The Influence of Orthotropy and Taper Angle on the
Compressive Strengths of Composite Laminates with
Scarfed Holes
Mitesh Patel, Stephen Daynes, Chun H. Wang
Sir Lawrence Wackett Aerospace Centre, RMIT University
11th International Fatigue Congress
2-7 March 2014, Melbourne, Australia
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Contents
• Context and motivation
• aims
• Methodologies
– Experiments
– Finite Element Modelling
• Results and analysis
• Conclusions
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Percentage of weight
Composite material usage trend in Aviation Industry
The growth of composite structure on major aircraft programs1
1Taylor, R. P. "Fibre composite aircraft - capability and safety." Australian Transport Safety Bureau, 2008, p. 6.
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Composite structure damage
Ethiopian Airlines Boeing 787 Dreamliner jet after fire incidence at
London’s Heathrow Airport, July 20132
2Source: BBC, http://www.bbc.co.uk/news/uk-23295109
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Metal repairs
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Composite repairs
• Bonded repairs works for
secondary load carrying
components (low strength)
Bonded repairs
• However, these methods do not
work to gain high strength (due to
low bond strength)
• Classical Laminate Theory will
work only for stiffness, not
strength, nor stress-strain curves
3D scarf
• Aerodynamic issue –flow
disturbance
2D scarf
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Scarf repair concept
Requirements to meet
DLL – scarfed panel
DUL – repaired structure
a) Schematic representation of Scarf Angle and Scarf
Strength with respect to the DLL and DUL
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b) Schematic of aircraft skin with scarf
repair patch and doubler.
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Importance of the research
• Requirement of Non-Quasi-isotropic Laminate Analysis
– Leads to new stacking sequences from traditional [0/45/90]xs
• Classical Laminate Theory – Doesn’t work for Strength nor stress-strain
curves
• Hence, New method is required
• NO data available for scarf angle influence
• NO data available for orthotropy influence (scarf geometry)
– Hence, different stacking sequences are considered
Composite Loading
• Static Testing
– Tension
– Compression
• Fatigue Testing
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Specimen design
a) Open Hole Compression (OHC)
Un-notched
Hole
Diameter
d [mm]
Scarf
Angle θ
[Degree]
Width W Length
[mm]
*L
[mm]
Number of
specimens
N/A
N/A
32
56
14
N/A
N/A
6°
10°
32
16
95
63
56
40
126
88
21
21
14
14
OHC specimen 6.35
3.175
SHC specimen 6.35
6.35
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b) Scarf Hole Compression (SHC)
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Specimen design
No. Lay-up
EYY/EX 0° Plies
X
1.
2.
3.
4.
5.
6.
7.
[-45/904/45/904/0]s
[-45/902/45/902/-45/902/45/0]s
[-45/902/45/0]2s
[-45/90/45/0]3s
[45/02/-45/90]2s
[45/02/-45/02/45/02/-45/90]s
[45/04/-45/04/90]s
0.26
0.37
0.66
1.00
1.52
2.70
3.87
2/22
2/22
4/20
6/24
8/20
12/22
16/22
% of 0°
Plies
9
9
20
25
40
55
73
EYY = loading RMIT
direction
direciton
University = 0 Degree ply Fatigue2014
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Test Setup
a)
b)
a) Anti-buckling plates & b) Test setup
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Experimental results
600
500
Strength (MPa)
400
300
200
100
Unnotched
OHC_6.35
OHC_3.175
SNC_6Deg
0
0.00
0.50
1.00
1.50
2.00
2.50
3.00
3.50
4.00
4.50
SNC_10Deg
Eyy/Exx
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Experimental results
Quasi (6
0° plies)
6.35 hole
6deg scarf
tension
900
800
Stiff 16
0° plies)
250
110
300
380
140
750
52%
27%
150%
OHC_6.35
SNC_6Deg
OHT_6.35
700
Strength (MPa)
600
500
400
300
200
100
0
0.00
0.50
1.00
1.50
2.00
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2.50
Eyy/Exx
3.00
3.50
4.00
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4.50
5.00
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Experimental results
Figure. Experimental load-time data of 10°SHC.
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Design strength analysis
Figure. Effect of scarf angle on residual compression strength of laminate.
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Failure Mechanism
Left to right => soft to stiff
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Failure mechanism
Figure. Microscopic images of specimen fracture; (a) OHC quasi-isotropic
laminate EYY/EXX=1, (b) OHC soft laminate EYY/EXX=0.26 and (c) OHC stiff
laminate EYY/EXX=3.87, and (d) SHC stiff laminate EYY/EXX=2.70.
