Durability and Damage Tolerance and Life Prediction of ...

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Durability and Damage Tolerance and Life Prediction of Composite & Metallic Components

Using GENOA Software

Mohit GargEmail: [email protected]

Alpha STAR Corporation, Long Beach, CA 90804www.ascgenoa.com

Java

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• Challenges in Composite Design • Why GENOA ?• Typical Applications• GENOA For Virtual Testing & Simulations • Technical Approach • GENOA Modules and Key Capabilities • Examples:Progressive Failure Static, Dynamic,

Material Qualification, Material Allowable, Fatigue,VCCT/DCZM, Filament Winding, and Reliability Prediction

Outline

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Material Modeling under hygral conditionsNonlinear Fiber/Matrix (Unidirectional, Bidirectional, and triaxial) propertiesSandwich structure (metals/composites/honeycomb/ceramics/foams)Filament Winding for Composite Over-wraped Pressure Vessels

Manufacturing (fabrication process)Fiber lay-up (Directional Uncertainty)Fiber & Void Volume RatioCuring and Springback

Life Assessment PredictionStatic, thermal and mechanical Fatigue (low, high, random, two stage, PSD)Buckling, Creep, thermal aging, low & high velocity impactDamage & Failure initiation loads and modes (delamination, shear, tensile, etc.,)Micro-crack Density & PermeabilityInspection Intervals & Repairs

Cost Effective Optimized DesignReliabilityDurability

Challenges In Composites Design

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FEM codes are challenged by complex mechanisms caused by pre-mature failure of composite structure– Typically failure is assessed at laminate/lamina scales– Actual failure has its source at micro scale fiber/matrix

Genoa was developed to meet the challenge posed by composite andceramic structural systems

Genoa is an augmentation to FEA– adding the necessary micro-scale analysis– Uses a building block approach from material development to design and

certification– Reduces amount of coupon, component and structure testing

Augmentation provides consistent, reliable and accurate (<10% error) predictions– Composite architectures are complex– Composite Micro crack formation during manufacturing and service– Helps to understand when (LOAD VALUE), where, and why failure

occurs in a composite structure

Why GENOA ?

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Feature GENOA ABAQUS MD NASTRANStrength

Durability & Damage Tolerance(linear and nonlinear)

FE Solver (linear and nonlinear) FE Solver (linear & nonlinear)

FTD/FCG

Fatigue (Quasi-Static, Spectrum, Two-stage)

Material Characterization (fiber/matrix)

A & B-Basis Allowables

Parametric Carpet Plots

Composite, Sandwich StructuresStrength (Damage Initiation,

Progression and Final Failure) and Stiffness

(limited damage initiation & final failure) Requires very advanced user

subroutine

(limited damage initiation & final failure)

Fiber/Matrix/Inter-phase/Ply Input (ply only) (ply only)

Creeplimited (several options available mainly

focused for metals)(TBD)

Dynamic (SOL 700) Invokes LS-DYNA and GENOA as library

Failure Criteria(several to chose from) (limited and interactive ply based) (Uses GENOA Library)

Solver Speed Comparatively Slower Comparatively Faster Comparatively Faster

Comparison with Other FE Solvers

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• Durability and Damage Tolerance• Impact, Fatigue and Creep• Crack Initiation, Accumulation and Growth• Structural reliability and reliability based optimization• Effects of temperature, cure, moisture, and

manufacturing defects on material/structural response

Typical Applications

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Equilib. After DamageInitiation~Cycle 14Node 549 Ply 11,12,Node 688 Ply 5Node 687 Ply 5,10

Last Equilib.Cycle 116

Cycle 91Peak

Fracture InitiationCycle 24

3.21 9497.66 18992.11 28486.55 37981.00

1098.30

823.73

549.15

274.58

0.00

Damage Energy Release Rate (psi)

Force

Full Scale Verification

Component Tests

Sub-Element Tests

Environmental

Coupon Tests

InspectionDamage Mechanisms

Constituent Verification

Sensitivity

GENOA For Virtual Testing & Simulations

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Material ConstituentVerification

Full Scale Verification

Coupon Verification

Sub-Element Verification

Requires Calibration

Process

Requires Configuration

Calibration Process

Requires Minimum

Verification Process

Defines Risk Mitigation

Benefits - Certification process, Reduced test plan

Genoa Is A Virtual Testing Tool – Follows ASTM Standards

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Organization Project Objective

Boeing X-37 Orbital Vehicle

Determination of strength of the composite sandwich fuselage

Im pactorFixture plate

BoeingDetermination of the residual strength after impact of composite panels

Launch Vehicle Panels

Boeing/NASA Wing Panels

Determination of the residual strength with discrete source damage

NASAEngineering

Safety Center

Detail RCC Impact

Response

Analysis of progressive failure mechanisms


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Piper Aircraft

virtual test, life assessment and verification

NASAColumbia Accident Inves-tigation Board

Shuttle RCC Impact and

Failure

Verification of the foam impact damage scenario and subsequent structural failure

Honeywell Engine System

Metallic Stiffened Structure

High CycleFatigue

Frequency degradation, life assessment and verification

NavyFailure of

Composite Joints

Determination of the strength and failure modes of composite joints for design improvements


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Honeywell,Siemens,

COI Ceramics

Ceramic Combustor

Liners

Determine creep behavior and fatigue life

Navy- ONR/ Carderock

Analysis of Deck-bulkhead

Structure

Fire resistance simulation of a deck-bulkhead assembly

Life assessment and Failure Analysis

Fatigue Analysis Turbofan

Blade

Hartzel

Navy Thermal Mechanical

Fatigue Analysis

Life assessment and Uncertainty Failure Analysis


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NASA/LaRCX-37 Sub-

Element Pi JointDesign

Simulating the failure and load capacity of the X-37 Honeycomb T-joint

Navy/ONRComposite Joint

Failure Bolted/Bonded

Determine strength and failure modes of composite joints & improved design Optimization

Aerospace Joint: Virtual Testing /Verification

Navy/ONRGeneral Dynamic

Composite Storage Module

(CSM)Bond Shear

Failure

Adhesive Bond progressive FailureMode I- II), void effect, surface polish effect

Navy/ONRComposite TAS

JointCore, Putty

Effect on Failure

Laminate damage at core tip and core cracking observed

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NASA/LaRC

(X-37)

