delamination damage of carbon fiber-reinforced … · 2019. 11. 6. · model (cczm). in this...
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DELAMINATION DAMAGE OF CARBON FIBER-REINFORCED POLYMER
COMPOSITE LAMINATES UNDER CYCLIC SHEAR-INDUCED LOADING
CONDITIONS
A thesis submitted in fulfilment of the
requirements for the award of the degree of
Doctor of Philosophy (Mechanical Engineering)
School of Mechanical Engineering
Faculty of Engineering
Universiti Teknologi Malaysia
OCTOBER 2018
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To my beloved
Mom, Dad, and Wife for
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ACKNOWLEDGMENT
Alhamdulillah, I have completed writing this thesis but of course with the help
and support from fantastic peoples around me. First and foremost, I wish to express
my heartfelt gratitude to my supervisor, Prof. Dr. Mohd. Nasir Tamin for his
professional guidance and support in academic and in real life. I am very indebted to
his patience and invaluable advices that inspired me to see things positively and felt
honoured with his confidence and trust on my ability.
I would also like to thank my co-supervisor, Dr. Wong King Jye and a special
thanks to Dr. Seyed Saeid Rahimian Koloor for providing me with valuable
suggestions and recommendations. My sincere thanks to all my friends and colleagues
for their cheerfulness and fun. Thank you for the great moments that we have shared
together especially Mahzan, Amirul, Sharul, Izzudin, Zal, Faizah, Marhaini,
Azisyahirah, Intan, Nurasykin and all members of Computational Solid Mechanics
Laboratory (CSMLab)
Grateful acknowledgement is also made for financial support by the Ministry of
Higher Education (MOHE) Malaysia and Universiti Tun Hussein Onn Malaysia.
Last but not least, a special thanks to my family for their continuous support
and encouragement especially to Mak, Baba and my wife, Hafizah Ibrahim. This
dissertation is the product of so much more than five years and my family has
supported me through it all. There are so many people to whom I am indebted for
support and encouragement.
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ABSTRACT
Interface delamination is a major failure mode induced by in-plane shear stress frequently encountered in carbon fiber-reinforce polymer (CFRP) composite laminates structures. This failure process under monotonic loading has been successfully described using cohesive zone model (CZM). Many previous CZM approaches for cyclic case were considering damage parameters based on a crack growth relation which has some disadvantages for predicting a non-linear crack growth. In addition, the previous CZM approach lacks ground in understanding the physics underlying the delamination process and effect of stress ratio. The objective of the study is to extend the existing CZM to account for the delamination damage evolution of CFRP composite laminates under cyclic shear-induced loading conditions named as cyclic cohesive zone model (CCZM). In this respect, the fatigue damage response and the residual interfacial properties associated with the development of the CCZM are established under cyclic shear-induced loading condition. A series of Mode-II-type tests were performed on pre-fatigued end-notched flexural (ENF) beams of CFRP composite laminates, [0]8 for different applied load ratio conditions (R = 0.1, 0.15 and 0.25) to induce only interlaminar damage at the pre-existing delamination interface crack front. Subsequent quasi-static test to catastrophic failure establishes the characteristic residual strength responses of the damaged specimen. A hybrid experimental-computational approach was introduced to obtain the residual interlaminar properties for all the loading cases. A normalized gradual degradation rule was used to present the degradation for interlaminar shear strength ( ), penalty stiffness ( ) and the critical Mode-II energy release rate ( ) which cover the wide range of interlaminar failure mode from wear out to sudden death. This interlaminar properties degradation model can describe the characteristic evolution of the interlaminar damage response and the degradation of CCZM properties under cyclic shear-induced loading case. The interlaminar properties degradation model together with the CCZM model is coded by using user-define material model (UMAT) subroutine to implement in ABAQUS finite element analysis (FEA) software. This model had been used to simulate under 3-point bending cases and compared with the experiment results. Result had shown that the comparison between the FE simulation and the experiment fatigue load-life cycles for CFRP composite laminate interfaces are close with the difference of less than 1% and shows a very successful verification of the modified CCZM model to simulate the interlaminar damage evolution and failure response. Besides that, an independent validation had been run to validate the performance of interlaminar properties which were obtained in the study. A load-deflection response under 3-point bending case was simulated based on [0]16 ENF specimen under identical load cycle parameters and compared with the measured experiment results. Result shows that the peak load differences between the experiment and simulation is less than 6%. From the study, the capability of CCZM model for cyclic case has been demonstrated by linking interlaminar properties degradation with damage mechanics approach. This will help in understanding the physics underlying the delamination process and effect of stress ratio.
