Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
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Transcript of Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
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Robust Design Optimization of Composite
Stiffened Panel With Discrete Source Damage
Frank Abdi (Ph.D), Cody Godines, Saber Dor-Mohammadi (Ph.D), Anil Mehta (DSC)
AlphaSTAR Corporation, Long Beach, California 90804, USA.
Robert Yancey (Ph.D), Harold Thomas (Ph.D);
ALTAIR Engineering Inc., Irvine, CA
2014 European Altair Technology Conference,
Munich, Germany
24- 26 June 2014
2
Agenda
•Objective/Benefits
•FAA Categories of Damage & Defect Considerations
•Technical Approach:
•Multi-Scale Modeling
•Multi-Scale Progressive Failure Analysis
•Multi Scale Failure Criteria
•Robust Design (Durability Damage Tolerance, Reliability) Consideration
•Experimental Methods
•Test Panel and Material Description
•Simulation
•Damage Simulation for Composite Stiffened Panels
•Parametric Robust Design Analysis of Compression and Tension panels
•Conclusion
3
Objective
OBJECTIVE
• Demonstrate Virtual Testing’s capability
• Maximize Durability, Damage Tolerance
• Meet B Basis Reliability requirements (95% confidence)
• Select among several design candidate
• Maintain Same Weight
BENEFITS
• Predict durability, damage tolerance and reliability in composite S/RFI
stiffened panels under DSD (Discrete Source Damage),
• Determine damage modes and their locations, critical failure events and
failure loads,
• Monitor the damage propagation and crack turning,
• Reduce design efforts and test costs of structures using S/RFI composites.
Robust Design Optimization of Stitched Stiffened Panel with DSD Damage
4
FAA Categories of Damage & Defect Considerations
Primary Composite Aircraft Structures
5
Technical Approach: Multi-Scale Modeling
Damage
Tracking Module
Progressive Failure Analysis (D&DT)
Optimization
Module
Composite
Micro -
Mechanics
Finite Element
Module
Probabilistic
Module
Manufacturing Anomalies
Damage
Assessment
FEA Results (u, s, e)
Material Properties
& Refined Mesh
Time Dependent Reliability Analysis
Progressive Failure Analysis
•Nonlinear static •Linear Static •Fatigue:
•LCF, HCF, •Random, PSD • 2 stage Fatigue
•Creep • Impact (low and high Velocity) •Lightning Strike
Composite mechanics •2-D/3-D fiber architecture
•Void, Residual Stress
•Anisotropic Matrix
•Waviness
• Interface coating
•Edge Effect Delamination
•Micro crack density
•Cool down process
•Fiber, matrix, interface model
Material Type • Metal
• Polymer Composite
•Thermosets
•Thermoplastic
•Chopped Fiber
• Fiber-Metal-Laminate
• Honeycomb
• Nano Composite
FEM Solvers
OPTISTRUCT,
NASTRAN, ANSYS
MHOST, ABAQUS
LS-DYNA, MARC
RADIOSS
Durability & Damage Tolerance, Reliability Software, Material Modeling
6
• Manufacturing defects
• Residual stresses
• Moisture
• Extreme temperatures
• Fiber/matrix interphase*
Input fibers
architecture
angles and
contents
Input matrix
properties and/or void
contents
Fibers Types
(e.g. filler,
warp, braid, etc.)
Composite laminate
properties
0
10
20
30
40
50
60
0.E+00 1.E-03 2.E-03 3.E-03 4.E-03Strain
Str
ess (
ksi)
• Stress-strain curve
• Strength, stiffness,
conductivity
• Moisture diffusivity
• Design failure envelope
• Allowables
• Moisture expansion
• CTE
• Recession*
• Global/local oxidation*
• Crack density
Woven Stitch Fabric/Weaves
Multi-Scale Material Characterization Micro-mechanics Modeling Considers Fiber Architecture, and Defects
• Micro-mechanics builds composites from the ground up • MCQ generates nonlinear “constituent” composite properties • Effect of defects, and scatter in properties are evaluated
Ref: M. Garg, G. Abumeri and D. Huang , “Predicting Failure Design Envelop for Composite Material System Using Finite Element and Progressive Failure
Analysis Approach”. Sampe 2008 Conference Paper, Long beach, CA, May 2008..