Joining MethodPanel-to-Panel

2D/3D C/SiC Assembled

Sub-Elements

Predict/Verify Durability, failure location/load, failure type Predict/Test=0.99

Composite Joint: Virtual Testing /Verification

NASA/LaRC

(X-37)

Joining MethodPanel-to-Tube

2D/3D C/SiC Sub-Elements

Predict/Verify Durability, failure load, location, failure type Predict/Test=2882/3045 lbs

NASA/LaRC

(X-37)

Joining MethodTube-to-Tube

2D/3D C/SiC CMC Sub-Elements

Predict loading effect on, failure load, location, failure type

Navy

(ONR)Composite Joint

Delamination failure

Predict/Verify Delaminatin Initiation, Growth, location/load

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Organization Project ObjectiveAerospace Composite Joint: Virtual Testing /Verification

Navy- ONR, General Dynamics , Boeing

Design Director’s Room Joint Blast Condition: Ply Drop off,Closure

Actual design: yielding in metallic gusset. Metallic Components thickened to force failure in composite material

Lockheed Martin

F-22

Failure under Shake ( PSD)

, harmonic Environment

Determine failure location, Natural frequency degradation

Static progressive damage analysis was conducted with contact interface between each component.

ONRL Bolted JointDesign Failure,

washer, insert type

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Technical Approach

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Technical Approach

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GENOA PFA (Progressive Failure Analysis) – Solution Hierarchy

U nit cell a t node

2D W oven

Lam inate

S liced un it ce ll

C om ponent FE M

V ehic le

M icro -S cale

Traditional FE M stops hereG E N O A goes dow n to m icro-scale

Lam ina

3D F iber

F E M resu lts carried dow n to m icro scale R educed properties propagated up to veh icle scale

PFA takes full-scale finite element model and breaks the material properties down to the microscopic level. Material properties are updated, reflecting any changes resulting from damage or crack

Technical Approach

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Multi Scale Multiple Failure Criteria

* Options: Tsai-Wu, Tsai-Hill, User defined criteria, Puck, SIFT, HOFF, HASH

** Wrinkling, Crimpling, Dimpling, Intra-cell buckling, Core crushing

Reference: D. Huang, F. Abdi, A. Mossallam, “Comparison of Failure Mechanisms in Composite Structure”. SAMPE 2003 Conference Paper.

Unit Celldamagecriteria

Delamcriteria

1. Matrix: Transverse tension2. Matrix: Transverse

compression3. Matrix: In-plane shear (+)4. Matrix: In-plane shear (-)5. Matrix: Normal compression6. Matrix: Micro crack Density

7. Fiber: Longitudinal tension8. Fiber: Longitudinal

compression• Fiber micro buckling• Fiber crushing• Delamination

11. Normal tension12. Transverse out-of-plane shear (+)13. Transverse out-of-plane-shear (-)14. Longitudinal out-of-plane shear (+)15. Longitudinal out-of-plane shear (-)16. Relative rotation criteria17. Edge Effect

9. Strain limit18. VCCT, DCZM-2d-3d (LEFM)19.Honeycomb**

10. Interactive*• MDE (stress)• SIFT (strain)

Damage, and Fracture Mechanics based

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Load Stepping

• User defined maximum damage per increment

• Reduced properties after damage

Technical Approach

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Adjustment of MaterialProperties

FEA Iteration

DamageInitiation/Growth

No Change In Damage

State

IncreaseLoad

STOP

Structural Failure

• Node displacements• Resultant laminate forces (shells)• Laminate stresses (solids)• Generalized laminate strains (shells)• Laminate strains (solids)

DamageAssessment

Technical Approach

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o o

oo

x

y

z

x

Lamina Micro-stress Theory*Stresses, strains and rotations at a node

z

l

Unit Cellat a node

Delamination at a Node

zl

Laminate TheoryFEM results at a node in each lamina

y

FEM Nodal –Based Resultsu, v, w x, y, xy

x, y, xy

zn

Sub-Divided Unit Cell at a Node

Interrogate Unit Cell

Consequences of Delamination•Loss of local stiffnesses at node•Loads transfer to other nodes or lamina possibly causing delam propagation

i

i+1

z

x

Occurs due to a combination of interlamina micro-stresses or relative

rotations

z

z

ll

nn

l

• Stresses and strains from micro-stress theory*

• Typically 7x7 subdivisions • 16 failure criteria per sub-division

l

i

Technical Approach

Delamination Prediction Process

22Answer 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 (use reversed Taguchi: regression analysis)• 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

ElementsCoupons

Non Generic Specimens

Generic Specimens

Structural Features

Data Base

Integration of Design

& Processes

Incr

easin

g Sa

mpl

e Si

ze

Virtual Testing Guides/Reduces Testing at Each Level

--

Configuration Validation Process

Requires Minimum Verification

Defines Risk Mitigation

Calibration Process

Analysis Processes

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PFA Provides Exhaustion Of Energy Tracking

10 100 1000

Percent Damage

Failure initiates at 9 cycles


0

3

6

9

1

10

10 1001

1000 10000

Total Damage Energy Release RateAn overall indicator of energy expanded in creation of damage• Ratio of total energy expanded during damage to total damage produced.•Usually crosses a minimum value, corresponding to a damage tolerance limit, prior to structural fracture

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TDERR

1

Percent of Damage Volume• Indicator of structural margin of

safety, and degradation• Ratio of the total damage volume in

composite to total volume of composite structure, multiplied by 100

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PFA Provides Exhaustion Of Energy Tracking

100

10

1

Cycles10 1001 1000 10000


Track Damage Energy Release RateFailure is consistently indicated by all critical parameters

Damage Energy Release Rate• Indicator of inspection criteria• Ratio of incremental work done by external forces to the incremental volume of damage created during a load increment that causes damage

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GENOAModular Structure

& Key Capabilities

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Modular Structure & Key Capabilities

GENOA is also Available on the Web

Graphical User Interface with Automatic Import Capability From NASTRAN, MARC, ABAQUS, and ANSYS