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ABSTRAK
Pelekangan antara lamina adalah mod kegagalan utama yang disebabkan oleh tegasan ricih sesatah yang seringkali dialami oleh struktur lapisan komposit polimer bertetulang-gentian karbon (CFRP). Proses kegagalan ini di bawah beban ekanada telah berjaya dihuraikan menggunakan model zon jelekat (CZM). Banyak pendekatan CZM terdahulu untuk kes berkitar telah mengambilkira parameter kerosakan berdasarkan hubungan pertumbuhan retakan yang mana mempunyai beberapa kelemahan dalam meramal pertumbuhan retak bukan lelurus. Selain dari itu, pendekatan CZM sebelum ini kurang mempunyai asas pemahaman fizik berdasarkan proses pelekangan dan kesan nisbah tegasan. Objektif kajian ini adalah untuk memperluaskan CZM sediaada bagi mengira evolusi kerosakan pelekangan pada lapisan komposit CFRP yang disebabkan oleh keadaan beban kitaran ricih-teraruh yang dinamakan model zon jelekat berkitar (CCZM). Dalam hal ini, tindak balas kerosakan lesu dan sifat-sifat sisa antara lamina yang dikaitkan dengan pembangunan CCZM dihasilkan berdasarkan keadaan beban kitaran ricih-teraruh. Siri ujian jenis Mode-II dijalankan pada rasuk lenturan takuk-hujung (ENF) daripada lapisan komposit CFRP, [0]8 yang dipra-lesu pada keadaan nisbah beban berbeza (R = 0.1, 0.15 and 0.25) untuk mengaruh kerosakan antara lamina pada hadapan retakan antara lamina yang sediaada. Seterusnya ujian kuasi-statik sehingga kegagalan bencana menetapkan ciri-ciri tindak balas sifat-sifat sisa kekuatan spesimen yang rosak. Kaedah hibrid pengkomputeran-eksperimen diperkenalkan bagi mendapatkan sifat-sifat sisa antaramuka pada kesemua kes pembebanan. Aturan penurunan beransur ternormal telah digunakan untuk mewakili kemerosotan kekuatan ricih antara lamina ( ), kekukuhan denda ( ) dan Mod-II kritikal kadar pelepasan tenaga ( ) yang merangkumi pelbagai mod kegagalan antara lamina dari haus sehingga kegagalan mengejut. Model sifat penurunan antara lamina ini akan menghuraikan evolusi ciri tindak balas kerosakan antara lamina dan kemerosotan sifat-sifat CCZM di bawah kes beban kitaran ricih-teraruh. Model sifat penurunan antara lamina bersama-sama dengan model CCZM dikodkan dengan menggunakan model bahan takrif-pengguna (UMAT) subrutin untuk dilaksanakan dalam perisian analisis unsur terhingga (FEA) ABAQUS. Model ini telah digunakan untuk mensimulasi kes lenturan 3-titik dan dibandingkan dengan keputusan eksperimen. Keputusan telah menunjukkan bahawa perbandingan antara simulasi FE dan eksperimen kitaran hayat-beban kelesuan untuk lapisan komposit CFRP dengan perbezaan kurang dari 1%. Ini menunjukkan pengesahan yang sangat berjaya terhadap model CCZM yang diubah suai untuk mensimulasikan evolusi kerosakan antara lamina dan tindak balas kegagalan. Selain itu, pengesahan bebas dijalankan untuk mengesahkan prestasi sifat antara lamina yang telah diperolehi dalam kajian ini. Tindak balas beban-permesongan di bawah kes lenturan 3-titik telah disimulasi berdasarkan spesimen ENF [0]16 di bawah parameter kitaran beban yang sama dan dibandingkan dengan data eksperimen yang diukur. Keputusan menunjukkan, perbezaan beban puncak di antara eksperimen dan simulasi adalah kurang daripada 6%. Dari kajian ini, keupayaan model CCZM untuk kes kitaran telah dibuktikan dengan menghubungkan sifat penurunan antara lamina dengan pendekatan mekanik kerosakan. Ini akan membantu memahami dasar fizik proses pelekangan dan kesan nisbah tegasan.
