1
UNIVERSITA DI PISA
DEPARTMENT OF ENGINEERING
MASTER’S DEGREE COURSE IN AEROSPACE ENGINEERING
Experimental shear testing campaign and FEM
analysis for characterization of aerospace
composite laminates
GUIDANCE
AUTHOR
Prof. Ing. Mario Rosario Chiarelli
Kinattingalakath Arun Joshi
Ing. Roberta Lazzeri
2 TABLE OF CONTENTS LIST OF FIGURES 5 LIST OF TABLES 13 NOTATIONS 16 ABSTRACT 18 ACKNOWLEDGEMENT 19 1. INTRODUCTION 20 1.1. DEFINITION 20
1.2. SIGNIFICANCE OF THE MATRIX AND REINFORCEMENT IN
COMPOSITES 21
1.3. REINFORCEMENT FIBRE: TERMINOLOGIES 22
1.3.1. Fibre Forms 22
1.3.2. Roving 22
1.3.3. Unidirectional (Tape) 23
1.3.4. Bidirectional (Fabric) 23
1.3.5. Nonwoven (Knitted or Stitched) 24 1.4 REINFORCEMENT FIBRE: TYPES OF FIBRE 25
1.4.1. Fibreglass 26 1.4.2. Kevlar® 27 1.4.3. Carbon fibres 27 1.4.4. Ceramic Fibres 28 1.5 MATRICES 28 1.5.1. Polymer matrices 28 1.5.2. Thermoset 28 1.5.3. Thermoplastics 29
1.6 THE AEROSPACE STRUCTURES AND FEATURES 30 1.7 USE OF COMPOSITES IN AEROSPACE STRUCTURE 31 2. DESIGNING IN COMPOSITE STRUCTURES
2.1.The design procedure 36
2.1.1. Design concepts and parameters 38 2.1.2. Verification of composite structures 39
2.1.2.1.Building block Approach 39
2.2.Certification of Composites in Aeronautical Structures 40
2.3.Manufacturing 41
2.3.1. Manufacturing Methods 41
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Prepeg 43
Filament Winding 45
Pultrusion 46
3. MECHANICAL TESTING OF COMPOSITES 47
3.1.Test Campaign on fibre-reinforced composites and related execution
Problems 47
3.2.Tensile Testing 49
3.3.Compressive Testing 50
3.4.Shear Testing 50
3.4.1. Purpose of shear testing 51
3.4.2. Tensile test [± 45] s laminate 51
3.4.3. ASTM standard D3039/D3039M 51
3.4.4. ASTM standard D3518/D3518M 52
3.4.5. Shear Data Analysis 53
4. LABORATORY TESTING EQUIPMENT AND INSTALLATION 57
4.1.Digital Caliper 57
4.2.Torque Wrench 58
4.3.Ohmmeter 58
4.4.Soldering Iron 59
4.5.Servo Hydraulic Testing Machine and Frame 59
4.5.1. Load Cell 63
4.5.2. LVDT (Linear Variable Differential Transducer) 64
4.5.3. Grips 65
4.5.3.1. Screw Action Grips 66
4.6.Output Device 67
4.7.Conditioning 67
4.8.Strain measurement for coupon testing: extensometers 69
4.8.1 Strain Gauges 70
4.8.2 Strain Gauge measurement 70
4.9.Gauge Bonding 74
4.10. Tabbing of composite test specimens 76 5. FINITE ELEMENT ANALYSIS
5.1.FEA Using ANSYS 18.1 78
5.2.Materials and procedures 79
5.3.ACP (Ansys Composite Pre/Post) 79
5.4.Ansys Results 96
5.4.1. E 1.1 RTDRY 5H 3.1 96
5.4.2. E 1.1 RTDRY 5H 3.2 104
5.4.3. E 1.1 RTWET 5H 3.1 110
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5.4.5. E 1.4 120WET 5H 3.1 122
5.4.6. E 1.6 120WET 5H 3.2 128
5.5.Conclusion 134
6. FAILURE CRITERIA 135
6.1.Tsai-Wu failure criteria 136
7. COMPARISON OF SAMPLES
7.1.Comparison of samples 138
7.2.Testing two samples 138
7.2.1. Hypothesis of a common parent population 138 7.2.2. Selection of Level of Significance 138
7.2.3. Test Procedure 139
7.3.Results 140
7.3.1. Hypothesis 1: Samples have a common parent variance 145 7.3.2. Hypothesis 2: Samples have a common parent mean 146
7.4.Conclusion 148
REFERENCES 149
APPENDIX
A. RESULTS 151
A.1. E1 Specimen tests 152
A.1.1. RTDRY Specimens 154
A.1.1.1. Series RTD - 5H31 154
A.1.1.2. Series RTD – 5H32 159
A.1.2. RTWET Specimens 164
A.1.2.1. Series RTW – 5H31 164
A.1.2.2. Series RTW – 5H32 169
A.1.3. 120WET Specimens 174
A.1.3.1. Series 120W – 5H31 174
A.1.3.2. Series 120W – 5H32 177
A.2. Summary of Series 5H 180
A.3. Summary of Dry, Wet and 120Wet conditions (Series 5H) 182 A.4. Comparison of current thesis results with previous thesis 183 A.4.1. Summary of E1 RTW 5(8)H 2(3)1 series 184 A.4.2. Summary of E1 RTW 5(8)H 2(3)2 series 185 A.4.3. Summary of E1 120W 5(8)H 2(3)1 series 187 A.4.4. Summary of E1 120W 5(8)H 2(3)2 series 188
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LIST OF FIGURES
Figure 1.1: Composition of composites 21 Figure 1.2: General classification of composite materials 22
Figure 1.3: Tape and fabric products 23
Figure 1.4: Typical fabric weave styles 24 Figure 1.5: Nonwoven material (stitched) 25 Figure 1.6: Uses of composite materials in various sectors 31 Figure 1.7: Comparing materials properties – Young's modulus versus density
graph of various materials 33
Figure 1.8: Property variation of materials 34 Figure 1.9: Materials distribution (in percentage) in Boeing 787 aircraft 34 Figure 1.10: The growth of composite structure on major aircraft programs
(1975-2010) as a percentage of weight 35 Figure 2.1: Design process of composite structures 37 Figure 2.2: The building block approach (Campbell, 2010b, p. 500) 39
Figure 2.3: Manufacturing process 42
Figure 2.4: Design processes of the various composite. (Source: Avalon
consultancy service ltd.) 43
Figure 2.5: Schematic diagram illustrating the production of prepreg tape 44 Figure 2.6: Schematic representations of helical, circumferential, and polar
filament winding techniques 45
Figure 2.7: Principle of pultrusion 46
Figure 3.1.1: Reduced sample technique 48 Figure 3.1.2: Specimen selection methodology of current thesis 48
Figure 3.2: Test Coupon dimensions 50
Figure 3.3: Illustration of chord shear modulus and 2% offset strength 55
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Figure 4.2: Torque Wrench 58
Figure 4.3: Ohmmeter 58
Figure 4.4: Soldering Iron 59
Figure 4.5: Servo hydraulic testing machine. (Aerospace structure Laboratory,
University of Pisa, Italy) 61
Figure 4.5.1: Management Computer 62
Figure 4.5.2: Schematic diagram of entire parts of servo hydraulic machine 62
Figure 4.5.3: Load cell 63
Figure 4.5.4: LVDT (linear variable differential transducer) 64 Figure 4.5.5: Screw action grips on experiment setup 66 Figure 4.7: Humidity and temperature control chamber 68
Figure 4.8.1: Strain gauge parts 70
Figure 4.8.2: Wheatstone Bridge 71
Figure 4.8.3: Schematic diagram of amplified Wheatstone bridge 72 Figure 4.8.4: Catman professional software interface 73
Figure 4.9.1: Surface preparation 75
Figure 4.9.2: Gauge mounting 75
Figure 4.9.3: Lead wire attachment 75
Figure 4.10.1: Tabs designed in ANSYS 18.1 76 Figure 4.10.2: Catman professional force Vs time graph 77 Figure 5.1: Physical couplings in an anisotropic stress-strain relationship 83 Figure 5.2: Section plot of testing coupon 85 Figure 5.3: Workflow ansys composite pre/post (ACP) 85 Figure 5.4: Single layer Laminate deign in ANSYS 18.1 Design modeler 86 Figure 5.5: Mesh input details applied in ANSYS 18.1 Workbench 86
Figure 5.6: Mesh element quality 87
Figure 5.7: Meshing in detail 87
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Figure 5.9: Polar plot of in plane laminate engineering constant 88
Figure 5.10: Rosette definition 89