Material Simulation Modules

Material Characterization Analysis

Material Uncertainty Analysis

Reliability & Optimization

Fracture Toughness Determination

Fatigue Crack Growth Determination

Progressive Failure Analysis

Random Fatigue

Creep

Power Spectrum Density

Virtual Crack Closure Technique

Discrete Cohesive Zone Modeling

Progressive Failure Analysis -

Virtual Crack Closure Technique (Metal Fatigue)

Static

Vibration

Buckling

Post -Buckling

Quasi -Static Fatigue

High Cycle Fatigue

(Harmonic Loading)IMPACT (High & Low

velocity

Graphical User Interface with Automatic Import Capability From NASTRAN, ABAQUS, ANSYS, and LS-DYNA

Material Simulation Modules

Material Certification & Qualification

Material Uncertainty Analysis

Reliability & Optimization

Fracture Toughness Determination

Fatigue Crack Growth Determination

Progressive Failure Analysis

Random Fatigue

Creep

Power Spectrum Density

Virtual Crack Closure Technique

Discrete Cohesive Zone Modeling

Progressive Failure Analysis -

Virtual Crack Closure Technique (Metal Fatigue)

Static

Vibration

Buckling

Post -Buckling

Quasi -Static Fatigue

High Cycle Fatigue

(Harmonic Loading)IMPACT (High & Low

velocity

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Material Modeling and Analysis

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Material Characterization Analysis (MCA)

• Predicts lamina (ply) and laminate composite properties of metal, polymers (braid, weave, stitched 2D/3D) and ceramic composites

• Predicts distribution of micro stresses in fibers and matrix caused by external loading

• Material properties calibration based on experimental data• Enables accounting for manufacturing defects (voids, residual stress-strain

fields, etc.) and environmental effects (temperature, moisture, etc.)• Predicts the following lamina/laminate properties:

– Modulus / Poisson’s Ratio / Strength– Thermal Expansion Coefficient / Heat Conductivity / Moisture Diffusivities

• Simulates material processing

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Stuffer

Wrap Weaver

Layer-to-Layer Angle Interlock Woven

Filler

0.5 cm

Stuffer

Through-The-Thickness Angle Interlock Wrap

Weaver

Filler

1.0 cm

Wrap Weaver

Stuffer

Filler

Orthogonal Interlock Woven

1.0 cm

HL

HT

HO

Composite Parameter Test Cox GENOA E11(msi) 12.3 13.23 12.35 E22(msi) 6.35 8.15 6.42 E33(msi) 2.32 1.75 1.7

HL1

G12(msi) 0.9 0.78 0.69 E11(msi) 11.6 11.77 11.68 E22(msi) 6.13 8 6.6 E33(msi) 2.03 1.48 1.46

HL2

G12(msi) 0.84 0.71 0.6 E11(msi) 11.45 12.84 11.49 E22(msi) 6.16 7.89 6.17 E33(msi) 2 1.85 1.78

HT1

G12(msi) 0.81 0.77 0.67 E11(msi) 10.44 12.34 10.62 E22(msi) 6.64 9.63 6.7 E33(msi) 2.01 1.64 1.56

HT2

G12(msi) 0.83 0.72 0.64 E11(msi) 12.76 13.5 12.79 E22(msi) 5.78 8.18 6.1 E33(msi) 2.23 2.5 2.85

HO1

G12(msi) 0.73 0.78 0.67 E11(msi) 10 12.15 10 E22(msi) 6.03 8.1 6.97 E33(msi) 3.23 2.95 3.33

HO2

G12(msi) - - 0.62

Heavily Compacted PMC Composites—Prediction vs. Test

GENOA MCA Results Compared to Test

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Calibration of lamina Mechanical Properties for Cryogenic Tank made of IM7/PETI-5

0.0

0.2

0.4

0.6

0.8

1.0

-423F -320F 75F 350F 400F 450FTemperature

Shea

r Mod

ulus

(msi

)

CalibrationTest

220

240

260

280

300

320


Long

itudi

nal T

ensi

le S

tren

gth

(ksi

)

CalibrationTest

22

23

24

25

26

27


Long

itudi

nal M

odul

us (m

si) Calibration

Test

0.0

0.2

0.4

0.6

0.8

1.0

1.2

1.4

1.6

1.8

2.0


Tran

sver

se M

odul

us (m

si) Calibration

Test

Longitudinal Modulus Transverse Modulus

Shear Modulus Longitudinal Tensile StrengthFrank Abdi, Xiaofeng Su, “Progressive Failure Analysis of RLV Laminates of IM7/PETI-5 at High, Room, and Cryogenic Temperatures”, SDM 44 Conference Paper.

Material Property Prediction Verification

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-5

0

5

10

15

20

25

30

35

-500 -400 -300 -200 -100 0 100 200 300 400 500

Temperature (F)

CTE

x10

-6(in

/in -

F)

CTE1 (longitudinal) - TestCTE1 (longitudinal) - SimulationCTE2 (transverse) -TestCTE2 (transverse) - Simulation

-423F -320F 75F 350F

450F

400F

Test (CTE2)

Calibration of lamina Mechanical Properties for Cryogenic Tank made of IM7/PETI-5

Frank Abdi, Xiaofeng Su, “Progressive Failure Analysis of RLV Laminates of IM7/PETI-5 at High, Room, and Cryogenic Temperatures”, SDM 44 Conference Paper.

Material Property Prediction Verification

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ECM: Equivalent Constraint Model

9 0

9 0

-4 5

-4 5

• Many composites failures initiate as a result of matrix cracking

• When matrix cracks it releases energy that causes accumulation of additional cracks. The process continues until reaching a saturation point

0

1

2

3

4

5

6

7

8

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8Applied Strain (%)

Cra

ck D

ensi

ty (c

m-1

)

Simulation - [45/-45/90/90]sTest - [45/-45/90/90]sSimulation - [0/0/-60/-60/60/60]sTest - [0/0/-60/-60/60/60]s

Prediction/Test verification of crack densities in 90° coupons subject to longitudinal tension

0

1000

2000

3000

4000

5000

6000

7000

8000

0.00 0.20 0.40 0.60 0.80 1.00 1.20 1.40

Strain (%)

Dam

age

Ener

gy (i

n-lb

)

450F400F350F75F-320F-423F

UN-RLV(90)

Permeability Study of Cryogenic Tank[-45/03/45/903/45/03/-45] IM7-977-2 at the range of

temperatures between 423°F and 450 °F

Ref: Frank Abdi, Xiaofeng Su “Composite Tank Permeation and Crack Density Prediction and Verification”, ASME Paper No. IMECE2003-4439, November 2003.