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TABLE OF CONTENTS
CHAPTER TITLE PAGE
DECLARATION ii
DEDICATION iii
ACKNOWLEDGEMENTS iv
ABSTRACT v
ABSTRAK vi
TABLE OF CONTENTS vii
LIST OF TABLES xii
LIST OF FIGURES xiii
LIST OF ABBREVIATIONS xviii
LIST OF SYMBOLS xix
LIST OF APPENDICES xxi
1 INTRODUCTION 1
1.1 Introduction 1
1.2 Background and Rationale 4
1.3 Statement of the Research Problem 6
1.4 Objectives of Study 6
1.5 Scope of Study 7
1.6 Significance of Study 8
1.7 Layout of Thesis 8
2 LITERATURE REVIEW 11
2.1 Introduction 11
2.2 Carbon Fiber-Reinforced Polymer (CFRP)
Composites
12
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2.3 CFRP Composite Laminates Construction 13
2.3.1 Modes of Failure in CFRP Composite
Laminates
14
2.3.2 Delamination in CFRP Composite
Laminates
15
2.3.3 Why Interlaminar Shear-Induce Loading Is
Critical in Composite Laminates
Structures?
17
2.4 Fatigue Phenomenon in CFRP Structures 18
2.5 Effect of Stress Ratio or Mean Stress on Fatigue
Delamination Growth in Composite Laminates
19
2.6 Development on Damage Model and Finite
Element Analysis for Composite Laminates
21
2.6.1 Continuum Damage Mechanics 23
2.6.2 Damage Mechanics-Based Models 24
2.7 Failure Criteria for Fiber Reinforced Composites 25
2.8 Cohesive Interface Elements in the Finite Element
Method
28
2.8.1 Cohesive Zone Model 29
2.8.2 Review of Cohesive Zone Model for Cyclic
Loading Application
35
2.9 Gradual Material Property Degradation Rules 38
2.9.1 Gradual Strength Degradation Model 39
2.9.2 Gradual Stiffness Degradation Model 41
2.9.3 Normalized Fatigue Life Model 41
2.10 Summary of the Literature Review 43
3 RESEARCH METHODOLOGY 44
3.1 Introduction 44
3.2 Research Framework 45
3.3 Part 1 - Formulation of CCZM for Cyclic Loading
Case
47
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3.4 Part 2 - Obtaining Interface Properties and Damage
Model Parameters
48
3.4.1 Specimen Preparation 50
3.4.2 3-point Bending Test Procedure 51
3.4.3 Identify Permissible Loading for Shear
Fatigue Test
52
3.4.4 Shear-Induced Fatigue Test Procedure 55
3.4.5 Hybrid Experimental-Computational
Process
58
3.4.5.1 Obtaining CCZM Parameters for
Quasi-Static Loading
59
3.4.5.2 Obtaining CCZM Parameters for
Fatigue Case
60
3.4.6 Obtaining Fatigue Damage Parameters 62
3.5 Part 3 - Implementation of The Extended Damage
Model into FE Simulation Model
62
3.5.1 User-defined Subroutine Code (UMAT for
CCZM
63
3.6 Part 4 Verification and Validation of The CCZM
Model by Predicting Interface Damage Evolution
64
3.6.1 Verification of The New CCZM Model for
Cyclic Shear-Induced Loading
64
3.6.2 CCZM Illustrative Model 64
3.6.3 Independent Validation 65
3.7 Summary of The Research Methodology 68
4 FORMULATION OF THE CYCLIC COHESIVE
ZONE MODEL FOR CYCLIC SHEAR-INDUCED
LOADING
68
4.1 Introduction 68
4.2 Extension of The Cohesive Zone Model for Cyclic
Loading
69
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4.2.1 Key Equations for the Cohesive Zone
Model
70
4.2.2 Formulations of the CCZM 73
4.3 CCZM Parameters 74
4.3.1 Quasi-Static Damage Result 75
4.3.2 Permissible Loading for Shear-Induced
Fatigue Test
77
4.3.3 Residual Load For Fatigue Damage
Specimen
82
4.3.4 Hybrid Experimental-Computational
Result
85
4.3.5 Cyclic Cohesive Zone Model Parameters 89
4.3.6 Performance of the Cyclic Cohesive Zone
Model
92
4.4 Normalized Properties Degradation Model 94
4.4.1 Interlaminar Property Degradation
Characteristics
96