Figure 5.11: Oriented element sets definition 89 Figure 5.12: Stacking direction of Plies in one direction 90 Figure 5.13: Stacking direction of plies in both directions (Symmetric) 90 Figure 5.14: 8 Plies with [+45/-45/+45/-45]2S configuration in ACP 91
Figure 5.15: Static structural Analysis workflow diagram 92 Figure 5.16: Fixed boundary condition of testing specimen in ANSYS 18.1 93 Figure 5.17: Deformation along loading axis of E1.1 RTDRY 5H 3.1 specimen 97 Figure 5.18: Reaction force on E 1.1 RTDRY 5H 3.1 97 Figure 5.19: E 1.1 RTDRY 5H 3.1 Force of reaction components on three axes 98 Figure 5.20: Comparison of Force Vs displacement curve of
E 1.1 RTDRY 5H 3.1 Specimen 99
Figure 5.21: Initial time steps Force Vs displacement curve of
E 1.1 RTDRY 5H 3.1 Specimen 99
Figure 5.22: Equivalent (Von mises) Stress plot of E 1.1 RTDRY 5H 3.1 100 Figure 5.23: Minimum and Maximum equivalent stress variation of
E 1.1 RTDRY 5H 3.1 100
Figure 5.24: [+45/+45/+45/-45]2S Ply wise equivalent stress distribution of
E 1.1 RTDRY 5H 3.1 102
Figure 5.25: Shear stress distribution of E 1.1 RTDRY 5H 3.1 specimen 102 Figure 5.26: FEM plot, values and graphical representation of equivalent
elastic strain of E 1.1 RTDRY 5H 3.1 103 Figure 5.27: FEM plot, values and graphical representation of shear strain of
E 1.1 RTDRY 5H 3.1 104
Figure 5.28: Deformation along loading axis of E 1.1 RTDRY 5H 3.2 104 Figure 5.29: E 1.1. RTDRY 5H 3.2 force of reaction components on three axes 105
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Figure 5.30: Comparison of force Vs displacement curve of E 1.1 RTDRY 5H 3.2
Specimen 105
Figure 5.31: Initial time steps Force Vs displacement curve of
E 1.1 RTDRY 5H 3.2 specimen 106
Figure 5.32: Equivalent (von mises) stress plot of E 1.1 RTDRY 5H 3.2 106 Figure 5.33: Minimum and Maximum equivalent stress variation of
E 1.1 RTDRY 5H 3.2 106
Figure 5.34: [+45/-45/+45/-45]2S ply wise equivalent stress distribution of
E 1.1 RTDRY 5H 3.2 108
Figure 5.35: Shear stress distribution of E 1.1 RTDRY 5H 3.2 108 Figure 5.36: FEM plot, values and graphical representation of equivalent
elastic strain of E 1.1 RTDRY 5H 3.2 109 Figure 5.37: FEM plot, values and graphical representation of shear strain of
E 1.1 RTDRY 5H 3.2 110
Figure 5.38: Deformation along loading axis of E 1.1 RTWET 5H 3.1 110 Figure 5.39: E 1.1 RTWET 5H 3.1 Force of reaction components on three axes 110 Figure 5.40: Comparison of Force Vs displacement curve of
E 1.1 RTWET 5H 3.1 Specimen 111
Figure 5.41: Initial time steps of force Vs displacement curve of
E 1.1 RTWET 5H 3.1 Specimen 112
Figure 5.42: Equivalent (von mises) stress plot of E 1.1 RTWET 5H 3.1 112 Figure 5.43: Minimum and Maximum equivalent stress variation of
E 1.1 RTWET 5H 3.1 112
Figure 5.44: [+45/-45/+45/-45]2S ply wise equivalent stress distribution of
E 1.1 RTWET 5H 3.1 114
Figure 5.45: Shear stress distribution of E 1.1 RTWET 5H 3.1 specimen 114 Figure 5.46: FEM plot, Values and graphical representation of equivalent
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Figure 5.47: FEM plot, Values and graphical plot of shear strain of
E 1.1 RTWET 5H 3.1 116
Figure 5.48: Deformation along the loading axis of E 1.1 RTWET 5H 3.2 116 Figure 5.49: E 1.1 RTWET 5H 3.2 Force of reaction components on three axes 117 Figure 5.50: Comparison of Force Vs displacement curve of
E 1.1 RTWET 5H 3.2 specimen 117
Figure 5.51: Initial time steps of Force Vs displacement curve of
E 1.1 RTWET 5H 3.2 specimen 118
Figure 5.52: Equivalent (von mises) stress plot of E 1.1 RTWET 5H 3.2 118 Figure 5.53: Minimum and Maximum equivalent stress variation of
E 1.1 RTWET 5H 3.2 118
Figure 5.54: [+45/-45/+45/-45]2S ply wise equivalent stress distribution of
E 1.1 RTWET 5H 3.2 120
Figure 5.55: Shear stress distribution of E 1.1 RTWET 5H 3.2 specimen 120 Figure 5.56: FEM plot, Values and graphical representation of equivalent
elastic strain of E 1.1 RTWET 5H 3.2 121 Figure 5.57: FEM plot, Values and graphical representation of shear strain of
E 1.1 RTWET 5H 3.2 122
Figure 5.58: Deformation along loading axis of E 1.4 120WET 5H 3.1 122 Figure 5.59: E 1.4 120WET 5H 3.1 Force of reaction components on three axes 123 Figure 5.60: Comparison of Force Vs displacement curve of
E 1.4 120WET 5H 3.1 specimen 123
Figure 5.61: Initial time steps of force Vs displacement curve of
E 1.4 120WET 5H 3.1 specimen 124
Figure 5.62: Equivalent (von mises) stress plot of E 1.4 120WET 5H 3.1 124 Figure 5.63: Minimum and Maximum equivalent stress variation of
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Figure 5.64: [+45/-45/+45/-45]2S ply wise equivalent stress distribution of
E 1.4 120WET 5H 3.1 126
Figure 5.65: Shear stress distribution of E 1.4 120WET 5H 3.1 specimen 126 Figure 5.66: FEM plot, Values and graphical representation of equivalent
elastic strain of E 1.4 120WET 5H 3.1 127 Figure 5.67: FEM plot, Values and graphical plot of shear strain of
E 1.4 120WET 5H 3.1 128
Figure 5.68: Deformation along loading axis of E 1.6 120WET 5H 3.2 128 Figure 5.69: E 1.6 120WET 5H 3.2 Force of reaction components on three axes 129 Figure 5.70: Comparison of force Vs displacement curve of
E 1.6 120WET 5H 3.2 specimen 129
Figure 5.71: Initial time steps force Vs displacement curve of
E 1.6 120WET 5H 3.2 specimen 130
Figure 5.72: Equivalent (von mises) stress plot of E 1.6 120WET 5H 3.2 130 Figure 5.73: Minimum and Maximum equivalent stress variation of
E 1.6 120WET 5H 3.2 130
Figure 5.74: [+45/-45/+45/-45]2S ply wise equivalent stress distribution of
E 1.6 120WET 5H 3.2 132
Figure 5.75: Shear stress distribution of E 1.6 120WET 5H 3.2 132 Figure 5.76: FEM plot, Values and graphical representation of equivalent
elastic strain of E 1.6 120WET 5H 3.2 133 Figure 5.77: FEM plot, Values and graphical representation of shear strain of
E 1.6 120WET 5H 3.2 133
Figure 6.1: Confidence level displayed by world-wide failure exercise theories 135 Figure 6.2: Industrial uses of composite failure criteria 135 Figure 7.1: Normal distribution curve for 10% level of significance 141 Figure 7.2: Normal distribution curve for 5% level of significance 142 Figure 7.3: Normal distribution curve for 1% level of significance 143
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Figure 7.4: Normal distribution curve for mean tests for 10%, 5% and 1%
levels of significance 144
Figure A.1: Nomenclature of specimens 151 Figure A.2: Definition of Specimen and Material Axes 152 Figure A.1.1: E1 RTDRY 5H-31 Specimens – Shear Stress Vs Shear Strain curves 155 Figure A.1.2: E1 RTDRY 5H-31 Specimens – Shear Modulus Slope Curves
and Mean value 155
Figure A.1.3: E1 RTDRY 5H-31 Specimens – Shear Stress Vs Cross head
Displacement 156
