybridisation. Bodur et al. [63] investigated the impact performance of hybrid glass/cotton/LDPE composite laminates. The interfacial interaction of glass fibre with matrix composite was improved by using maleic anhydride coupling agent. Impact tests on fabricated composite laminates were conducted according to EN ISO 179 standard. Both hybrid cotton/glass/LPDE composite laminates and maleic anhydride treated hybrid cotton/glass/LPDE composite laminates showed good impact properties. Muthu et al. [64] added the carbon nanofibre to improve the impact property of the glass/coir/polyester hybrid composite laminates. The glass and coir fibre used were chopped strand mat. The carbon nanofibres were synthesised from waste coal fly ash. The glass and coir fibres contents were varied. The hybrid glass/coir composite laminates reinforced with carbon nanofibre was fabricated by vacuum assisted resin transfer moulding method. The notched impact tests were conducted according ISO 179-1 standard. The coir fibre mat was surface treated with sodium hydroxide. The

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carbon nanofibre surface was functionalised by hydroxyl and carboxyl groups. Addition of carbon nanofibre to the glass/coir hybrid composites improved the impact performance because the carbon nanofibre improvised the load transfer efficiency from matrix to fibre. Hazzard et al. [65] studied the impact behaviour of ultrahigh molecular weight polyethylene fibre reinforced composites for different stacking sequences such as cross-ply, quasi-isotropic and rotational helicoidal. All the stacking sequences were subjected to LVI using a drop weight impact machine at an impact velocity of 3.38 m/s. Digital image correlation approach was used to capture full field displacement of the specimen using the LIV. It was observed that the cross-ply laminate showed highest back face deflection. The in-plane shear and buckling were the dominant failure modes observed in cross-ply and quasi-isotropic laminates respectively. The damage area of quasi-isotropic laminate design was less than the cross-ply stacking sequence. The rotational helicoidal laminate design showed bend-twist and extension twist coupling. It was concluded that the stacking sequence significantly influenced the deformation mechanisms in Dyneema composite laminates subjected to LVI. Sisal (S), carbon (C), glass (G) and hybrid laminate properties were investigated by Arulkumar et al. [66]. Laminates were fabricated by hand lay-up method with hybrid configuration [G/S/G/S/G], [S/G/S/G/S], [C/S/C/S/C] and [C/S/G/S/C] cured at room temperature under 350 psi for 24 h. Mechanical properties such as tensile, flexural and impact tests were conducted according to the ASTM D638, D790 and D256, respectively. From experimental results, it was evident that the tensile and flexural properties for [C/S/C/S/C] was found to be the highest among the other hybrid configuration. However, the impact strength was lowest for the same laminate. [S/G/S/G/S] configuration absorbed the highest impact energy (≈7.7 J) among other hybrid configurations. Ramesh et al. [67] investigated the impact strength, water uptake properties such as tensile strength and flexural strength of hybrid laminate composite made of banana and carbon fibres with epoxy matrix (Fig. 6.5). Maximum tensile strength, flexural Fig. 6.5 Samples in the solution for soaking [67]

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strength, impact strength and water uptake for CFRP laminates were 288.03 MPa, 3.12 kN, 4.58 J and 62.3%, respectively. Carbon and banana fibre hybrid composite can withstand the maximum tensile strength, flexural strength, impact strength and water uptake that were found to be 277.06 MPa, 3.07 kN, 4.36 J and 70%, respectively. ANSYS 15.0 was used to carry out numerical analysis to predict mechanical properties. Experimental and numerical analysis results were in good agreement. Impact loading behaviour of basalt and flax hybrid composite under dry and saltwater conditions was studied by Zivkovic et al. [68]. Eight-layered laminates were prepared by wet hand lay-up technique with a vinyl ester matrix. Then the samples were cured at room temperature for two hours and post-cured at 100 °C for three hours. Saltwater treatment of the fibres was carried out in 35 ppt saltwater at 80 °C for 912 h. Low velocity impact tests were carried out using a drop weight tower according to D7136 standards. Basalt and flax fibres absorbed the minimum and maximum percentage of water. Conditioned and dry flax fibres absorbed highest and lowest energy (mean absorbed energy 24.85 J and 10.39 J, respectively). In the case of hybrid composites, 2% enhancement and 7% reduction was observed at dropping heights of 2.5 and 3 m, respectively. Finally, it was concluded that hybrid laminate could be used for the betterment of impact resistance rather than using individual fibres separately. El-baky et al. [69] adopted the two-parameter Weibull statistical approach to examine the impact performance of glass/polypropylene/epoxy hybrid composite laminates. The hybrid composite laminates were fabricated using hand layup process. Experimental results showed that the stacking sequence of hybrid glass/polypropylene composite laminates has no significant effect on impact strength under edgewise impact while it showed significant influence under flatwise impact. Further, those hybrid composite laminates consisting polypropylene plies at impact site showed comparatively high impact strength because the flexible polypropylene ply experiences higher deformation. Furthermore, as the polypropylene content in the hybrid glass/polypropylene composite laminate was increased, the edgewise and flatwise impact strength was also increased. The intraply hybrid composite laminate showed higher impact strength than the inter and intraply hybrid composite laminates. Finally, it was concluded that the polypropylene composite laminates showed good impact strength compared to GF/epoxy composite laminates. Chaudhary et al. [70] investigated the tensile, flexural, impact and hardness properties of jute, hemp, flax and their hybrid laminates. Hand lay-up technique was used to fabricate the jute/epoxy, hemp/epoxy, flax/epoxy, jute/hemp/epoxy, hemp/flax/epoxy and jute/hemp/flax/epoxy laminates. Hardness test was conducted using shore D hardness tester. Tensile and flexural tests were conducted according to ASTM D3039 and D790 standards, respectively. The Charpy method was used to evaluate the impact strength of the laminate. Scanning electron microscopy (SEM) was used to examine the interfacial bonding between the fibre and matrix. Flax/epoxy showed the highest tensile and hardness, i.e. 98 and 46.2 MPa, respectively. Meantime, hemp/epoxy and jute/epoxy yielded maximum flexural and impact strength, i.e. 85.59 MPa and 7.68 kJ/m2 , respectively, whereas, the jute/hemp/flax showed the highest flexural strength (86.6 MPa). From SEM analysis, it was evident that the strong interfacial bonding

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between the matrix and fibre was the main reason behind the improved properties of the laminates. New construction material coconut fibre reinforced concrete (CFRC) wrapped around plain-woven flax fibre/epoxy laminates flexural behaviour under static, repeated and single impact loading was investigated by Wang et al. [71]. The flax fibre laminates (FFRP) were fabricated using the hand lay-up method. CFRC was prepared by mixing varying percentage of coconut fibre, i.e. 1, 35 and 5% of cement mass. Static test was conducted according to C78/78 M-10 standard to study the central deflection of the composite. Impact test was conducted on a drop weight impact machine using flat face impactor of mass 48 kg and diameter of 100 mm. It was noticeable that FFRP-CFRC laminate flexural strength was three times greater than CFRC composites, i.e. 13 and 4.92 MPa, respectively. Further, the failure criterion for both the composites was different. During impact testing, 3% CFRC-FFRP laminate showed better impact resistance than 1 wt.% and 5% wt.% of CFRC-FFRP composites. Finally, it was concluded that 3 wt% CFRC-FFRP yielded better impact and static characteristics than 1 and 5 wt.% CFRC-FFRP laminates and controlled FFRP composite. Carrillo et al. [72] fabricated the fibre metal laminate made of aluminium/aramid fibre/polypropylene subjected to LVI. Two different fibre metal laminates were fabricated. The first design includes two layers of aluminium and two layers of aramid/polypropylene, whereas the second design included three designs of aluminium and four layers of aramid/polypropylene layers. The fibre metal composite laminate with three layers of aluminium and four layers of aramid/polypropylene exhibited the highest impact resistance. The fabricated fibre metal laminate failed by aluminium layer plastic deformation and tear along with fibre breakage. Ricciardi et al. [73] examined the influence of stacking sequence on the impact performance of hybrid flax/basalt/epoxy composite laminates. Figure 6.6 shows the stacking sequence of hybrid flax/basalt/epoxy composite laminate. The intercalated stacking sequence (Laminate No, 3) showed gradual damage propagation at high impact energies. The flax fibre at skin area resulted in more damage absorption compared to other two stacking sequence because the flax fibre at skin area absorbs the resin material which results in fibre swelling. The fibre swelling was resulted into increase in fibre diameter which reduced the porosity.

Fig. 6.6 Different stacking sequence of hybrid flax/basalt/epoxy composite laminate (B-Basalt fibre; F-Flax fibre) [73]

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Fig. 6.7 Schematic representation of stacking sequences used and their corresponding fibre content [74]

Kumar et al. [74] conducted the impact test on different stacking sequences of hybrid flax/glass/vinyl ester composite laminates. The hybrid composite laminates were fabricated using vacuum assisted resin transfer moulding process. Figure 6.7 shows different stacking sequence of hybrid flax/glass/vinylester composite laminates used. It was found that in hybrid composite laminates, the glass fibre plies as outer skins exhibited good impact resistance properties compared to other stacking sequences. Finally, it was concluded that the positioning of glass fibre plies in the hybrid composite laminates significantly influenced the impact strength. Further, careful positioning of glass fibre ply in the hybrid composite laminate yielded impact strength similar to glass fibre composite laminates resulting overall reduction in weight. Giridharan et al. [75] examined the impact strength of hybrid galss/ramie/epoxy composite laminate. Pendulum type impact machine was used for impact tests. The experimental results showed that the hybrid ramie/glass/epoxy composite laminate showed highest impact strength compared to glass/epoxy and ramie/epoxy composite laminates. It was concluded that the good impact resistance property of glass fibre and good interfacial strength between ramie fibre and epoxy were the major reasons for the good impact strength of hybrid ramie/glass composite laminates. Bunea et al. [76] studied the LVI behaviour of hybrid composite laminates consisting of Kevlar and carbon fibre reinforced in stratified epoxy matrix (Fig. 6.8).

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Fig. 6.8 Hybrid composite laminates consisting of aramid and carbon fibre [76]

The stratified epoxy was made of using following ingredients with different combination such as potatoes starch, carbon black, aramid powder, glass whiskers, carbon whiskers and barium ferrite. It was found that the matrix properties strongly affected the impact resistance properties. Further, the impact resistance properties of the hybrid composite laminates were also affected by the aramid and carbon fibre content along with fibre orientation. It was concluded that the hybrid composite laminate with highest carbon content showed lowest energy absorption and the damage severity with increase in ply orientation in the form of delamination. In an investigation, Barouni et al. [77] compared the impact strength of flax/vinylester and hybrid flax/glass/vinylester composite laminates. Both the specimens were subjected to LVI with 25 and 50 J of impact energy. The LVI tests revealed that the flax/vinylester composite laminate absorbed more energy than the hybrid flax/glass/vinylester composite laminates. Further, the flax/vinylester composite laminates at lower impact energy (25 J) showed greater damage extent. It was concluded that the hybridisation of flax/vinylester composite laminates with the glass fibre plies improved the impact strength of flax/vinylester composite laminates. Wu et al. [78] tried to improve the impact strength of silk/epoxy composite laminate by introducing flax fibres. Both intra and interhybrid flax/silk/epoxy composite laminates were fabricated by vacuum assisted resin transfer moulding process. Experimental results revealed that the addition of flax fibre increased the peak impact load of silk/epoxy composite laminate, whereas the silk fibre deflected the crack propagation. Overall addition of flax fibre increased the impact strength of silk/epoxy composite laminate.

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6.3 Low Velocity Impact Impact Test on Nanomaterials Doped Other Fibre Reinforced Polymer Composites Laminates Liu et al. [79] tried to enhance the mechanical and thermal properties of nylon by doping multiwalled carbon nanotubes (MWCNTs). Fabrication of MWCNTs doped with nylon 6 composite was carried out by melt compounding method assisted by twin-screw mixer at 250 °C under 150 bar pressure then quenched in ice/water bath. Tensile tests were conducted according to ASTM standard D638 with a crosshead speed of 5 mm/min, and the hardness test was conducted with a three-sided pyramidal diamond indenter. At 2wt.% doping of MWCNTs elastic modulus was enhanced by 214% from 396 to 1242 MPa, meanwhile yield strength betterment was about 162%, i.e. from 18 to 47 MPa. However, the ductility of the 2 wt.% reinforced nylon 6 composite was less than that of nylon 6 composite. Hardness enhancement was achieved by about 83%. From differential scanning calorimeter (DSC) and X-ray diffraction (XRD) results, it was found that the crystallisation and melting behaviour of nylon 6 was greatly influenced by MWCNTs incorporation. From experimental results, it was proved that the enhancement was significantly due to the uniform dispersion and better interfacial bonding confirmed by transmission electron microscopy. Veedu et al. [80] MWCNTs were grown directly onto silicon carbide fibre fabric to overcome the poor out-of-plane characteristics of FRP laminates. Three-point flexural testing was carried out according to ASTM D790-0017 and found an enhancement in strength, modulus and toughness by 140%, 5% and 424%, respectively. From this method, uniform dispersion of MWCNTs achieved directly. From the hardness test, it was observed the enhancement in indentation modulus and hardness were 30% and 37%, respectively. It was concluded that enhancement in properties was due to the interlocking between fibre and the matrix material established by the MWCNTs. Salekeen et al. [81] studied the role of MWCNTs in enhancing the energy absorption capacity of the ceramic plate (Silica) under ballistic impact. Hand lay-up method assisted by vacuum bagging was used to fabricate the samples. Curing and postcuring of laminate were done at 120 °C for 4 h and 180 °C for 2 h, respectively. Gas gun and LS-DYNA were used to carry out the ballistic impact and numerical simulation. From results, it was observed that 0.5 wt.% of silica doping yielded an improvement of 67.48% and 28.7% in energy absorption and ballistic limit and, for 1 wt.% of MWCNTs doping yielded an improvement of 45.8% and 19.77% in energy absorption and ballistic limit when compared with the neat ceramic epoxy armour. In this research investigation, Morka et al. [82] carried out numerical analysis of aluminium doped with carbon nanotubes when subjected to ballistic impact. The composite made of two aluminium plates reinforced with CNTs. In this study, a steel core projectile was used. Numerical analysis showed the significance of CNTs as reinforcement material and its role in absorbing the impact energy. In this study, Laurenzi et al. [83] used the functionalised MWCNTs to improve the impact strength of Kevlar/epoxy composite laminates. The MWCNTs used were functionalised by

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acid treatment (mixture of sulphuric and nitric acids) process. Experimental results showed that the impact resistance property of Kevlar/epoxy composite laminates was improved after inclusion of functionalised MWCNTs. Alomari et al. [84] used the nanoclay material with Kevlar/epoxy composite laminate to increase the impact resistance. It was concluded that within a certain optimal weight percentage, addition of nanoclay reduced the delamination while beyond that optimal weight percentage addition the delamination was aggravated. Taraghi et al. [85] studied the impact resistance of MWCNTs doped Kevlar/epoxy composite laminates under LVI tested at 27 °C and -40 °C. The fabricated specimens were subjected to 40 J of impact energy. The doping of MWCNTs improved the impact resistance of Kevlar/epoxy composite laminate and reduced the damage size irrespective of the testing temperature. At room temperature LVI test, the 0.5 doped Kevlar/epoxy composite laminate showed better impact performance while at −40 °C LVI test, the 0.3 doped Kevlar/epoxy composite laminate showed better impact resistance. It was concluded that both impact resistance and damage formation both were strongly affected by the testing temperature and MWCNTs content. Ibrahim et al. [86] studied the enhancement in tensile, impact strength and water absorption capacity of sisal fibre (SF) laminates reinforced with nanoclay. Along with that the effect of maleic acid polypropylene treated (MAPP) sisal fibres and recycled polypropylene (rPP) were also investigated. Initially, fibre and recycled polypropylene were mixed, and then the nanoclay was mixed and poured into a mould and cured at 200 °C under 1800 psi pressure for 8 min. Fibre content was varied from 10, 20, 30 and 40 wt.% and nanoclay wt.% was varied from 1, 3 and 5 wt.%. Tensile and impact tests were carried out according to ASTM D638 and D256 on universal and notched Izod impact testing machines, respectively. Young’s modulus for maleic acid treated polypropylene laminate with 50% maleic acid polypropylene grafted (MAPP), 5% nanoclay and 40wt.% of SF/rPP showed maximum tensile strength, Young’s modulus and impact properties. These properties were enhanced by 32.80, 37.62 and 5.48% compared to untreated SF/rPP composite laminates, respectively. As the fibre content of the sisal fibre increased, corresponding water absorption also increased. It was maximum for 40 wt.% of untreated sisal fibre with polypropylene. Finally, it was concluded that the sisal fibres were hydrophilic, but the incorporation of nanoclay reduced the water absorption capacity, and the mechanical properties were also improved. Shao et al. [87] tried to improve the cryogenic behaviour of CNT/Epoxy composites. Fragmentation and universal testing machines were used to evaluate the mechanical properties. FESEM, TEM, TG, and Raman spectroscopy characterisations were used to characterise the purity of a few walled carbon nanotubes (FWCNTs). FWCNTs were dispersed into epoxy resin and poured into a dog bone shaped mould and cured at 50 °C for ten hours. Cryogenic treatment was carried out from room temperature to −196 °C at a decreasing rate of 2 °C/min and maintained for 12 h. The tensile test was carried out on a universal testing machine according to ISO 527 standards with 2 mm/min crosshead speed. The electrical resistance was measured by the fragmentation method using a two-probe instrument. From the tensile test, it was observed that the Young’s modulus of CNT/Epoxy composite was increased by 51%

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than the neat epoxy composite at cryogenic treatments. Meantime, 31% improvement in interfacial shear strength was observed in the case of CNT/Epoxy, which was higher than neat composite. Finally, it was concluded that strength enhancement was mainly due to the different shrinkage properties of CNT and epoxy. From Weibull distribution parameters, it was found that the strength of CNTs was not affected significantly, but the epoxy strength improved after cryogenic treatment. In an experimental investigation, Rahman et al. [88] used the nanoclay and graphene as the nanofiller material to Kevlar/epoxy composite laminate to improve the impact tolerance property. The fabricated specimens were conditioned to ultraviolet light and moisture for 4 and 8 days. Then the specimens were impacted at 32 J of impact energy using an impactor mass of 6.27 kg. It was concluded that the nanoclay filled Kevlar/epoxy composite laminates were more effective against impact loading than the graphene filled Kevlar/epoxy composite laminates. Further, both nanoclay and graphene addition enhanced the ultraviolet light and water absorption resistance. Yaghoobi et al. [89] tested the impact performance of MWCNTs doped polypropylene/kenaf/polypropylene grafted maleic acid biocomposite laminates. The MWCNTs doped content was 0.5, 1.0, 1.5 and 2 wt.% (Fig. 6.9). The melt blending process was used to fabricate the biocomposite laminates. The 0.5 and 1.0 doped biocomposite showed 31.11 J/m and 33.21 J/m of impact resistance respectively. It was concluded that the MWCNTs bridge linkage mechanism, fibre pull-out and fibre fracture were the major reason for increased impact performance of MWCNTs doped polypropylene/kenal/polypropylene grafted maleic acid biocomposite than the neat polypropylene/kenal/polypropylene grafted maleic acid biocomposite. In an experimental investigation, Nor et al. [90] examined the influence of CNTs doping on the impact strength of hybrid bamboo/glass/epoxy composite laminates. The damage was analysed by using ultrasonic wave propagation imaging. The high shear mixer was used to disperse the MWCNTs in epoxy matrix. From experimental results, it was found that the addition of MWCNTs into the epoxy matrix improved the impact strength of hybrid bamboo/glass/epoxy composite laminates. Further, the MWCNTs doped bamboo/glass/epoxy composite laminate absorbed less energy than the neat bamboo/glass/epoxy composite laminate resulting in low damage area.

6.4 Summary The impact performance of all these other fibre reinforced polymer composite laminates decreased as the impact energy or impact velocity increased. Further, the impact damage mechanism of these other fibre reinforced polymer composite laminate is also influenced by the impactor geometry. The natural fibres such as basalt, jute, flax, hemp, and kenaf show good impact properties comparable to glass fibre. However, the natural fibres are limited by their water absorption property and behaviour in moisture environment. Further, these limitation can be overcome by hybridisation of

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Fig. 6.9 Scanning electron micrographs of MWCNTs doped biocomposites, a, b 0.5wt.% doped biocomposite, c 1 wt.% doped biocomposite, d 1.5 wt.% doped biocomposite, e 2 wt.% biocomposite [89]

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two or more fibres together. The hybridisation can be interply or intraply. The hybridisation resulted in increased impact performance of other fibre reinforced polymer composite laminates. However, in hybridisation the ply material facing the impact influence the impact resistance properties. Furthermore, addition of nanofillers such as CNTs, graphene, nanoclay and other nanomaterials enhances the impact strength of these other fibre reinforced polymer composite laminates.