7
Initiation of Fracture Near Crown
Blade Joint
Fracture of 1st Fork Blade & Transfer
of Load to 2nd Fork Blade Wheel Tilts Due to Instability of Both
Composite Fork Blade Failures
Possible Carbon Fiber Bike Failure
8
Possible Carbon Fiber Bike Failure
Joint region of Carbon Fiber Fork Blade to Crown
Close Up Cross Section
FEM Analysis Results: high stress (red)
Close-up View of FEA Region Where Machining Defect Exists
Analysis Conditions & assumptions
Carbon Fiber Stress Distribution (normal ride weight Load)
Leading edge (Compression)
Trailing edge (Tension)
9
Fiber “Waviness” Fork Blade Laminate
Voids and Delaminations Fork Blade Laminate
Defects Example - Carbon Fiber Bike Example of Defects
Leading Edge Voids and Delaminations
Fiber Fork Blade
10
Fiber Waviness (Inside) & Voids/
Delaminations Fork Blade Leading Edge Large Void and Numerous
Delamination/Porosity (Leading Edge)
Defects Example - Carbon Fiber Bike SEM Photo of Defects
Extensive Compression Failure in
Fork Blade Leading Edge Fiber “Waviness” and Extensive Voids in Fork Blade
11
Fatigue Life Cycles % Void Content Test/Prediction: Effect of Fatigue life vs. Void content
Reference: V. Kunc, L. Klett., Z. Qian F. Abdi, B. Knouff “The Prediction of Fatigue Sensitivity to Void Content for 3D Reinforced Composites”. SAE, 2006,
Detroit, MI. 06M-265.
E-glass fiber, Dion 9800 matrix
0
5
10
15
20
25
30
35
40
45
1.E+00 1.E+01 1.E+02 1.E+03 1.E+04 1.E+05 1.E+06 1.E+07
Cycles to failure, Nf
Ma
xim
um
Str
es
s L
ev
el
(ks
i)
high void content - DION 9800 resin with clay filler
low void content - DION 9800 resin no clay filler
low void content - Reichold 31638/31100 blend
GENOA - 2% void content
GENOA - 10% void content
runout
(3D composites)
(3D composites)
345
1,400
8,050
Test Ave.
12% void
number of cycles to failure
314
1,350
6,990
GENOA
N/A
N/A
70,056
Test
2% void
476
2,340
66,900
GENOA
1.5270%
1.7350%
9.5730%
Life
increase
(times)
load
345
1,400
8,050
Test Ave.
12% void
number of cycles to failure
314
1,350
6,990
GENOA
N/A
N/A
70,056
Test
2% void
476
2,340
66,900
GENOA
1.5270%
1.7350%
9.5730%
Life
increase
(times)
load
(2D composites: 0/90)
(2D composites: 0/90)
12
Technical Approach
PFA takes full-scale FE model & breaks material properties down to microscopic
level. Material properties are updated, resulting from damage or crack
Unit cell at node
2D Woven
Laminate
Sliced unit cell
Component
FEM Vehicle
Micro - Scale
Traditional FEM stops here
GENOA goes down to micro - scale
Lamina
3D Fiber
FEM results carried down to micro scale
Reduced properties propagated up to vehicle scale
Multi-Scale Progressive Failure Analysis
(Cont’d)
13
Technical approach: Multi Scale Failure Criteria
* Options: Tsai-Wu, Tsai-Hill, User defined criteria, Puck, SIFT, HOFF, HASH ** Wrinkling, Crimpling, Dimpling, Intra-cell buckling, Core crushing
*** Environmental: Recession, Oxidation (Global, Discrete)
Reference: D. Huang, F. Abdi, A. Mossallam, “Comparison of Failure Mechanisms in Composite Structure”. SAMPE 2003 Conference Paper.