Figure A.1.4: Specimens of E 1 – RTD 5H 3.1 Series 158 Figure A.1.5: E1 RTDRY 5H-32 Specimens – Shear Stress Vs Shear Strain curves 160 Figure A.1.6: E1 RTDRY 5H-32 Specimens – Shear Modulus Slope Curves and
Mean value 160
Figure A.1.7: E1 RTDRY 5H-32 Specimens – Shear Stress Vs Cross head displacement 161 Figure A.1.8: Specimens of E 1 – RTD 5H 3.2 Series 163 Figure A.1.9: E1 RTWET 5H-31 Specimens – Shear Stress Vs Shear Strain curves 165 Figure A.1.10: E1 RTWET 5H-31 Specimens – Shear Modulus Slope Curves and
Mean value 165
Figure A.1.11: E1 RTWET 5H-31 Specimens – Shear Stress Vs Cross head
Displacement 166
Figure A.1.12: Specimens of E 1 – RTW 5H 3.1 Series 168 Figure A.1.13: E1 RTWET 5H-32 Specimens – Shear Stress Vs Shear Strain curves 170 Figure A.1.14: E1 RTWET 5H-32 Specimens – Shear Modulus Slope Curves and
Mean value 170
Figure A.1.15: E1 RTWET 5H-32 Specimens – Shear Stress Vs Cross head
Displacement 171
Figure A.1.16: Specimens of E 1 – RTW 5H 3.2 Series 173 Figure A.1.17: E 1.4 – 120W 5H 3.1 Specimen 174
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Figure A.1.18: E1.4 120WET 5H-31 Specimen – Shear Stress Vs Shear Strain curve 175 Figure A.1.19: E1.4 120WET 5H-31 Specimen – Shear Modulus Slope Curve 175 Figure A.1.20: E1.4 120WET 5H-31 Specimen – Shear Stress Vs Cross head
Displacement 176
Figure A.1.21: E 1.6 – 120W 5H 3.2 Specimen 177 Figure A.1.22: E1 120WET 5H-32 Specimen – Shear Stress Vs Shear Strain curve 178 Figure A.1.23: E1 120WET 5H-32 Specimen – Shear Modulus Slope Curve 178 Figure A.1.24: E1 120WET 5H-32 Specimen – Shear Stress Vs Cross head
Displacement 179
Figure A.2.1: 0.2% Offset Tensile Strength summary of 5H Series 180 Figure A.2.2: Ultimate Load of 5H Series specimens 180 Figure A.2.3: Shear Chord Modulus (GPa) of 5H Series specimens 181 Figure A.3.1: Shear stress Vs Shear strain summary of Dry, Wet and 120Wet
Specimens in 3.1 panels 182
Figure A.3.2: Shear stress Vs Shear strain summary of Dry, Wet and 120Wet Specimens
in 3.2 panels 182
Figure A.4.1.1: Shear Chord Modulus summary of E1 RTW 5(8)H 2(3)1 series 184 Figure A.4.1.2: 0.2% Offset Tensile Strength summary of E1 RTW 5(8)H 2(3)1 series 184 Figure A.4.1.3: Ultimate Load summary of E1 RTW 5(8)H 2(3)1 series 185 Figure A.4.2.1: Shear Chord Modulus summary of E1 RTW 5(8)H 2(3)2 series 185 Figure A.4.2.2: 0.2% Offset Tensile Strength summary of E1 RTW 5(8)H 2(3)2 series 186 Figure A.4.2.3: Ultimate Load summary of E1 RTW 5(8)H 2(3)2 series 186 Figure A.4.3.1: Shear Chord Modulus summary of E1 120W 5(8)H 2(3)1 series 187 Figure A.4.3.2: 0.2% Offset Tensile Strength summary of E1 120W 5(8)H 2(3)1 series 187 Figure A.4.3.3: Ultimate Load summary of E1 120W 5(8)H 2(3)1 series 188 Figure A.4.4.1: Shear Chord Modulus summary of E1 120W 5(8)H 2(3)2 series 188 Figure A.4.4.2: 0.2% Offset Tensile Strength summary of E1 120W 5(8)H 2(3)2 series 189 Figure A.4.4.3: Ultimate Load summary of E1 120W 5(8)H 2(3)2 series 189
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LIST OF TABLES
Table 1.1: Advantages and disadvantages of reinforced fibres 25 Table 1.2: Properties of Composite Reinforcing Fibres 28 Table 1.3: Commonly used matrices in composites 29 Table 1.4: Features of aircraft structures 30 Table 2.1: Classification of manufacturing processes 42 Table 3.1: Standard dimension of Test specimen 52 Table 4.1: Characteristics of the servo-hydraulic machine 60 Table 4.2: Standard dimensions of tabs 76 Table 5.1: Mechanical properties of the testing coupon 92 Table 5.2: Mechanical strength values of the testing coupon 93 Table 5.3: Displacement input with respect to time of E 1.1 RTD 5H 3.1 specimen
for Ansys 18.1 analysis 94
Table 5.4: Displacement input with respect to time of E 1.1 RTD 5H 3.2 specimen
for Ansys 18.1 analysis 94
Table 5.5: Displacement input with respect to time of E 1.1 RTW 5H 3.1 specimen
for Ansys 18.1 analysis 94
Table 5.6: Displacement input with respect to time of E 1.1 RTW 5H 3.2 specimen
for Ansys 18.1 analysis 95
Table 5.7: Displacement input with respect to time of E 1.4 120W 5H 3.1 specimen
for Ansys 18.1 analysis 95
Table 5.8: Displacement input with respect to time of E 1.6 120W 5H 3.2 specimen
for Ansys 18.1 analysis 95
Table 5.9: Time steps input for FEM analysis of E 1.1 RTDRY 5H 3.1 96 Table 5.10: Analysis settings for FEM on all the specimens 97 Table 5.11: E 1.1 RTDRY 5H 3.1 Force of reaction Values from FEM analysis 98 Table 5.12: Time steps input for FEM analysis of E 1.1 RTDRY 5H 3.2 104
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Table 5.13: E 1.1 RTDRY 5H 3.2 force of reaction values from FEM analysis 105 Table 5.14: Time step input for FEM analysis of E 1.1 RTWET 5H 3.1 110 Table 5.15: E 1.1 RTWET 5H 3.1 Force of reaction values from FEM analysis 111 Table 5.16: Time steps input for FEM analysis of E 1.1 RTWET 5H 3.2 116 Table 5.17: E 1.1 RTWET 5H 3.2 Force of reaction values from FEM analysis 117 Table 5.18: Time steps input for FEM analysis of E 1.4 120WET 5H 3.1 122 Table 5.19: E 1.4 120WET 5H 3.1 Force of reaction values from FEM analysis 123 Table 5.20: Time steps input for FEM analysis of E 1.6 120WET 5H 3.2 128 Table 5.21: E 1.6 120WET 5H 3.2 force of reaction values from FEM analysis 129 Table 7.1: Ratio of estimated population variances based on shear chord modulus of
E1 RTD 5H 31 and 32 Specimens 145
Table 7.2: Ratio of estimated population variances based on 0.2% offset shear
strength of E1 RTD 5H 31 and 32 Specimens 145 Table 7.3: Ratio of estimated population variances based on shear chord modulus of
E1 RTW 5H 31 and 32 Specimens 145
Table 7.4: Ratio of estimated population variances based on 0.2%offset tensile
strength of E1 RTW 5H 31 and 32 Specimens 145 Table 7.5: Ratio of estimated population means based on shear chord modulus of
E1 RTD 5H 31 and 32 Specimens 146
Table 7.6: Ratio of estimated population means based on 0.2% offset shear
strength of E1 RTD 5H 31 and 32 Specimens 146 Table 7.7: Ratio of estimated population means based on shear chord modulus of
E1 RTW 5H 31 and 32 Specimens 146
Table 7.8: Ratio of estimated population means based on 0.2% offset shear
strength of E1 RTW 5H 31 and 32 Specimens 146 Table 7.9: Upper and Lower limit values of acceptance of 𝐹 under variance test 147 Table 7.10: Upper and Lower limit values of acceptance of 𝐹 under mean test 147 Table A.1.1: E1 RTD – 5H31 specimen series dimensions and gauge properties 155