Micro-crack Formation Prediction Verification

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Through-the-thickness crack density distributions in the layers of 25-inch composite disk made of quasi-isotropic laminate [45/90/-45/0]s subject to various levels of external pressure

(Cont’d)Micro-crack Formation Prediction Verification

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Material Uncertainty Analysis (MUA)

• Allows for probabilistic design of composite materials• Calculation of the lamina/laminate allowables and design variables based on the

scatter in the test/simulated data• Predicts critical design variables prior to the start-up of the test program• Combines composite micro-mechanics and probabilistic analysis to evaluate the

reliability of a composite material/structure based on the scatter of fundamental primitive variables (fiber/matrix stiffness and strength, fiber volume ratio, void volume ratio and others)

• Predicts the sensitivity to constituent material properties and other design variables including fiber architecture, as-built manufacturing processes and defect content

• Outputs Cumulative Distribution Function (CDF) and Probability Density Function (PDF) of the response and primitive variables over the entire failure probability

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GENOA ModuleMaterial Uncertainty Analyzer

(MUA)

PDF

CDF

Ef1

1

Ef2

2

Nuf

12

Nuf

21

Gf1

2

Sf11

T

Sf11

C

Em

Num

SmT

SmC

Sm12

S11T S1

1C S12

S22T S2

2C

0 .00 .10 .20 .30 .40 .50 .60 .70 .80 .91 .0

Sens

itivi

ty

C o n stitu en t P ro p ertyL am in a

P ro p ertyFive Lamina Properties

Twelve Fiber and MatrixConstituent Properties

Normalized Sensitivity

INPUTFiber/Matrix or Lamina

Uncertainty Data

Sensitivity Data Probability and cumulativedensity functions for

lamina/laminate

100%

ExptAnalysis

ExptAnalysis

Establish A-B Base Allowables(Cont’d)

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Material Characterization Optimization (MCO]

• Allows for calibration of fiber/matrix properties of a selected composite system while simulating the ASTM coupon tests (stress-strain fields) until they are closely reproduced

• Applicable for short- and long-fiber composites• Allows for simulating multiple tests at once• Instrumental for developing/validating standards for material/coupon testing• Reliable for virtual simulation of material testing• Ideal tool for identifying fabrication process variables that maximize material

performance

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Simulation of 96-oz 3TEX 3Weave E-glass/Dion 9800TM composite system

Longitudinal tension of cross-ply coupon


38

Transverse tension of cross-ply coupon



39

Shear Coupon Test



40

MCO Predicted Effective Matrix SS Curve for the Composite Systemvs. original stress-strain curve for Dion 9800TM matrix

“Original” stands for Dion Matrix alone



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Test shows Composite thickness Effects Strength Property

Fiber Micro-Buckling, Fiber Volume Ratio, Void Volume Ratio Fiber Waviness/ Misalignment

Reference: Soutis, C., and Lee, J., (2007). A study on the compressive strength of thick carbon fibre-epoxy laminates. Composite Science & Technology 67, pp. 2015-2026.

%Void Vs. thickness Failure stress Vs. Fiber Waviness

Test Vs. Simulation Un-notched Coupons: Combined Effect

0.0

0.2

0.4

0.6

0.8

1.0

1.2

2 4 8

Laminate Thickness [mm]

Nor

mal

ized

Stre

ngth

[-] Test

Simulation

Combined Effect: VVR, FVR, FMB, WavinessU n n o t c h e d C o m p r e s s io n [ P ly - L e v e l S c a l in g ]

01 0 02 0 03 0 04 0 05 0 06 0 07 0 08 0 09 0 0

0 .0 0 0 0 .0 0 5 0 .0 1 0 0 .0 1 5 0 .0 2 0

S t r a in [ m m /m m ]

Stre

ss [M

Pa]

t = 2 m mt = 4 m mt = 8 m m

t 1=2mm

t 2=4mm

t 3=8mm

Thickness Effect on Stress-Strain CurveVVRFVR

WAVINESS

Probabilistic Sensitivity Factors

THICKNESSFMB_D11C

Laminate Property Reduction Factors:

42

State-of-the art durability and damage tolerance capability by means of progressive failure analysis• Predicts thermo-mechanical limit loads of composite/metallic structural

components while taking into consideration:– Material degradation

• Matrix plasticity and micro-cracking• Environmental effects and manufacturing defects

– Material Nonlinearity– Changes in structural geometry

• Predicts crack initiation and growth• Computes various damage and failure modes and locations at different material

scales beginning with the micro-cracking in fiber, matrix, fiber/matrix interface• Predicts the post-buckling response• Predicts the limit time to failure and time-dependent crack initiation and growth • Effective for permeability and damage tolerant design• Effective for reducing the number of tests and certification

Progressive Failure Analysis (PFA): Static

43

1 inch

0.229 inches19 plies

302 lbs 343 lbs

Delamination Process

(Cont’d)

Ref: V. S. Sokolinsky, J. Housner, Jonas Surdenas, and F. Abdi, “Progressive Failure Analysis of Shuttle Reinforced Carbon-Carbon Plate Specimens”, Proceedings of the 47th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference, Newport, Rhode Island, May 1–4, 2006, AIAA–2006–1781.

Evaluation of RCC


44Computation time: 25 min

(Cont’d)


Evaluation of RCC

45

Crack turns parallel to loading

Specimen Cross-Section

Crack turns

Crack growth is initially normal to loading

Simulation

Test

Photo-elasticity

Ref: D. Moon, F. Abdi, B. Davis, “Discrete Source Damage Tolerance Evaluation of S/RFI Stiffened Panels”, SAMPE 1999 Symposium

(Cont’d)

Predict Ultimate Load of 3 Stringer Panel


46

(Cont’d)

• Uses an explicit transient dynamic algorithm through the integration with LS-DYNA software.