Unit Cell
damage
criteria
Delam
criteria
1. Matrix: Transverse tension
2. Matrix: Transverse
compression
3. Matrix: In-plane shear (+)
4. Matrix: In-plane shear (-)
5. Matrix: Normal compression
6. Matrix: Micro crack Density
7. Fiber: Longitudinal tension
8. Fiber: Longitudinal
compression
9. Fiber micro buckling
10. Fiber crushing
11. Delamination
12. Fiber Probabilistic
15. Normal tension
16. Transverse out-of-plane shear (+)
17. Transverse out-of-plane-shear (-)
18. Longitudinal out-of-plane shear (+)
19. Longitudinal out-of-plane shear (-)
20. Relative rotation criteria
21. Edge Effect
13. Strain limit
18. LEFM: VCCT (2d-3d)
19. Cohesive: DCZM (2d-3d)
19. Honeycomb**
20. Environmental***
14. Interactive*
• MDE (stress), SIFT (strain)
Damage, and Fracture Mechanics based (Cont’d)
MATRIX
FIBER
INTERACTION
DELAMINATION
FRACTURE
14
Answer Critical Design Questions Determine: when, why, where, and how to fix failure
Role of Analysis in FAA Building-Block Verification
Role of Analysis • Guides the integration and design processes • Identifies causes when failure to meet performance requirements occurs • Benefits certification process by establishing:
• Reduced test plan at each Level • Consider Scatter in Geometry, manufacturing, and material levels
GENOA reduces
testing at each
level of the
Building Block
Process
SDD_03_0102
Components
Sub - Components
Details
Elements
Coupons
Generic Specimens
Generic Specimens
Structural Features
Data Base
Integration of Design and Processes
Increasing Sample Size
SDD_03_0102
Components
Sub-Components
Details
Elements
Coupons
Non Generic
Specimens
Generic
Specimens
Structural Features
Data Base
Virtual Testing Guides/Reduces Testing at Each Level
- -
Configuration
Validation Process
Requires Minimum
Verification
Defines Risk
Mitigation
Calibration
Process
Analysis
Processes
15
Technical Approach
Generate FEA Model
from Initial Design
and Assign Loads
and Boundary
Conditions
Generate FEA Model
from Initial Design
and Assign Loads
and Boundary
Conditions
Select/Define
Objective Function
to be Minimized
Select/Define
Objective Function
to be Minimized Select
Constraints
Select
Constraints
Define Performance
Reliability Function
To Determine
Probability of Failure
Define Performance
Reliability Function
To Determine
Probability of Failure
Define Composite
Architecture and
Properties
Define Composite
Architecture and
Properties
Select Random
Variables for
Probabilistic
Analysis
Select Random
Variables for
Probabilistic
Analysis
Assign Random
Variables COVs and
Distribution Types
Assign Random
Variables COVs and
Distribution Types
Start with Initial Design
Configuration Based on
Engineering Knowledge
User Interface
Generate FEA Model
from Initial Design
and Assign Loads
and Boundary
Conditions
Generate FEA Model
from Initial Design
and Assign Loads
and Boundary
Conditions
Select/Define
Objective Function
to be Minimized
Select/Define
Objective Function
to be Minimized Select
Constraints
Select
Constraints
Define Performance
Reliability Function
To Determine
Probability of Failure
Define Performance
Reliability Function
To Determine
Probability of Failure
Define Composite
Architecture and
Properties
Define Composite
Architecture and
Properties
Select Random
Variables for
Probabilistic
Analysis
Select Random
Variables for
Probabilistic
Analysis
Assign Random
Variables COVs and
Distribution Types
Assign Random
Variables COVs and
Distribution Types
Start with Initial Design
Configuration Based on
Engineering Knowledge
User InterfaceFinite Element
Model Import
and Translation
Finite Element
Model Import
and Translation
Deterministic
Design
Optimization
Deterministic
Design
Optimization
FEA
Solver
FEA
Solver
1. Obtain New Design
2. Calculate Objective Function
3. Evaluate Constraints
1. Obtain New Design
2. Calculate Objective Function
3. Evaluate Constraints
Probabilistic*
Progressive
Failure Analysis
Probabilistic*
Progressive
Failure Analysis
Constraints
Violated?
yes no
Calculate
Sensitivities,
Probability of
Failure (Pf), and
Most Probable
Location of Failure
Calculate
Sensitivities,
Probability of
Failure (Pf), and
Most Probable
Location of Failure
Pf
Constraint
Violated?
Final
Design
no
yes
FEA Solver,
Damage Tracking
and
Probabilistic
Engine
FEA Solver,
Damage Tracking
and
Probabilistic
Engine
Finite Element
Model Import
and Translation
Finite Element
Model Import
and Translation
Deterministic
Design
Optimization
Deterministic
Design
Optimization
FEA
Solver
FEA
Solver
1. Obtain New Design
2. Calculate Objective Function
3. Evaluate Constraints
1. Obtain New Design
2. Calculate Objective Function
3. Evaluate Constraints
Probabilistic*
Progressive
Failure Analysis
Probabilistic*
Progressive
Failure Analysis
Constraints
Violated?
yes no
Calculate
Sensitivities,
Probability of
Failure (Pf), and
Most Probable
Location of Failure
Calculate
Sensitivities,
Probability of
Failure (Pf), and
Most Probable
Location of Failure
Pf
Constraint
Violated?