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Table A.1.2: E1 RTD – 5H31 Test results 155 Table A.1.3: Mean, SD and COV of E1 RTD – 5H31 specimens 155 Table A.1.4: E1 RTD – 5H32 specimen series dimensions and gauge properties 159 Table A.1.5: E1 RTD – 5H32 Test results 159 Table A.1.6: Mean, SD and COV of E1 RTD – 5H32 specimens 159 Table A.1.7: E1 RTW – 5H31 specimen series dimensions and gauge properties 164 Table A.1.8: E1 RTW – 5H31 Test results 164 Table A.1.9: Mean, SD and COV of E1 RTW – 5H31 specimens 164 Table A.1.10: E1 RTW – 5H32 specimen series dimensions and gauge properties 169 Table A.1.11: E1 RTW – 5H32 Test results 169 Table A.1.12: Mean, SD and COV of E1 RTW – 5H32 specimens 169 Table A.1.13: E1.4 120W – 5H31 specimen dimensions and gauge properties 174 Table A.1.14: E1.4 120W – 5H31 Test results 174 Table A.1.15: E1.6 120W – 5H32 specimen dimensions and gauge properties 177 Table A.1.16: E1.6 120W – 5H32 Test results 177 Table A.4.1: Authors name and Specimen Series 183
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NOTATIONS
E Young’s modulus
𝜌 Density
𝜏 Maximum in plane shear stress in MPa 𝜏 Shear stress at 𝑖 data point in MPa
𝑃 Maximum load at or below 5% shear strain in N 𝑃 Load at 𝑖 data point, N
A Cross sectional of specimen defined by ASTM D3039 in 𝑚𝑚 𝛾 Shear strain at 𝑖 data point, 𝜇𝜀
𝜖 Longitudinal normal strain at 𝑖 data point, 𝜇𝜀 𝜖 Lateral normal strain at 𝑖 data point, 𝜇𝜀 𝛾 Maximum shear strain, 𝜇𝜀
𝐺 Shear chord modulus of elasticity, GPa
∆𝜏 Difference in applied shear stress between the two shear strain points, MPa.
∆𝛾 Difference between the two shear strain points 𝑥̅ The sample mean, Average
𝑆 Sample standard deviation
𝐶𝑉 Sample coefficient of variation, in %
𝑛 Number of specimens
𝑋 Measured or derived property
K Gauge factor
R Resistance, Ω.
𝑉 Excitation voltage or input voltage 𝑉 Output voltage of the bridge
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𝐹 Ratio of estimated population variances 𝜎 𝜎 𝐹 Limit of acceptance of 𝐹
𝑛 Number of observations in sample 1 𝑛 Number of observations in sample 2
𝑡( ) Relative deviation between means of two samples 𝑡( ) Limit of acceptance of 𝑡( )
𝑥̅ Mean of sample 1
𝑥̅ Mean of sample 2
𝛽 Level of significance
𝜎 Estimated standard deviation of population 1 𝜎 Estimated standard deviation of population 2
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ABSTRACT
The key reasons behind the continuous research regarding the material science in the aerospace industry are to find the better materials. These materials have suitable characteristics such as low density, good stiffness, good resistance to abrasion, impact and good corrosion resistance. The objective of this research is the discovery of new materials that offer specific qualities that differ from typical traditional metal alloys, ceramics and polymers. One specific kind of materials where these required appreciable properties can be found is called composites. The composite materials used in engineering are typically constituted by continuous fibres embedded in a matrix, whose function is to hold together the fibres that make up the resistant element. In nature there are many examples of this type of materials are visible. For example, in trees that are made of long cellulose fibres held together by lignin: just as reinforced concrete, created by the expert hands of man, is also defined as composite material. The main cause for the introduction of composite materials in aerospace engineering lies in the possibility of constructing very resistant, highly rigid and low weight structures.
This thesis dealt with the work of certification of the mechanical properties of materials used in aerospace, through a campaign of experimental tests in accordance with the guidelines to follow the regulations issued by the main certification bodies like ASTM. And finally, analyse the specimens which tested in the laboratory used by ANSYS software and verify the numerical and experimental accuracy of mechanical properties and characteristics. There are many other tests were carried out from Aerospace Structure Laboratory, University of Pisa on carbon fibre reinforced polymer composites. So, this thesis also performed a statistical comparison between the results obtained from current thesis with previous author’s thesis of shear tests on various batch/panel composites.
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ACKNOWLEDGEMENT
My endeavour stands incomplete without dedicating my sincere gratitude to everyone who has contributed a lot of guidance towards the successful completion of my thesis and masters course in space engineering. First of all, I would like to thank my parents for their blessings and constant support throughout my life. I am indebted to God Almighty for blessing me with his grace and taking my efforts to a successful culmination.
I deeply obliged to express my gratitude to Mr Mario Rosario Chiarelli, Associate professor, Department of Civil and Industrial Engineering, University of Pisa, Italy and Mrs. Roberta Lazzeri, Assistant professor, Department of Civil and Industrial Engineering, University of Pisa, Italy for both of your guidance during the entire thesis with patience, motivation, enthusiasm and immense knowledge. Despite the length of this thesis, you both were never hesitated to answer my questions, to help me out with numerical and experimental tests and analysis, and a number of other things. I am really thankful for your boundless support. Next, I would like to thank my fellow laboratory in-charges, Gabriele Adorni Fontana, Claudio Lippi, Luca Lombardi and Fabio Pioli for their stimulating discussions, help and useful comments. I never felt language as a barrier for communicating with all of you sirs.
I express my sincere thanks to all the professors in the Aerospace department at the University of Pisa. During the last three years, I have learned a lot regarding different subjects from my beloved professors.
I thank my fellow classmates Andrea Ferrara, Dimitri, Pasquale, Mauro and Dario for sharing your vast knowledge of space engineering and helping me to learn a lot about this new country and its culture. My hearty thanks also go to my colleagues Anuraj, Titto, Arjun, Lan, Lineesh, Avantika, Noemi, Helen, Rahul, Albie, Blesson, Pankaj, Harman, Giovanni, Ermes, Sreepali, Sai and abhishek for providing me such an appreciable mental and practical support in these last few years. You dears always hold my hands and brought me up when I was struggling.
I am really grateful to Mr Rafael Fernando Heitkoetter, Capt. Brazilian Airforce. You were my classmate but moreover my best friend, who helped me a lot during our masters course and also in our Space Structure project. I was really wondered about your practical knowledge in the space industry. Your family including Beatriz, Helizabat, Leo always gave me a support. I never too much missed my family because of them and will never forget about all the fun we all have had in the last three years.
