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Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage
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

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A Robust Design Optimization (RBO) computational approach has been developed to maximize the Durability and Damage Tolerance (D&DT) and Reliability of composite stiffened panel structure with 2 bay …

A Robust Design Optimization (RBO) computational approach has been developed to maximize the Durability and Damage Tolerance (D&DT) and Reliability of composite stiffened panel structure with 2 bay cracks. Robust design optimization is integrated in GENOA D&DT by means of coupling Multi-Scale Progressive Failure Analysis (MS-PFA), Reliability, Optistruct FE Optimizer, and Hypermesh software. The integrated software capability is utilized to minimize weight, maximize peak load and the residual strength, and enforce the directionality of cracks growth turning for safe design. The random variable parameters considered are: a) skin/stiffener load ratio, stiffeners height/width, and ply angle orientation in the stiffener and skin. The capability provides engineers with the predictive computational technology to characterize and qualify advanced composites materials and structures while considering manufacturing anomalies. First, the D&DT analysis prediction was performed using stiffened panels with imbedded 2-bay skin cracks under compression and tension loading. Results are compared with test: a) Load-displacement curve, b) Crack Growth direction; and c) Strain Gauge and photoelastic measurements. Second, we performed risk mitigation utilizing the Reliability Based Design Optimization to improve the failure evolution (translaminar, and interlaminar) with the main objective to increase the load displacement and turn crack(s) for confinment and safety.

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  • 1. 1 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. 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. 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. 4 FAA Categories of Damage & Defect Considerations Primary Composite Aircraft Structures
  • 5. 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. 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-03 Strain Stress(ksi) • Stress-strain curve • Strength, stiffness, conductivity • Moisture diffusivity • Design failure envelope • Allowables • Moisture expansion • CTE • Recession* • Global/local oxidation* • Crack density Woven StitchFabric/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. 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. 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. 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. 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. 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 MaximumStressLevel(ksi) 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. 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. 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. 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. 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 Interface 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 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. 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 HeightTension 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 HeightCompression 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. 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. 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. 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 EdgeArea shown in Testing a) Damage/Fracture Propagation: 805 kN b) Damage to Stiffeners: 1188 kN (Cont’d)
  • 20. 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. 21 Damage Simulation vs Test for Tension Panels Shear Cracks
  • 22. 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. 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. 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. 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. 26 3-Stringer Panel Damage Mechanisms
  • 27. 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. 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. 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. 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 panel Crack 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. 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) Load(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) Load(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. 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. 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. 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. 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.

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