Final
Design
no
yes
FEA Solver,
Damage Tracking
and
Probabilistic
Engine
FEA Solver,
Damage Tracking
and
Probabilistic
Engine
*With Optional
Parallel Processing
Robust Design (Durability Damage Tolerance, Reliability) Consideration
1
2
3
16
Experimental Methods
Test groups for 3-Stringer S/RFI composite panels with DSD
(Discrete Source Damage)
Panel Loading Direction Specimen
No. Panel / Stiffener Description DSD Geometry
1 541 plies and 45.72mm(1.8”) Stringer Height
2 362 plies and 58.42mm(2.3”) Stringer Height
Tension
3 541 plies and 58.42mm(2.3”) Stringer Height
Saw cut
4 541 plies and 45.72mm(1.8”) Stringer Height
5 362 plies and 58.42mm(2.3”) Stringer Height
Compression
6 541 plies and 58.42mm(2.3”) Stringer Height
Diamond shape slot
Compression
Test Tension
Test
1: [45/-45/02/90/02/-45/45]9
2: [45/-45/02/90/02/-45/45]4
(Cont’d)
17
Test Panel and Material Description
Six panels are divided evenly into tension and compression groups
All panels are constructed from orthotropic stacks of warp-knit fabric
having a carbon fiber orientation of [45/-45/02/90/02/-45/45]
Compression panels are fabricated with the standard modulus HTA
5131 fibers from Tenax
Tension panels fabricated with intermediate modulus IMS 5131
fibers from Tenax in zero-degree direction and HTA 5131 fibers in
other 3 orientations.
19.05mm
(3/4 in)
177.8mm (7 in)
R=2.38mm (3/32 in)
4 places
Compression Diamond Configuration
R=2.38mm (3/32 in)
4 places
Tension Slotted Configuration
177.8mm (7 in)
19.05mm
(3/4 in)
177.8mm (7 in)
R=2.38mm (3/32 in)
4 places
Compression Diamond Configuration
R=2.38mm (3/32 in)
4 places
Tension Slotted Configuration
177.8mm (7 in)
DSD (Discrete Source Damage) geometry of compression
and tension panels
(Cont’d)
18
Damage Simulation for Composite Stiffened Panels
FE models for composite stiffened panels under DSD
A. Compression B. Tension
DSD crack
geometry
FE model
Overview
Dimensions 1.00 x 0.61 x 0.07 m (Length x Width x
Height)
1.02 x 0.61 x 0.07 m (Length x Width x
Height)
Loads &
Boundary
conditions
Nodes 2076 2387
Elements 2140 2060
Elements
Type 4-node QUAD (shell) 4-node QUAD (shell)
19
Damage Simulation for Compression Panels
Damage Propagation
Final Failure: 1299 kN Simulation Test
A panel with 54 plies and
58.42mm stringer height
The crack growth stop when the
damage reaches edges of
stiffeners
Little damage is accumulated at
the edge
The panel fails after cracks
cross stringers at a higher load
Slot Radius
Panel Edge Area shown in Testing
a) Damage/Fracture Propagation: 805 kN
b) Damage to Stiffeners: 1188 kN
(Cont’d)
20
Damage Simulation for Tension Panels
Damage Propagation
Crack Turning and Ultimate Load: 2552 kN
Simulation Testing
A tension panel with 54 plies
and 58.42 mm stringer
height
The crack growth stop when
the damage reaches edges
of stringers
More damage accumulate
along the edge
Crack turning take place in
the panel
Shear Cracks
a) Damage/Fracture Propagation: 805 kN
b) Damage to Stiffeners: 1735 kN
21
Damage Simulation vs Test for Tension Panels
Shear Cracks
22
Simulation Results for Composite Stiffened Panels
Summary of Damage and Damage Tolerance (D&DT) of Stiffened Panels
COMPRESSION TENSION Damage
initiation Final Failure
Damage
initiation Final Failure *
Load (kN) Load (kN)
ACT
Panel
Configuration Genoa Genoa Test
Error
% Genoa Genoa Test
Error
%
36 plies (4 stacks)
45.72 mm Stringer
Height
195 890 920.26 -3.3 215 1401 1326.10 +5.4
36 plies (4 stacks)
58.42 mm Stringer
Height
207 1065 1005.25 +5.6 227 1486 1374.43 +7.5
54 plies (6 stacks)
45.72 mm Stringer
Height
258 1205 1209.86 -0.4 253 1691 1716.93 -1.6
54 plies (6 stacks)
58.42 mm Stringer
Height
267 1299 1307.71 -0.7 234 1935 1899.30 +1.8
* Tension panels were considered to have failed when damage started to propagate along the outer stringers.