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73. Ricciardi MR, Papa I, Lopresto V et al (2019) Effect of hybridization on the impact properties of flax/basalt epoxy composites: influence of the stacking sequence. Compos Struct 214:476–485. https://doi.org/10.1016/j.compstruct.2019.01.087 74. Naga Kumar C, Prabhakar MN, Song J (2019) Effect of interface in hybrid reinforcement of flax/glass on mechanical properties of vinyl ester composites. Polym Test 73:404–411. https:// doi.org/10.1016/j.polymertesting.2018.12.005 75. Giridharan R (2019) Preparation and property evaluation of glass/ramie fibers reinforced epoxy hybrid composites. Compos Part B Eng 167:342–345. https://doi.org/10.1016/j.compositesb. 2018.12.049 76. Bunea M, Cîrciumaru A, Buciumeanu M et al (2019) Low velocity impact response of fabric reinforced hybrid composites with stratified filled epoxy matrix. Compos Sci Technol 169:242– 248. https://doi.org/10.1016/j.compscitech.2018.11.024 77. Barouni AK, Dhakal HN (2019) Damage investigation and assessment due to low-velocity impact on flax/glass hybrid composite plates. Compos Struct 226:111224. https://doi.org/10. 1016/j.compstruct.2019.111224 78. Wu C, Yang K, Gu Y et al (2019) Mechanical properties and impact performance of silk-epoxy resin composites modulated by flax fibres. Compos Part A Appl Sci Manuf 117:357–368. https://doi.org/10.1016/j.compositesa.2018.12.003 79. Liu PIY, Shen L et al (2004) Morphology and mechanical properties of multiwalled carbon nanotubes reinforced nylon-6 composites. Macromolecules 37:7214–7222. https://doi.org/10. 1021/ma049132t 80. Veedu VP, Cao A, Li X et al (2006) Multifunctional composites using reinforced laminae with carbon-nanotube forests. Nat Mater 5:457–462. https://doi.org/10.1038/nmat1650 81. Salekeen S, Khan MGK, Jeelani S (2011) High velocity impact properties characterization of nano-phased bi-layered body armor, pp 227–238. https://doi.org/10.1115/IMECE2011-63284 82. Morka A, Jackowska B (2011) Ballistic resistance of the carbon nanotube fibres reinforced composites—numerical study. Comput Mater Sci 50:1244–1249. https://doi.org/10.1016/j.com matsci.2010.03.046 83. Laurenzi S, Pastore R, Giannini G, Marchetti M (2013) Experimental study of impact resistance in multi-walled carbon nanotube reinforced epoxy. Compos Struct 99:62–68. https://doi.org/ 10.1016/j.compstruct.2012.12.002 84. Alomari A, Aldajah S, Hayek S et al (2013) Experimental investigation of the low speed impact characteristics of nanocomposites. Mater Des 47:836–841. https://doi.org/10.1016/j. matdes.2012.12.079 85. Taraghi I, Fereidoon A, Taheri-Behrooz F (2014) Low-velocity impact response of woven Kevlar/epoxy laminated composites reinforced with multi-walled carbon nanotubes at ambient and low temperatures. Mater Des 53:152–158. https://doi.org/10.1016/j.matdes.2013.06.051 86. Ibrahim ID, Jamiru T, Sadiku RE et al (2017) Dependency of the mechanical properties of sisal fiber reinforced recycled polypropylene composites on fiber surface treatment, fiber content and nanoclay. J Polym Environ 25:427–434. https://doi.org/10.1007/s10924-016-0823-2 87. Shao Y, Xu F, Liu W et al (2017) Influence of cryogenic treatment on mechanical and interfacial properties of carbon nanotube fiber/bisphenol-F epoxy composite. Compos Part B Eng 125:195–202. https://doi.org/10.1016/j.compositesb.2017.05.077 88. Rahman AS, Mathur V, Asmatulu R (2018) Effect of nanoclay and graphene inclusions on the low-velocity impact resistance of Kevlar-epoxy laminated composites. Compos Struct 187:481–488. https://doi.org/10.1016/j.compstruct.2017.12.054 89. Yaghoobi H, Fereidoon A (2019) Preparation and characterization of short kenaf fiber-based biocomposites reinforced with multi-walled carbon nanotubes. Compos Part B Eng 162:314– 322. https://doi.org/10.1016/j.compositesb.2018.11.015 90. Nor AFM, Sultan MTH, Jawaid M et al (2019) Analysing impact properties of CNT filled bamboo/glass hybrid nanocomposites through drop-weight impact testing, UWPI and compression-after-impact behaviour. Compos Part B Eng 168:166–174. https://doi.org/10. 1016/j.compositesb.2018.12.061

Chapter 7

Low Velocity Impact Study on Symmetric and Asymmetric Fibre Reinforced Polymer Composite Laminates

Abstract This chapter explores the influence of symmetric and asymmetric (unsymmetric) laminate stacking sequence design on the impact resistance property of FRP composite. Traditionally, the symmetric FRP laminate has the zero extensionbending coupling matrix, whereas the asymmetric FRP laminate has nonzero extension-bending coupling matrix. The nonzero extension-bending coupling matrix is not desired unless intentionally required. Because the presence of nonzero extension-bending coupling matrix induces warping or curling of the laminate after the fabrication. However, until recently, few authors have investigated the impact behaviour of asymmetric laminate with zero extension-bending stiffness matrix which showed very similar impact resistance properties of symmetric stacking sequences. This chapter explores the basic difference between symmetric and asymmetric laminate. Further, in this chapter, few works from the literature where symmetric and asymmetric laminates subjected to impact loading are covered. Keywords Symmetric laminate · Asymmetric laminate · Extension-bending coupling matrix · Low velocity impact

7.1 Introduction to Symmetric and Asymmetric Composite Laminates Any laminate consisting of plies oriented in different direction can be categorized as symmetric or asymmetric laminate. If the plies oriented in the first half of the laminate about the mid-pane is mirror image of the second half of the laminate or vice-versa, then the laminate design is called as symmetric laminate design. Further, if the plies oriented in the first half of the laminate about the mid-plane are not mirror image of the second half of the laminate or vice-versa, then the laminate design is called as asymmetric or unsymmetric laminate design. Furthermore, all the laminate stacking designs can be broadly categorized into symmetric and asymmetric laminate designs. For example, a balanced type laminate stacking design can be categorized under symmetric, asymmetric, and antisymmetric laminate design. If a laminate stacking sequence is asymmetric and the extension-bending coupling matrix is not zero, then © The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_7

221

222

7 Low Velocity Impact Study on Symmetric … Ɵn

ply

ɸn

Ɵn

Ɵn

ɸn

ɸn

ɸn

Ɵn

laminate

ɸn

Ɵn ɸn

(a)

Ɵn

ɸn

mid-plane

Ɵn

Ɵn

ɸn

asymmetric

symmetric

transformation of co-coordination according to the ply orientation y'

y

(b) x'

x x, y – original co-ordinates x’, y’ – Transformed co-ordinates

Fig. 7.1 Schematic illustration of a Symmetric and asymmetric laminate design, b transformation of coordination

the laminate may get curled or warped after curing. Furthermore, while designing a laminate stacking sequence consisting of different weaving ply architectures such as satin, then also care should be taken to avoid the warping or curling of the laminate after curing. Figure 7.1 represents the generalized asymmetric and symmetric laminate design. Where ‘8’ and ‘F’ indicates the ply orientation; ‘n’ indicates the ply number. Further, the generalized laminate designs showed in Fig. 7.1 have the same stiffness matrices. According to the classical laminate plate theory, the relationship between applied in-plane loads, bending moment, mid-plane strains and curvature is given by Eq. 7.1 

N M



 =

A:B B:D



ε0 κ

 (7.1)

where [A] = extension stiffness matrix in GPa-mm; [B] = extension-bending coupling stiffness matrix in GPa-mm2 ; [D] = bending stiffness matrix in GPa-mm3 ; N = in-plane applied loads in N; M = bending moment in N-mm; ε = mid-plane strain; κ = mid-plane curvature. According to the classical laminate plate theory, the bending coupling stiffness matrix is given by Eq. 7.2 

n  1 2

Q i j k h k − h 2k−1 Bi j = 2 k=1

(7.2)

7.1 Introduction to Symmetric and Asymmetric …

223

where [B] = extension-coupling matrix in GPa-mm2 ; Q i j = transformed reduced stiffness matrix of the kth ply; i, j = co-ordinate or cell location or cell position in the bending stiffness matrix; k = ply position about the mid-plane; h = thickness of the corresponding kth ply about the mid-plane; In symmetric laminate design, the extension-bending coupling matrix is zero which can be derived using classical laminate plate theory. Figure 7.2 shows the schematic illustration of a symmetric laminate design consisting of ‘n’ plies. The thickness and material properties of corresponding plies about the mid-plane in the laminate are same. Now, according to the classical laminate plate theory, the extension-bending coupling stiffness matrix is given as



1  1 ¯  2 Q i j θn h 1 − h 20 + · · · + Q¯ i j θ2 h 27 − h 26 2 2

1  2

1 ¯  2 2 + Q i j θ1 h 8 − h 7 + Q¯ i j θ1 h 9 − h 28 2 2



1 ¯  2 1  + Q i j θ2 h 10 − h 29 + · · · + Q¯ i j θn h 2n − h 2n−1 . 2 2

[B] =

h0

t n

t – ply thickness

h1 h2 h3

– ply orientation

h4 h5

mid-plane

3

h6

2

h7 h8

1 1

h9

2

h10

3

h11 h12

n

hn

Fig. 7.2 Schematic illustration of symmetric laminate consisting of ‘n’ plies

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7 Low Velocity Impact Study on Symmetric …

In the above equation, thickness values are incorporated



1  

1 ¯  Q i j θn ((n − 1)t)2 − (nt)2 + · · · + Q¯ i j θ2 t 2 − (2t)2 2 2

1  2

1 ¯  2 + Q i j θ1 0 − (t) + Q¯ i j θ1 (t) − 0 2 2



1  1  + Q¯ i j θ2 (2t)2 − (t)2 + · · · + Q¯ i j θn (nt)2 − ((n − 1)t)2 ) . 2 2

[B] =

Now, rearrange the terms 1   Q i j θn ((n − 1)t)2 − (nt)2 + (nt)2 − ((n − 1)t)2 2 1    1   

+ Q i j θ2 t 2 − (2t)2 + (2t)2 − t 2 + Q i j θ1 −(t)2 + (t)2 . 2 2

[B] =

Eliminating the negative and positive terms [B] = 0 Same procedure cannot be used for asymmetric laminate because the corresponding ply in the laminate about the mid-plane has the different orientation. Thus, in an asymmetric laminate design, the extension-bending coupling matrix will not become zero. In the subsequent section, few literature works are presented which explored the effect of symmetric and asymmetric laminate design influence on the low velocity impact performance of FRP composite laminates.

7.2 Low Velocity Impact Test on Symmetric and Asymmetric Composite Laminate Design Singh et al. [1] conducted experimental investigation on symmetric and asymmetric unidirectional GFRP composite laminate under impact loading. Impact test on GFRP composite laminate was conducted using a mild steel striker. The impact energy and impact velocity applied onto the FRP composite were 12 J and 50 m/s, respectively. To further enhance the delamination, a circular precracks of diameter 12 mm and thickness 0.15 mm was introduced at the mid-plane. The circular precrack was made of Teflon. The GF/epoxy composite laminates were fabricated using hydraulic press machine under a pressure of 1 MPa. The fabricated laminates were cured at 150 °C. The symmetric and asymmetric laminate configurations were [04 /904 /904 /04 ]S and [04 /904 /904 /04 //904 /04 /04 /904 ]. Experimental investigation was also extended for the asymmetric laminates bonded by epibond 1590 A/B and XNBR. The bonded symmetric and asymmetric laminates were fabricated in two stages. In the first phase, half of the laminate is prepared and cured. In the second stage, remaining half of

7.2 Low Velocity Impact Test on Symmetric …

225

Fig. 7.3 Failure at the unlike interfaces in GFRP laminates, a symmetric laminate design, b asymmetric laminate design, c asymmetric laminate bonded by epibond adhesive, d asymmetric laminate bonded by epibond and XNBR mixture adhesive [1]

the prepregs were placed over the cured half of the laminate and bonded using mentioned adhesive glues. The configuration of the bonded asymmetric configuration was [04 /904 /904 /04 //I//904 /04 /04 /904 ], where ‘I’ represents the adhesive interface which is the mixture of epibond 1590 A/B and XNBR. The XNBR portion in the mixture was 7.5%. Experimental showed that the delamination was present at unlike ply interfaces (Fig. 7.3). Further, use of XNBR in epibond reduced the damage area extension by 23%. The damage area extension due to the precrack was 1.83 cm2 in asymmetric laminate configuration, whereas it was 0.73 cm2 for symmetric laminate design. David-West et al. [2] studied the behaviour of balanced symmetric, balanced antisymmetric and balanced asymmetric stacking sequence under repeated impact. The balanced symmetric stacking sequence was [±453 / ± 603 ]S . Similarly, the balanced antisymmetric laminate design was [456 /606 /−606 /−456 ] and balanced asymmetric was [456 /606 /−456 /−606 ]. The mentioned sacking sequences were fabricated using vacuum autoclave method. Each laminate contained 24 layers of carbon fibre plies. The balanced asymmetric CFRP composite laminate showed warpage or out-ofplane deformation after curing (Fig. 7.4a). The warpage was mainly attributed by the presence of nonzero bending coupling stiffness matrix. The fabricated CFRP composite laminates were placed over a simply supported circular boundary condition. A hemispherical impactor with 12.1 mm nose diameter and mass of 30 kg was impacted onto the CFRP composite laminates. The symmetric stacking sequence showed good impact resisting properties than the asymmetric and antisymmetric CFRP composite laminates. It was concluded that the delayed failure and dissimilar

226

7 Low Velocity Impact Study on Symmetric … (a)

(c)

(b)

(d)

Fig. 7.4 a Warpage in the asymmetric CFRP laminate due to nonzero extension-bending coupling stiffness matrix, b–d energy absorbed vs number of repeated impacts for, b symmetric laminate, c asymmetric laminate, d antisymmetric laminate [2]

interfaces were the reason for improved impact resistance properties of symmetric CFRP composite laminates. Figure 7.4b–d shows the variation in absorbed energy for different stacking sequences of CFRP composite laminates at different number of repeated impacts. Rosa et al. [3] examined the impact and post impact residual strength properties of symmetric and asymmetric E-glass/basalt hybrid composite laminates. The hybrid composite laminates were fabricated using resin transfer moulding process. The stacking sequences used were [14B], [14 V], [3 V/8B/3 V], [3B/8 V/3B], and [1B/1 V/1B/1 V/1B/1 V/1B]S where V represents the E-glass fibre ply and B represents the basalt fibre ply. The fabricated laminates were subjected to 7.5, 15, and 22.5 J of impact energy using a hemispherical impactor with nose diameter of 12.7 mm. Experimental results showed that the hybrid E-glass/basalt composite laminate with basalt fibre ply as face sheet showed gradual failure. It was concluded that the symmetric stacking sequence consisting of glass fibre as core and basalt fibre plies as outer skin showed good impact resistance properties. Figure 7.5 represents damage formation in symmetric and asymmetric laminate design of E-glass/basalt hybrid composite laminates.

7.2 Low Velocity Impact Test on Symmetric …

227

Fig. 7.5 Infra-red thermography of impacted E-glass/basalt hybrid composite laminates [3]

Singh et al. [4] investigated the low velocity impact behaviour of quasi-isotropic symmetric and asymmetric GF/epoxy composite laminates. In this investigation, the asymmetric laminate stiffness characteristics were similar to that of symmetric laminate where the extension-bending coupling matrix was zero. The asymmetric and symmetric lay-up used were [(0°, 90°)/(+45°, −45°)/(+45°, −45 + )/(0°, 90°)//(+45°, −45°)/(0°, 90°)/(0°, 90°)/(+45°/−45°)] and [(0°, 90°)/(+45°, −45°)/(+45°, -45°)/(0°, 90°)//(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)], respectively. The required laminate designs were fabricated using vacuum bagging process assisted with wet hand lay-up technique. The fabricated laminates were impacted by hemispherical impactor of mass 5.23 kg. Three different impact velocities 3, 4, and 5 m/s were considered for testing. At impact velocities 3 and 4 m/s, the symmetric and asymmetric laminate showed no fibre breakage. However, the symmetric laminate absorbed the energy in the form of strain energy, whereas the asymmetric laminate absorbed the energy as delamination. Both symmetric and asymmetric laminates showed fibre breakages when impacted at 6 m/s, where the symmetric laminate design absorbed 59.7 J of impact energy while asymmetric laminate absorbed 45.98 J of impact energy (Fig. 7.6). Further, both damage height and width for asymmetric lay-up design were higher than the symmetric lay-up. Shin et al. [5] explored the influence of symmetric and asymmetric stacking sequences on the impact properties of CF/epoxy composite laminates. The symmetric stacking sequences considered for investigations were (0°/0°), (−15°/ + 15°), (−30°/30°), (−45°/ + 45°), and (−90°/ + 90°). The asymmetric stacking sequences considered were (0°/15°), (0°/30°), (0/45°), and (0°/90°). The unidirectional CF/epoxy laminates were fabricated using vacuum bagging method. The impact tests onto the symmetric and asymmetric CF/epoxy laminates were conducted using drop weight impact machine. ASTM D7136 was adopted to conduct the drop weight

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7 Low Velocity Impact Study on Symmetric …

Fig. 7.6 Energy–time response curve for symmetric and asymmetric lay-up impacted at a 3 m/s, b 4 m/s and c 5 m/s [4]

impact tests. In symmetric laminate design as the ply orientation increased, the strain of the laminate decreased except of symmetric (0°/0°) and (0°/90°) laminate. In asymmetric laminate design also, the strain in the laminate decreased as the ply orientation increased except (−90°/ + 90°) laminate. The symmetric CF/epoxy laminate was absorbed more impact energy than the asymmetric CF/epoxy composite laminate. Figure 7.7 shows the variation in damage formation and progression in symmetric and asymmetric CF/epoxy composite laminates. Hu et al. [6] conducted finite element simulation of low velocity impact on symmetric and asymmetric CF/epoxy bumper beams (Fig. 7.8). The symmetric and asymmetric lay-up used for the bumper were [45°/−45°]3S and [45°/−45°]6 , respectively. The numerical simulation results indicated that the CFRP car bumper did not show significant difference in impact resistance between symmetric and asymmetric stacking sequence. Ginzburg et al. [7] investigated the symmetric and asymmetric stacking sequences inspired from bio-inspired (mantis shrimp) helicoidal structure under low velocity impact. The CF/epoxy composite laminates were manufactured using vacuum bagging process. The symmetric stacking sequences were [0°/90°]8S , [+45°/−45°/0°/90°]4S , and [0°/12°/24°/36°/…/180°]S —with 6° increment. The asymmetric stacking sequences were [0°/6°/12°/18°/…/180°]—with 6° increment

7.2 Low Velocity Impact Test on Symmetric …

229

Fig. 7.7 Damage formation in CF/epoxy laminate under impact loading for different stacking sequences, a symmetric stacking sequence, b asymmetric stacking sequence where CD—crack direction; FD—fibre direction [5]

Fig. 7.8 Finite element analysis model of the CFRP car bumper beam [6]

and [0°/12°/24°/36°/…/360°]S —with 12° increment (Fig. 7.9). All the laminates were made of 32 plies. The low velocity impact tests on fabricated composite were conducted according to ASMT D7136 standard using drop weight impact machine. The impactor mass was 12.864 kg and the impactor had a hemispherical geometry with 10.25 mm nose radius. All the laminates were subjected to 40 J and 80 J impact energy. It was concluded that the both symmetric and asymmetric stacking sequences with 6° and 12° increment showed better impact resistance properties than other stacking sequences. Bozkurt et al. [8] investigated influence of symmetric and asymmetric lay-up design in basalt/aramid hybrid composite laminates under impact loading. Impact resistance of symmetric and asymmetric laminates was measured using Charpy

230

7 Low Velocity Impact Study on Symmetric …

[00/900]8S

[+450/-450/00/900]4S

[00/120/240/360/…./1800]S

[00/60/120/180/…./1800]

[00/120/240/360/…./3600]

Fig. 7.9 Symmetric and asymmetric stacking sequences used in the investigation [7]

impact test. The basalt fibre was unidirectional, and aramid fibre was twill weaved. The areal density of both basalt and aramid fibre was 300 and 170 g/m2 , respectively. The epoxy resin matrix material was used to fabricate the hybrid composite laminates. The basalt/aramid hybrid composite laminates were fabricated using vacuum assisted resin transfer moulding process. Total stacking sequence configurations prepared were 11, where each laminate had 10 plies in it. Out of these 11 configurations, five stacking sequences were symmetric, and four stacking sequences were asymmetric. Remaining two laminate configurations were made of completely aramid and basalt fibre. In asymmetric hybrid composite laminates, the 2, 4, 6 and 8th positioned basalt plies were replaced by aramid plies. Further, in symmetric hybrid composite laminates were made of eight aramid and two basalt fibre plies. The fabricated hybrid aramid/basalt composite laminates were subjected to Charpy impact test according to ISO 179/92 standard. Experimental test results showed that as the aramid plies were replaced by basalt fibre, the impact energy of the hybrid aramid/basalt composite laminates was increased. This increase in impact energy was due to the formation of delamination between basalt plies. The formation of delamination between basalt plies was mainly due to basalt fibre stiffness and weak interfacial bonding among basalt plies. Among asymmetric stacking sequences [A4 /B6 ], hybrid composite laminate (here, A—aramid ply; B—basalt ply) showed an impact energy of 4.94 J, where the basalt plies were facing the impact. For asymmetric stacking sequence [B4 /A6 ] hybrid laminate, the impact energy was 5.30 J where the aramid plies faces the impact loading. These two configurations showed highest impact energy among other asymmetric configurations. Further, the asymmetric [A8 /B2 ] hybrid composite laminate showed minimum impact energy. Among symmetric hybrid configurations, the

7.2 Low Velocity Impact Test on Symmetric …

231

[A1 /B1 /A3 ]S showed highest impact energy of 5.10 J, whereas the [A4 /B1 ]S showed minimum impact energy. It was concluded that the number of basalt fibres and position of the basalt fibre in the hybrid composite laminate greatly influenced the impact properties of hybrid FRP composite laminates. Singh et al. [9] investigated the symmetric and asymmetric GFRP laminates impact behaviour when doped with multiwalled carbon nanotubes (MWCNTs). The low velocity impact test on GF/epoxy/MWCNTs laminates was conducted at an impact energy of 94.14 J and impact velocity of 6 m/s. The MWCNTs doping wt.% used was 1, 2, 3, 4 and 5. Bidirectional plain woven 600 GSM glass fibre was reinforced in epoxy polymer matrix. The asymmetric and symmetric stacking sequences were [(0°, 90°)/(+45°, −45°)/(+45°, −45 + )/(0°, 90°)//(+45°, −45°)/(0°, 90°)/(0°, 90°)/(+45°/−45°)] and [(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)//(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)], respectively. The MWCNTs were mixed into the epoxy matrix using ultrasonication and then a curing agent was added to the MWCNTs/epoxy mixture. The hand lay-up techniques assisted by vacuum bagging process were used to fabricate the MWCNTs reinforced GFRP composite laminates. The low velocity impact tests were performed using a drop weight impact machine. The symmetric lay-up design without MWCNTs absorbed more impact energy than the asymmetric lay-up without MWCNTs. This behaviour was attributed by the presence lesser delamination prone interfaces in symmetric laminate. Further, the 2 wt.% doped symmetric GF/epoxy laminate lay-up design absorbed 59.71 J of energy and asymmetric lay-up design absorbed 44.6 J of energy. Whereas, at 5 wt.% doping, the symmetric lay-up absorbed 46.06 J of energy, while the asymmetric lay-up absorbed 38.23 J of energy. Figure 7.10 represents the energy–time curve for symmetric and asymmetric GFRP composite laminate with and without MWCNTs.