4.4.2 Effect of Mean Stress on Fatigue Life of the
Interfaces
98
4.4.3 Interlaminar Fatigue Damage
Characteristics
99
4.5 Subroutine UMAT 101
4.5.1 Unloading/reloading Paths 102
4.5.2 Viscous Regularization 103
4.6 Chapter Summary 103
5 SIMULATING INTERFACE DAMAGE PROCESS
UNDER CYCLIC SHEAR-INDUCED LOADING
CONDITIONS
105
5.1 Introduction 105
5.2 Interlaminar Damage Evolution of CFRP
Composite Laminates
106
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5.2.1 Interlaminar Damage Evolution Under
Cyclic Shear-Induced Loading Condition
106
5.2.2 CFRP Composite Laminates Interface
Fracture Process
108
5.2.3 Comparison Between Simulation and
Experimental Results
113
5.3 CCZM Illustrative Model Results 114
5.4 Independent Validation Result 117
5.5 Chapter Summary 122
6 CONCLUSIONS AND RECOMMENDATIONS 123
6.1 Conclusions 123
6.2 Recommendations 125
REFERENCES 126
Appendices A - C 137 - 158
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LIST OF TABLES
TABLE NO. TITLE PAGE
3.1 Elastic properties and damage model parameters of lamina
54
3.2 Elastic properties and damage model parameters of interface for unidirectional CFRP composite laminates
54
3.3 Fatigue loading conditions used to induce different levels of interlaminar damage
56
4.1 Load cycle-dependent residual interface properties of the CFRP composite laminates
90
5.1 Load cycle-dependent properties for the CCZM 120
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LIST OF FIGURES
FIGURE NO. TITLE PAGE
1.1 Typical use of composite in aircraft Boeing 787 2
1.2 The building block approach applied to aerospace composite structures
3
2.1 CFRP laminated composite construction 13
2.2 Failure scale in composite laminates 15
2.3 Modes of delamination failure 16
2.4 Nomenclature for constant stress amplitude loading
19
2.5 Mechanics of solid materials and analysis domains classification
25
2.6 Cohesive zone laws (a) polynomial, (b) piecewise linear, (c) exponential and (d) rigid-linear
30
2.7 Bi-linear traction-separation curve for CZM 31
2.8 Pure mode constitutive equations: (a) Mode II or Mode III; (b) Mode I
32
2.9 Schematically gradual strength degradation in laminar composite
40
3.1 Flow chart of research methodology 46
3.2 Flowchart of hybrid experimental-computational process
49
3.3 (a) Actual ENF specimen, b) Location of insert, made of Teflon tape as pre-crack, c) The specimen geometry and dimensions in millimeters
51
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3.4 (a) 3-point bending test set-up in millimeters, (b) Actual set-up of the ENF fracture test
52
3.5 FE model construction of unidirectional CFRP composite (a) representing the Multi-layer FE model-based construction (b) FE model of the ENF specimen of CFRP composite laminate, illustrating the element mesh
53
3.6 Location of cohesive layer in between 4th and 5th
the initial crack)
55
3.7 Illustration of the loading applied for fatigue test for different load ratio (a) R = 0.1 (b) R = 0.15 and (b) R = 0.25
57
3.8 a) Simulated global load-deflection responses, b) simulated shear traction-displacement of the cohesive zone
60
3.9 a) Simulated global load-deflection responses, b) simulated shear traction-displacement of the cohesive zone under fatigue loading
61
3.10 Flowchart for subroutine UMAT 63
3.11 (a) Cyclic 3-point bending test set-up in millimeters, (b) Actual set-up cyclic 3-point bending for CFRP composite laminates with the composite code of [0]16.