Factors impact D&DT of composite stiffened panels
Stringer height
Number of stacks in the skin panel
Percentage of [0/±45/90] fiber
23
Predict Failure process of Stitched NASA ACT 3-Stringer Panel Durability & Damage Tolerance Under Static Load
Crack turns
parallel to
loading
Simulation
SEALED ENVELOP PREDICTION
Photoelasticity
TEST
Prediction
Crack
turns
Crack
growth is
initially
normal to
loading
Test
Crack growth stiffened panel
23
Reference: D. Moon, F. Abdi, B. Davis, “Discrete Source Damage Tolerance Evaluation of S/RFI Stiffened Panels”, SAMPE 1999 Symposium
24
GENOA Prediction of 3-Stringer Panel Failure Modes
Environment RTD ETD CTD ETW
1.0 0.97 0.94 0.93 Normalized Failure Load
Max Load
Failure Load
For this example, the stiffener ratio is not significantly affected by environmental conditions as as to cause crack turning
Ref: J. M. Housner., Rose. Ragalini, "Design of Composite Stiffened Panels for D&DT and reliability without weight penalty", Journal of Society of
Allied Weight Engineering, Volume 69, Spring 2010, No. 5.
25
3-Stringer Panel Damage Mechanisms
Ref: J. M. Housner., Rose. Ragalini, "Design of Composite Stiffened Panels for D&DT and reliability without weight penalty", Journal of Society of
Allied Weight Engineering, Volume 69, Spring 2010, No. 5.
26
3-Stringer Panel Damage Mechanisms
27
Parametric Robust Design Analysis
Design Variables and Parameters Considered in GENOA Parametric
Robust Design Module
Random variables Designation Unit Initial value
Lower bound
Upper bound
Geometry
Stringer Height h mm 52.07 45.72 58.42
Skin Number of Stacks Skin thickness
/ t
/ mm
5 6.89
4 5.49
6 8.23
Manufacturing Uncertainties
Skin fiber content FVR % 55.3 49.8 60.8
Skin void content VVR % 1.150 1.035 1.265
Skin fiber orientation Angle ° +/-45;0;90 +/- 5°
Stringer fiber content FVR % 55.3 49.8 60.8
Stringer void content VVR % 1.150 1.035 1.265
Stringer fiber orientation Angle ° +/-45;0;90 +/- 5°
Material Uncertainties
Fiber Longitudinal Modulus Ef11 MPa 2.27E+05 2.04E+05 2.50E+05
Fiber Shear Modulus Gf12 MPa 1.38E+05 1.24E+05 1.52E+05
Fiber Compressive Strength Sf11C MPa 2.12E+03 1.90E+03 2.33E+03
Fiber Shear Strength Sf12S MPa 4.21E+02 4.62E+02 3.79E+02
Matrix Normal Modulus Em MPa 4.14E+03 4.55E+03 3.72E+03
Matrix Compressive Strength SmC MPa 2.07E+02 1.86E+02 2.28E+02
Down selected to % 0, +/- 45, 90 distribution in skin and stiffener
DV
DV
28
Compression Panel Performance
A. Compression
DSD crack
geometry
FE model
Overview
Dimensions 1.00 x 0.61 x 0.07 m (Length x Width x
Height)
Loads &
Boundary
conditions
FE model of the panel
Compression Panels: Effect of Stringer Height on Load Displacement
29
Compression Parametric Analysis: Stringer Height (Cont’d)
Compression Panels : Weight, Damage Initiation, and Final Failure Load
Random variables Unit Initial Design Design # 1 Design # 2 Design # 3
Geometry
Stringers Height mm 52.07 45.82 47.38 55.19
Skin number of Stacks / 5 (45 plies) 6 (54 plies) 6 (54 plies) 6 (54 plies)
Mechanical Results
Compression Strength kN 1121.4 1282.8 1290 1298
Damage Initiation Load kN 233.7 260.8 261.6 268.9
Damage Volume Ratio at Final Failure % 0.56 0.34 0.33 0.45
Constraints
Weight Kg 17.47 19.19 19.30 19.80
All designed panels give improved D&DT performance
Numbers of stack in the skin panel is critical for D&DT
performance
For the skin panel with a fixed stack number, the stringer
height has a little influence on D&DT performance
30