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CHAPTER 1 1. INTRODUCTION
Fibre-reinforced polymer composite materials are among the most sort after materials for various industries in particular to that of the aerospace industry. In the recent years, the benefits of these materials have been proved relevant in various industrial sectors in particular to that of the aerospace industry. The progress has been elevated to an extent as such these materials are now considered to be the primary materials for the construction of spacecraft.
The materials range can be classified into various categories like Metals, Polymers, Ceramics and inorganic glasses and composites. Lack of strength of metals at elevated temperatures is a critical reason for their limited use in certain applications. High polymer materials, in general, can withstand to a greater extent in the lower temperature levels. Ceramics outperforms metals and polymers in their favourable melting points, ability to withstand high temperatures, strength and thermal expansion properties. But due to their brittleness, they are often deemed unsatisfactory as structural materials. This led to the concept of composites which would have the merits of the both while eliminating their respective disadvantages.
1.1 DEFINITION
Composite materials belong to those material system groups which are formed by the mixture or combination of two or more microcomponents which are mutually insoluble and dissimilar in form and/or material composition. Commonly there are two types of materials are concretized in composites i.e. a stronger one and the other a weaker one. The former material is commonly referred to as the reinforcement and the later material is generally treated as the matrix. Every structure that is made by utilizing composite materials requires in general, its own strength and rigidity which helps it to support as well as withstand the overall structural load imposed upon the structure, and this structural integrity is provided by the reinforcement component. Some of the commonly existing composites are steel reinforced concrete (metals + ceramics), vinyl-coated steel (metals + polymers), fibre reinforced plastics (ceramics + polymers).
The matrix in the composite material is the binder component that helps it to maintain the structural orientation and the shape of the reinforcement material. This component is comparatively more brittle than the reinforcement material.
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1.2 SIGNIFICANCE OF MATRIX & REINFORCEMENT MATERIALS IN A COMPOSITE:
The matrix yields the continuous phase of the composite. Its principal role is to give and maintain the required shape of the overall structure. Therefore, the matrix materials ought to be that of which that can be easily shaped and be able to maintain the overall structural orientation in an effective manner. It should also be noted that the matrix component of the composite is a primary component of a structure that interacts with the external forces that are being experienced upon the structure.
The composition of composites and its roles are pictorially described in figure 1.1,
Figure 1.1: Composition of composites (Advanced composite materials of the future in aerospace industry, Maria Mrazova), [3]
Various ongoing researches have opened up opportunities towards numerous stronger and stiffer reinforcements like carbon fibre materials. These advancements along with developments in polymer science have enabled us to produce various high performing resins as matrix materials which have equipped us to meet the various challenges that are posed by the complex nature of aircraft design requirements. The large-scale use of variously advanced composites in multiple developmental and manufacturing programs related to a broad set of advanced military fighter aircraft, varied civil transport aircraft, helicopters, satellites, launch vehicles and missiles all around the world is perhaps the most evident example of the numerous positive aspects of such composite materials.
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General classification of composite materials is,
Figure 1.2: General classification of composite materials, [2]
1.3 REINFORCEMENT FIBRE: TERMINOLOGIES
This section discusses about the various terms as well as their definitions which are used with regard to composites.
1.3.1 Fibre Forms:
Primarily, almost all product formation begins with spooled unidirectional raw fibres packaged as continuous strands. An individual fibre is called a filament. The word strand is also used to identify an individual glass fibre. Bundles of filaments are identified as tows, yarns, or rovings. Yarns like fibreglass are twisted, while Kevlar® yarns are not. Tows and rovings do not have any twist. Most fibres are available as dry fibre that needs to be impregnated (impreg) with a resin before use or prepreg materials where the resin is already applied to the fibre.
1.3.2 Roving
A roving can be defined as an individual grouping of fibre ends or filaments, such as 20-end or 60-end glass rovings. All these filaments are aligned in the same direction without any twist. For example, carbon rovings are usually identified as 3000, 6000, or 12000 rovings. Mandrels are utilized for filament winding which leads to curing of resin to a final configuration for most of the rovings products applications.
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1.3.3 Unidirectional (Tape)
For many years, in the aerospace industry unidirectional prepreg tapes have been treated as a typical form, along with thermosetting resins impregnated fibres. The most frequent manufacturing method for tape is to draw parallel raw (dry) strands into the machine which used for impregnation where we applied heat and pressure to combine hot melted resins with the strands. Unidirectional products have high strength in the direction of fibres and negligible or virtually no strength across the fibres. Compared to woven fabrics, unidirectional (tapes) shows higher strength, [1].
Figure 1.3: Tape and fabric products (Advanced composite materials eBook, chapter 7), [1]
1.3.4 Bidirectional (Fabric)
It is to be noted that varied fabric constructions provide much more ease and flexibility to facilitate complex shapes than what a straight unidirectional tape could offer. Fabrics enable the possibility of resin impregnation either through a process like hot melt or by that of a solution. Usually, fabrics used for structural needs mostly use like fibres or strands of the similar weight or yield in both the warp (longitudinal) and fill (transverse) directions. With regard to aerospace structures, tight woven fabrics are mostly chosen to save reducing resin void size, weight and withholding the fibre orientation during the fabrication process.
Woven structural fabrics are usually created with reinforcement tows, strands, or yarns interlocking upon themselves with over/under placement during the weaving procedure. The most common fabric styles are prevalent is that o the plain or satin weaves, [1].
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Figure 1.4: Typical fabric weave styles (Advanced composite materials eBook, chapter 7), [1]
1.3.5 Nonwoven (Knitted or Stitched)
Knitted or stitched fabrics can provide similar mechanical advantages as that of the unidirectional tapes. Fibre placement can exist either as straight or unidirectional without the need of any over/under turns of woven fabrics. The fibres are held together by stitching using fine yarns or threads after pre-selected orientations of one or multiple layers of dry plies. These kinds of fabrics provide a wide range of multi-ply orientations. Although there may have been some added weight penalties or loss of some substantial reinforcement fibre properties, some gain with respect to that of interlaminar shear and toughness properties may be obtained. Some of the most common stitching yarns that are currently in use are polyester, aramid, or thermoplastics, [1].
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Figure 1.5: Nonwoven material (stitched), (Advanced composite materials eBook, chapter 7), [1]
1.4 REINFORCEMENT FIBRE: TYPES OF FIBRE
Carbon fibres have high strength and stiffness properties. The same is found only in boron and carbon that have similar combined properties; the other types of fibres usually have either high strength or high stiffness properties.
Table 1.1 shows advantages and disadvantages of some common fibres.
Fibre Advantages Disadvantages
E-glass, S- glass High strength Low stiffness Low cost Short fatigue life
High-temperature sensitivity Aramid (Kevlar) High tensile strength Low compressive strength
Low density High moisture absorption
Boron High stiffness High cost
High compressive strength
Carbon (AS4, T300, C6000) High strength Moderately high cost High stiffness
Graphite (GY-70, Pitch) Very high stiffness Low strength High cost Ceramic (Silicon carbide,
Alumina) High stiffness High use temperature Low strength High cost
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The fibres can be of both short as well as that of a continuous one. Short fibres, also termed as discontinuous fibres, can be either all oriented along the single direction or in random directions. When it comes to stiffness and strength continuous fibres composites are the most capable and efficient one. The reinforcement in those composites consists of long continuous fibres, where the configuration can be of a parallel manner, oriented at right angles to each other, or oriented along various angle directions.
In the perspective of materials, it can be either isotropic or anisotropic. An isotropic material has the same properties in all directions. On the other hand, the anisotropic material has different properties in all directions at a given point in the material. Bulk material like metal and polymers are normally considered as isotropic material, while composites itself is the best example for exhibiting anisotropic properties. A composite material is predominantly anisotropic if the fibres are oriented along one direction.