• Predicts complicated large deformation behavior in composite structures including prediction of any extensive material degradation

• Simulates low/high speed impact event, including damage initiation and propagation and ultimately final collapse of the composite structure

• Detects damage in composite laminates at each time interval at micro-level using the advanced micro-mechanics

• Tracks material degradation at the fiber/matrix level enables accurate simulation of the global response of a composite structure

• Supports multiple element types (shell and solid)• Enables simulation of the components subjected to large deformation,

material nonlinearity, and nonlinear boundary conditions

Progressive Failure Dynamic Analysis (PFDA):Dynamics

47

Back

Top Top Layer

Back Layer

I m p a c t o rF i x t u r e p l a t e5 inch by 5 inch panel

Impact Speed: 3.01 ft/secImpact Energy: 7.58 ft-lbsTotal Time: 19.75 msecClamped boundary condition on all four sidesSpherical Steel Impactor (Diameter 1 inch)G30-500/45 R6367: /-45/0/90/0/90/0/90/0/90/-45/45

Impact of Steel Ball on Composite Plate

-2.00E+00

0.00E+00

2.00E+00

4.00E+00

6.00E+00

8.00E+00

1.00E+01

-5.00E+00 0.00E+00 5.00E+00 1.00E+01 1.50E+01 2.00E+01 2.50E+01

Time (msec)

Load

(lb)

*100

simulationexperiment

Load Vs. Time Curve

Low Velocity Impact of Composite Panel (Boeing)

Reference:1- D. Huang, F. Abdi, M. Khatiblou “Impact, and Tension After Impact of Composite Launch Space Structure” , SAMPE 2001, Long Beach, CA2- F. Abdi, “Impact Damage Propagation In Composite Structures”. JEC 2006 Journal Publication.

Experimental Simulated

Foot Print Foot Print

48

Low Velocity Impact(GENOA Predicts Residual Strength After Impact)

ImpactorFixture plate

0

100

200

300

400

500

600

700

800

900

0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 0.4 0.45Deflection in.

Forc

e L

b

AnalysisTest

TestGENOA

0

20

40

60

80

100

120

0 0.2 0.4 0.6 0.8 1 1.2 1.4Damage Size in.

Stre

ss P

si

TestGenoa

TestGENOA

Back

Front

6 layers of G30-500/R3676 fabric• (45,-45), 4x(0,90),(45,-45) • Ply thickness - 0.014 in • Total thickness - 0.084 in

• Impactor mass - 53.75 lbs • impact velocity - 3.01 ft/sec • impact energy - 7.58 ft-lbs

Front

Back

Ref: D. Huang, F. Abdi, M. Khatiblou “Impact, and Tension After Impact of Composite Launch Space Structure” , SAMPE 2001, Long Beach, CA

49

High Velocity Impact

30 mph

825 lbsTest Comparison of Composite Response at 48 km/hr to

GENOA-LS-DYNA and LS-DYNA Stand Alone Solutions

Damage Initiation Damage Progression

Pattern of Bulging Side Walls of Boxed Beam Is Consistent with Test Observation

50

GENOA Impact Simulation(With Compression After Impact Compared to Test)

Impact on Sandwich Composite Panel

Ref: M. Garg1, F. Abdi1, A. Zammit2, J. Bayandor2 , Y. Suh3 , S. Song (2009). IMPACT DAMAGE RESISTANCE AND COMPRESSION-AFTER-IMPACT STRENGTH OF SANDWICH COMPOSITES. SAMPE 2009.

51

GENOA Impact Simulation - Damage Propagation

Time Dependent Damage Progression & Contributing Failure Mechanisms Simulation of Impacted Composite Foam

52

GENOA Impact Simulation – Load vs Time & Energy vsTime

Load v.s. Time Energy v.s. Time

0

5

10

15

20

25

30

0 5 10 15 20 25

Time [ms]En

ergy

[ft-l

bf]

Test1Test2

Sim (Fiber/Matrix)Sim (Ply)

53

GENOA Results For Compression After Impact(Compared to Test)

Compression Load = 0 lb Compression Load = 24,680 lb

Test Simulation

Impact Energy 25 lbf-ft 25 lbf-ft

Impact Peak Load 1,400 lb 1,530 lb

CAI Residual Strength

23,280 lb 24,136 lb

Failure Mechanism Asymmetric Crippling of bottom skin

Crippling of skin (Normal Shear Failure)

Summary of Results

Load v.s. Displacement

54

Time=3.715×10-4 SecondAluminum

Composites

Reference: F. Abdi, D. Xie, H. Bhugaloo, J.L. Bozec “PROGRESSIVE FAILURE SIMULATION OF AN AIRCRAFT FAN BLADE IMPACT ONTO ENGINE INLET”. Proceedings of COMP07: 6th International Symposium on Advanced Composites, 16-18 May, 2007, Corfu, Greece.

Hard Body Impact: Engine Blade Out

Composite Blade Out

Titanium Blade;Composite Housing

Inner Layer: Aluminum 6061-T6Out Layer: Kevlar 29 Dry Fiber

Titanium Ti-6Al-4V

Blade initial speed: 4000 in/sec

55

Comparison of Impact Simulations and TestComparison of Impact Simulations and Test

RCC6 RCC7 RCC8

18.6 inches

10.2 inches

Final State Due to Impact

RCC6 RCC7 RCC8

Prior to Final State Due to ImpactShown to Display Fracture Progression

From Simulations

RCC8 looking outboard.Cracking of panel 8 extends all the way to and including the outboard rib of panel 8.

RCC8 looking up.Cracking of panel 8 extends from apex of RCC8 curvature to before the representative carrier panels. RCC7

RCC8

RCC9

From TestsDamage From Foam Impact on RCC Panel 8Damage From Foam Impact on RCC Panel 8

Reference: K. Bowcutt, D. Picetti, K. Yun, F. Abdi “A High-Fidelity Aero-Thermal-Structural Analysis Of The STS-107 Columbia Reentry With Postulated Wing Leading Edge Damage”. JANNAF Conference Paper 2003, Colorado Spring, Colorado.