Parametric Robust Design Analysis of Tension Panel (Cont’d)
Percentage of Fibers on Structural Performance
Original designed panel gives an ultimate load of 2551.77 kN and a
residual load of 1252.41 kN
Optimized structural performance enhances residual strength and
Ultimate Load, using increased percentage of 00 fibers in laminates
Much higher percentage of 00 fibers in the skin panel could cause the
skin panel failure right after reaching the maximum loading
00
±450
900
00
±450
900
Maximum Residual
44 44 12 2551.77 1252.41 Yes
80 10 10 5165.46 0 Yes
10 10 80 2559.29 1246.42 Yes
10 80 10 2329.15 1107.64 No
25 25 50 2913.97 1439.02 Yes
25 50 25 2667.69 1239.12 No
44 44 12 2940.93 1418.78 Yes
50 25 25 4041.50 2031.4 Yes
80 10 10 7150.60 0 Yes
Fiber orientation in skin panelCrack Turning
44 44 12
Fiber orientation in stringer Load (kN)
80 10 10
Maximize residual strength for Baseline Panel using Skin and Stringer lay up configuration without Weight Increase
31
Parametric Robust Design Analysis (Cont’d)
Tension Panels : Load-Displacement Behavior of Designed
The tension panel with 54 plies and 58.42 mm stringer height
The percentage of [0/±45/90] fibers
Original design: (44/44/12) in the skin panel and stringers
Robust design: (50/25/25) in the skin panel and (80/10/10) in stringers
Load-Displacement Curve of the ACT
Tension Panel
0
500
1000
1500
2000
2500
3000
3500
4000
4500
0 2 4 6 8 10
Displacement (mm)
Lo
ad
(kN
)
original_design
robust_design
4042 kN
2031 kN
2552 kN
1252 kN
Load-Displacement Curve of the ACT
Tension Panel
0
500
1000
1500
2000
2500
3000
3500
4000
4500
0 2 4 6 8 10
Displacement (mm)
Lo
ad
(kN
)
original_design
robust_design
4042 kN
2031 kN
2552 kN
1252 kN
B. Tension
DSD crack
geometry
FE model
Overview
Dimensions 1.02 x 0.61 x 0.07 m (Length x Width x
Height)
Loads &
Boundary
conditions
32
Parametric Robust Design Analysis for Compression Panels (Cont’d)
Comparison of Damage Volume Ratio before and after Optimization
Optimized panel has a smaller damage volume ratio (percent), and
higher ultimate load
33
Parametric Robust Design Analysis (Cont’d)
Compression Panel: Probabilistic Sensitivities at ultimate load under
geometric, manufacturing and materials uncertainties
Material properties have greatest influence on load carrying capability of
compression panel.
Skin stack number, stringer height and percentage of fibers also affect the
ultimate load in a significant way
34
Parametric Robust Design Analysis (Cont’d)
Cumulative Probability of Failure before and after optimization and
Reliability Improvement of compression panel
Before optimization, few panel will fail under a load of 640 kN and few
panel will remain if the load exceeds 1640 kN
After optimization, few panel will fail under a load of 840 kN and few
panel will remain if the load exceeds 1840 kN
35
Conclusion
Damage propagation and crack turning phenomena in tension
panels are fully monitored
For ACT panels, stringers height, skin stack number and
percentage of fibers are key variables which influence the
structural performance
Discrete source damage (DSD) analysis methodology is
coupled with optimization and probabilistic analysis
Evaluates durability and damage tolerance of S/RFI
composite structures under DSD events
Methodology can reduce excessive experimental costs & time
Robust design methodology can accelerate certification
process and reduce iterative design cycles.
Conform to FAA certification requirement.