Fig. 7.10 Energy–time curves of GFRP laminates with and without MWCNTs a symmetric lay-up and b asymmetric lay-up [9]

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7 Low Velocity Impact Study on Symmetric …

Rawat et al. [10] investigated the behaviour of symmetric and asymmetric stacking sequence with and without precrack under low velocity impact loading for GF/epoxy composite laminates. The symmetric and asymmetric stacking sequences were [(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)//(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)] and [(0°, 90°)/(+45°, −45°)/(+45°, −45°)/(0°, 90°)//(+45°, −45°)/(0°, 90°)/(0°, 90°)/(+45°, −45°)], respectively. The GFRP laminates were impacted at 5 m/s. The GFRP laminates were simply supported on a circular boundary condition. The stress patterns obtained from the numerical simulation showed that the symmetrical laminate showed good impact resistance properties than the asymmetric laminate design with and without precrack the mid-plane. Further, the asymmetric laminate with precrack showed maximum penetration of the impactor along with maximum fibre damage. The symmetric laminate absorbed more energy and showed 24.06% less damage area than the asymmetric laminate with precrack at the mid-plane. Uddin et al. [11] studied the effect of ultraviolet light and moisture exposure on the low velocity impact resistance properties of CF/epoxy composite laminates. Three types of CF/epoxy composite laminates unidirectional, plain woven and non-crimp type of fibres were used. The CF/epoxy composite laminates were fabricated by vacuum bagging process. The stacking sequence of non-crimp CF/epoxy composite laminate was [45/0/−45/90] with 16 plies. Similarly, the sacking sequence of the woven CF/epoxy composite laminate was [(+45°/−45°)/(0°/90°)]. ASTM D7136 was followed to conduct the low velocity impact tests on fabricated CF/epoxy composite laminates. Experimental results showed that the impact resistance property of non-crimp symmetric and asymmetric laminates was not significantly affected by ultraviolet light and moisture exposure when compared to unidirectional and woven CF/epoxy composite laminates. Further, the non-crimp asymmetric CF/epoxy composite laminate absorbed 4.33 kJ/m of impact energy and the non-crimp symmetric CF/epoxy composite laminate absorbed 3.75 kJ/m of impact energy. Anjaneyulu et al. [12] evaluated the unidirectional symmetric and asymmetric GFRP composite laminate behaviour under impact loading. Hand lay-up process was used to fabricate the symmetric and asymmetric GFRP composite laminates. The symmetric and asymmetric composite laminates fabricated were [0°/90°/90°/0°] and [0°/90°/0°/90°], respectively. The Charpy impact tests were conducted according to ASTM D256-04 standards. Experimental results showed that both symmetric and asymmetric GFRP composite laminates did not show significant difference in the impact strength. Pinto et al. [13] fabricated the carbon and hemp hybrid composite laminates and subjected to low velocity impact. The areal density of hemp and carbon fibre was 160 and 200 g/m2 . The hand lay-up with vacuum bagging process was used to fabricate the carbon/hemp hybrid composite laminates. The symmetric carbon/hemp hybrid composite laminate designs were [CHC5 HC5 HC], where ‘C’ and ‘H’ represent the carbon and hemp fibre ply. Similarly, the asymmetric carbon/hemp hybrid composite laminate designs were [CH3 C11 ] and [C11 H3 C]. The fabricated carbon/hemp hybrid composite laminates were subjected to impact energies of 5, 10 and 20 J. The impact

7.2 Low Velocity Impact Test on Symmetric … 5J

10 J

5J

5J

15 J

(a)

233

10 J

(b)

15 J

10 J

15 J

(c)

Fig. 7.11 Damage formation in hybrid carbon/hemp composite laminates impacted at different energy levels, a symmetric hybrid carbon/hemp composite laminate ([CHC5 HC5 HC]), b asymmetric hybrid carbon/hemp composite laminate ([CH3 C11 ]), c asymmetric hybrid carbon/hemp composite laminate ([C11 H3 C]) [13]

mass was 2.66 kg. The delamination in symmetric carbon/hemp hybrid composite laminate increased with increase in impact energy. Further, the hemp fibre ply near the mid-plane enhances the impact resistance property of carbon/hemp hybrid composite laminates. Figure 7.11 shows the detailed difference in failure modes between symmetric and asymmetric carbon/hemp hybrid composite laminates. May et al. [14] studied the damage resistance behaviour of asymmetrical unidirectional CF/epoxy composite laminates subjected to impact loading. Four stacking sequences were selected which showed identical extension stiffness matrix along with zero extension-bending coupling matrix. The stacking sequences used were [453 /03 /−453 /903 ]S , [45/0/−45/90]3S , and [−452 /902 /452 /03 /45/90/−45/0/45/90/−45/0/45/90/−45/0/45/90/−45] [45/90/−45/0/45/90/−45/0/45/90/−45/0/45/90/−45/03 /−452 /902 /452 ]. It was concluded based on the finite element simulation results that both asymmetrical and symmetrical stacking sequences did not show much difference in the impact resistance properties.

7.3 Summary From the above literature, it is clear that the symmetric laminate design shows good impact resistance properties than the asymmetric laminate design. Because, the extension-bending coupling matrix for the symmetric laminate is zero which allows the laminate to be not warp or curl. Further, the number of mismatching interfaces also plays crucial role in asymmetric laminate design for absorbing impact energy. In hybrid symmetric and asymmetric laminate designs along with extension-bending coupling matrix, the type of fibre ply material which faces the impact also plays

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7 Low Velocity Impact Study on Symmetric …

important role in impact energy absorption. Furthermore, recent studies have shown that it is possible to fabricate an asymmetric laminate keeping extension-bending coupling matrix as zero. Moreover, very few studies have also reported insignificant influence of symmetric and asymmetric composite laminate design on the energy absorption capacity of FRP composite laminate.

References 1. Singh KK, Singh RK, Chandel PS, Kumar P (2008) An asymmetric FRP laminate with a circular precrack to determine impact-induced damage. Polym Compos 29:1378–1383. https://doi.org/ 10.1002/pc.20422 2. David-West OS, Nash DH, Banks WM (2008) An experimental study of damage accumulation in balanced CFRP laminates due to repeated impact. Compos Struct 83:247–258. https://doi. org/10.1016/j.compstruct.2007.04.015 3. De Rosa IM, Marra F, Pulci G, et al (2011) Post-impact mechanical characterisation of Eglass/basalt woven fabric interply hybrid laminates. Express Polym Lett 5. https://doi.org/10. 3144/expresspolymlett.2011.43 4. Singh KK, Singh NK, Jha R (2015) Analysis of symmetric and asymmetric glass fiber reinforced plastic laminates subjected to low-velocity impact. J Compos Mater 50:1853–1863. https://doi. org/10.1177/0021998315596594 5. Shin H-J, Kwa L-K, Lee M-S, Kim H-G (2015) Influence of laminated orientation on the mechanical and thermal characteristics of carbon-fiber reinforced plastics. Carbon Lett 16:241– 246. https://doi.org/10.5714/CL.2015.16.4.241 6. Hu Y, Liu C, Zhang J, et al (2015) Research on carbon fiber–reinforced plastic bumper beam subjected to low-velocity frontal impact. Adv Mech Eng 7:1687814015589458. https://doi. org/10.1177/1687814015589458 7. Ginzburg D, Pinto F, Iervolino O, Meo M (2017) Damage tolerance of bio-inspired helicoidal composites under low velocity impact. Compos Struct 161:187–203. https://doi.org/10.1016/ j.compstruct.2016.10.097 8. Bozkurt ÖY, Erkli˘g A, Bulut M (2018) Hybridization effects on charpy impact behavior of basalt/aramid fiber reinforced hybrid composite laminates. Polym Compos 39:467–475. https:// doi.org/10.1002/pc.23957 9. Kishore Singh N, Singh KK (2017) Impact analysis of multi-wall carbon nano tubes doped symmetric and asymmetric glass fiber reinforced polymer laminates. Mater Today Proc 4:8059– 8068. https://doi.org/10.1016/j.matpr.2017.07.145 10. Rawat P, Singh K, Singh N (2017) Numerical investigation of low-velocity impact in symmetric and asymmetric GFRP laminate with and without pre-crack. Adv Mater Proc 2:152–155. https:// doi.org/10.5185/amp.2017/304 11. Uddin MN, George JM, Patlolla VR, Asmatulu R (2019) Investigating the effects of UV light and moisture ingression on low-impact resistance of three different carbon fiber–reinforced composites. Adv Compos Hybrid Mater 2:701–710. https://doi.org/10.1007/s42114-019-001 17-4 12. Anjaneyulu J, Moizuddin M, Chandra Kumar P (2020) Evaluation of mechanical behaviour of glass fibre-epoxy composite laminates. Mater Today Proc 22:2899–2905. https://doi.org/10. 1016/j.matpr.2020.03.423 13. Pinto F, Boccarusso L, De Fazio D et al (2020) Carbon/hemp bio-hybrid composites: effects of the stacking sequence on flexural, damping and impact properties. Compos Struct 242:112148. https://doi.org/10.1016/j.compstruct.2020.112148

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14. May M, Arnold-Keifer S, Haase T (2020) Damage resistance of composite structures with unsymmetrical stacking sequence subjected to high velocity bird impact. Compos Part C Open Access 1:100002. https://doi.org/10.1016/j.jcomc.2020.100002

Chapter 8

Compression After Impact on Fibre Reinforced Polymer Composite Laminates

Abstract This chapter is dedicated to compression after impact because it helps in defining the damage tolerance and post-impact residual compressive properties of FRP laminates. Thus, in this chapter few experimental works conducted by various authors have been included. The chapter can be divided into three parts. The beginning of the chapter is related to the after impact compressive properties of carbon fibre reinforced composite laminates. This section also covers the hybridization of carbon fibre composite laminate followed by which doping of nanomaterials into the carbon fibre reinforced polymer composites is also covered. Further, the second section of this chapter emphasizes the post-impact compressive strength of glass fibre reinforced composite laminates and their related hybrids. This section also covers the glass fibre composite laminates doped with nanomaterials. The final section of this chapter deals with the residual compressive strength of natural and other synthetic fibre reinforced polymer composites. Keywords Compression after impact · Carbon fibre · Glass fibre · Flax fibre · Nanomaterials

8.1 Compression After Impact Test on Carbon Fibre Reinforced Polymer Composite Laminates Compression after impact test on controlled carbon fibre reinforced polymer composites laminates Hitchen et al. [1] studied the influence of stacking sequence on compressive strength after impact in CF/epoxy composite laminates. The composite laminates of CF/epoxy were fabricated using autoclave process. The stacking sequence adopted for experimental investigation was [(±45, 02 )2 ]S , [(±45)2 , 04 ]S , [(+45, 0, −45, 0)2 ]S , [(02 , ± 45)2 ]S , [04 , (±45)2 ]S and [(0, + 45, 0, −45)2 ]S . The impact damage was introduced onto the fabricated specimen using an instrumented falling weight impact machine. The specimens were struck with 7 J of impact energy where the impactor mass was 2 kg and impactor velocity was 2.5 m/s. The impact-induced damage was analysed © The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_8

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using an ultrasonic c-scan process. It was found that the energy responsible for delamination was increased as the 45° moved towards the laminate periphery. Further, the delamination initiation energy also increased when the number of dissimilar interface in the stacking was increased. It was concluded that the stacking sequence not only influenced the compressive strength of impact-induced CF/epoxy composite laminates but also affected the failure mechanisms. In a study, Sanchez-Saez et al. [2] investigated the residual compressive strength of thin CF/epoxy composite laminates after LVI. The thickness of the laminate was varied between 1.6 and 2.2 mm. The stacking sequences considered for experimentation were [0/90]3S (cross-ply) and [45/0/90]S (quasi-isotropic). The delamination propagation during compressive loading was majorly found to be perpendicular to loading direction. The woven laminate and quasi-isotropic stacking sequence showed good residual compressive strength than other stacking sequences. It was concluded that the woven reinforcement architecture controlled the shear cracking and delamination which resulted in improved residual compressive strength of woven CF/epoxy composite laminates. Aoki et al. [3] investigated the effect of water absorption and thermal exposure on the residual compressive strength after impact of CF/epoxy composite laminates. For experimentation, the CF/epoxy composite laminates were immersed in water maintained at 71 °C for 10, 000 h. Further, the CAI tests were conducted on the specimens which were exposed to −54, 22, 82, 149 and 177 °C temperatures. The stacking sequence of the CF/epoxy composite laminate was [45/0/−45/90]4S . The experimental results have shown that the residual compressive strength was higher for wet CF/epoxy composite laminate with smaller delamination area than the CF/epoxy specimens exposed to 22, 82 and 121 °C. However, at 149 and 177 °C the wet CF/epoxy composite laminates showed slight and drastic reduction residual compressive strength. It was concluded that the glass transition temperature was the major reason for the drastic reduction in residual compressive strength at 149 °C and 177 °C. Moreover, water absorption induced more instabilities due to the chemical changes that occur during the water absorption. Williams et al. [4] experimentally recovered more than 90% residual compressive strength of CF/epoxy composite laminate via self-healing process. The hollow glass fibres were used for self-healing process. The hollow glass fibres were filled with resin. The glass fibres were distributed at targeted interfaces within the CF/epoxy composite laminates. The reference CF/epoxy composite laminate was [(−45°/90°/45°/0°)]2S . The stacking sequence of CF/epoxy composite laminate with hallow glass fibre was [−45°/90°/45°/90°/HGF1 /−45°/90°/45°/HGF3 /0°/0°/HGF4 /45°/90°– 45°/HGF2 /0°/45°/90°/−45°]. Experimental results showed that the CF/epoxy composite laminates showed excellent recovery and recovered more than 90% of residual strength when compared to undamaged specimens. In a study, Gonzalez et al. [5] have investigated the influence of ply clustering on the residual compressive strength of CF/epoxy composite laminates. The stacking sequence used was [(45/0/−45/90)4 ]S , [(452 /02 /−452 /902 )2 ]S and

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[454 /04 /−454 /904 ]S . Based on experimental results, it was concluded that the residual compressive strength of CF/epoxy laminates was not affected by the ply clustering. Ghelli et al. [6] studied the residual compressive strength properties of CF/epoxy composite laminates. A quasi-isotropic stacking sequences [0/0/90/90/45/−45/45/−45]S and [90/90/0/0/−45/45/−45/45]S were considered for CAI testing. Two different geometries rectangular and circular were also considered during the investigation. The circular geometries were not affected by the stacking sequence. However, the stacking sequence in rectangular plate influenced both residual strength and the failure mechanism. It was concluded that the thin impactinduced CF/epoxy composite laminates were characterized by global buckling under compressive loading. Rivallant et al. [7] examined the low velocity impact and compression after impact properties of highly oriented and quasi-isotropic CFRP laminates. Low velocity impact was carried out on drop weight tower machine with 0–30 J of impact energy and 2 kg of impactor mass with 16-mm diameter of hemispherical shape in accordance with airbus test method AITM 1–0010 standards. Crack played a crucial role in the case of highly oriented laminate in deciding the final failure compared to quasiisotropic laminate. Finally, it was concluded that both types of laminate designs showed buckling failure. In a study, Vieille et al. [8] examined the effect of matrix toughness on the residual compressive strength of woven CF composite laminates. Two different types of matrix materials (thermoplastic and thermoset) were used. The thermoplastic matrix material was PEEK and PPS, while the thermoset matrix material was epoxy. The stacking sequence for CF/PEEK and CF/PPS composite laminate was [(0°, 90°)/(±45°)/(0°, 90°)/(±45°)/(0°, 90°)/(±45°)/(0°, 90°)], whereas the stacking for CF/epoxy was [(0°, 90°)/(±45°)/(±45°)/(0°, 90°)/(0°, 90°)/(±45°)/(±45°)/(0°, 90°)]. The fabricated CF composite laminates were subjected to 2, 6, 10.5, 17 and 25 J of impact energy. The 3D digital image correlation was adopted to capture the failure mechanism in impacted CF specimens under compressive loading. It was found that the residual compressive strength of CF/PEEK composite laminate was 10% and 40% greater than the CF/epoxy and CF/PPS, respectively. It was concluded that the matrix toughness was not the majorly influencing parameter concerning the residual compressive strength after the impact; rather, it was ductility which slows the crack propagation under compressive loading. Hence, the toughening of a thermoplastic matrix material does not result in significant improvement in residual compressive strength of FRP composite laminates compared to FRP composite laminates made of thermoset matrix material. Shi et al. [9] conducted compression after impact tests on FRP composite laminates made of recycled carbon fibre and epoxy polymer. For comparison, three types of CF/epoxy composite laminates were used: (i) virgin CFRP, (ii) recycled CFRP and (iii) treated recycled CFRP. The weaving architecture of the virgin and recycled CF was maintained same. The experimental results showed that as the impact energy increased the damage in the CF/epoxy also increased and correspondingly the residual compressive strength was reduced for all three types of CF/epoxy

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composite laminates. Further, the impact damage-induced CF/epoxy composite laminates showed shear failure under compressive loading for three types of CF/epoxy composite laminates. However, the shear failure angle was varied where the virgin CF/epoxy composite laminate showed 45°, the recycled CF/epoxy composite laminate showed 16°, and treated-recycled CF/epoxy composite laminate showed 24°. Finally, it was found that the surface cleaning improved the damage resistance of virgin CF/epoxy composite laminates by 80%, while it was 50% for recycled CF/epoxy composite laminates. Shufeng et al. [10] studied the residual compressive strength of scarf-repaired CF/epoxy composite laminates after being subjected to LVI. The CF/epoxy composite laminate stacking sequence was [45/02 /−45/90/45/02 /−45/0]S . The CF/epoxy composite laminates were subjected to 4.45 and 6.67 J of impact energy. The CAI test results showed that the residual compressive strength of scarf-repaired CF/epoxy composite laminates was decreased when impacted at adhesive edges. It was concluded that the scarf-repaired adhesive bond line was more sensitive to the impact and affected majorly to the residual compressive strength of scarf-repaired CF/epoxy composite laminates. Remacha et al. [11] developed a new device for testing compression after impact properties of thin FRP laminates, i.e. less than 4-mm thickness. Autoclave method was used to fabricate the CFRP laminates with [45°/0°/−45°/90°]ns ply orientation. Impact loading was applied using a drop weight tower machine according to ASTM D7136 with a hemispherical impactor of mass 5.6 kg and nose of 15.9 mm. Experimental and numerical simulation results showed good agreement. From the results, it was evident that there was no global buckling and laminates were completely failed due to compressive strength. Han et al. [12] studied the failure mechanism of carbon fibre composite laminates under compressive loading after being subjected to impact loading. Two types of carbon fibres (CCF300 and CCF800) and two types of resin systems (epoxy and bismaleimide) were used to fabricate the CF composite laminates. The fabricated CF composite laminates were subjected to impact loading using drop weight impact where the applied impact energy was 4.45 and 6.67 J/mm. Experimental results showed that the CF/epoxy composite laminates had greater residual compressive strength after the impact compared to CF/bismaleimide composite laminates. Moreover, the CCF300/epoxy and CCF800/epoxy showed no significant difference in residual compressive strength after impact, whereas the CCF300/bismaleimide and CCF800/bismaleimide composite laminates showed significant difference in residual compressive strength after impact. Further, the CCF300/bismaleimide composite laminate showed higher residual compressive strength compared to CCF800/bismaleimide composite laminate. In this investigation, Martins et al. [13] have examined the influence of throughthickness tufting angle on compression after impact. Carbon fibre reinforced with epoxy laminates were fabricated by vacuum-assisted resin transfer moulding process under −1 bar pressure at room temperature. It was cured at room temperature for 24 h and post-cured at 80 °C for 16 h. Tufting was done by using a six-axis robot at a various angle. Samples were bifurcated based on tufting density and angle.