Hard Body Impact: Shuttle Foam Impact Accident Re-construction

56

Progressive Failure Analysis (PFA): Quasi-Static (Low-Cycle) Fatigue

• Provides an effective evaluation of the fatigue (service) life of the composite/metallic structures subjected to constant amplitude static thermo-mechanical loads.

• Predicts the damage initiation site in structural components under applied cyclic loading

• The degree of cumulative damage incurred is calculated from the SN curve at each stress level; the SN curve can be obtained from tests or literature

• Can be used in conjunction with efficient Multi-factor interaction model (MFIM) to account for a broad range of factors that affect the fatigue life of a structure

• Can be used in conjunction with GENOA’s Probabilistic Analysis module to account for the probabilistic nature of fatigue

• Accounts for SN curve degradation for a considered material

57

Lamina Ply # Mat # OprTmp CureTmp Moisture, angle Thickness

PLY 1 1 70.00 70.000 0.0000 0.00 0.01268

PLY 2 1 70.00 70.000 0.0000 90.00 0.02056

PLY 3 1 70.00 70.000 0.0000 0.00 0.01268

PLY 4 1 70.00 70.000 0.0000 90.00 0.01014

PLY 5 1 70.00 70.000 0.0000 0.00 0.01268

PLY 6 1 70.00 70.000 0.0000 90.00 0.02056

PLY 7 1 70.00 70.000 0.0000 0.00 0.01268

MATCRD 1EGKGDION 0.55000 0.09655EGKGDION 0.00000 0.55000 0.09655 1.00000

BRAID 1EGKGDION 60.00 90.00 30.00 0.0250

BRAID 2EGKGDION -60.00 90.00 30.00 0.0250

Progressive Failure Analysis (PFA): Quasi-Static Fatigue

E-glass fiber; Dion 9800 matrix

Force

Experimental Strain-Stress curve of resin

Strain-Stress Curve of DION

0

3

6

9

12

15

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5

Strain (%)St

ress

(ksi

)

original experimental data representative points picked

Ref: Kunc, V.; Lynn K.; Abdi, F.; Qian, Z.; Knouff, B. The Prediction of 3 Dimensional Fabric Composite Fatigue Sensitivity to Void Content. 2006 SAE World Congress, Detroit, MI, April 3-6, 2006

3-D Woven Composite, Material: 3-Tex

58

Results for 3-D Woven coupon Average 10% Void : Fiber Micro Buckling Longitudinal compression damage (12,500 Cycles)

Longitudinal compression damage of plies after fatigue cycle 12500 (stage 3)

(30% of static ultimate load)

0°

90°

0°

0°

0°

90°

90°

After Matrix damages, Matrix’s Shear Modulus changes to Zero. It comes with fiber microbuckling, which results the composite damage.


59

Longitudinal tensile damage after fatigue cycle 12500 (stage 4) (30% of static ultimate load)

Results for 3-D Woven coupon, Average 10% Void : Longitudinal tensile damage (12,500 Cycles)

0°

90°

0°

0°

0°

90°

90°


60

Failure Progression for 3-D Woven coupon: Matrix (Transverse Tensile), 90 deg Fiber (fiber buckling), and 0 deg fiber (Long tensile)

Cycle 12400

Cycle 12500

After Cycle 12500

No Matrix damage

before cycle 12400

Matrix damage Initiation

Transverse tensile damage

Longitudinal compression

damage

Fiber Micro buckling

Longitudinal tensile damage

No damage


Matrix damage Propagation



61

Comparison between simulation and test fatigue cyclic lives

40

50

60

70

80

90

100

110

120

130

100 1000 10000 100000 1000000

Cycles to Failure

Max

Cyc

lic S

tres

s (k

si)

Test - RTSimulation - RTTest - CryoSimulation - Cryo

Ref: Xiaofeng Su, Frank Abdi, Ron. Kim, “Prediction of Micro-crack Densities in IM7/977-2 Polymer Composite Laminates under Mechanical Loading at Room and Cryogenic Temperatures”, SDM 46, Austin, Texas, 2005.

(Cont’d)


62618.25

2,831.5

13,158.5

Test Ave.

10% void

number of cycles to failure

422

2925

13,000

GENOA

540,000

Test

2% void

466

3,900

534,400

GENOA

1.170%

1.350%

41.130%

Life increase (times)

load

618.25

2,831.5

13,158.5

Test Ave.

10% void

number of cycles to failure

422

2925

13,000

GENOA

540,000

Test

2% void

466

3,900

534,400

GENOA

1.170%

1.350%

41.130%

Life increase (times)

load

10

15

20

25

30

35

40

45

50

100 1000 10000 100000 1000000number of cycles to failure

stre

ss (k

si)

axial-Experimental transverse-ExperimentalGENOA predicted (10% void) GENOA predicted (2% void)Experimental

~40 times life

2% Void

10% Void


Comparison of life cycles between Test and Predictions: Void Effect

63

Progressive Failure Analysis (PFA): Harmonic (High-Cycle) Fatigue

• Provides an effective evaluation of the fatigue (service) life of the composite/metallic structures subjected to constant amplitude dynamic thermo-mechanical loads.


• The degree of cumulative damage incurred is calculated from the SN curve at each stress level; the SN curve can be obtained from tests or literature




64

Progressive Failure Analysis (PFA):Spectrum (Random) Fatigue

• Provides an effective evaluation of the fatigue (service) life of the composite/metallic structures subjected to a sequence of external cyclic thermo-mechanical excitations with variable amplitude and period.