66
3.12 FE model for 16-ply unidirectional CFRP composite laminate with ENF test set-up, illustrating element mesh
66
4.1 (a) Traction-relative displacement softening laws for the interface under shear-mode loading condition; (b) Traction and displacement components acting on interface plane.
71
4.2 Fatigue degradation parameters of cohesive interface
74
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4.3 Load-center-span deflection responses for ENF quasi-
static
76
4.4 CFRP specimen micrographs before test at the crack tip for 20× magnification
76
4.5 CFRP specimen micrographs after test at the crack tip for 20× magnification
77
4.6 Finite model construction of CFRP composite (a) FE model of ENF specimen under 3-point bending setup, (b) Multi-layer FE model-based construction
77
4.7 Measured and FE-calculated load-deflection responses of the ENF beam specimen.
78
4.8 FE-calculated HCriterion at 300 N load responses; (a) Fiber Compression, (b) Fiber Tension, (c) Matrix Compression, (d) Matrix Tension
80
4.9 Contour of matrix tension initiation criterion illustrated for the 4th layer at the point load of 300 N
80
4.10 Evolution of (a) stress and (b) damage initiation variable at CFRP composite middle interface crack front (QUADSCRT)
81
4.11 Load-center-span deflection responses for ENF specimen of CFRP composite beam at various levels of interlaminar damage
82
4.12 CFRP specimen micrographs at 20× magnification (a) before test, (b) after 250 k cycle fatigue test, (c) after Mode-II fracture test
84
4.13 Measured and FE-calculated load-deflection responses of the ENF beam specimen
86
4.14 (a) Shear stress contour at cohesive layer, (b) Shear stress distribution along the crack front
87
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4.15 (a) Relative tangential displacement at the crack tip (b) Resultant shear stress evolution at the edge location
88
4.16 Measured and FE-calculated load-deflection Exp.
Fatigue 50 k
90
4.17 Characteristics of the residual shear fatigue properties of the interface in CFRP composite laminates (a) Penalty stiffness, (b) Interface shear strength
91
4.18 Residual load-deflection responses of ENF beam specimens following the respective number of load cycles ( measured; --- FE-calculated)
93
4.19 Characteristic log-normal degradation of the interlaminar shear strength for the loading condition of = 30 MPa, R = 0.1
97
4.20 Residual interlaminar properties, normalized with their respective static values for the loading condition of = 30 MPa, R = 0.1
98
4.21 Fatigue load-life relationship for CFRP composite laminate interfaces under Mode-II loadings
99
4.22 Unified residual fatigue property of the interface under fluctuating shear stress: (a) Interlaminar shear strength ( =0.6564, =18.2375), (b) Penalty stiffness ( =2.1123, =12.3425), and (c) Critical Mode-II strain energy release rate ( =2.1123, =8.3425)
101
4.23 Unloading and reloading path 102
5.1 (a) Applied load at the loading roller; Shear stress during the peak of first cycle (b) contour (c)
contour
107
5.2 (a) Damage onset evolution at selected number of cycles, (b) Damage onset evolution variable legend
109
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5.3 (a) Evolution of damage at selected number of cycles, (b) Damage evolution variable legend
110
5.4 (a) Critical material points ahead of the crack tip, (b) Evolution of shear stress and (c) Evolution of damage at CFRP composite middle interface (damage initiation variable, SDV1 and propagation variable, SDV2)
112
5.5 Number of cycles to failure comparison between FE simulation and fatigue load-life relationship graph for CFRP composite laminates interfaces
113
5.6 Evolution of (a) shear stress, (b) damage at CFRP composite middle interface (damage initiation variable, SDV1 and propagation variable, SDV2) and (c) critical material point ahead of the crack tip for the case Pmax = 330 N (R = 0.1)
115
5.7 Evolution of (a) shear stress, (b) damage at CFRP composite middle interface (damage initiation variable, SDV1 and propagation variable, SDV2) and (c) critical material point ahead of the crack tip for the case Pmax = 150 N (R = 0.2)
116
5.8 Fatigue load-life comparison between FE simulation and calculated life parameters of CFRP composite laminates interfaces
117
5.9 CFRP specimen micrographs at 20× magnification (a) before test, (b) after 50 k cycle fatigue test, (c) after Mode-II fracture test
119
5.10 FE model for 16-ply unidirectional CFRP composite laminate with ENF test set-up