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Transversal square pattern with 10*10 and 5*5 mm, density tufting pattern with +30, −30 angle were considered. Drop weight impact machine was used to introduce the impact at 25 and 60 J. Impact and compression after impact tests were carried out according to ASTM D7136 and D7137 standards. Through thickness and the tufting effect were analysed by ultrasonic c-scan and digital image correlation. The optical micrograph was used to analyse crack propagation. From the analysis, it was concluded that tufting enhanced the damage resistance property and 5*5 mm transversal square pattern reduced the damaged area by four times compared to the samples without tufting. Caminero et al. [14] examined the carbon fibre ply thickness and plied sequencing under compressive loading after subjecting to impact loading. Low velocity impact was carried on a drop weight impact machine according to ASTM D7136 standard, and compression was done on universal electromechanical testing machine. Ultrasonic c-scan and destructive methods were used to examine the damage mechanism involved. From experimental results, it was observed that as the impact energy increased, its corresponding compressive strength decreased. Further, it observed the delamination as primary failure mode along with ply cracking and translaminar failure. It was also found that angle ply exhibited better compression after impact property than cross-ply laminate. Ply clustering influenced the damage tolerance property greatly because it delayed the buckling phenomenon in the composite. Sun et al. [15] investigated the compression after impact and compression after indentation behaviour of unidirectional CF/epoxy composite laminates for two similar quasi-isotropic stacking sequences. The stacking sequences considered were [452 /902 /02 /−452 ]2S and [45/90/0/−45]4S . Full field 3D digital image correlation was used to capture the failure mechanism of CF/epoxy composite laminates during compression testing after being subjected to low velocity impact and indentation test. Experimental results showed that the stacking sequence [45/90/0/−45]4S showed better impact tolerance compared to [452 /902 /02 /−452 ]2S . Compression after impact test on hybrid carbon fibre reinforced polymer composite laminates Naik et al. [16] conducted the CAI experimental investigation on glass–carbon/epoxy composite laminates for different stacking sequences. The weaving pattern for glass and carbon fibre was plain and twill, respectively. The stacking sequences adopted for experimentations were [G4 /C4 ]S , [G/C]4S , [C4 /G4 ]S . For reference, complete CF/epoxy and GF/epoxy composite laminates were also fabricated. The post-impact compressive strength tests were conducted on NASA 1142 test fixture. Experimental results showed that the [C4 /G4 ]S hybrid composite laminate showed highest residual compressive strength among all stacking sequences. It was concluded that the glass/carbon hybrid composites were less notch sensitive. Further, the [C4 /G4 ]S laminate showed least sensitivity to notch and hence showed highest after impact compressive strength. Gonzalez et al. [17] investigated the residual compressive strength of interply hybridization of woven carbon, woven glass and unidirectional carbon tapes after being subjected to impact loading. The ply configurations selected for

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CAI tests were [C-G], [C-UC] and [G-UC] where ‘C’, ‘G’ and ‘UC’ represent the carbon, glass and unidirectional carbon fibre plies. The stacking sequences selected for [C-G] configuration were [0G /60G /-60G /(0C /45C )2 ]S , [(0C /45C )/0G /60G /-60G ]S and [0G /(45C /0C )2 /60G /-60G ]S . The stacking sequences selected for [C-UC] configuration were [0UC /60UC /-60UC /(0C /45C )2 ]S and [(0C /45C )2 /0UC /60UC /-60UC ]S . Similarly, the stacking sequences for [G-UC] configuration were [0UC /45UC /90UC /-45UC /(0G /45G )2 /0G ]S , [(0G /45G )2 /0G /45UC /0UC /45UC /90UC ]S , [0UC /45UC /0G /45G /90UC /-45UC /0G /45G /0G ]S and [0G /45G /0G /90UC /45UC /0G /45G /0UC /45UC ]S . These laminates were subjected to impact loading where the applied impact energies were 30, 40 and 50 J. Experimental results showed that the presence of glass fibre plies around the mid-plane significantly enhanced the residual compressive strength of the hybrid composite laminates. Among all the composite laminate made of only glass fibre plies showed highest residual compressive strength, while the composite laminate made of complete carbon fibre showed poor residual compressive strength among all the stacking sequences. Rhead et al. [18] conducted compression after impact tests on carbon/glass hybrid composite laminates. The complete CF composite laminate stacking sequence was [(±45C )4 /(0C /90C )4 ]S . The hybrid CF/GF composite laminate stacking sequences were [±45C /45C /(±45G )/−45C / ± 45C /(0C /0G )4 ]S , [(±45G )2 /(±45C )2 /((0C )3 /0G )2 ]S , [(±45C )4 /(0C /0G )4 ]S , [(±45C )4 /(0G /0C )4 ]S and [45C /−45C /0C /0G ]4S . It was found that the carbon/glass hybrid composite laminates showed better structural integrity of 51% and 41% when subject to 12 J and 18 J of impact energy, respectively, than the full GF/epoxy and CF/epoxy composite laminates. It was concluded that presence of glass fibre at outer side of the laminate resists the delamination formation under impact loading and sublaminate buckling which in turn enhances the residual compressive strength of the carbon/glass hybrid composite laminates. Liu et al. [19] studied CAI behaviour of CF/epoxy hybrid composite laminates. The CF hybrid composite laminates were designed using unidirectional CF and 5HS woven CF. The stacking sequences used were L1[5HS/−45/+45/90/0/−45/+45/90/+45/−45/90/+45/−45/0/90/+45/−45/5HS] and L2-[5HS/0/0/+45/−45/0/0/0/−45/+45/0/0/0/+45/−45/0/0/5HS]. The applied impact energies were 10, 17 and 25 J. For same impact energy, the L2 stacking sequence showed better residual compressive strength than the L1 stacking sequence. It was concluded that the presence of woven CF ply as outer layer had insignificant effect on residual strength of CF/hybrid composite after being subjected to LVI. Figure 8.1 shows different failure modes occurred in the hybrid composite laminates during the CAI tests. Ismail et al. [20] investigated the residual compressive strength of MWCNTsdoped flax/carbon and flax/glass hybrid nanocomposites. The hybrid laminates were fabricated from wet lay-up process. The MWCNTs were dispersed into the epoxy polymer using shear mixing and mechanical stirring process. The MWCNTs content doped into the epoxy was 1.0 wt.%. The stacking sequences of the flax/carbon and flax/glass hybrid composite laminates were [F/C], [C/F], [F/G] and [G/F] where F, C and G represent the flax, carbon and glass fibre ply, respectively. Further, the first letter in the bracket faces the impact loading. Each laminate consists of 6 plies with

8.1 Compression After Impact Test on Carbon …

(a)

(c)

[5HS/-45/+45/90/0/-45/+45/90/+45/45/90/+45/-45/0/90/+45/-45/5HS]

(b)

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[5HS/0/0/+45/-45/0/0/0/45/+45/0/0/0/+45/-45/0/0/5HS]

(d)

Fig. 8.1 Failure modes observed on the edges of the CF hybrid composite laminates subjected to compression loading after being impacted. a, b Failure modes observed in different stacking sequences subjected to different impact energy levels. c Failure morphology on the front and back surface observed on CF hybrid composite laminates subjected to 25 J of impact energy. d Marking representing different failures observed in (a, b) [19]

[(0°/90°)]3 ply orientation. The fabricated composite laminates were subjected to impact loading. The applied impact energies were 5 and 20 J. Experimental results showed that the [F/G] hybrid composite with glass fibre ply facing impact showed better residual compressive strength than the [F/C] hybrid composite laminates. Farhood et al. [21] examined the residual compressive strength of FRP pipes after being subjected to impact. The pipes were made of glass–carbon hybrid composite laminates. The pipes were fabricated from filament winding process. The stacking sequence or orientation of the fibres in the pipe was [±55/902 /±55/902 ]. The fabricated specimens were subjected to 100 J impact and then applied compressive load to investigate the residual compressive strength. The presence of carbon and glass fibres in the stacking sequence were [C8 ], [G8 ], [C4 G4 ], [G4 C4 ], [G4 C4 ], [C2 G4 C2 ], [C6 G2 ], [G2 C6 ] and [C2 G2 C4 ] where ‘C’ indicates the carbon fibre, ‘G’ indicates the glass fibre and subscript numerical value represents the number of plies. The glass–carbon hybrid composites with carbon were placed at periphery or alternatively exhibited the better residual compressive strength after the impact. Further, the [C4 G4 ] and [C6 G2 ] hybrid configurations showed poor residual compressive strength after the impact. Compression after impact test on nanomaterial-doped carbon fibre reinforced polymer composite laminates Yokozeki et al. [22] investigated the effect of cup-stacked carbon nanotubes (CSCNTs) doping in CFRP composite laminates on the compressive strength

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after subjecting to low velocity impact. The aspect ratio of CSCNTs maintained was 10 to achieve homogeneous dispersion. The dispersion of CSCNTs into the epoxy was done by planetary mixing followed by wet mixing. Autoclave method was used to fabricate the CSCNTs-doped CF/epoxy composite laminates. The stacking sequence of the fabricated CSCNTs-doped CF/epoxy nanocomposite was [0°/90°/45°/−45°]3S . The Suppliers of Advanced Composite Materials Association (SACMA) standard was followed to conduct the CAI tests. The impactor nose diameter was 15.9 mm. Impact energy applied to create the damage was 6.67 J/mm. Experimental results showed that the CSCNTs doping had no significant improvement on the CAI strength of CF/epoxy composite laminates. In a study, Iqbal et al. [23] have used the nanoclay as nanoreinforcement to enhance the residual compressive strength of unidirectional CF/epoxy composite laminates. The nanoclay content doped was 3 and 5 wt.%. The nanoclay was dispersed into the epoxy matrix using high-speed shear mixing and ultrasonication processes. The stacking sequence of the laminate was [0/90]3S . The nanoclay-doped CF/epoxy laminates were fabricated by vacuum hot press machine. The fabricated composite laminates were tested according to the ASTM D7137 standard. The CF/epoxy composite laminate with nanoclay 3 wt.% showed higher residual compressive strength followed by 5 wt.% doped CF/epoxy composite laminate. The enhancement in residual compressive strength after doping nanoclay was due to increase in shear strength and stiffness of the epoxy matrix. The enhancement in shear strength in turn enhanced the buckling of fibre under compressive loading. Further, the nanoclay also enhanced the interfacial bonding between CF fibre and epoxy matrix. Finally, it was concluded that the failure mode of CF/epoxy composite laminates was changed from brittle buckling mode to more ductile nature along with multi-layer delamination mode after the addition of nanoclay. Among all the doping weight percentages of nanoclay, the 3 wt.% was found to be the optimal content. Kostopoulos et al. [24] used the carbon nanotubes to enhance the compressive strength of CF/epoxy composite laminates after subjecting to low velocity impact. The MWCNTs were dispersed into the epoxy matrix using high shear mixing process. The MWCNTs content used was 0.5 wt.% by epoxy weight. The stacking sequence of CF/epoxy composite laminate was a quasi-isotropic [0, +45, 90, −45]2S . The MWCNTs reinforced CF/epoxy composite laminates were fabricated using a vacuum bagging autoclave process. The ASTM D5628-07 standard was followed for impact testing followed by which ASTM D7137 M-07 was followed to conduct CAI testing. The impactor was made of aluminium with mass of 3.01 kg. The geometry of the impactor was hemispherical with nose diameter of 20 mm. Experimental results showed that addition of MWCNTs improved the residual compressive strength of CF/epoxy composite laminates after subjecting to the low energy impacts for all impact energies. It was concluded that the MWCNTs pull-out and breakage were the main reasons for the improvement in the residual compressive strength of the MWCNTs-doped CF/epoxy composite laminates (Fig. 8.2). Ashrafi et al. [25] used single-walled carbon nanotubes (SWCNTs) to improve the residual compressive strength of the CF/epoxy nanocomposites. For effective dispersion, the SWCNTs were reduced. The content of SWCNTs used 0.1 wt.% (by

8.1 Compression After Impact Test on Carbon … Fig. 8.2 SEM image showing MWCNTs in CF/epoxy composite laminate [24]

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MWCNTs

epoxy weight). Experimental results showed that the use of 0.1 wt,% of SWCNTs improved the residual compressive strength of CF/epoxy composite laminates by 3.5% compared to CF/epoxy composite laminate without SWCNTs. It was concluded that the high aspect ratio of SWCNTs and enhancement in the interaction between epoxy and carbon fibre after the addition of SWCNTs were the major reason the improvement in residual compressive strength of the CF/epoxy composite laminates. Mannov et al. [26] tried to improve the residual compressive strength of CF/epoxy composite laminates by doping thermally reduced graphene oxide to the epoxy. The three-roll mill was used to disperse the graphene oxide into the epoxy matrix. The graphene oxide weight percentage used was 0.3 and 0.5. The stacking sequence used was [0/90]S . The instrumented drop weight impact machine was used to induce the impact damage. ASTM D7137 was followed to conduct the CAI tests. The 0.3 wt.% graphene oxide-doped CF/epoxy showed highest residual compressive strength than the neat CF/epoxy and 0.5 wt.% doped graphene oxide CF/epoxy composite laminates. Siegfried et al. [27] studied the use of CNTs as nanoreinforcement on the residual compressive strength of CF/epoxy composite laminates. Three different types of epoxy systems were used. First is the liquid bisphenol A epoxy with high concentration of CNTs. The other two epoxy systems had different storage times where one epoxy had three-month storage time while the other epoxy system has threeyear storage time. The other type of CF/epoxy composite laminate was made with amino group functionalized CNTs. The stacking sequence used in the study was [(+45°/−45°)/(0°/90°)]4S . Experimental results of functionalized CNTs-doped CF/epoxy nanocomposites and pristine CNTs-doped CF/epoxy nanocomposites have shown degraded residual compressive strength compared to CF/epoxy composite laminate. However, the aged CNTs-doped CF/epoxy composite laminates have shown 1% improvement in the residual compressive strength compared to CF/epoxy composite laminate which was insignificant.

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Fig. 8.3 Schematic representation of thin film polyetherketone cardo/dichloromethane in CF/bismaleimide composite laminates [28]

Xu et al. [28] inserted the MWCNTs-doped polyetherketone cardo/dichloromethane thin thermoplastic films to improve the residual compressive strength of CF/bismaleimide composite laminates (Fig. 8.3). The ball milling process was used to disperse the MWCNTs into the polyetherketone cardo/dichloromethane. The MWCNTs contents used were 5 wt.% and 10 wt.%. The stacking sequence of the laminate was [45°/0°/−45/90°]4S . The laminates were fabricated using autoclave process assisted by vacuum bagging. The residual compressive strength of the CF/bismaleimide composite laminates with 10 wt.% interleaved MWCNTsdoped polyetherketone cardo/dichloromethane thin films was approximately improved by 33% compared to the CF/bismaleimide composite laminates without MWCNTs-doped polyetherketone cardo/dichloromethane thin films. Nikfar et al. [29] added the silica nanoparticle and rubber particles to the CF/epoxy composite laminates. The stacking sequence of CF/epoxy composite laminate was [45°/0°/−45°/90°]S . Experimental results showed that the residual compressive strength of CF/epoxy composite laminate was improved by 30% after the addition of nanosilica particles and rubber particles. Addition of silica nanoparticles and rubber particle enhanced the fracture toughness of the epoxy material which in turn enhanced the residual compressive strength of the CF/epoxy composite laminate. However, addition of nanosilica particles and rubber particles did not influence the residual compressive strength of CF/epoxy composite laminates impacted at high velocity because the residual compressive strength of FRP composite depends on the ply lay-up and fibre architecture. Pantelakis et al. [30] studied the effect of MWCNTs and glycidyl polyhedral oligomeric silsesquioxanes (GPOS) doping in CF/epoxy composite laminate on the residual compressive strength after subjecting to impact. Experimental results showed that the addition of MWCNTs and GPOS deteriorated the residual compressive strength of CF/epoxy composite laminates. It was concluded that agglomeration of MWCNTs and GPOS was the major possible reason for the deterioration of residual compressive strength in CF/epoxy composite laminates (Fig. 8.4).

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Fig. 8.4 Formation of agglomeration of MWCNTs and GPOS in CF/epoxy composite laminate [30]

Yourdkhani et al. [31] studied the compression after impact behaviour of CFRP laminate fabricated by vacuum-assisted resin film infusion method. Two lay-up designs were considered where carbon fibre and CNTs-modified resin films were grouped separately, and in another case, modified resin films and carbon fibres were stacked alternately. Quasi-isotropic stacking sequence [(45/−45)/(0/90)]2S was considered for the investigation. Grouped laminate showed poor compression after impact strength than interleaved laminate because in interleaved laminate filtration effect was less compared to grouped laminate due to the less movement of CNTs. Nevertheless, void content was more in interleaved laminate than grouped laminate design due to restricted movement of trapped air inside the laminate. Finally, it was concluded that resin film infusion method improved the compression after impact strength by improving the dispersion of CNTs in laminate and the improvement was more in case of interleaved design than grouped lay-up when compared to neat CFRP laminate. Hsieh et al. [32] introduced the nanocarbon aerogels (Fig. 8.5) to improve the residual compressive strength of CF/epoxy composite laminates. The nanocarbon aerogels were functionalized with amine groups. The ultrasonication and three roll mixing processes were used to disperse the nanocarbon aerogels into the epoxy matrix. The stacking sequence used was [0/90/90/0]5S . Experimental results showed that the inclusion of nanocarbon aerogels into the CF/epoxy composite laminates improved the residual compressive strength at 0.3 wt.% of nanocarbon aerogel.

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Fig. 8.5 Transmission electron microscope of nanocarbon aerogels [32]

8.2 Compression After Impact Test on Glass Fibre Reinforced Polymer Composite Laminates Compression after impact test on controlled glass fibre reinforced polymer composite laminates Davies et al. [33] studied compressive strength of thick woven GF/polyester composite laminates after being subjected to impact loading. A flat nosed impactor was used to induce the damage into the GF/polyester composite laminates. Experimental results showed that the residual compressive strength of GF/polyester reduced rapidly as the applied impact energy associated increased. Kim et al. [34] carefully examined the failure of unidirectional cross-ply GF/epoxy and woven GF/epoxy composite laminates under compression after impact test. Investigation was also extended to study the influence of silane coupling agent influence on damage and delamination resistance under low velocity impact. Experimental results showed that the residual compressive strength strongly influenced by the damage area formed after the impact rather than damage propagation during the compressive loading. The silane coupling agent concentration also strongly influenced the damage resistance property of GF/epoxy composite laminate. Finally, it was concluded that the silane concentration strongly affected the interfacial bonding strength between glass fibre and epoxy which in turn affected the damage resistance as well as residual compressive strength properties of GF/epoxy composite laminates. The woven GF/epoxy composite laminate showed smaller damage area than the unidirectional cross-ply GF/epoxy composite laminate. Further, the woven GF/epoxy composite laminate also showed higher residual compressive strength than the unidirectional GF/epoxy composite laminate. Khondker et al. [35] examined the compression after impact testing behaviour of weft knitted GF/vinyl ester reinforced polymer composite laminate. Three different

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Fig. 8.6 Different weft knitting styles. a Plain. b Rib. c Milano [35]

kinds of weft knitting plain, Milano, and rib knit were used for the experimentation (Fig. 8.6). The GF/vinyl ester composite laminates were fabricated using resin transform moulding process. Among all the GF/vinyl ester composite laminates, the plain knitted laminate showed better retention of compressive strength than the Milano and rib knitted GF/vinyl ester composite laminates where both laminates lost their original compressive strength by 22% after subjecting to impact loading. It was concluded that loop length and stitch densities greatly influenced the damage tolerant properties of GF/vinyl ester composite. Ibekwe et al. [36] conducted the CAI tests on GF/epoxy composite laminates at low temperatures. The testing temperatures were 0, −10, −20, 10 and 20 °C. The stacking sequences considered for the CAI tests were [08 ]S and [(0/90)4 ]S . For lowtemperature CAI tests, an environmental controlled chamber was used. The liquefied nitrogen was used to achieve the required room temperature. Experimental results of CAI tests showed that the reduction in after impact compressive strength for specimens tested at -10 °C was less than the specimen tested at −20 °C because at −20 °C the epoxy matrix became more brittle than the matrix tested at -10 °C. It was also found that though the cross-ply laminate showed better impact resistance than the unidirectional laminate, however, the unidirectional GF/epoxy composite laminate showed better after impact compressive strength than the cross-ply laminate for all testing temperatures. Yin et al. [37] tried to regain the after impact compressive strength of GF/epoxy composite laminates by doping a healing agent microcapsule. These microcapsules were filled with epoxy and latent hardener CuBr2 (2-MeIm)4 . The microcapsules were dispersed into the epoxy by stirring and ultrasonication. Then, the laminate with stacking sequence [45/0/−45/−90]S was fabricated by compression moulding

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process assisted by hand lay-up. For self-healing, the GF/epoxy composite laminate consisting of microcapsule was heated to 150 °C for 0.5 h. During self-healing process, the microcapsule fills the matrix cracks formed when the laminate was subjected to impact loading (Fig. 8.7). The healing efficiency of the GF/epoxy composite laminates was dependent on the content of microcapsule in the laminate.