Fibre composites commonly consist of multiple layers, and the fibres in continuous-fibre composites laminated materials are usually oriented along those directions that will improve the strength in the primary load direction. Unidirectional laminates fibre reinforced polymers have a fibre direction of 0° and is very strong and stiff along this direction but are significantly weak in the perpendicular direction as the load must be carried by the polymer matrix that is much weaker. For achieving specific stiffness and values of strength by rotating the fibre orientation angles. The orientation of the fibres in the matrix is the factor that primarily influences the fibre performance characteristics. The most efficient utilisation is achieved when the fibres are placed along the direction of the load, even a small variation can reduce the strength and stiffness of the composite. The way of placing the fibres work for some structures, but it is usually necessary to place them in varied directions, e.g. 0°, +45°, -45°, 90°, to balance the load-carrying capacity. When a laminate has a similar number of layers in the 0°, 90°, 45°, -45° directions it is called as a quasi-isotropic laminate. It carries loads in all four directions and is, therefore, the most preferred orientation. There are also other factors that influence the performance of the composite, such as fibre volume fraction, the stiffness of the matrix, damage tolerance, single ply thickness, voids, fibre matrix interface, moisture & media, temperature, holes and cut-outs.
1.4.1 Fiberglass
Fiberglass is primarily used for the secondary structure on aircraft, such as fairings, radomes, and wing tips. Fiberglass is also used for the construction of helicopter rotor blades. There exist various types of fiberglass that are used in the aerospace industry. Electrical glass, or E-glass, is considered as such for various electrical applications. It exhibits a high resistance to the current flow. E-glass is constructed from borosilicate glass. S-glass and S2-glass exhibits structural fiberglass that have a higher strength than E-glass. S-glass is obtained from magnesia-alumina-silicate. Advantages of fiberglass are cheaper cost than other composite materials, chemical or galvanic corrosion
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resistance, and electrical properties (fiberglass is a non-conductor of electricity). Fiberglass is white in colour and is available as a dry fibre fabric or pre-preg material, [1].
1.4.2 Kevlar®
Kevlar® is DuPont’s given name for classifying aramid fibres. Aramid fibres have characteristics of being light weight, strong, and tough. There exist two types of Aramid fibres that are used in the aviation industry. Kevlar® 49 has the advantage of high stiffness and Kevlar® 29 has a low stiffness value. High resistance to impact damages make Kevlar a unique selection in the aerospace industry. So, they are often utilized in those areas that are prone to impact damage. The main demerit of the aramid fibres is their general weakness in compression and hygroscope. Service reports are suggestive that some parts made from Kevlar® absorb up to 8 percent of their weight in water. Therefore, parts constructed using aramid fibres needs to be protected from the environment. A similar disadvantage is that Kevlar® is difficult to drill and cut. The fibres fuzz easily, and special scissors are required to cut the material, [1].
1.4.3 Carbon fibres
The properties of carbon fibres are similar to that of steel, but with lesser density. Carbon fibres are usually combined with a polymer matrix and are then often referred to as CFRP or carbon fibre reinforced plastics. It is possible to obtain a high-performing material with a significant weight reduction of over 50 per cent compared to high strength steel by utilising CFRP.The better fatigue and creep resistance make CFRP a special material, and by using laminate orientation the material can be designed to be more tougher and damage tolerant than other metals. The chemical and corrosion resistance are also good with comparison to other metals, CFRP has exhibits unique properties of dimensional stability and vibration damping. Also, this material shows lower electrical resistivity as well as a higher thermal conductivity. On the same matter, composites based on carbon fibres have low energy-absorbing capabilities,low resistance to transverse impact loading and the plies tend to separate from each other in the laminate.
Carbon fibres and graphite are similar terminologies that are used to describe carbon fibres. The major difference between carbon fibres and graphite is the content of carbon in them. Carbon fibres is composed of typically 95% of carbon, and are carbonised at about 1000 to 1500°C, while graphite fibres contain about 99% of carbon, and are first carbonised and then graphitised at temperature that ranges between 2000°C and 3000°C.
As seen in Table 1.2, the fibres used in modern composites have strength and stiffness factors far above than those of traditional bulk materials. The high strengths of the glass fibres are due to processing that eliminates both the internal or any surface flaws which usually weakens glass, and the strength and stiffness of the polymeric aramid fibre is a result of the near perfect alignment of the molecular chains within the fibre axis.
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Table 1.2:Properties of Composite Reinforcing Fibres (F.P. Gerstle, “Composites,” Encyclopedia of Polymer Science and Engineering, Wiley, New York, 1991), [4]
1.4.4 Ceramic Fibres
Ceramic fibres are used specifically for high-temperature applications, such as that of the turbine blades in a gas turbine engine. The ceramic fibres can be used to temperatures up to 1,205 °C.
1.5 MATRICES
The matrix maintains the position as well as the orientation of the individual fibres and protects them from any possible degrading environments. In polymer and metal composites, the matrix transmitting loads from the matrix to the fibres occurs through the shear loading at the interface. Matrices also exist in various form as that of the fibres. A few of them has been briefly described below,
1.5.1 Polymer matrices
Polymers have low strength and stiffness values, but by combining polymers with fibres it is possible to obtain a higher strength and stiffness than that of the polymer. Due to high strength and low density, fibre-reinforced polymers are the most dominating among various polymer structural materials. Polymers are divided into two different categories, thermoset or thermoplastic depending upon on the matrix.
1.5.2Thermoset
Thermosetting matrices are the most popular matrix system used for composite materials. The material has good mechanical properties while operating in hot and moist environments. It is easy to process and have good adhesive features with many fibres and have a relatively low material cost. These types of plastics are impossible to reshape as they are rigid cross-linked material. So, at higher temperatures it gets degraded rather than to melt. Before curing, the thermoset resin has low viscosity that allows easy impregnation of the fibre. When the curing is done in the material,
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then there is no longer the possibility for melting. There are different types of thermosetting matrices; some of them are detailed down below:
Polyester Epoxy Vinyl Ester Polyimide Phenolic resins 1.5.3 Thermoplastics
Thermoplastics can be reshaped as they become soft at high temperatures. This is however only possible in a very limited number of times, as multiple reprocessing can degrade the respective resin. Many thermoplastic composites (e.g. carbon-reinforced PEEK) have good resistance to impact loading and are therefore apt for use in high performance engineering applications. Because of thermoplastic’s greater ductility and processing speed compared to thermosetting matrices, and the fact that thermoplastics can withstand very high temperature levels, the material becomes ever more important. The reason for higher processing speed is because thermoplastics softens quickly when heated above a given temperature, and the soften material is easy to shape.
There are several thermoplastics that are used as matrices, few of them are described below, Polyetheretherketone (PEEK)
Polyetherimide (PEI)
Polyphenylene sulphide (PPS) Polyamide (PA)
The below table 1.3 shows properties of some matrices those are commonly used in modern composite industry.
Polymer Density (g cm-3) Young’s
Modulus (GPa) Tensile strength (MPa) Failure strain (%) Thermosets Epoxy resins 1.1 - 1.4 3 - 6 35 - 100 1 - 6 Polyesters 1.2 – 1.5 2.0 – 4.5 40 – 90 2 Thermoplastics Nylon 6.6 1.14 1.4 – 2.8 60 – 70 40 - 80 Polypropylene 0.9 1.0 – 1.4 20 – 40 300 PEEK 1.26 – 1.32 3.6 170 50
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1.6 THE AEROSPACE STRUCTURES AND FEATURES
The most important requirements of an aerospace structural component and their effect on the overall design of the structure are presented below table 1.4, [2].
Table 1.4: Features of aircraft structures (Applications of composite materials in aerospace, International journal of science technology and management Vol.No.4, November 2015.), [2]
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Further, the structure has to meet certain requirements of fuel sealing and provide access for easy maintenance of equipments. Passenger carriage usually requires safety standards that are to be followed and these put special criteria of fire-retardance and crash-worthiness on the materials and the related designs used. In respect to a spacecraft, the space environment–vacuum, radiation and thermal cycling must be considered, and specifically developed materials are required for ensuring their durability.