120
5.11 Load-central deflection response of ENF specimens with different cyclic interlaminar damage induced
121
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LIST OF ABBREVIATIONS
ASTM - American society for testing and materials
CAE - Computer aided engineering
CCZM - Cyclic cohesive zone model
CDM - Continuum damage mechanics
CFRP - Carbon fiber-reinforced polymer
CLS - Crack lap shear
CZM - Cohesive zone model
ENF - End-notched flexure
FE - Finite element
FEM - Finite element method
FRP - Fiber-reinforced polymer
LEFM - Linear elastic fracture mechanics
QUADSCRT - Quadratic traction damage initiation variable
SDEG - Scalar stiffness degradation
SDV - Solution-dependent state variables
UD - Unidirectional
UHMW - Ultra high molecular weight polyethylene
UMAT - User-define material model
VCCT - Virtual crack closure technique
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LIST OF SYMBOLS
a - Crack length
b - Width
d - Fatigue degradation parameters
- Frequency
- Critical energy release rate
- Mode-I (crack opening) critical energy release rate
- Mode-II (in-plane shearing) critical energy release rate
- Mode-III (out-of-plane shearing) critical energy release rate
- Mode-I energy release at failure
- Mode-II energy release at failure
- Total energy release rate
- Penalty stiffness
- Mode-I Penalty stiffness
- Mode-II Penalty stiffness
- Penalty stiffness in shear direction, fatigue loading
L - specimen span
- Tensile strength
- Number of cycles
- Fatigue life
- Load
- Maximum cyclic level
- Minimum cyclic level
- Mean load
- Load amplitude
- Load ratio
- In-plane shear strength
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- Maximum stress
- Minimum stress
- Stress amplitude
- Mean stress
- Out-of-plane shear strength
t - thickness
- Exponent for B-K criterion
- Deflection
- Mode-I relative displacement at damage onset
- Mode-II relative displacement at damage onset
- Mode-I relative displacement at fracture
- Mode-II relative displacement at fracture
- Normal stress
- In-plane shearing stress
- Out-of-plane shearing stress
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LIST OF APPENDICES
APPENDIX TITLE PAGE
A Panel Fabrication Process and Test Report 137
B User-Written Subroutine Code for Implementing
CCZM In FE Simulation
146
C User Subroutine Interface (UMAT), Variables to
Be Defined, Variables That Can Be Updated,
Variables Passed in For Information
153
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CHAPTER 1
INTRODUCTION
1.1 Introduction
Fiber-reinforced polymer (FRP) matrix composites consist of thermoset or
thermoplastic resin matrix reinforced by much stronger and stiffer fibers such as
carbon, glass and ultra-high-molecular-weight (UHMW) polyethylene fibers. Carbon
fiber-reinforced polymer (CFRP) composites are typically hired in advanced
applications which includes the aerospace, automotive, marine, sport and construction
applications. Modern aircraft design like the Boeing 787 were among the aircraft
representing the maximum use of this material in the primary structures such as wing
skins and fuselage which are made of CFRP composite. The increment usage of CFRP
composites is derived from their advantages such as high strength, light weight,
corrosion resistance, chemical resistivity, electrical conductivity and many more. In
this case, it is very necessary to predict the failure response of CFRP composite during
the design and analysis processes of the structures. There are two approaches to predict
the failure response which are by using the experimental and numerical approaches.
Most of the experimental procedure is very costly and time consuming, especially
when dealing with complex loading. For the industry applications, CFRP structural
components are usually subjected to complex fatigue loading histories [1] specified by
various amplitude, stress ratio, frequency and waveform of the stress cycles during
service. However, deficiencies in current lifetime prediction methodologies for these
materials often require large factors of safety that need to be adopted to ensure safety
of the materials in the application. Therefore, composite structures are usually
overdesigned and require an extensive prototype testing that need to be used for a
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proper lifetime prediction. By using the finite element (FE) tools, the lifetime
prediction of CFRP material could be done to improve the design by using less number
of expensive physical tests.