• Rainflow analysis is used to reduce the complex loading to a series of simple cyclic loadings

• The degree of cumulative damage incurred is calculated from the SN curve at each stress level; the individual contributions are combined using Miner’s rule




65

Progressive Failure Analysis (PFA):Fatigue with Fracture Mechanics (PFA / FATIGUE / VCCT)

• Provides an effective evaluation of the fatigue life of a metallic structure by combining the PFA with VCCT and FCG modules

• Analyzes the development of the damage initiation site along a predetermined path under the applied cyclic loading

• The approach is fracture mechanics based and requires:– A predetermined fracture path– Fatigue Crack Growth curve (da/dN vs. ∆K data); can be obrained from GENOA/FCG– SN curve


• The approach is restricted to linear elastic materials and shell elements

66

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00

1 10 100

DK (psi in^0.5)

da/d

N (i

n/cy

cle)

Simulation of Boeing IAS (Integral Airframe Structures) using da/dN vs ∆K input and VCCT utilization

FCG curve (da/dN v.s. ∆K) for T7457 is generated by GENOA/FTD/FCG

Material: T7475-T7351, Initial crack: 2.5in, Load: pressure load on internal surfaces

Fuselage panel (8.6psi)

Test GENOA

Cycle to Fracture

10,333 10,230

Static Loading (ultimate)

10.30 psi ~10.48 psi

Ref: B. Farahmand, C. Saff, De Xie and F. Abdi, “Estimation of Fatigue and Fracture Allowables For Metallic Materials Under Cyclic Loading”. AIAA-2007-2381, Honolulu, Hawaii, April, 2007

Progressive Failure Analysis (PFA): Fatigue with Fracture Mechanics (PFA / FATIGUE / VCCT)

68

Progressive Failure Analysis (PFA):Virtual Crack Closure Technique [VCCT]

• Virtual Crack Closure Technique (VCCT) is a fracture mechanics based approach for PFA integrated into GENOA-PFA:

– The technique is based on linear spring elements and is insensitive to FEM mesh size– Avoids the use of singular crack elements (extensive mesh preparation work not

required)– Computationally efficient due to the use of the node-based displacements and forces– Only Fracture toughness data is required as input and can be obtained from

the following sources:• Experimental testing/Material Handbooks• GENOA Fracture Toughness Determination (FTD) Module

– Requires fracture path to be pre-determined based on the following:• Experimental testing/Experience• Preliminary Progressive Failure Analysis

– Computes strain energy release rate in linear elastic materials– Provides accurate failure analysis of interfaces and adhesively bonded joints– Able to detect critical crack propagation and arrest

0

10

20

30

40

50

60

70

0 5 10 15

D e f l e c t i o n ( mm)

69

X

Y

1 2

3 4

Δa

Ref:Rybicki EF, Kanninen MF. A finite element calculation of stress intensity factors by a modified crack closure integral. Engng

Fract Mech 1977;9:931-938.Xie D and Biggers, Jr. SB, Strain energy release rate calculation for a moving delamination front of arbitrary shape based on

virtual crack closure technique, Part 1 and Part 2, Engineering Fracture Mechanics, accepted.Xie D and Biggers, Jr. SB, Progressive crack growth analysis using interface element based on the virtual crack closure

technique, Finite Elements in Analysis and Design, submitted.

Requires pre-defined crack path,

Needs fracture toughness test data.

If ply configuration is changed, repeated toughness test required.

Orthotropic Material model


70

0

10

20

30

40

50

60

70

0 2 4 6 8 10 12 14

Deflection (mm)

Load

(N)

ABAQUS/UELMesh 400, increment 0.04 mmMesh 400, increment 0.4 mmMesh 1200, increment 0.4 mmCrisfield

a=2.832L=10

b=1

2h=2×0.1924P, δ

P, δ

x

y

z

Ex = 20.6 MSI, Ey = Ez = 1.13 MSIGxy = 0.58 MSI, Gxz = 0.58 MSI, Gyz = 0.37 MSIVxy = 0.34, Vxz = 0.34,Vyz = 0.53GIC = 1.89 lb/in

Unit: Inch

PFA–VCCT offers• PFA offers damage tracking approach no crack path definition • PFA detects the failure mechanisms involved at Micro, and

macro-level• PFA-VCCT offers load displacement contour plots• PFA –VCCT can be simulated with NASTRAN

PFA Defined Propagation Path


71

Multi-Site cracks - plate containing two parallel cracks

F2

Crack a

Crack bF1

Multi-cracks - plate containing two perpendicular cracks

Progressive Failure Analysis (PFA): Virtual Crack Closure Technique [VCCT]

72

Progressive Failure Analysis (PFA):Discrete Cohesive Zone Modeling [DCZM]

• Discrete Cohesive Zone Model (DCZM) is a fracture mechanics based approach for progressive crack growth analysis in GENOA-PFA:

– The technique is based on bi-linear spring elements and is insensitive to FEM mesh size– Avoids the use of singular crack elements– Computationally efficient due to the use of the node-based displacements and forces– Requires fracture toughness, cohesive strength, and cohesive stiffness as

input:• Fracture toughness can be obtained from Experimental testing/Material Handbooks/GENOA

Fracture Toughness Determination (FTD) Module• Cohesive strength and stiffness can be obtained via test calibration

– Requires fracture path to be pre-determined based on the following:• Experimental testing/Experience• Preliminary Progressive Failure Analysis

– Enables modeling material softening– Provides accurate failure analysis of interfaces and adhesively bonded joints– Predicts face-sheet core delamination in sandwich materials– Able to detect critical crack propagation and arrest

73

Discrete Cohesive Zone Model (DCZM) Continuum Cohesive Zone Model (CCZM)

X

Y

1 2

Δa

X

Y

1 2

3 4σ(MPa)

σc

GIC δ(mm)δc δm

Cohesive Law

Nonlinear Spring Continuum

ICmc G=δσ21

Ref:Ungsuwarungsri T and Knauss WG, The Role of Damage-Softened Material Behavior in the fracture of composites and adhesives, International Journal of Fracture, 35(1987): 221-241. See also Ph.D. thesis, 1985, Aeronautics Department, Caltech, Pasadena, CA.

Song SJ and Waas AM, Energy-based mechanical model for mixed-mode failure of laminated composites, AIAA Journal, 33 (1995): 739-745

Shahwan KW and Waas AM, Non-self-similar decohesion along a finite interface of unilaterally constrained delaminations, Proceedings of the Royal Society of London. Series A: Mathematical and Physical Sciences, 453 (1997): 515-550.