Two key developments in the scientific & technological world have had a tremendous influence on the present generation and satisfaction of the demands required by the aerospace community: one, the advances in the computational power and the other, composites technology utilizing fibre reinforced polymeric materials.
1.7 USE OF COMPOSITES IN AEROSPACE STRUCTURE
It is to be realized that in order to meet the demands of various sectors, it is necessary to have materials with a peculiar property-set. The use of composites has been motivated largely by such considerations.
Over the past year, there have been several changes in the application areas for carbon fibre. Aerospace & defence applications have grown significantly and are now the largest consumers of carbon fibre – 13,900 t or 30% based on a total of 46,500 t. An approximate interpretation is given below graphically. (Fig. 1.6).
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The composites offer several of these features as given below:
1. Light-weight: Composites are light in weight, compared to most metals. Their lightness is important in aircraft, where less weight means better fuel efficiency (more miles to the gallon). This light weight due to high specific strength and stiffness
2. Fatigue-resistance and corrosion resistance: Composites resist damage from the weather and from harsh chemicals that can eat away at other materials. Outdoors, they stand up to severe weather and wide changes in temperature.
3. The capability of a high degree of optimization: tailoring the directional strength and stiffness
4. Capability to mould large complex shapes in small cycle time reducing part count and assembly times: Good for thin-walled or generously curved construction. This gives designers the freedom to create almost any shape or form.
5. The possibility of low dielectric loss in radar transparency: Radar signals pass right through composites, a property that makes composites ideal materials for use anywhere radar equipment is operating, whether on the ground or in the air. Composites play a key role in stealth aircraft, such as the U.S. Air Force’s B-2 stealth bomber, which is nearly invisible to radar.
6. The possibility of achieving low radar cross-section.
7. Durable: Structures made of composites have a long life and need little maintenance. We do not know how long composites last, because we have not come to the end of the life of many original composites. Many composites have been in service for half a century. It is capable to maintain dimensional and alignment stability in the space environment.
These composites also have some inherent weaknesses:
1. Laminated structure with weak interfaces: poor resistance to out-of-plane tensile loads 2. Susceptibility to impact-damage and the strong possibility of internal damage going
unnoticed
3. Moisture absorption and consequent degradation of high-temperature performance 4. The multiplicity of possible manufacturing defects and variability in material properties. A common reason for the use of composites is that they deliver outstanding mechanical properties at a lower weight when compared to conventional materials – for example, to the aluminium alloys used in aerospace. This is illustrated in Figure 1. Such property comparisons are essential when selecting materials or, conversely, when marketing materials for particular applications. These comparisons require the right property data – not just about technical properties, but about economic and environmental properties, and process history. There is also a need for tools to analyse this data and present the results – for example, the CES Selector™ software2 used to create Figure 1.7 and 1.8.
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Properties variation of a various selection of matrices in composites and other materials are shown below,
Figure 1.7: Comparing materials properties – Young's modulus versus density graph of various materials. composites provide high strength at low density (materials property chart created
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Figure 1.8: Property variation of materials. (Source: Avalon consultancy service ltd.), [12]
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Fibre composites have been used in an ever-increasing percentage of jet airliner structures for several decades (Figure 1.9). Boeing began using composites over 30 years ago in 737 spoilers; composites have now replaced light alloys to create significantly lighter and lower-maintenance control surfaces and empennages in the - 5 - 737 Classic (-300, -400 and -500) and Next Generation (-600, -700, -800 and -900) models, 757, 767 and 777 product lines. The 787 Dreamliner is a defining aircraft in the use of fibre composites – it will be the first airliner that is primarily composite, with a fully composite skin, fuselage, wing box and empennage (Werfelman 2007). This is a quantum leap when compared to the current generation of airliners containing composite components (the Boeing 777, which is 9% composite by weight) (Sater, Lesieutre & Martin 2006). Such a large increase in composite use brings numerous production and safety challenges to aircraft manufacturers.
Black: civil aircraft, Blue: military aircraft
Figure 1.10: The growth of composite structure on major aircraft programs (1975-2010) as a percentage of weight. (Source: Teal Group, Boeing, Airbus, Composite Market Reports.), [12]
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CHAPTER 2:
COMPOSITE STRUCTURES DESIGNING
2.1 The design procedure
Several choices need to be made when designing composite materials, which both raise the freedom of designing as well as make the process more complex. The matrix and fibres, fibre orientation and the lay-up have to be determined, and a properly suited manufacturing method must be selected. Compared to a structural design which based on static strength and stiffness, are much more complex to design than composites based on fatigue. The reason behind this uncertainty is that the material’s properties are alter during loading. The property variation with respect to loading is not linear; its rate of variation depends on how far the material is from failure, i.e. the loading conditions and instantaneous material status.
One of the most crucial causes of structural failures in composite materials are fatigue and related phenomena and those are therefore necessary to consider, but these are highly dependent on the critical design parameters and the type of load which are applied on the specimen. Factors such as the fibre orientation, type of material, loading etc. are more common in composite materials design than the effects like fatigue. The conclusion that has been made from this concept is that fatigue is important to have in mind when designing, but other factors are more critical and are therefore necessary to prioritize. The focus in the design development and procedures will therefore not lie solely on fatigue, but always give an attention on all those issues which is identified as critical, and how those may affect the design in composite materials.
The design process for structures in CFRP has been divided into four major steps: 1. material selection
2. manufacturing method 3. material design 4. structural design
and these four steps are necessary to carry out to design and develop composite parts.
At the point when the design is set, it is costly to return and change it, and the greater part of the expenses are set in the principal phases of the process. Preparations, before the last plan begins, is accordingly imperative. It is relatively difficult to settle one stage before beginning the following. The chose material chooses which manufacturing process need to utilize, and the multifaceted nature of the design, in the meantime as the manufacturing procedure influences those different other steps. The information gathered in the variety steps needs, therefore, to be utilized as a part of alternate steps in the entire designing stage. It is, notwithstanding, conceivable that one parameter is set from the earliest starting point, e.g. a manufacturing technique is utilized inside the organization and is an appropriate strategy to keep on using. At that point the manufacturing
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procedure must be utilized as an input to the whole desigining procedure itself. Figure 2.1 shows one way to illustrate the process.
Figure 2.1: Design process of composite structures, [31]
The prerequisites on the products should be set before beginning the design technique and to know the parameters that influence the design process. The following factors are essentially to determine:
1. The type of loading, e.g. axial, bending, torsion or a combination 2. Mode of loading, e.g. static, fatigue, impact, etc.
3. Service life
4. Operating or service environment, e.g. temperature, humidity, the presence of chemicals etc.
5. Other structures or components which the considered design is required to interact with 6. Manufacturing processes that can be used to produce the structure or component.
7. Cost, both material cost and the costs to transform the selected material into a final product, e.g. manufacturing, machining and assembly costs.
It is important to calculate which stresses and strains that will be excerted to the composite, and what it ought to be intended to withstand. It should be possible by using design allowables and safety factors to get the ultimate design limit; it is the breaking point the composite ought to be planned by.
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Equation: 1
Design allowables are the point of confinement of stress, strain or stiffness that is allowed for a particular material, design, application and ecological condition. Another technique for designing additional strength into a composite structure is by applying a margin of safety. At the point when the safety factor has been applied to the design load and the laminate is intended to withstand the loading condition and additional load carrying capacity and it is seen as a margin of safety.
2.1.1 Design concepts and parameters
There are at least two alternative design concepts for anticipating the fatigue life for structural components made by composite material: damage tolerant (or fail-safe) and safe-life design concepts. For damage tolerant concepts it is assumed that a damage metric, e.g. crack length, delamination area, or residual strength or stiffness, can be correlated to fatigue life via a valid criterion. The damage is permitted if it is not critical and cannot prompt sudden failure. Structures based on safe-life design can operate if no measurable cracks are initiated. Cyclic stress or strain in safe-life designs is straightforwardly associated with operational life by means of the S-N or ε-N curves.