Fig. 8.7 T-scan ultrasonic images of GF/epoxy composite laminates consisting of impact-induced damages subjected to different impact energy levels: a–c impact energy level 1.5 J, d–f 2.5 J and g–i 3.5 J; a, d, g impact-induced GF/epoxy composite laminate without microcapsule, b, e, h selfhealed GF/epoxy composite laminate under no external pressure and c, f, i self-healed GF/epoxy composite laminate under external pressure [37]

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Further, the self-healing efficiency was decreased with increase in impact energy because as the impact energy increased the dominant failure mode shifts from matrix cracks to fibre failure. It was concluded that the addition of self-healing microcapsules not only repaired the matrix cracks formed due to the impact loading but also retained after impact compressive strength. Aktas et al. [38] investigated the after impact compressive strength properties of GF/epoxy composite laminates which were impacted at room temperature, 40, 60, 80 and 100 °C. Further, the tests were also conducted for two different impact damages oriented along vertical and horizontal directions. The hot press method assisted by hand lay-up method was used to fabricate the GF/epoxy composite laminates consisting of two stacking sequences [0°/90°/0°/90°]S and [0°/90°/45°/−45°]S . Experimental results showed that the after impact compressive strength properties of GF/epoxy composite laminates were decreased with increase in impact energy and impact test temperature. The specimen impact tested at 100 °C and subjected to 70 J of impact energy showed the maximum reduction in after impact compressive strength for both the stacking sequence and impact damage orientations. However, among both the stacking sequences the [0°/90°/45°/−45°]S laminate configuration showed good after impact compressive properties. Berketis et al. [39] investigated the residual compressive strength of GF/polyester composite laminates exposed to water immersion for long periods. The water temperature was kept at 65 °C. The GF/polyester laminates were immersed in water and kept for 30 months. The lay-ups selected for experimental investigation were (±45°, 0°), (±45°)2 , (0°, ±45°). The composite laminates were subjected to two impact energies 5 and 10 J. The chemical integrity of the GF/polyester composite laminates was reduced after exposure to the water. Experimental results showed that the residual compressive strength of water aged GF/polyester laminates was reduced. Further, a local plateau for residual compressive strength was observed for GF/polyester specimens after 24 and 30 months of water immersion. Yan et al. [40] subjected the GF/vinyl ester composite laminates to compression after impact tests to study the failure behaviour. The GF/vinyl ester composite laminates were fabricated using vacuum-assisted resin transfer moulding process. The CAI tests were conducted according to ASMT D7137 standard. The drop weight impact machine was used to induce impact damage into the fabricated specimens. The experimental results showed that the delamination was the critical failure mode which majorly affected the buckling strength of the GF/vinyl ester composite laminates. Further, the constituent cracks were observed at the back face of the GF/vinyl ester composite laminates and propagated along the through-thickness direction towards the impact face (Fig. 8.8). The propagation of the crack resulted in failure of the sample in compressive loading. Aktas et al. [41] conducted CAI tests on 1D plain weave, 2D double plain-weaved and 3F triple plain-weaved GF/epoxy composite laminates (Fig. 8.9). The laminates were fabricated by hot press process assisted by hand lay-up. The shape of the impactor was hemispherical with 12.7-mm of nose diameter mass of 5.027 kg. The specimens were simply supported onto a circular boundary condition consisting of circular hole of diameter 76.2 mm. The CAI tests were conducted according to

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(a)

(b)

Fig. 8.8 a Impact-induced damage at front and back face of the GF/vinyl ester composite laminates. b Stress–strain curve of compression after impact test on GF/vinyl ester composite laminate [40]

Fig. 8.9 Glass fibre woven architectures used in the experimentation. a Plain woven 1D, b double woven 2D and c triple woven 3D [41]

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ASTM D7137 standard. The residual compressive strength of GF/epoxy composite laminates was decreased with an increase in impact energy irrespective of woven fabric layers. Further, among all the 1D plain architecture showed highest after impact compressive strength than others for impact energy rage 5–22.5 J. However, for impact energy range 22.5–52.5 J the 3D woven fibre architecture showed highest after impact compressive strength. Andrew et al. [42] examined the residual compressive strength of thin repaired GFRP composite laminates after being subjected to repeated impact loading. The GF/epoxy composite laminates were fabricated by compression moulding process assisted by hand lay-up. Initially, the fabricated composite laminates were subjected to repeated impact (one, three and five times) loadings. Then, the damaged area was removed and filled it with short Kevlar/epoxy mixture. The CAI experimental results showed that for specimen impacted once showed abrupt reduction in after impact compressive strength for both repaired and unrepaired specimens. However, the abrupt reduction observed for specimen impact three and five times was much less compared to the specimen impact one time. Moreover, the repaired GF/epoxy composite laminates showed less delamination progression due to the cut-outs and Kevlar/epoxy filling. Hart et al. [43] studied the post-impact compressive strength of 2D and 3D woven GF/epoxy composite laminates. Figure 8.10 shows the schematic illustration of 2D and 3D woven GF fabric. Experimental results showed that the post-impact compressive strength of 2D woven GF/epoxy composite laminate was insignificantly sensitive to impact energy. Furthermore, the after impact compressive strength of 3D woven GF/epoxy composite laminate showed even much less sensitivity to the impact energy compared to 2D woven GF/epoxy composite laminate. Based on experimental observation, it was concluded that the 3D woven GF/epoxy composite laminate showed greater after impact compressive strength than the 2D woven GF/epoxy composite laminate. It was concluded the presence of through-thickness Z-tow in 3D woven GF/epoxy composite laminate restricted the delamination growth which enhanced the post-impact compressive strength of 3D woven GF/epoxy composite laminate. Balikoglu et al. [44] studied the impact and compression after impact properties of pinewood and ash wood sandwich composite with PVC foam core and glass/epoxy vinyl ester laminate as face sheets. Sandwich composites were fabricated using

Fig. 8.10 Schematic illustration of 2D and 3D woven architecture of GF fabric [43]

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vacuum-assisted resin transfer moulding method. The sandwich design consists of PVC as foam core with and without groove sandwiched between ash wood, which in turn sandwiched between glass/epoxy vinyl ester face sheets, and the pinewood sandwich composite followed the same design. The impact was made by hemispherical and conical impactor under 30 and 60 J of impact energy, and CAI was conducted according to ASTM C364/C364M standard under column compression loading condition. From CAI results, it was concluded that ash wood sandwich structure with grooved core performed better than pinewood and ash wood composites without grooved PVC foam core. Saravankumar et al. [45] added the milled glass fibre as a filler material to the GF/epoxy composite laminate to improve the after impact compressive strength. The filler content used was 5% by epoxy weight. The fabricated laminates were impacted at 10, 15, 20, 25 and 30 J of impact energies. The milled glass fibre was dispersed into the epoxy matrix using mechanical stirring and ultrasonication process. A combination of hand lay-up and compression moulding processes was used to fabricate the milled glass fibre reinforced GF/epoxy composite laminates. The ply stacking sequence used for experiment was [0/90]4S . The experimental results showed that the residual compressive strength of milled glass fibre reinforced GF/epoxy composite laminates was improved by 19%, 25%, 30%, 28% and 26% for specimens impacted at 10 J, 15 J, 20 J, 25 J and 30 J, respectively. It was concluded that the addition of milled glass fibre delayed the buckling which in turn increased the residual compressive strength of GF/epoxy composite laminates. Gliszczynski et al. [46] investigated the residual compressive strength of GFRP composite laminates after being subjected to corner impact. Figure 8.11 illustrates the experimental fixture used for corner impact. The stacking sequences of the GFRP composite laminates adopted to the investigation were (quasi-isotropic), (quasiorthotropic) and (angle ply). The corner impacts were conducted at three different styles: (i) perpendicular to the flange, (ii) at an angle 45° to the web and (iii) perpendicular to the web. The residual compressive strength of GFRP composite laminate was decreased due to the corner impact. Among all the GFRP specimen corner impacted perpendicular to the web showed highest reduction in the compressive strength of the GFRP composite laminate.

Fig. 8.11 Schematic illustration of fixtures used for corner impact [46]

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(a)

(b)

(c)

Fig. 8.12 Glass fibre architecture used: a Plain woven. b Milano. c Rib knotted [47]

Compression after impact test on hybrid glass fibre reinforced polymer composite laminates In a study, Aktas et al. [47] used woven and knitted glass fibre to study residual compressive strength after subjecting to LVI. The knotted fabric architecture was rib and Milano (Fig. 8.12). The stacking sequences used were [2D2 R2 ]S , [2D2 M2 ], [R2 2D2 ]S , [M2 2D2 ]S , [M2 R2 ]S and [R2 M2 ]S where ‘2D’, ‘M’ and ‘R’ represent the woven, Milano and rib knitted glass fibre ply. Experimental results showed that the residual compressive strength after impact reduced as the impact energy increased for all the hybrid composite laminates. Among all the stacking sequences, the glass hybrid composite laminate with woven fabric as face sheet ([2D2 R2 ] and [2D2 M2 ]S ) showed highest residual compressive strength compared to other stacking sequences. Further, among knitted hybrid composite laminates the Milano fabric as outer layer performed better under CAI test than the glass hybrid composites with rib knitted plies as outer layers. Andrew and Ramesh [48] explored the post-impact compressive strength of unidirectional glass/basalt hybrid composite laminates. The glass/basalt hybrid composite laminates were fabricated by compression moulding process assisted by hand layup. The hybrid composite laminates were made of 16 plies. The stacking sequences used for CAI testing were [G], [B] and [G3 /B3 /G4 /B3 /G3 ]. The hybrid composite laminates were subjected to impact loading at 2.17 J of impact energy. The [G] composite laminate showed good post-impact compressive strength than the [B] and

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[G3 /B3 /G4 /B3 /G3 ] composite laminate. Further, the after impact compressive properties of the hybrid composite laminate were not superior than the completely glass and basalt composite laminates. Nor et al. [49] incorporated the MWCNTs into bamboo/glass fibre hybrid composite laminates and then conducted CAI tests. Initially, the MWCNTs were dispersed into the epoxy using high shear mixer; then, the short bamboo fibres were added to the MWCNTs/epoxy mixture. The hand lay-up technique was used to fabricate the bamboo/glass hybrid composites. Experimental results showed that addition of MWCNTs (0.5 wt.%) into the bamboo/glass hybrid composite increased the residual compressive strength by 23.67%. The delaminations were formed perpendicular to the loading direction during CAI test. It was concluded that the addition of MWCNTs into the bamboo/glass hybrid composites increased the transverse strength which in turn improved the residual compressive strength of bamboo/glass hybrid composites. Further, MWCNTs also acted as stress distributor which also helped in improving the residual compressive strength of bamboo/glass hybrid composites. Safri et al. [50] examined the post-impact compressive strength of sugar palm/glass hybrid composite laminates. The sugar palm fibres were treated with benzoyl. The matrix material used was epoxy. The hand lay-up process was used to fabricate the sugar palm/glass hybrid composite laminates. The combination of epoxy, sugar palm fibre and glass fibre is as follows: (i) EP/GF, (ii) EP/TSPF, (iii) EP/70TSPF/30GF, (iv) EP/50TSPF/50GF, (v) EP/30TSPF/70GF, (vi) EP/UTSPF, (vii) EP/70UTSPF/30GF, (viii) EP/50UTSPF/50GF and (ix) EP/30UTSPF/70GF. Here, ‘EP’, ‘TSPF’, ‘GF’ and ‘UTSPF’ represent the epoxy, benzoyl-treated sugar palm fibre, glass fibre and untreated sugar palm fibre, respectively. Experimental results showed that the post-impact compressive strength was decreased for all the laminates as the impact energy was increased. However, it was found that the laminate EP/30TSPF/70GF showed poor after impact compressive strength than other laminates. It was concluded that treatment of sugar palm fibre with benzoyl along with addition of glass fibre enhanced the post-impact compressive strength of sugar palm composite. Further, it was also concluded that the specimen which sustained more impact damage showed poor compressive strength. Compression after impact test on nanomaterial-doped glass fibre reinforced polymer composite laminates Lal et al. [51] doped the nanosilica to epoxy matrix to enhance the post-impact compressive strength of GF/epoxy composite laminate. Doping of nanosilica improves the toughness of the epoxy matrix. The nanosilica content used was 0.25, 0.5, 0.75 and 1 wt.%. The mechanical stirring was used to disperse the nanosilica into epoxy matrix. Experimental results showed that the 1 wt.% nanosilica-doped GF/epoxy composite laminate showed highest after impact compressive strength. It was concluded that the addition of nanosilica contained matrix micro-cracks and also enhanced the load transferring efficiency by improving the fibre-matrix interfacial strength. These were the major reasons which attributed to the enhancement in post impact compressive strength of GF/epoxy composite laminate.

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8.3 Compression After Impact Test on Other Fibre Reinforced Polymer Composite Laminates Compression after impact test on controlled other fibre reinforced polymer composite laminates Liang et al. [52] studied the post-impact residual compressive strength of two quasiisotropic lay-ups of flax/epoxy composite laminates. The quasi-isotropic composite laminate stacking sequences were [0/90/45/−45]2S and [90/0/−45/45]2S . The hot pressing assisted by hand lay-up method was used to fabricate the flax/epoxy composite laminate. For flax/epoxy composite laminate a 15–30% reduction in compressive strength was observed when impacted at 10 J of impact energy. Lopez et al. [53] studied the after impact compressive strength of fully biodegradable flax/polylactic acid (PLA) composite laminate. The flax/PLA composite laminates were fabricated using compression moulding process. The impact tests were conducted with two different impactors of diameter 12.7 and 20 mm. The residual compressive strength after impact was reduced as the impact energy increased. Further, the residual compressive strength of flax/PLA composite laminate was also reduced as the impactor diameter increased. It was found that the major failure mode observed in flax/PLA specimens during CAI tests was fibre failure while no delaminations were absent. Thorsson et al. [54] studied the effect of edge impact on residual compressive strength of FRP composite laminates. For experimentation, two different impact angles 0° and 45° with respect to the edge of the laminate were selected. A specially designed fixture was used to conduct the edge impact on FRP composite laminates (Fig. 8.13). The stacking sequence of the FRP laminate was [45/90/−45/0/0/−45/0/45/0/45/−45/0/45/0/0/−45/90/45]S . The digital image correlation was used to analyse the strain and displacement fields during the edge impact and CAI tests. During CAI test, the edge-impacted region was neither on loading side nor on towards fixed side. It was placed along the sideways so that the damage propagation can take place. It was found that the post-edge impact compressive strength of FRP composites was not sensitive to the damage induced when CAI tests conducted according to Boeing experimental standards or ASTM D7137 standards. However, the post-edge impact compressive strength of FRP composite laminate was sensitive to the damage inducted when CAI tests were conducted according to ASTM D6641 standard. The FRP composite laminates edge-impacted at 0° showed highest reduction in compressive strength than the specimens edge-impacted at 45°. Habibi et al. [55] examined the post-impact compressive strength of the unidirectional flax/epoxy composite laminates. The conical impactor nose angle was varied with 15, 30, 45 and 60°. Figure 8.14 shows the variation in effect of impact energy and impact angle on the residual compressive strength of the flax/epoxy composite laminates. It was found that the post-impact compressive strength of the flax/epoxy composite laminate did not show significant reduction for impact energy less than 3 J. Further, the residual compressive strength was strongly dependent on the impact angle. The buckling failure of the flax composite laminates was seriously affected

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a

b

Fig. 8.13 Specially designed fixture to conduct the edge impact on FRP composite at different impact angles: a 0° and b 45° [54]

Fig. 8.14 Effect of impact angle and impact energy on normalized residual compressive strength of flax/epoxy composite laminate [55]

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by the impact energy. Testing condition with impact angle 60° and impact energy 5 J showed approximately 40% reduction in the residual compressive strength of flax/epoxy composite laminate. Habibi et al. [56] studied the flax/epoxy composite laminates after impact compressive behaviour. The flax/epoxy composite laminates showed a 5% decrease in compressive strength when impacted less than 3 J. More linear reduction in normalized residual compressive strength was observed for impact angle 15 and 60°. The digital image correlation showed that the damage propagation under compressive loading depends on the damage extent induced in the laminate during the impact loading. In a study, Prasath et al. [57] studied the effect of basalt fibre powder addition into the flax/polyester composites on the post-impact compressive strength. The basalt fibre powder content used was 5% to 30% with an increment. A total of ten layers of flax fibre were used. The compression moulding process was used to fabricate the flax/polyester composite laminates consisting of basalt fibre powder. The optimal value of basalt powder addition in flax/polyester composite laminate was 10% at which good bond strength between flax layers was achieved. Li et al. [58] explored the after impact compressive behaviour of flax/epoxy composite laminates. Different stacking sequences [0/90]6S (cross-ply), [0/45/90/−45]3S (quasi-isotropic) and [0/30/60/90/−30/−60]2S (multi-directional) were investigated during the experimentation. Among all the stacking sequences, the cross-ply laminate design showed highest residual compressive strength, while the multi-directional laminate design showed least residual compressive strength. It was found that the residual compressive strength of flax/epoxy composite laminates was lower than the GF/epoxy composite laminates. The failure modes under CAI loading observed in flax/epoxy composite laminates were localized cracking and delamination, while the GF/epoxy showed global delamination. Compression after impact test on other hybrid fibre reinforced polymer composite laminates Dehkordi et al. [59] studied the after impact residual compressive strength of intraply hybrid basalt/nylon composite laminates. The stacking sequence of the hybrid composite laminate was [(+45, −45)/(0, 90)]S irrespective of fibre content of basalt and nylon. Upon CAI test, the composite laminate with 100% basalt fibre laminate showed highest failure strength, while 100% nylon fibre showed least failure strength. Further, it was observed that as the nylon content in the hybrid composite laminate was increased more ductile failure nature of laminate was observed when failed. It was concluded that among all the hybrid laminate with 50% basalt and 50% nylon showed highest buckling strength. Prasath et al. [60] investigated the post-impact compressive strength of flax/basalt hybrid composite laminates. The matrix material used was polyester. The compression moulding process was used to fabricate the flax/basalt hybrid composites. Each hybrid composite laminates had 10 layers. The stacking sequences used [F10 ], [F5 /B5 ], [F2 /B2 /F2 /B2 /F2 ], [B2 /F2 /B2 /F2 /B2 ], [F3 /B2 /F3 /B2 ], [F4 /B2 /F2 /B2 ], [B3 /F2 /B3 /F2 ], [B4 /F2 /B2 /F2 ] and [B10 ], where ‘F’ and ‘B’ indicate the flax and basalt

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fibre plies, respectively. Further, the numerical value represents the corresponding ply number. Experimental results showed that among all the stacking sequences the hybrid composite laminate with basalt and flax fibre as alternate stacking sequences with more flax fibre plies showed highest post-compressive strength. Amir et al. [61] studied the post-impact compressive strength of gamma irradiated Kevlar/oil palm empty fruit bunch (OPEF) hybrid composites. The stacking design of the hybrid composite was [K/OP/K], where ‘K’ and ‘OP’ indicate the Kevlar and oil palm empty fruit bunch fibre plies. The hybrid composite laminates were fabricated using hand lay-up process. The hybrid composite laminates subjected to 10 and 35 J of impact energy showed no penetration while impacted at 40 J of impact energy perforated the gamma irradiated hybrid composite laminates. The hybridization and irradiation by gamma rays enhanced the post-impact compressive strength. It was concluded that the hybrid composite laminates could be used for medical application and military industries where FRP materials are exposed to gamma radiations. Hassan et al. [62] examined the post-impact compressive strength of banana fibre/epoxy composite laminates sandwiched by different synthetic fibre laminates (carbon, glass, Kevlar). The banana fibre/epoxy composite with Kevlar, glass and carbon fibre composite as skin layers showed 22%, 63% and 136% residual stresses, respectively. It was evident that the banana fibre/epoxy composite laminate sandwiched by carbon/epoxy composite laminate showed highest residual compressive strength because the CF/epoxy composite showed high resistance to shear crack formations. Tabrej et al. [63] studied the after impact compressive strength of kenaf/jute hybrid composite laminates. These natural fibres were treated with sodium hydroxide solution. The stacking sequence of hybrid composite laminate was kenaf/jute/kenaf. The experimental results revealed that the residual compressive strength after impact of hybrid composite laminates were decreased as the impact energy was increased.

8.4 Summary From the literature, it is clear that the post-impact compressive strength of FRP composite laminates reduces with increase in impact energy. The extent of reduction in the compressive strength of FRP composite materials depends on the type of fibre, matrix material toughness, stacking sequence of the laminate. Use of two different fibres enhances the post-impact compressive strength of the FRP composite laminates. Further, the extent of reduction depends on the outer layer fibre material. It was also found that the toughness of the matrix material can be improved by filler material such as nanomaterials (MWCNTs). Further, addition of nanomaterial also reduced the extent of reduction in the post-impact compressive strength. The failure under the compression loading always started from the impact site and propagated towards the edges. The failure modes for natural fibre consist of less delamination than the synthetic fibre.

