Robust Design Optimization of Composite Stiffened Panel with Discrete Source Damage

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

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


To Determine


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


To Determine


Define Performance


To Determine


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




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







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


Tension


Saw cut



Compression


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


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


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


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


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.