Figure 1.1: Typical use of composite in aircraft Boeing 787 [2].
The knowledge of Computer Aided Engineering (CAE) is used to evaluate the
material's resistance under different loading conditions in simulated service
environment. It is well known from the aerospace industry that composite structures
are sized based on the building block approach [3]. This methodology is described in
Figure 1.2, with the pyramid concept. The main idea is to build the knowledge on the
material and structural behavior step by step, starting from the fundamental stage at
the coupon level up to the full-scale structures. It has been observed over the years that
simulation, and especially models based on the finite element method, are more and
more used in the different stages of the pyramid, trying to become a companion of the
physical tests. It is indeed evident that tests can be expensive when repeated several
times for different material configurations (e.g. different stacking sequences) or when
changes in the components geometry or loading conditions. Therefore, by using virtual
testing can help reduce the product development costs and time. To fulfil this
requirement, finite element analyses must be predictive. If this condition is satisfied,
simulation can then replace some physical tests.
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Figure 1.2: The building block approach applied to aerospace composite
structures.
Currently, the CFRP composite simulation under monotonic and quasi-static
loading has been well established in the literature. However, under cyclic or fatigue
failure is still considered as an open topic for further investigation. There is different
mode of failure that can be occur during the operational time such matrix yielding,
matrix cracking, fiber/matrix interface debonding, fiber pull-out, fiber fracture and
interface delamination. Interface delamination is one of the most critical issues face by
the CFRP composite laminates under fatigue loading since most of the structures have
a relatively weak ply-to-ply interface strength. Therefore, this thesis had described a
damage mechanics concept for predicting the fracture of CFRP composite laminates
under cyclic shear-induced loading conditions. Delamination under cyclic shear-
induced loading condition had been considered in this research due to the fact that
many composite structures are far more sensitive of being loaded in shear rather than
in tension [4-8]. Besides, Mode-II delamination in composite laminates is a major
matrix-controlled failure mode induced by out-of-plane flexural loading. The
methodology which had been described in this thesis will be a useful guidance to
predict a larger scale of specimen or structure using finite element software which had
been integrated with proper user-written subroutine.
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1.2 Background and Rationale
CFRP matrix composites belong to a new advanced material developed that are
strong, lightweight, low densities, and not easily corroded. In the transportation
industry, by reducing the vehicle weight will help to reduce the fuel consumption and
has become one of a main goal in this sector since the fuel price in this world are being
increased rapidly. This can be achieved by substitution of metal-based alloys,
commonly used in aerospace and automotive structures with lighter weight material
such as CFRP. The rapid increment in CFRP usage from 20% in 1990 (A320) to over
60% in 2010 (A380) for aircraft structures and high lift components is demonstrated
[3]. In military vehicles, composite structures have an advantage in stealth application
since this material are transparent to radar. Beside transportation industry, this CFRP
composite laminates have helped to develop construction sector for bridges and precast
concrete. In addition, CFRP composites offer flexibility in design through sequencing
of pre-impregnated laminates for tailored strength and stiffness properties in particular
loading direction. The relative low consolidation or curing temperature further lowers
the manufacturing cost of the part. The challenge is to consider the design tradeoffs in
choosing the lightest material that still meets the strength requirements of the part
while maintaining the cost effectiveness. The reduced through-life support cost of the
composite structures must also be considered. This is what the industry need which is
to evaluate the resistance of the material under various loading conditions such as
quasi-static, fatigue and impact loading with evolution of mechanical damage under
simulated service environment. This calls for evaluation of the material's resistance to
quasi-static, impact and fatigue loading along with mechanical damage evolution in
simulated service environment. Simple mechanical tests of the composite laminates
coupons, under various loading conditions such as compression, tension, shear and
flexural loading are beneficial to measure intrinsic properties of the composites [9-11].
Failure process is a very complex phenomenon for CFRP composite laminates.
Different mode of failure, such as matrix yielding, matrix cracking, fiber/matrix
interface debonding, fiber pull-out, fiber fracture and interface delamination could
happen during the operational time. Interface delamination is one of the most critical
issues face by the CFRP composite laminates due to the fact that most of composite
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