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52. Liang S, Guillaumat L, Gning P-B (2015) Impact behaviour of flax/epoxy composite plates. Int J Impact Eng 80:56–64. https://doi.org/10.1016/j.ijimpeng.2015.01.006 53. Rubio-López A, Artero-Guerrero J, Pernas-Sánchez J, Santiuste C (2017) Compression after impact of flax/PLA biodegradable composites. Polym Test 59:127–135. https://doi.org/10. 1016/j.polymertesting.2017.01.025 54. Thorsson SI, Sringeri SP, Waas AM et al (2018) Experimental investigation of composite laminates subject to low-velocity edge-on impact and compression after impact. Compos Struct 186:335–346. https://doi.org/10.1016/j.compstruct.2017.11.084 55. Habibi M, Selmi S, Laperrière L et al (2019) Experimental investigation on the response of unidirectional flax fiber composites to low-velocity impact with after-impact tensile and compressive strength measurement. Compos Part B Eng 171:246–253. https://doi.org/10.1016/ j.compositesb.2019.05.011 56. Habibi M, Selmi S, Laperrière L et al (2020) Post-impact compression behavior of natural flax fiber composites. J Nat Fibers 17:1683–1691. https://doi.org/10.1080/15440478.2019.158 8829 57. Prasath KA, Arumugaprabu V, Amuthakkannan P et al (2020) Low velocity impact, compression after impact and morphological studies on flax fiber reinforced with basalt powder filled composites. Mater Res Express 7:15317. https://doi.org/10.1088/2053-1591/ab611e 58. Li Y, Zhong J, Fu K (2020) Low-velocity impact and compression-after-impact behaviour of flax fibre-reinforced composites. Acta Mech Solida Sin 33:431–448. https://doi.org/10.1007/ s10338-019-00158-8 59. Tehrani Dehkordi M, Nosraty H, Shokrieh MM et al (2013) The influence of hybridization on impact damage behavior and residual compression strength of intraply basalt/nylon hybrid composites. Mater Des 43:283–290. https://doi.org/10.1016/j.matdes.2012.07.005 60. Prasath KA, Amuthakkannan P, Arumugaprabu V, Manikandan V (2019) Low velocity impact and compression after impact damage responses on flax/basalt fiber hybrid composites. Mater Res Express 6:115308. https://doi.org/10.1088/2053-1591/ab43f4 61. Muhammad Amir SM, Hameed Sultan MT, Md Shah AU et al (2020) Low velocity impact and compression after impact properties on gamma irradiated kevlar/oil palm empty fruit bunch hybrid composites. Coatings 10:646. https://doi.org/10.3390/coatings10070646 62. Hassan MZ, Sapuan SM, Rasid ZA, et al (2020) Impact damage resistance and post-impact tolerance of optimum banana-pseudo-stem-fiber-reinforced epoxy sandwich structures. Appl. Sci. 10:684. https://doi.org/10.3390/app10020684 63. Tabrej K, Sultan MTH, Jawaid M, et al (2021) Low velocity impact, ultrasonic c-scan and compression after impact of Kenaf/jute hybrid composites. In: Impact studies of composite materials. Springer, pp 73–85. https://doi.org/10.1007/978-981-16-1323-4_6

Chapter 9

Numerical Analysis of Low Velocity Impact and Compression After Impact on Fibre Reinforced Composite Laminates

Abstract Finite element analysis enables better understanding of damage behaviour of FRP composite laminate. Because the FRP composite laminate is highly anisotropic and heterogeneous, thus the damage mechanism associated becomes highly complex. In this chapter, the numerical analysis of FRP composite laminate under low velocity impact and compression after impact from the literature are presented. Controlled, hybrid and nanomaterial-doped fibre reinforced composite materials are covered. The first section of the chapter covers the finite element analysis of FRP composite laminate under LVI, and the second part of the chapter covers the finite element analysis of FRP composite laminate under compression after impact. Keywords FRP composites · Low velocity impact · Compression after impact · Finite element analysis

9.1 Numerical Analysis of Low Velocity Impact on Fibre Reinforced Polymer Composite Laminates Numerical analysis of low velocity impact on carbon fibre reinforced polymer composite laminates Hosseinzadeh et al. [1] studied the impact damage behaviour of thin GF/epoxy, thick GF/epoxy, CF/epoxy and hybrid CF/GF/epoxy composite laminates using finite element analysis. All the composite plates were meshed with shell elements. The Chang-Chang progressive failure criteria was used to model the damage in composite plates. The finite element results showed that the CF/epoxy and hybrid CF/GF/epoxy composite laminates showed no impact damage under 30 J of impact energy. Further, the damage shape showed by the finite element analysis was not in agreement with the experimental observation. Minak et al. [2] developed finite element code using Fortran 90 language and studied the influence of boundary condition and diameter on the impact response of CF/epoxy composite laminate. Two-dimensional flat shell elements obeying Hughes and Tezduyar’s formulation were used for numerical analysis. Numerical simulation © The Author(s), under exclusive license to Springer Nature Singapore Pte Ltd. 2022 K. K. Singh and M. Shinde, Impact Behavior of Fibre Reinforced Laminates, Materials Horizons: From Nature to Nanomaterials, https://doi.org/10.1007/978-981-16-9439-4_9

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results showed that both boundary condition and diameter significantly influenced the impact performance of CF/epoxy composite laminate. Lopes et al. [3] accounted the physical progressive damage of fibre, matrix and interfaces between plies to conduct the finite element analysis of CF/epoxy composite laminates under LVI. The matrix and fibre damages were modelled using 3D continuum damage mechanics, and the delamination between the plies was modelled using cohesive damage model. The damage initiation was modelled based on the LaRC04 failure criteria, while the damage evolution was modelled based on the exponential damage evolution. The impactor was modelled as a rigid body. The composite plates were meshed with C3D8R solid elements. The contact between the impactor and laminate was modelled using general contact algorithm. The coulomb friction model was implemented to simulate the friction between all the contacting surfaces. It was concluded that the simulation result reliability was influenced by impact energy and the number of angle mismatching interfaces. Wang et al. [4] used the Hashin and Yeh failure criteria to understand the damage mechanism in CF/epoxy composite laminates under LVI. The CF/epoxy composite plate was modelled using C3D8R 8-node linear reduced integration solid elements. The impactor was modelled as a rigid body and meshed 4-node linear tetrahedron continuum elements. The progressive damage model was adopted for the finite element analysis. The Hashin failure detected the fibre and matrix failures in both tension and compression. The finite element results showed an average of 3.3% divergence from the experimental results. Bouvet et al. [5] conducted finite element simulation to capture the permanent indentation in CF/epoxy composite laminates under LVI using plastic like model. The delamination was modelled using interface elements worked based on the fracture mechanics. The intralaminar damages such as matrix and fibre failure were modelled using interface elements and degradation in the element volume, respectively. The numerical simulation results showed good corroboration with the experimental observations. Xu et al. [6] presented the shear stress continuity between angle mismatching interfaces in CF/epoxy composite laminates under LVI using 3D finite element analysis. Stress-based failure criteria was used to model the laminate failure. The cohesive initial damage was modelled based on the empirical formula. The impactor and CF/epoxy composite plates were meshed using C3D8R and COH3D8 elements. General contact algorithm was used to define all the contacts during the finite element analysis. The delamination size and shape predicted using the finite element approach showed good agreement with the experimental results. However, the percentage variation in the predicted damage area with the experimental observations was affected by the stacking sequence of the composite laminate. It was concluded that the error or variation observed in the predicted value was attributed to the non-consideration of delamination damage initiation and propagation. Hongkarnjanakul et al. [7] investigated the effect of ply position switching in the reference laminate stacking sequence [02 , 452 , 902 , −452 ] on the impact performance using finite element analysis. The fibre compressive failure law was implemented during the finite element analysis to capture the fibre failure. The mesh

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Fig. 9.1 Comparison of damage formation between experimental and finite element analysis for CF/epoxy composite laminate under LVI [7]

shape was square for 90° and parallelogram for 45° plies. Different shapes were used so that each node coincides with the adjacent plies. The interfaces were meshed using COH3D8 zero thickness cohesive elements, whereas the plies were modelled using C3D8 solid elements. The fibre damage initiation was modelled based on the strain failure. The fibre damage propagation was modelled based on strain softening. The matrix cracking was modelled using the Hashin failure criteria. The delamination was modelled using power law criteria based on the mixed mode. Figure 9.1 demonstrates the comparison of failure modes observed between experimental and numerical analyses. The predicted finite element results were in good correlation with the experimental results. Maio et al. [8] used the progressive damage model to simulate the LVI on CF/epoxy composite laminate. The progressive damage model considers three fibre and three matrix damages. The three fibre failures include (i) fibre failure under uniaxial tension and transverse shear, (ii) fibre damage under compression and (iii) damage under transverse compression loading. The three matrix failures include (i) matrix failure under transverse compression, (ii) matrix failure under tensile and shear stresses and (iii) matrix failure due to through-thickness tensile and shear stresses. The predicted results showed good agreement with experimental observations considered from the previous literature. Shi et al. [9] developed the finite element model to define the intra- and interlaminar crack initiation and evolution based on stress and fracture mechanics, respectively. The nonlinear shear behaviour was also considered in the numerical simulation. The delamination was modelled using interface cohesive elements, whereas the cohesive elements were used along the fibre direction to model the splitting and matrix crack. The Puck failure model was used to predict the damage, and the Hashin failure criteria were used to define the damage initiation in fibre and damage initiation in matrix under tension. The delamination between the ply interfaces was modelled based on the stress-based failure criteria. The matrix nonlinear shear response was accounted by using the semi-empirical equations [19, 20]. The 8-node brick elements with zero thickness cohesive elements were used to mesh the composite laminates. General contact algorithm was used to define the contact between laminate and

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the impactor. The adopted finite element analysis captured the force, and energy data which were in good agreement with the experimental observations. Further, it was concluded that the matrix cracks and damage evolution were successfully and accurately captured by the finite element method showing good correlation with the non-destructive observations. Zhang et al. [10] used the quasi-static load model along with surface-based cohesive contact to simulate the LVI damage formation in CF/epoxy composite laminates (Fig. 9.2). Further, the compressive through-thickness stress on delamination was also considered for the numerical simulation. The through-thickness stresses were defined by introducing the contact friction along the shear direction. The composite laminate was meshed using C3D8 solid elements. The impactor was considered as solid rigid body. The contact between the laminate and impactor was defined by using surfaceto-surface contact. The matrix failure was modelled by using Hashing failure criterion and the matrix failure evolution defined by using stiffness degradation model. The surface-based cohesive elements were only defined for the angle mismatching interfaces because the interfaces with no angle mismatch show no delamination. The finite element results showed good correlation with the experimental observations. It was found that the delamination initiation and propagation in upper interfaces were dominated by through-thickness compression stresses, while the lower interfaces were dominated by the cohesive behaviour. The numerical results also showed that the delamination area was not significantly affected by the interlaminar friction coefficient when the value was greater than 0.6. Long et al. [11] conducted finite element analysis of CF/epoxy composite laminates under LVI using cohesive contact-based damage model and focused on efficient modelling of delamination damage. The cohesive elements were only for the bottom plies, while the tied contact definition was implemented for the plies towards the impact side (Fig. 9.3). The Hashin damage failure criteria were used to model the intralaminar damage. The cohesive contact method based on quadratic damage criteria was adopted to model the interlaminar damage in the composite laminates. The linear damage evolution law was used to model both intralaminar and interlaminar damage progression. The adopted damage model showed good correlation with

Fig. 9.2 Schematic illustration of load transfer for surface-based cohesive contact model [10]

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Fig. 9.3 Schematic illustration of contact definition used during the finite element simulation [11]

the observed experimental results. It was found that the delamination area formed at the top two layers whose contact was defined by cohesive elements was accurate. It was concluded that the damage (delamination) area formed was symmetric about the impact cite. However, the damage shape was dependent on the adjacent plies interface orientation. Qu et al. [12] tried to calculate the energy dissipation in CF/epoxy composite laminate under LVI using finite element analysis. The finite element analysis was emphasized to capture the evolution of delamination by considering normal crack and tangential slip. The intralaminar damage or failure mechanism in CF/epoxy composite laminate was modelled using the Hashin failure criteria. The laminates were meshed with 4-node shell elements. The cohesive zone model was used to define the interlaminar delamination. The cohesive zone modelling was based on the energy dissipation mechanism. The impactor was modelled as a rigid body. The contact between adjacent plies in the CF/epoxy composite laminate was modelled using tiebreak contact definition. The numerical model developed using cohesive zone model and continuum damage mechanics accounted the energy responsible for damage evolution as well as the energy dissipated. Finally, it was concluded that the finite element model showed good agreement with experimental results along with stable simulation. Antonucci et al. [13] validated the experimental penetration energy, maximum force and indentation depth results obtained by LVI on CF/epoxy composite laminates. The cohesive elements were used to indicate the interlaminar damage formation and progression. These cohesive elements followed the bilinear traction–separation law. Both numerical and experimental results showed good correlation with each other. Czarnocki et al. [14] tried to numerically simulate the sequential failure modes that occur during the low velocity impact (Fig. 9.4). The plies contact was defined by using CONTA173 and Targe170 contact element. The composite laminate was meshed using SOLID185 elements. The damage mechanism in CF/epoxy was modelled using cohesive zone model which worked on the principle of bilinear traction–separation law. The velocity–time (rising portion of the curve) plot showed good correlation between numerical and experimental results. The numerical simulation results suggested that the first kink in the laminate was due to the composite plate inertia. The numerical and experimental results were not consistent in terms of damage extent and amount of energy dissipated due to the coarser mesh.

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Fig. 9.4 The following schematic representation illustrated the force history curve for CF/epoxy with sequential damage mechanism. a, b Divergence in nodal velocity in nodes present beneath the impactor. c Numbering of nodes. d Intralaminar failure initiation in the bottom layer. e Intralaminar failure evolution in the bottom plies. f–h Delamination initiation in the bottom and middle plies of the composite laminate. i, k Overall intralaminar damage in the middle and bottom plies. j, m Overall delamination failure in the top-middle and middle-bottom plies [14]

Penettieri et al. [15] conducted the finite element analysis of LVI on CF/epoxy showcasing complete experimental (Fig. 9.5). The C3D8 elements were used to mesh the composite laminates. The numerical simulation results predicted that the imperfect tightening of internal bolt sliding resulted in energy dissipation. Then, the results were used to correct the impactor set-up. Farooq et al. [16] developed finite element model based on the ply-by-ply failure of the CF/epoxy composite laminate. For numerical simulation, both flat and round nosed impactors were considered along with [0/45/−45/90]ns stacking sequence with n = 1, 2 and 3. The models were modelled using linear and nonlinear shell elements. From the results, it was found that the 8-node iso-parametric meshing was well

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Fig. 9.5 Finite element model of the complete experimental set-up of drop weight impact used during the experimentation [15]

suited to model the experimentation. Further, the composite laminates were meshed using conventional shell elements. The damage in composite laminates was modelled using progressive damage model which worked on the Hashin failure criteria. The numerical simulation results showed that the Hashin failure criteria predicted the ply failure successfully and showed good agreement with the experimental results. Jagtap et al. [17] used the LS-DYNA finite element software to simulate the LVI on CF/epoxy composite laminates for different boundary conditions and mesh sizes. The 3D solid elements were used for the numerical simulation. The simply supported and clamped boundary conditions were considered for the analysis. The composite laminate was modelled by MAT_059 (orthotropic material model) which used the stress failure criteria. The contact definitions were defined by surface to surface and node to surface contact. The experimental results were converged for mesh size 2 mm. The boundary condition strongly influenced the energy absorption capacity of the laminate. The simply supported boundary condition absorbed more energy than the clamped boundary condition. Chen et al. [18] investigated the damage behaviour of CF/epoxy-honeycombsandwiched composite laminates using finite element analysis. General contact algorithm was used to define the contact. The interlaminar delamination was modelled using cohesive elements. The CF/epoxy composite skin stacking sequence was [(45/−45), (0/90), (0/90), (−45/45)]. The S4R shell elements were used to model the composite laminate. The continuum damage model was used to depict the damage mechanism. The proposed finite element results showed a good correlation with the experimental results. The peak force predicted was within an error of 4.2%. Elias et al. [19] reported the finite element analysis of 3D woven CF/epoxy composite laminate under LVI (Fig. 9.6). The Onera damage model developed on

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Fig. 9.6 3D woven architecture of carbon fibre fabric reinforced with epoxy [19]

continuum damage mechanics was used to model the 3D CF/epoxy composite laminate. Since the used laminate was 3D, the macroscopic damage variables such as fibre yarn failure and interyarn debonding were accounted to reduce the simulation instabilities. The impactor was modelled using isotropic elastic behaviour law. Unlike experimental results, the simulation results did not show any large-scale oscillations on the force–time curve. The shear matrix cracks and interyarn debonding were in good correlation with the experimental results. Further, the impact dent depth was accurately predicted by the used finite element model. Alzeanidi et al. [20] conducted finite element analysis to analyse the damage mechanism of repaired CF/epoxy composite laminates under impact loading. The stacking sequences considered for numerical simulation were [(0/90)/(±45)/(0/90)/(±45)/(0/90)/(±45)/(0/90)/(±45)/(0/90)/(±45)/(0/90)] and [(±45)/(0/90)/(0/90)/(±45)/(0/90)/(0/90)/(±45)]. The impactor was modelled using MAT_ELASTIC_PLASTIC_HYDRO material card. Further, a 4-node Belytschko-Tsay element along with one through-thickness integration point was also considered to model the impactor. The composite laminate was modelled using solid elements which represents the thickness of the ply. The MAT_COMPOSITE_FAILURE_SOLID_MODEL was considered to model the composite laminate. The numerical simulation results showed good agreement with experimental results in terms of kinetic energy, damaged area and damage shape. Sun et al. [21] conducted numerical analysis of CF/epoxy composites by combining 3D solid and 2D shell elements to understand the damage created by LVI (Fig. 9.7). The plies of the CF/epoxy composite laminate was modelled using single integration point brick elements. The matrix cracks were modelled using intralaminar cohesive elements. The segment-based contact definition was used to define the adjacent plies surfaces. Three different types of stacking sequences were considered for numerical analysis: (i) ply-blocked scaling (Ps)—[452 /02 /902 /−452 ]2S , (ii) sublaminate scaling (Ss)—[45/0/90/−45]4S and (iii) large-scale laminate (Ls)— [452 /02 /902 /−452 ]2S . Both Ss and Ps laminates showed significant difference in

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Fig. 9.7 a, b, c Solid/shell model in a global/local approach for impact modelling; d modelling strategies for integrating high-fidelity solid and shell part into solid/shell impact model [21]

the critical or peak load. Further, there was no clear difference in delamination area between Ps and Ss. The high-fidelity 3D finite element analysis showed good efficiency as well as accuracy. Liu et al. [22] numerically modelled the CFRP/aluminium sandwiched composite laminates under LVI for different impactor shapes (hemispherical, flat and conical). The corrugated aluminium core geometry was trapezoidal. The composite plates were modelled using Hashin failure criteria and Yeh delamination failure criteria. The composite laminates were meshed using 8-node linear brick reduced integration elements. The aluminium alloy core was modelled using elastic–plastic material model. The damage progression was modelled using solution-dependent variables based on maximum strain criterion which eliminates the solution instabilities. Figure 9.8 illustrates the damage formed in the sandwiched composite laminates under LVI for different impactor shapes. According to the numerical simulation results, the flat nosed impactor showed higher maximum force, shorter contact duration and higher initial slope. The investigation results of finite element analyses showed good agreement with the experimental results.