When designing in fibre-reinforced composite materials usually to use the same design criteria as for metals. The accompanying criteria are used for designing essential structural parts in an aeroplane, whether if it is produced using aluminium or CFRP:
1.The structure must withstand the ultimate design load in static testing. 2.The fatigue life must be equivalent to or surpass the projected vehicle life.
3.Deformations that outcome from the applied cyclic loadings and limit design shall not meddle with the mechanical operation of the aeroplane, unfavourably influence its aerodynamic characteristics, or require repair or substitution of parts.
Other critical parameters when designing in composite materials are crashworthiness and durability. Crashworthiness is particularly essential when designing for the automotive industry. If composites are properly designed the material can offer great crash performance, where the specific energy absorption are superior to for metals. The durability effects of cyclic loads is likely to be disregarded if durability is assessed on the basis of static strength calculations.
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2.1.2 Verification of composite structures
There are a few different ways to verify that a composite structure can oppose the applied loads and parameters that may influence it. Design allowables and safety factors are both used to guarantee that the composite would not fail. High coefficients of safety are generally used when designing composite structures, due to the difficulty to precisely model the material behaviour. With exception of strategies that already are used, composite structures are over-dimensioned due to the stochastic nature of fatigue loadings. It is important to describe each material and to appropriate model the quasi-isotropic and fatigue behaviour to develop theories that help the design process.
2.1.2.1 Building block Approach
The building block approach can be used to ensure that the component and assembly can withstand the applied loads and forces, Figure 2.2. The approach is based on several levels including different numbers of tests excessively complex structures. In the first levels, large number of simple specimens are tested, and the tests become lesser and the structures get more complicated for each level. The top level generally includes testing of the final design. The fundamental material properties are determined on the lower levels by using large numbers of specimens from several batches of material. The collected data from the lower levels are then used to anticipate failure modes and loads on higher levels. The approach gives the chance to create and refine the tooling and processing approaches that later can be used in the production. The test method is costly, as many tests need to be done, however, it has been effective for design and builds e.g. aeroplane systems and frameworks, where safety is extremely significant.
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2.2 Certification of Composites in Aeronautical Structures
As far as the necessity required by the aeronautical standards are concerned, there are no differences for any materials used in an aeronautical structure. In certification there are contrasts, in certainty this procedure with a composite structure is much more complex than for the metallic solution for the same structure. It is mainly due to the fact that the composites have an extensive range of properties and the fact that a complete acquirement of the design techniques has not yet been achieved. However, structures made of composite materials must guarantee the same safety as the respective metal solutions.
In 1978 the Federal Aviation Administration (FAA) issued the advisory circular AC20-107 on the certification of aeronautical structures in composite material (on the same topic, the FAA issued two subsequent advisory circulars, the AC20-107A in 1984 and the AC20-107B in the 2009 any legislation issued repeals the antecedents). This is a short document specifying that the composite design must achieve a level of safety at least equal to that required by metal structures. The advisory circular emphasizes the need to determine the mechanic properties of the material taken into consideration through the conduct of targeted experimental tests; these tests must be carried out in climatic and environmental conditions as close as possible to the operative ones, [11]. The tests typically required for the certification of a composite structure are the following:
Static test, in which the structure is subjected to 150% of the Design Limit Loading (DLL), ie, to the ultimate load.
Fatigue tests on primary structures
Damage tolerance compliance and impact resistance on primary structures.
Civil and military aircrafts are certified according to separate procedures. Certification takes place by certifying authorities, which have national or international dimensions. The regulations issued by the US institutions are of greater importance and are considered the point of reference for all the other authorities.
Civil Aircraft: Federal regulations require all aircraft operating in US territory to receive a certificate of airworthiness. This certificate is issued if the FAA determines that the aircraft in question was constructed in accordance with regulatory requirements,
Inspection, maintenance and repair operations allow the aircraft to maintain airworthiness status, [12].
The regulations of interest to aircraft manufactures are as follows:
a) FAR 23 - Normal, utility, acrobatic and commuter category airplanes b) FAR 25 - Transport category airplanes
c) FAR 27 - Normal category rotorcraft d) FAR 29 – Transport category rotorcraft
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Military aircraft: The certification of composite on aeronautical structures for military aircraft requires the fulfilment of the specifications contained mainly in the following documents:
a) Military specifications MIL-A-8860A and MIL-A-8870A; b) Military specifications MIL-A-8860B and MIL-A-8870B;
c) Military specifications MIL-A-87221, General specification for aircraft structures;
d) Military specifications MIL-STD-1530A – Aircraft Structural Integrity Program, Airplane requirements;
e) Joint Service Specification Guide JSSG-2006 (1998) – Aircraft Structures;
The certification of a military aeronautical structure is a process characterized by a continuous debate between the user and the manufacturer and is based mainly on the regulations listed above, [12].
From the above, we can understand the importance of performing experimental tests for the certification of materials and of production processes in order to create an aircraft that meets the requirements imposed by the certifying authorities.
2.3 Manufacturing
The manufacturing process is either decided in the requirements and works as an input to the designing procedures or is chosen in parallel with the alternate stages. The selected materials are often an essential factor of which procedure to choose, as all materials are not appropriate for all procedures.
The manufacturing method influences the composite’s mechanical properties and its performance. It is therefore vital to choose a manufacturing method that can ensure the quality of the component. Damages like wrinkles and voids are generally caused by the manufacturing process and can act as sites for failure due to fatigue. The selection of a perfect manufacturing method to use is also dependent on the material that will be used, as some procedures are not appropriate for all types of composite material, e.g. compression moulding is mainly used for thermoplastic resins and not thermoset. The manufacturing volume and complexity of the structures are other variables to consider.
2.3.1 Manufacturing Methods
The entire procedure for manufacturing a structure with composite materials consists of four major steps, Figure 2.3. These steps are:
1. Fibre production.
2. The pre-formed material in shape of prepregs or dry fibre performs are manufactured, this step is not necessary to do before manufacturing the composite, as matrix and fibres can be impregnated during the manufacturing.
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3. Manufacturing of composite parts, here is the fibres reinforcing the matrix.
4. Parts are assembled to create a whole product. It is also in this step the parts are processed in different ways, e.g. surface treatment.
Figure 2.3: Manufacturing process
The manufacturing process can be divided into three major steps. Fibre placement, application of matrix and curing. Different types of methods used for manufacturing in each step are given below.
Fiber Placement Matrix Application Curing Prepreg
Dry fibre preforms
Injection moulding Filament winding Pultrusion
Vacuum infusion
Resin film infusion Hand lay-up & spray-up
Resin Transfer Molding (RTM) Vacuum Assisted RTM (VARTM) Vacuum Assisted Processes (VAP) Compression moulding
Autoclave curing Microwave oven Convection oven
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A wealth of design process of various composites listed below,
Figure 2.4: Design processes of the various composite. (Source: Avalon consultancy service ltd.) 2.3.1.1 Fibre placement
There are several techniques for placing the fibres they can, for example, be woven into weaves and pre-impregnated with the matrix, so-called prepregs that shall be combined with other manufacturing methods. Dry fibre preforms and the prepregs are placing the fibres. Injection moulding and filament winding are also a typical method for the placement of matrix, and pultrusion consists of all three steps.
PREPREG:
Prepreg is the composite industry’s term for continuous fibre reinforcement pre-impregnated with a polymer resin that is only partially cured. This material is conveyed in tape form to the manufacturer, who then directly moulds and fully cures the product without adding any resin. It is probably the composite material form most widely used for structural applications. The prepregging method for thermoset polymers are represented schematically in Figure 2.5. It begins by collimated release of a series of spool-wound continuous fibre tows. Using heated rollers these tows are then sandwiched and pressed between sheets of release and carrier paper. And this process termed as “calendaring.” The release paper sheet has been covered with a thin film of relatively low viscosity heated resin solution to provide for its thorough impregnation of the fibres. A “doctor blade” distributes the resin into a film of uniform thickness and width. The final prepreg product— the thin tape consisting of continuous and aligned fibres inserted in a partially cured resin—is