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Failure mechanism
500 µm
0 Degree plies
[-45/90/45/0]3s
Figure. OHC quasi-isotropic laminate EYY/EXX=1 ; d=6.35mm
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Failure mechanism
500 µm
04 plies
[45/04/-45/04/90]s
Figure. OHC stiff laminate EYY/EXX=3.87 ; d = 6.35mm
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500 µm
[45/04/-45/04/90]s
Figure. OHC stiff laminate EYY/EXX=3.87 ; d = 6.35mm
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500 µm
0 Degree plies
[-45/902/45/0]2s
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FE Models
lch =
EGtotal
σ 02
σ 0 = EεULT
XC
=E
E11
Figure. Abaqus model for (a) OHC and (b) SHC
Figure. Single-shell model with zero stiffness elements.
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Material Properties for Abaqus FE Models
Properties (damage)
Hashin’s damage initiation criteria
Longitudinal tensile, XT [MPa]
Longitudinal compression, XC [MPa]
Transverse tensile, TT [MPa]
Transverse compression, TC [MPa]
Longitudinal shear, XS [MPa]
Transverse shear, TS [MPa]
Hashin’s damage evolution criteria
Longitudinal tensile (Gft) [kJ/m2]
Longitudinal compression (Gfc) [kJ/m2]
Transverse tensile (Gmt) [kJ/m2]
Transverse compression (Gmc) [kJ/m2]
* Gfc value is calibrated using experimental
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Quantity
Properties (elastic) Quantity
2575
1235
40
181.7
85.7
85.7
E11 [MPa]
E22 [MPa]
G12 [MPa]
G13 [MPa]
G23 [MPa]
υ12 [-]
ρ [kg/mm3]
118000
9100
3940
3940
2960
0.33
1.19×10-6
80
8.5*
0.15
0.84
results
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Model calibration
Figure. Fiber compression fracture energy Gfc calibration for
different laminates.
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EXP vs FEA data
Figure. Abaqus FE normalized strength vs. laminate
stiffness ratio.
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Failure mechanics (Exp vs FE)
Figure. SHC(6° Scarf) soft laminate EYY/EXX=0.66 specimen failure; a) test
specimen failure, b) Abaqus analysis, showing the fiber failure damage
index in a 0° ply; c) matrix damage index; and d) shear damage index.
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[-45/90/45/0]3s – Quasi-isotropic – 6 degree SHC
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[45/04/-45/04/90]s –Stiff – 10 degree SHC
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Conclusions
• Experimental
– Laminate strength with respect to scarf angle – soft laminates reach DLL at
shallow angels. However, stiff laminates need steep angles. In other words, steep
scarf angles are required for highly loaded structures.
– Blocked plies (more than two 0 degree plies) stacked together is not useful (due to
delamination). However, in tension testing, blocked plies give much higher strength
• FEM
– ABAQUS data agrees with the ply cracking mechanism
– Compression energy release rate for Hashin’s criteria of laminates are highly
influenced by volume of 0Deg plies stacked together.
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Findings & further tasks
• More angles (suggested: 15;20;25)
• Investigation for other loading cases
– Tension
– Shear stress
– Bending
• Cohesive zone modelling for failure mechanism
– Delamination
– Micro-buckling of fibres
– Fibre splitting
– Matrix cracking
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Test Setup
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Test Setup
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Experimental Results
Quasi-isotropic
Figure 7. Effect of orthotropy on residual strength of laminate.
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No. Lay-up
EYY/EXX
SHC 6°
Scarf [MPa]
1.
2.
3.
4.
5.
6.
7.
0.26
0.37
0.66
1.00
1.52
2.70
3.87
87 ± 15.1
110 ± 6.0
117 ± 3.8
116 ± 2.1
113 ± 7.6
132 ± 1.7
145 ± 1.5
[-45/904/45/904/0]s
[-45/902/45/902/-45/902/45/0]s
[-45/902/45/0]2s
[-45/90/45/0]3s
[45/02/-45/90]2s
[45/02/-45/02/45/02/-45/90]s
[45/04/-45/04/90]s
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SHC 10°
Scarf [MPa]
OHC
d3.175mm
Hole [MPa]
OHC d6.35mmUn-notched
Hole [MPa] strength [MPa]
97 ± 0.7
122 ± 5.2
136 ± 11.3
139 ± 15.9
165 ± 0
165 ± 13.2
164 ± 9.5
178 ± 5.5
171 ± 7.0
239 ± 5.2
296 ± 0.8
365 ± 9.2
465 ± 32.8
483 ± 43.7
136 ± 7.3
168 ± 4.3
228 ± 13.8
251 ± 10.8
292 ± 30.3
363 ± 16.9
382 ± 21.2
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266 ± 16.4
283 ± 5.7
363 ± 42.5
426 ± 26.1
475 ± 22.4
529 ± 20
547 ± 10.4
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Name of the Project: Construction of Shitalakhya Bridge Project

Rotation measurements of a 6U payload dispensed in reduced