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Fig. 9.8 Finite element analysis of damage formation in CF/epoxy/aluminium sandwiched composite laminates under LVI (10 J of impact energy) for different impactor shapes [22]

Moumen et al. [23] simulated the low velocity impact on CF/epoxy composite laminate reinforced with CNTs. The progressive failure model based on the continuum damage model was used to model the damage formation and progression in CF/epoxy/CNT nanocomposite. The Hashin failure criteria were used to model the intralaminar failure criteria. The numerical results showed good corroboration with the experimental results. Topac et al. [24] simulated the LVI on unidirectional CF/epoxy composite beams (Fig. 9.9). The impactor shape was cylindrical head. A 3D finite element analysis with intraply matrix damage based on continuum damage mechanics was used to model the damage in the CF/epoxy composite beams. The damage initiation in the composite laminate was modelled using LaRC04. The cohesive interface elements were used to define the delamination failure. The damage evolution was modelled using linear softening response. The 2D model proposed by Matzenmiller was used to model the stiffness degradation of CF/epoxy composite laminate as the damage

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Fig. 9.9 a Schematic illustration of finite element model of CF/epoxy composite beam. b Representative volume element of the finite element model showing cohesive layer and CF/epoxy ply [24]

evolution takes place. The simulation results showed that the delamination at the top interfaces was initiated by matrix cracking. The mesh size affected the matrix crack distribution and total delamination area. Finally, it was concluded that the simulation results of matrix cracking and delamination were in good agreement with the experimental results. Wang et al. [25] conducted numerical simulation of LVI on unidirectional CF/epoxy composite laminates. The intralaminar damage initiation was modelled using Hashin failure criteria along with linear damage evolution law. Each interface in the laminate was defined by cohesive contact. The delamination shape and size predicted by the numerical simulation were in good corroboration with the experimental results. Dubary et al. [26] used the discrete ply model technique to capture the failure modes such as permanent indentation, delamination, fibre breakage and matrix cracks in CF/epoxy composite laminates generated under LVI. Figure 9.10 illustrates the discrete ply mode principles. The delamination between the plies was simulated using classical interface elements. The Hashin failure criteria were used to model the intraply matrix cracking, whereas the fibre failure was modelled based on the continuum damage mechanics. The permanent indentation in the laminate created after the LVI was modelled using pseudo-plastic non-return law. The numerical model determined the best impact damage tolerant laminate configuration. Finally, it was found that even after optimization based on the numerical model results the total numerical simulation time was 36 h for single impact. Pashmforoush et al. [27] combined the multi-objective genetic algorithm with the numerical model to optimize the stacking sequence for CF/epoxy composite laminate under LVI. Three objective functions (i) Hashin failure criterion, (ii) interlaminar shear stress failure and (iii) tensile stress were used for stacking sequence optimization. The CF/epoxy composite laminates were considered as homogenous material with orthotropic linear elastic behaviour prior to the damage initiation. After the damage initiation, a 3D Hashin failure model was used to model the damage

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Fig. 9.10 Discrete ply model illustrating the principles of elements and their associated failure modes [26]

progression. Both numerical and experimental results showed good agreement. The optimized results were as follows: (i) [90/90/90/0]S and [0/90/90/0]S showed least fibre damage value under compression; (ii) [45/−60/60/45]S and [60/90/0/−45]S stacking sequences showed least fibre damage value in tension; (iii) [45/−60/0/90]S and [45/−60/60/45]S stacking sequences showed least matrix damage value under compression; (iv) the stacking sequences [45/−45/45/−45]S and [30/−60/60/−30]S showed low delamination damage; (v) the stacking sequence [45/0/90/−45]S and [60/90/00/-45]S yielded least out-of-plane shear stresses; and (vi) [90/90/90/90]S and [0/90/90/0]S showed highest out-of-plane strength. Overall, the stacking sequence [60/90/0/−45]S was considered as the optimized stacking sequence. Liu et al. [28] conducted a finite element analysis on hybrid CF/epoxy composite laminates subjected to LVI. Here, the hybridization was done by using two different CF fabric architectures such as unidirectional and satin-weaved fabrics. The delamination (interlaminar damage) was modelled using cohesive elements. From simulation results, it was observed that the delamination propagation was influenced by the lower ply. The simulated force and absorbed energy results showed good correlation with the experimental results. Mahmoud et al. [29] adopted the semi-continuous strategy (Fig. 9.11) to carry the finite element simulation of LVI on hybrid unidirectional/woven CF/epoxy composite laminates. In this analysis, the woven fabric architecture was modelled similar to a truss structure. The contact definition was modelled using shell-to-shell interface elements. The elements used were eight nodes where each element had three translational and three rotation degrees of freedom. The fibres and matrix failure damage mechanism were successfully modelled by the semi-continuous strategy process. Chen et al. [30] used the continuum damage mechanism method and cohesive elements to describe the intraply and interply failure. Here, the carbon/glass/basalt/epoxy hybrid composite laminates were considered for the

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Fig. 9.11 Semi-continuous strategy to model the woven CF/epoxy composite laminate [29]

numerical analysis. The penalty-based contact definition was used for modelling the contact between plies and impactor. The delamination was modelled by using surface-based cohesive contact. Here, the delamination behaviour of the composite laminate was modelled using bilinear traction–separation law. The force–displacement curve obtained from numerical simulation showed good agreement with the experimental force–displacement curve. Numerical analysis of low velocity impact on glass fibre reinforced polymer composite laminates In a study, Li et al. [31] developed a finite element model based on the 9-node Mindlin plate element. The finite element method predicted peanut-shaped damage at the angle mismatching interfaces in the GF/epoxy composite laminate. As the composite plate size was increased correspondingly, the maximum peak force also decreased; however, the maximum deflection of the plate increased. Further, as the impact energy was increased then correspondingly the delamination area was also increased. The finite element results also revealed that the boundary condition does not influence the impact properties significantly, whereas the impactor mass significantly influenced the impact performance of the GF/epoxy composite plate. Overall it was concluded that the used finite element approach showed good agreement with the experimental observations. Sevkat et al. [32] adopted the 3D dynamic nonlinear finite element analysis to analyse the impact performance of GF/epoxy composite laminate. The orthotropic elastic material card was used to assign the material properties to the GF/epoxy composite laminate. The adopted material card worked on the Chang-Chang failure criteria. The delamination failure was defined by using CONTACT_AUTOMATIC_SURFACE_TO_SURFACE_TIEBREAK. Further, the contact between impactor and composite plates was defined by using

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ERODING_SURFACE_TO_SURFACE contact model which worked on strainbased failure criteria. It was found that the predicted results from the adopted finite element method were in good correlation with experimental results. Menna et al. [33] used the finite element method to study the impact performance of GF/epoxy composite laminates. The plies were modelled using 3D 8-node elements. The MAT_059 material card was used to model the composite plate. The MAT_059 models the progressive damage in the composite plate based on the 3D stress-based failure criteria. The contact definition among the plies was modelled using a spring-based model. The delamination damage created in the GF/epoxy composite plate was modelled using SURFACE_TO_SURFACE contact card which works on the normal and shear strength. The impactor was considered as a rigid body, and the contact between impactor and composite plate was modelled using AUTOMATIC_SURFACE_TO_SURFACE. The developed finite element model results showed good agreement with experimental observation in terms of force–displacement curve along with ply-by-ply delamination. It was found that the delamination was more at the mid-section of the laminate. Finally, it was concluded that the interlaminar strength influences the delamination formation in the composite plate under LVI during finite element modelling. Barbero et al. [34] tried to establish a methodology to determine the material properties required to conduct the progressive damage model analysis in ABAQUS. The damage initiation criteria implemented were based on the Hashin and Rotem failure criteria. These failure models consider apparent stresses. The strain softening method was used for damage evolution. Apart from progressive damage model, the ply discount method was also used to study the damage evolution in GF/epoxy composite laminate. From numerical results, it was found that the progressive model adequately predicted the Mode I matrix cracking while showed inefficiency in determining the mixed mode matrix cracking. Further, the ply discount procedure overestimated the results than the progressive damage model. It was concluded that the adopted finite element approach was not suitable for CF/epoxy composite laminate because the degradation of elastic modulus due to transverse matrix cracks was very small compared to the carbon fire elastic modulus. Perillo et al. [35] studied the thick GF/epoxy composite laminate behaviour under LVI using finite element analysis. The damage at the plies interfaces was studied by implementing the cohesive elements which were defined based on the bilinear traction–separation law. The delamination initiation and evolution were defined based on the cohesive zone model. The progressive damage model was adopted to define the intraply failure of the GF/epoxy composite plate. The numerical results showed good agreement with experimental results in terms of impact time, the peak impact force, the impactor displacement, delamination (size and shape) and matrix cracking. However, as the impact energy increased the accurate predictability by the finite element method reduces. It was found that the numerical analysis complexity increases with the increase in number of interfaces in the composite laminate. It was concluded that the model adopted here with simple material properties can successfully predict the impact behaviour of the GF/epoxy composite laminate.

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Kursun et al. [36] conducted a finite element analysis of LVI on preloading GF/epoxy composite plate. Figure 9.12 represents the complete flow chart of the UMAT subroutine used in ABAQUS for conducting finite element analysis of LVI. Three preloading conditions (i) biaxial tension, (ii) biaxial compression and (iii) tension and compression were considered for the analysis. The composite plate was meshed using the Hashin failure criteria. The composite plate was modelled using 8-node C3D8 solid elements with 4-integration point. The impactor was meshed using 4-node elements with bilinear discrete quadrilateral R3D4 elements. The finite element results showed that the preloading increased the rigidity of the composite plate. Further, the preloaded specimens showed higher damage degree than the unloaded specimens because the preloaded specimens absorbed more energy than the unloaded specimens. Finally, the finite element results were in line with the experimental observations with respect to force–time and energy–time histories. However, the damage shape and size predicted by the finite element analysis were not in line with the experimental results when the preloading increased. In a study, Hassan et al. [37] used the MSC.MARC® tool to conduct the finite element analysis of LVI on GF/epoxy composite laminate. The plies and impactor were modelled using 3D 8-node solid elements. The orthotropic model was adopted to define the material properties to the composite plate. The impactor was considered as a rigid body. The automatic contact definition was used to define the contact

Fig. 9.12 Flow chart of the UMAT subroutine of the ABAQUS® used during the finite element analysis [36]

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between impactor and composite plate. The finite element results showed that the contact force was slightly lower than the experimental observations. Yang et al. [38] carried the numerical analysis on hybrid GF/CF/polymerized poly(butylene terephthalate) composite laminates subjected to LVI. The Hashin failure criteria with progressive damage model were used to define the impact damage in hybrid composite plates. The impactor was considered as a rigid body. Both composite plate and impactor were solid meshed using C3D8R elements. The surface-based cohesive behaviour contact definition was used to define the contact between glass and carbon plies. In this approach, the delamination initiation at the interface occurs when the quadratic interaction function (stress ratio) reaches the value one. The failure modes predicted by the developed finite element model were in good correlation with the failure modes observed in experimentation. Meybodi et al. [39] conducted the finite element analysis of LVI on GF/epoxy composite laminate beams doped with MWCNTs and nanoclay. Further, the analytical model was also developed using Euler–Bernoulli beam theory and Hertz’s contact law as governing equations. Ritz’s approximation approach was used to obtain the nonlinear equations with respect to time domain which were solved using Runge–Kutta numerical method. Figure 9.13 shows the finite element model of the composite beam with impactor. The plies were meshed using 3D 8-node solid elements. The composite laminates were modelled using ORTHOTROPIC_ELASTIC material card. The impactor was modelled as rigid body. The interaction among the composite plies was modeled using CONTACT_AUTOMATIC_SURFACE_TO_SURFACE_TIEBREAK. The impactor and composite laminate contact was defined by using CONTACT_ERODING_SURFACE_TO_SURFACE. The element deletion was defined by ADD_EROSION card. It was found that the analytical model predicted the lower contact force than the finite element model prediction compared to experimental observations. In a study, Singh et al. [40] investigated the impact performance of symmetric and asymmetric composite plates. The impactor was modelled using MAT_020 card which defined impactor as a rigid body. The composite plate was modelled Fig. 9.13 Finite element model of GF/epoxy nanocomposite beam with impactor [39]

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Fig. 9.14 Schematic illustration of the integrated multi-scale model of the FRP composites under LVI [41]

using MAT_059 (COMPOSITE_FAILURE_SOLID_MODEL) material card. The contact between impactor and GF/epoxy composite plate was defined using surfaceto-surface contact, whereas the contact among the plies was defined using tiedsurface-to-surface contact. The symmetric laminate showed lower damage height and width compared to asymmetric composite design which were in close agreement with experimental observations. Yang et al. [41] presented an integrated multi-scale model for composite plates under LVI (Fig. 9.14). The multi-scale model was based on the embedded cell model. The cohesive elements were used to model the delamination. From finite element analysis, it was found that the matrix cracks were at the bottom side of the laminate as the first damage form. Then, these matrix cracks were moved at the plies interfaces to form delaminations as a second damage form. The final failure mode observed was fibre pull-out or the fibre breakage. The numerical simulation results showed good agreement with the experimental observations available in the literature. Rawat et al. [42] studied the influence of inserting pre-crack in symmetric and asymmetric GF/epoxy composite laminates on the impact performance under LVI. The composite plate was modelled using MAT_SOLID_COMPOSITE_FAILURE_SOLID material card. The finite element results showed that the symmetric laminate impact performance was better compared to asymmetric composite laminate with and without pre-crack. However, the precrack insertion reduced the impact strength of both symmetric and asymmetric composite laminates. In another study, Rawat et al. [43] used the same material cards to model the composite and impactor to study, the effect of impactor nose shape on the damage and impact performance of GF/epoxy composite laminate (Fig. 9.15). It was observed that the impactor shape that exhibited maximum contact area with the composite laminate created maximum damage area. However, the oval-shaped impactor created fibre damage. It was also observed that the flat nosed impactor did not penetrate the composite laminate unlike spherical, hemispherical and oval-shaped impactor. Rawat et al. [44] used the finite element analysis approach to investigate the impact performance of GF/epoxy composite laminate. Both impactor and composite plates were meshed by solid elements. The impactor and composite plates were modelled

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Fig. 9.15 Different impactor shapes modelled: a hemispherical, b spherical, c oval and d flat [43]

using MAT_RIGID and MAT_SOLID_COMPOSITE_FAILURE_SOLID_MODEL material cards. Further, the impactor and composite plate contact was defined using AUTOMATIC_SURFACE_TO_SURFACE, whereas the contact among the plies was defined by TIED_SURFACE_TO_SURFACE contact model. The finite element analysis showed good correlation with the experimental force–time history along with damage size and shape. In an investigation, Mars et al. [45] used the finite element analysis to understand LVI behaviour of GF/polyamide composite laminates. Viscoelastic constitutive equations were used for the formulation of finite element analysis. These equations consist of Helmholtz free energy equations to account the thermodynamic potential. Further, the classical plasticity yield condition and quadratic Hill criterion were also considered for the numerical analysis. The impactor was meshed by 4-node C3D8 linear tetrahedron elements. The viscoelastic model predicted the stress–strain behaviour of GF/polyamide composite laminate. The results obtained from the finite element analysis showed consistency with the experimental observation. Rawat et al. [46] conducted numerical simulation on GF/epoxy composite laminate under oblique LVI. The oblique impacting angles were 0°, 15°, 30° and 45°. The impactor was considered as rigid structure. The composite laminate was modelled using MAT_SOLID_COMPOSITE_FAILURE_SOLID_MODEL material card. The contact among the plies was modelled using TIED_SURFACE_TO_SURFACE contact definition, while the contact between impactor and laminate was defined by AUTOMATIC_SURFACE_TO_SURFACE contact definition. The simulation results showed that the energy absorption by the composite laminate was increased with an increase in impact angle. The stress response of the composite laminate was also influenced by the impact angle. Complete failure of the composite plate was observed when impacted at 0° impact angle. The deflection of the composite laminate was also strongly influenced by the impact angle. In a study, Mahesh et al. [47] also used the same material card (as that of Rawat et al. [46]) to conduct the finite element analysis on GF/epoxy composite laminate under LVI with different edge boundary conditions. The finite element results showed

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that the composite plate constrained along the long edge absorbed more energy than the composite laminate constrained along the short edge. Further, the long edge constrained along one side showed more deflection than the composite plate constrained along one short edge. It was concluded that the space availability of the impact affects the damage pattern, deflection of the plate and energy absorption. Mahesh et al. [48] conducted finite element simulation of edge impact on GF/epoxy composite laminate under different boundary conditions. The composite plate and impactor were modelled using MAT_RIGID and MAT_COMPOSITE_FAILURE_SOLID_MODEL material cards. The finite element results showed that the impact force was more for edge constrained boundary condition and was more compared to the corner constrained boundary condition. It was concluded that for edge impact the boundary condition applied strongly affected the damage area, and energy absorption by the composite plate. In another finite element analysis, Mahesh et al. [49] used the same material cards (as that of Rawat et al. [46]) to investigate the effect of area availability for impact in a fully constrained circular boundary condition under LVI. The material considered for the analysis was GF/epoxy composite laminate. The finite element results showed that as the boundary condition area reduced correspondingly the damage area was also reduced along with composite plate deformation energy absorption capacity of the composite plate. Ahmadi et al. [50] developed an analytical model and conducted finite element simulation as well for grid stiffened composite panel. Here, the skin panel was made of GF/epoxy composite laminate. The smeared method was used to develop the analytical model. The mass spring model was used to conduct the finite element analysis (Fig. 9.16). Both analytical and finite element results were in good correlation with the experimental results. In an investigation, Berton et al. [51] have developed a multi-scale damage model for finite element analysis to study the LVI behaviour of GF/epoxy composite laminate. Figure 9.17 shows the representative volume element based on the synergistic Fig. 9.16 Schematic illustration of spring mass [50]

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Fig. 9.17 Representative volume element of the synergistic damage mechanics [51]

damage mechanics representing primary features such as uniform spacing of ply cracking, cracks run parallel to the fibre in each ply, crack formation in multiple off-axis orientation and loading scenarios in multiple directions. It was observed that the developed model was less sensitive to the rate dependency. Numerical analysis of low velocity impact on other fibre reinforced polymer composite laminates Goo et al. [52] used the 3D penalty-based finite element approach to simulate the LVI on graphite/epoxy composite laminates. The predicted results were then compared with the results obtained from modified Hertz contact law. The composite plates were meshed using 8-node elements. The discretized governing equations were coupled together by Newmark’s constant average acceleration approach. The developed finite element method also considered the laminate thickness for the analysis. It was found that the results obtained from the finite element analysis were in good agreement with the experimental observation. Roy et al. [53] carried finite element analysis on graphite/epoxy, Kevlar/epoxy and hybrid graphite/Kevlar/epoxy composite laminates under LVI. Each composite plate was meshed using 3D 8-noded iso-parametric solid elements. The finite element analysis was conducted based on the Newmark model along with the Hertzian contact law. Further, the stresses at the interfaces were estimated using Hinton and Campbell least square formulation. During finite element analysis, the damping effect was neglected. The stacking sequences of the composite laminates were as follows: (i) graphite/epoxy—[0G/−45G/45G/90G]2S where ‘G’ represents the graphite ply; (ii) Kevlar/epoxy—[0K/−45K/45K/90K]2S where ‘K’ represents the Kevlar ply; and (iii) hybrid graphite/Kevlar/epoxy—[0K/−45G/45G/90G/0G/-45G/45G/90G]S and [0K/−45K/45K/90K/0G/−45G/45G/90G]S . The finite element results showed that the delamination damage was more in the Kevlar/epoxy composite plates than the graphite/epoxy or the hybrid graphite/Kevlar/epoxy composite laminates. Further,

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the contact time was also much higher for Kevlar/epoxy composite laminates than the graphite/epoxy composite plate. Furthermore, among hybrid composite laminate the hybrid composite plate with Kevlar plies at top and bottom while graphite plies being inner of the laminate design showed much improved performance. Khalili et al. [54] investigated several parameters which affect the accuracy and predicting capability of finite element analysis for a composite plate under LVI. These parameters considered were element type, solution method, impactor modelling method, meshing pattern and contact modelling. Two types of elements such as 4-node shell elements and 8-node continuum shell elements were considered for the analysis (Fig. 9.18). Further, impactor was assumed to be a rigid body. Furthermore, both implicit and explicit were used to conduct the finite element analysis. Based on the finite element analysis, it was that the 8-node shell showed more accuracy than the 4-node shell elements. However, the 8-node shell elements consumed more computational run time than the 4-node shell elements. It was concluded that for thin plate and shells the 4-node shell elements were more suitable whereas for thick plate and shells the 8-node shell elements were more suitable. The consideration of impactor as a rigid body was appropriate as the computational time was less. However, consideration of deformable impactor leads realistic and more accurate results at the cost of computational time. Finally, it was found that for plates with large degree of freedoms the explicit solver was appropriate with good accurate results and low computational time. However, for small degree of freedom the computational time may increase if explicit method was adopted. Thus, it recommended the use of explicit solver for solving impact problem. Feng et al. [55] used the progressive damage model to predict the impact damage in graphite/epoxy composite laminates subjected to LVI. The progressive damage model was developed based on the continuum damage mechanics along with cohesive interface elements. The primary aim of the work was to predict the through-thickness internal damage distribution. The developed finite element predicted the impact

Fig. 9.18 Finite element model of graphite/epoxy composite plate modelled with different element types: a 4-node shell element and b 8-nose continuum shell element [54]

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results with good agreement with the experimental results. Further, the developed model predicted agreeable results with experimental results such as delamination shape, delamination size and delamination orientation. The intralaminar damage was model based on the energy-based continuum damage mechanics. The fibre damage initiation was modelled using a coupled tension–compression damage evolution law. The Schurmann and Puck failure criteria were used to model the matrix transverse failure under compression. The nonlinear shear failure modes were modelled from cubic polynomial stress–strain curve. The interface cohesive elements were used to model the interlaminar failure or delamination at the plies interfaces. These interface cohesive elements were defined based on the traction–separation constitutive law. The initial damage initiation was based on the linear elastic stage after the linear softening approach was used to model the damage evolution. Based on the predicted results, it was concluded that the addition of in-ply damage modes in the finite element analysis is required for accurate prediction of the results. However, the use of in-ply damage modes increased the computational complexity of the model used. Ansari et al. [56] used the AUTODYN hydro code to conduct the finite element analysis for Kevlar/epoxy composite laminate under low to hyper-velocity impact. In finite element analysis, the effect of shock was also considered. Different parameters such as boundary condition, impact velocity, thickness-to-span ratio, span of the composite plate and composite plate thickness were considered during the finite element analysis. The finite element analysis results showed that the ease of penetration during the impact into the plate was easier under fully constrained boundary condition. As the impact velocity increases, then the damage occurrence in the laminate concentrates around the impactor. Further, the damage area in thick composite plate was more compared to thin composite plate. Overall, it was concluded that the adopted finite element analysis results showed good agreement with the experimental results in terms of ballistic limit, damage pattern and residual velocity. Fan et al. [57] used the self-consistent model that was used to conduct the finite element analysis of LVI on flax/carbon/epoxy hybrid composite laminates doped with CNTs. Figure 9.19 represents the finite element model of the hybrid composite layer reinforced with CNT layer. The visco-Pasternak foundation was also used to investigate the environmental temperature influence on the impact performance of CNT reinforced carbon/flax composite laminate. The modified Hertz model was adopted to define the contact between hybrid nanocomposite and impactor. The damage in the composite plate due to LVI was established using higher-order shear deformation and von Karman nonlinear strain displacement relationship. It was concluded that the viscoelastic model can be used to predict the impact performance of hybrid composite plates reinforced with CNTs layer. Liao et al. [58] studied the graphite/epoxy composite laminates under LVI via finite element analysis with two different approaches. In the first approach, the interlaminar damage was modelled using plastic damage model. The plastic damage model was based on Puck’s failure criteria and strain-based damage evolution law both of which defined the fibre and matrix failure. The delamination was modelled using bilinear cohesive model. In the second approach, a coupling between intralaminar

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Fig. 9.19 Finite element model representing the hybrid composite laminate with CNT layer under LVI [57]

plastic deformation and damage evolution was coupled together by strain equivalence hypothesis. Here, the effective stresses and strains were estimated first using backwards Euler algorithm. After that, the nominal effective stresses were updated independently. The finite element results predicted that the matrix tension damage was observed at the bottom plies of the graphite/epoxy composite laminates and propagated upwards. Further, the matrix compressive damage was observed at top side of the composite plate and the fibre damage was observed in the bottom plies. The finite element results showed good agreement with the experimental results in terms of force–time curve, force–displacement curve and dissipated energy. It was concluded that the plastic damage model was efficient to predict the elastic damage features in graphite/epoxy composite plate. Fragassa et al. [59] conducted a benchmark study on finite element analysis of basalt/vinyl ester composite laminates under LVI. The impactor was considered as a rigid body and modelled using MAT_RIGID material card. The contact between indenter and composite plate was defined by CONTAC_AUTOMATIC_SURFACE_TO_SURFACE contact card, and the interface contact among the adjacent plies was defined by CONTACT_AUTOMATIC_ONE_WAY_SURFACE_TO_SURFACE_TIEBREAK card. The composite plate was meshed in three different ways: (i) single-layered shell approach, (ii) mixing of stacked and layered approach with two shells and (iii) four shells. Figure 9.20 shows the 1, 2 and 4 shells’ meshing. The MAT_054 material card was used to define the Chang-Chang failure criteria to the composite plate. In this material card, different parameters such as ALPH (elastic region coefficient), BETA (tensile fibre coefficient in the Chang-Chang failure criteria), SOFT (regulates the strength degradation), FBRT (represents the fibre tensile strength reduction) and YCFAC (progressive degradation in the fibre compressive strength) were optimized. These parameter adjustments had insignificant influence

288 Fig. 9.20 Schematic illustration of finite element with impactor and composite plate meshed with 1, 2 and 4 shells [59]

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2D

4D

on the 1 shell model. However, these parameters showed noticeable influence on the results particularly BETA and FBRT for 2 shell models. For 4 shell models, all the parameter adjustments showed considerable influence on the finite element results. It was found that the parameters ALPH with low, BETA of 0.5, SOFT with high and FBRT with high value bet for good and accurate output results. Bogenfeld et al. [60] conducted a benchmark study on finite element analysis of LVI on FRP composite plates. Here, different models were presented to examine their differences. These adopted methods include spring mass model, plate model, layered shell model, stacked shell model, stacked shell model and ply splitting model. Based on the finite element results, the following points were made: (i) for quick prediction and faster simulation, adaptation of macroscale models along with layered shell elements and implicit time integrations were recommended. (ii) Use of mesoscale modelling with solid elements and cohesive surfaces was recommended for better damage capturing at cost of computational efforts. (iii) Advanced damage models should be used to capture the damage more accurately along the accuracy. The layered shell model should be the choice because to adopt the high-fidelity damage models the number of strength parameters required is high. For structurelevel impact analysis, the combination of layered shell model and stacked layered model was suggested. It was also found that the computational efforts were reduced as the impactor mass reduced and impactor velocity increased. Choi et al. [61] conducted finite element analysis on composite plates under LVI using cylindrical shell elements with concave and convex shapes (Fig. 9.21). The shell elements were defined according to the von Karman’s large deflection theory. Both convex and concave shell elements showed similar contact force and central deflection. The strain distribution contour plot obtained from the convex shell elements was similar to the bottom layer made of concave shell. Jalon et al. [62] considered the strain rate effect to conduct the finite element analysis on flax/PLA composite laminate under LVI. The constitutive model based on the rheological approach was used to account the strain rate effect (Fig. 9.22). The Yeoh model was used to describe the nonlinear elastic branch. The viscoelastic branch was modelled using Maxwell model. The plastic branch was also modelled using Maxwell model along with the help of friction. The impactor was modelled considering linear elastic behaviour. The numerical model developed showed good

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Fig. 9.21 Cylindrical convex and concave shell used for the composite laminate [61]

Fig. 9.22 Schematic representation of rheological approach: a nonlinear elastic branch, b viscoelastic branch and c plastic branch [62]

agreement with the experimental observations. It was concluded that unlike synthetic fibres the fibre failure up to linear elastic was considered neglecting strain rate effect. However, the natural fibres show plastic region and strain rate sensitivity. Thus, the adopted model can be used to understand the LVI behaviour of natural composite laminates. Bogenfeld et al. [63] validated the scaling method of composite laminates under LVI through finite element method. The scaling method adopted was developed based

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on the spring mass model. The finite element results showed good correlation with experimentation as well as their usefulness in early design phase of the composite laminates. Uncertainties in the results increase as the scaling of the composite material increases. It was also observed that the location of the stiffening element around the impact site influenced the damage pattern and size.

9.2 Numerical Analysis of Compression After Impact on Fibre Reinforced Polymer Composite Laminates Numerical analysis of compression after impact on carbon fibre reinforced polymer composite laminate Ghelli et al. [64] carried the finite element simulation of CAI tests of pre-impacted CF/epoxy composite laminates. The major aim of the numerical simulation was to analyse the buckling behaviour of CF/epoxy composite laminates under compressive loading for pre-impacted specimens. In finite element simulation, no material model was used. The out-of-plane shear strains were modelled by defining the nodal properties directly. Gonzalez et al. [65] conducted the finite element simulation of CAI tests on CF/epoxy composite laminates. Two standard stacking sequences selected according to ASTM D7136 standard were [454 /04 /−454 /904 ]S and [(452 /02 /−452 /902 )2 ]S . The interlaminar damage was modelled using zero thickness cohesive elements. The element size was calculated by equating elastic energy of the element with the energy dissipated by fracture to avoid snap-back of the constitutive softening. Numerical simulation results suggested that the out-of-plane stresses should be considered for damage activation especially under LVI. Further, the shear component parallel to the fibre should be determined by using coupled plasticity and damage model instead of constitutive law when used for unidirectional composite laminate. Biao et al. [66] conducted the numerical simulation of CAI tests on CF/epoxy laminates using solid elements and cohesive elements. The stacking sequence used for numerical simulation was [45/0/−45/90/0/0/45/0/−45/−45]. 8-node C3D8R solid elements were used to model the CF/epoxy lamina. For numerical simulation, stress-based failure criteria were modified into strain-based failure criteria. Further, use of explicit method will avoid the solution convergence problem. The cohesive elements successfully simulated the delamination failure, and the predicted postimpact compressive strength values were in good agreement with the experimental results. Rivallant et al. [67] used discrete interface elements to simulate the CAI tests. The fibre failure was modelled based on GIC fracture toughness. The matrix cracking was introduced using zero thickness 3D cohesive elements at the ply interface along the fibre direction. These zero thickness at ply interfaces was also used to model the delamination which was based on quadratic criterion which was also used for modelling failure criteria of matrix. Both experimental and numerical simulation

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Fig. 9.23 Numerical simulation of CFRP composite laminates under CAI test [67]

results showed good correlation among themselves. Figure 9.23 shows the failure modes observed in CFRP composite laminate during numerical simulation. Suemasu et al. [68] used the analytical model to study buckling phenomenon in impacted FRP composites under compressive loading. Buckling in FRP composite laminates during compressive loading for impacted specimens was modelled using double spiral shape damage model and multi-annular damage model. It was observed that the damage propagation was unstable along the transverse direction with respect to loading direction. The adopted analytical models approximately estimated the post-impact compressive strength. Han et al. [69] carried computational simulation of CAI test on CCF300/epoxy, CCF800/epoxy, CCF300/bismaleimide and CCF800/bismaleimide composite laminates. The intralaminar multi-scale failure approach and interlaminar cohesive element methods were used to simulate the failure modes in CF composite laminate during CAI tests. The intralaminar failure of fibre and matrix was based on stress-based multi-scale failure criterion. The interlaminar cohesive elements were used to simulate the delamination, where the method was based on linear traction– separation law. To reduce the computational time, the global–local approach was used. The eight-node linear brick reduced integration elements were used to mesh the intralaminar region, while the eight-node 3D cohesive elements were used to mesh the interlaminar region. The mesh size was 2.5 mm × 2.5 mm. The impactor was considered as a rigid body. The simulation results showed good correlation

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with the experimental results of CAI tests. Further, the simulation also successfully analysed the failure mechanisms. Rozylo et al. [70] developed a simplified damage model (SDM) to access the stiffness of FRP laminate after subjecting it to impact via compressive loading. Here, CFRP laminate was considered for numerical simulation. SDM used progressive damage criterion with Hashin failure criterion to investigate the failure of the laminates. The damage evolution was modelled based on fracture energy dissipation criterion. Results obtained from numerical simulation were compared with experimental results available in the literature and found a good agreement with each other. Sun et al. [71] examined the compression after the behaviour of CFRP laminate by introducing barely visible damage via low velocity impact and static indentation. Two stacking sequences [452 /902 /02 /−452 ]2S and [45/90/0/−45]4S were considered, and digital image correlation was used to analyse the damage mechanism growth. Numerical models were developed and found that both experimental and numerical results were in functional correlations. Finally, it was concluded that the [45/90/0/−45]4S yielded better impact resistance properties than other lay-up design, and it showed the highest resistance to small deformations. Gonzalez et al. [72] used to shell elements and cohesive surfaces to simulate the CAI tests on unidirectional and woven CF/epoxy composite laminates. The shell and surface elements were connected by tie connectors which transferred the kinematics between these two elements. The cohesive elements were used to model the delamination between the interfaces of plies. The defined contact definition was based on the penalty stiffness method. 4-node general purpose shell elements (S4R) were used for element modelling. Further, at refined boundaries oriented at 45° the 3-node shell elements (S3R) were used. The stacking sequences used for the analysis were [(0, 45, 90, −45)2 ]S and [02 , 452 , 902 , −452 ]S for unidirectional CF/epoxy composites while [(0, 45)3 ]S for woven CF/epoxy composite laminates. The simulation results showed that the use of conventional shell elements and cohesive surfaces reduced the computational time. Tuo et al. [73] investigated the compression after impact behaviour of CF/epoxy composite laminates. A 3D damage model was developed based on continuum damage mechanics. Both interlaminar delamination and intralaminar damages were included in the developed damage model. The interlaminar delamination was modelled using cohesive elements. The matrix damage initiation was modelled using maximum strain failure criterion. Similarly, the fibre damage initiation was modelled using improved 3D Puck failure criterion. The damage evolution was modelled using bilinear damage constitutive relation. The proposed numerical simulation and experimental results showed good agreement with each other. Tuo et al. [74] conducted finite element analysis of CAI for thin CF/epoxy composite laminates. A 3D damage model was used for finite element analysis. The model considered both intralaminar and interlaminate failure modes. The intralaminar damage model was based on an energy-based continuum damage mechanics. Here, the fibre damage was modelled based on maximum strain criterion, while the matrix damage was modelled based on Puck and Schurmann matrix failure criterion.

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The interlaminar damage was modelled using cohesive elements. Here, a quadratic stress-based criterion was used to account the delamination initiation. The results of the adopted finite element method showed good agreement with the experimental results for thin CF/epoxy composite laminates. Numerical analysis of compression after impact on glass fibre reinforced polymer composite laminates Mendes et al. [75] used the single shell and split shell model to simulate CAI tests. Only shell elements were used in single shell approach where delamination failure was neglected. However, the split shell approach considered the delamination where the delamination in the FRP composite was defined by contact logic. The interlaminar damage was modelled using progressive failure which was formulated using smeared cracking formulation. The interlaminar damage was modelled using mixed mode failure criteria. The numerical predictions by split shell method were better compared to single shell element approach. However, both approaches were failed to predict the perforation of FRP composite laminates at high impact energies. It was concluded that the use of shell elements with solid elements reasonably accounted delamination failure mode and post-impact compressive strength of FRP composite laminates. Perillo et al. [76] carried the numerical simulation of CAI tests on thick GF/epoxy composite used in wind or marine industries. The through-thickness elements were modelled by C3D8R, whereas the COH3D8 cohesive elements were used between those interfaces where the two adjacent plies were oriented differently. The cohesive elements were used to model the damage development at the interfaces or interlaminar failure. Further, progressive failure method was adopted to model the intraply failure. The cohesive elements were worked on the principle of classical energy-based bilinear traction–separation law. The matrix cracks were modeled by using Puck’s failure criteria, whereas Hashin failure criteria was used to model the fibre failure. The contact definition was modelled based on penalty approach. It was concluded that the use of these simple approaches can be used to successfully conduct the numerical simulation. The numerical simulation results were in good correlation with the experimental results. Thin-walled structure behaviour under post-impact compressive loading was investigated by Debski et al. [77] and developed a simplified damage model to examine the compressive behaviour of post-impacted thin-walled FRP laminates at different positions. Damage initiation was predicted using Hashin failure criteria, and damage evolution was predicted by energy criteria. For experimentation, GFRP laminates were considered, and the impact test was conducted using a drop weight machine according to ASTM D7136 standard, 16 J of impact energy and impactor of 16 mm nose diameter. According to the proposed model, numerical analysis and experimental results, low energy impact at a different position on thin-walled structures affected the post-impact compressive strength of the laminate at a very low scale.

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Numerical Analysis of Compression after Impact on other Fibre Reinforced Polymer Composite Laminates Yan et al. [78] conducted numerical modelling and simulation of CAI tests of GF/vinyl ester composite laminates. The schematic representation of numerical mode is shown in Fig. 9.24. For simplicity, the impact-induced damage in the GF/vinyl ester laminate was assumed to be circular. Four failure modes are (i) matrix cracking, (ii) delamination within the impact-induced damage, (iii) cohesive zone outside the impact zone for delamination and (iv) matrix and fibre failure within the damage zone. The after impact compressive buckling strength was 173 MPa which was significantly higher than the experimental value which was 117 MPa. This significant difference was observed because the delamination was restricted only within the impacted zone. However, when the delamination zone was extended to the complete specimen then the numerical simulation underestimated the after impact buckling strength which was 103 MPa. From numerical simulation, it was also found that the delamination propagation was strongly influenced by loss in interfacial strength and less dependent on the critical energy release rate. In a study, Dang et al. [79] used the finite element analysis approach to study the post-impact compressive behaviour of variable angled tow FRP composite laminates. The interface elements were modelled using bilinear cohesive law. The laminates were modelled using C3D8R solid elements. Further, the interface region between two plies was modelled using COH3D8 elements. Finite element analysis showed

Fig. 9.24 Schematic representation of numerical modelling of CAI test specimen. a Boundary condition. b Impact-induced damage along through-thickness direction. c Meshed model of CAI test specimen [78]

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that the delamination failure was the primary for significant reduction in post-impact compressive strength. Further, the delamination was propagated along perpendicular to the loading direction and dependent on the lamina orientation. Esrail et al. [80] developed a mathematical model to predict the compression after impact properties. The model considered concentric elliptical inclusions which represented the damage. These inclusions were incorporated with progressive failure criterion to investigate the damage evolution. Obtained numerical results were compared with the experimental results available in the literature. Tan et al. [81] conducted finite element analysis of CAI tests using continuum shell elements and cohesive elements for stitched CF/epoxy composite laminates. The Hashin failure criterion was used with continuum shell elements, whereas the traction–separation law was used with cohesive element to define the delamination at interfaces. The stitched CF/epoxy composite laminates were modelled using 8-node quadrilateral shell elements (SC8R) with hourglass control. The cohesive interfaces were modelled using 8-node linear 3D cohesive elements. The stitching of the plies was defined by spring elements. The finite element simulation results were in excellent agreement with the experimental results. Tan et al. [82] developed a composite damage mathematical model to analyse the impact and compression after impact properties of FRP laminate using physically based tensile and compressive failure for matrix and fibre failure. While the interlaminar or delamination damage was modelled using cohesive zone method. These developed models were then inserted into ABAQUS for numerical simulation. Results obtained from numerical simulation compared with experimental results were available in the literature. From numerical simulation results, it was evident that the developed model captured both inter- and intralaminar failure mechanisms, both qualitatively and quantitatively. These results were in good agreement with the experimental results. Romano et al. [83] compared the residual compressive strength of impacted specimens with the specimens containing open hole (Fig. 9.25). The hole size was also varied to study the effect of hole size on the residual compressive strength of the FRP composite laminates. Two different stacking sequences [45/−45/45/−45/0/0/90/0/0/45/−45/0/90/0]S and [45/−45/0/90]3S were used for the numerical simulation. The hole size in the FRP laminate was ¼ inch. The impact damage was induced by impacting the specimens with 30 and 50 J for [45/−45/45/−45/0/0/90/0/0/45/−45/0/90/0]S , while the [45/−45/0/90]3S laminate was impacted at 50 and 100 J of impact energies. The Hashin failure criteria were used to detect the lamina failure. The Newton–Raphson iterative process was used to achieve the nonlinear finite element solution. Based on numerical simulation results, it was concluded that the residual compressive strength of a FRP composite laminate can be evaluated by simply modelling the FRP composite laminates with a hole whose size must be equal to the size of the barely visible impact damage. Abir et al. [84] developed a mathematical model and incorporated into ABAQUS to study the residual compressive strength after subjected to impact loading. Obtained numerical results were compared with existing experimental results from the literature. Continuum damage model was used to study the failure between the fibre

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Fig. 9.25 Finite element model. a Impacted specimen. b Specimen with hole mimicking the impact induced damage [83]

and matrix interfaces, i.e. intralaminar failure, whereas cohesive interface elements were used to model delamination behaviour, i.e. interlaminar failure. Numerical results compared with the obtained experimental results from the literature were in good corroborative. Finally, it was concluded that damage growth in CAI was not a continuous process but occurred rapidly. Fibre compressive failure and interlaminar fracture toughness were the key ingredients which greatly influenced the compressive strength of the laminate after subjecting it to the impact loading. Abir et al. [85] investigated the ply lay-up sequence, sublaminate scaling, ply blocking, boundary conditions to establish a relationship between various failure mechanisms with respect to compression after impact using continuum damage mechanics for fibre–matrix interface failure and cohesive zone modelling for delamination. Developed mathematical models were integrated with ABAQUS for numerical simulation. From obtained results, it was evident that change in lay-up directly influenced the damage position, pattern and compressive residual strength. Increase in damage size and decrease in CAI were found in the case of ply blocking parameter. Finally, as the thickness of the laminate was increased fibre compressive failure was observed. Soto et al. [86] developed a mathematical model to evaluate the damage mechanism associated with spread tow fabric under low velocity impact (LVI) and compression after impact (CAI) loading condition. Furthermore, the authors also evaluated fibre constitutive law shape and tried to enhance the computational efficiency of the model. Experimental testing, i.e. LVI and CAI, was conducted according to ASTM D7136 and D7137, respectively. Numerical simulation results suggested that for accurate prediction of damage mechanism, the constitutive law shape was more important even if the matrix cracking was neglected entirely. However, the damage

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mechanism of CAI was utterly different because unlike in LVI, CAI did not show any damage mode until the final collapse occurred. Finally, the authors concluded that experimental and numerical results were in good agreement, but numerical simulation lacked the robustness. Liu et al. [87] developed a mathematical model to predict compression after impact damage and strength reduction of a hybrid laminate under low velocity impact. Continuum damage model and linear elastic fracture mechanics models were used, which accounted for inter- and intralaminar damage. Hybrid laminates were prepared by unidirectional and woven carbon fibre, and developed damage model was defined using VUMAT user-defined tool using ABAQUS. From numerical and experimental results, it was found that in all the cases, hybrid laminate performed better when compared to pristine laminate. Thorsson et al. [88] used the continuum shell-based modelling approach to carry the finite element analysis of FRP composite laminates under compressive loading after being subjected to LVI. In-plane progressing damage and discrete cohesive elements were used to model the failure mechanisms. Further, enhanced Schapery theory was used to model the nonlinear matrix microcracks. Both Schapery theory and discrete cohesive elements were used to model the intralamina and interlamina failure. For CAI numerical simulation, the impacted model was imported. The finite element simulation results for the post-impact compressive strength of FRP specimen containing 25% barely visible impact damage showed 22% reduction where the experimental results showed only 8% reduction. However, good correlation was observed between numerical simulation and experimental results for specimens with 75% of barely visible impact damage. Overall the finite element predictions showed 7.2% to 14.4% variations in compressive after impact with the experimental results.

9.3 Summary Finite element analysis of LVI and CAI on FRP composite laminates helps in better understanding the associated damage mechanism and failure modes. It is observed that the computational time for LVI and CAI on FRP composite laminate primarily depends on the mesh type, element type, material models and contact formulation. Furthermore, the accuracy of the finite element depends on the same parameters mentioned above. However, there is a trade-off between accuracy and computational time along with computational cost. It is evident that the successful finite element analysis also depends the careful selection of the material models and contact formulation which depends on the extent to which the FRP laminate undergoes damage.

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