PFA-DCZM can model material softening that VCCT can notSame Input requirements as VCCTThis Unique capability does not exist with Abaqus

Progressive Failure Analysis (PFA):Discrete Cohesive Zone Modeling [DCZM]

74

GENOA-DCZM Benchmarks: CC, IC, CT

8.0

w2a0° Fiber

(a) CC2

16.0

w2a

0° Fiber

(b) CC4

8.0

w2a’

0° Fiber

(c) IC

1.70

1.75w

a

0.385

0°Fiber

(d) CT

Ref: Masters JE, Trans-laminar Fracture Toughness of a Composite Wing Skin Made of Stitched Warp-knit Fabric, NASA Contractor Report 201728, November 1997.

Illustrate the procedure for calibration Calibrate cohesive strength using CC2 specimenPredict load displacement curve for CC4, IC, and CT specimens

75

GENOA-PFA Predicts The Crack Path And Residual Strength: Benchmark: CC2

[+45/-45/02/90/02/-45/+45] with 44% as 0° plies

44% as ±/45° pliesand 12% as 90° plies

Test PFA/NASTRAN

64.5Ksi 67.2Ksi

W = 2.0in, a = 0.5in, L = 4.0in

Ref: J. Mastres., NASA Contractor Report 201728 “ Translaminar Fracture Toughness of a Composite Wing Skin Made of Stitched warp-Knit Fabric”

Number of Mindlin Shell Elements: 1936

Fracture starts at load: 43.7Ksi

Final Fracture load: 67.2 Ksi

AS4/3501

PFA shows that the crack propagates perpendicular to the load direction

76

GENOA-DCZM Calibration ProcessBenchmark: CC2

0.0E+00

1.5E+04

3.0E+04

4.5E+04

6.0E+04

7.5E+04

9.0E+04

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2

COD (inch)

Stre

ss (p

si)

Test

VCCT

DCZM,100%

DCZM, 50%

DCZM, 70%

DCZM,65%

0.0E+00

1.5E+04

3.0E+04

4.5E+04

6.0E+04

7.5E+04

9.0E+04

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2

COD (inch)

Stre

ss (p

si)

Test

VCCT

DCZM,100%

DCZM, 50%

DCZM, 70%

DCZM,65%

Test: 64.5ksi errorVCCT: 52.5ksi; –19%. DCZM: 64.6ksi; +0.2%.

Calibration: the cohesive strength is calibrated as the percentage of material axial tensile strength

Cohesive Strength ~ 65% of TS

8.0

2.0

2a

0° Fiber

Load Displacement Curves for Cohesive Strength Selection

Ref:Masters JE, Translaminar Fracture Toughness of a Composite Wing Skin Made of Stitched Warp-knit Fabric, NASA Contractor Report 201728, November 1997.

77

GENOA-DCZM Corrects Load Displacement Curve And Residual Strength Benchmark: CC4

0.0E+00

1.5E+04

3.0E+04

4.5E+04

6.0E+04

7.5E+04

9.0E+04

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2

COD (inch)

Stre

ss (p

si)

testVCCTDCZM , 65%

0.0E+00

1.5E+04

3.0E+04

4.5E+04

6.0E+04

7.5E+04

9.0E+04

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2

COD (inch)

Stre

ss (p

si)

testVCCTDCZM , 65%

Test: 58.8ksiVCCT: 49.5ksi; –16%. DCZM: 57.6ksi; –3%.

16.0

4.02a

0° Fiber

CC4

The calibrated Cohesive Strength (~ 65% of TS) is used to make the prediction for wide central crack panel (4.0inch wide).

Ref: Masters JE, Translaminar Fracture Toughness of a Composite Wing Skin Made of Stitched Warp-knit Fabric, NASA Contractor Report 201728, November 1997.

78

GENOA-PFA Predicts The Crack Path And Residual Strength Benchmark: IC

W = 2.0in, a = 0.5in, L = 4.0in

[+45/-45/02/90/02/-45/+45]with 44% as 0° plies


Ref: J. Mastres., Nasa-Contractor Report 201728 “ Translaminar Fracture Toughness of a Composite Wing Skin Made of Stitched warp-Knit Fabric”

Test PFA/NASTRAN

61.3ksi 65.5 Ksi

Number of Mindlin Shell Elements: 4488

Fracture starts at load: 24.0 Ksi

Final Fracture load: 65.5 Ksi

AS4/3501


79

GENOA-DCZM Corrects Load Displacement Curve And Residual Strength Benchmark: IC

0.0E+00

1.5E+04

3.0E+04

4.5E+04

6.0E+04

7.5E+04

9.0E+04

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2

COD (inch)

Stre

ss (p

si)

VCCT

DCZM, 65%

0.0E+00

1.5E+04

3.0E+04

4.5E+04

6.0E+04

7.5E+04

9.0E+04

0 0.02 0.04 0.06 0.08 0.1 0.12 0.14 0.16 0.18 0.2

COD (inch)

Stre

ss (p

si)

VCCT

DCZM, 65%

Test: 61.3ksi errorVCCT: 45.0ksi; –26%. DCZM: 60.0ksi; –2%.

8.0

2.0

2a’

0° Fiber

IC

The calibrated Cohesive Strength (~ 65% of TS) is used to make the prediction for inclined central crack panel.

Ref: Masters JE, Translaminar Fracture Toughness of a Composite Wing Skin Made of Stitched Warp-knit Fabric, NASA Contractor Report 201728, November 1997.

80

GENOA-PFA Predicts the Crack Path and Residual Strength Benchmark: CT

Test PFA/NASTRAN

2.31 Kip 2.47 Kip

W = 1.4in, a = 0.5in, L = 4.0in

Ref: J. Mastres., Nasa-Contractor Report 201728 “ Translaminar Fracture Toughness of a Composite Wing Skin Made of Stitched warp-Knit Fabric”

[+45/-45/02/90/02/-45/+45]with 44% as 0° plies


Fracture start at load:2.04Kips

Final Fracture load:2.47Kips

AS4/3501


81

Fracture Toughness Determination (FTD)

• Computes fracture toughness based on the extended Griffith energy balance theory

• Predicts the fatigue life and performance of metallic structures (aluminum, titanium, steel) with minimal experimental testing

• Cost and time effective approach for sophisticated fracture testing--eliminates high standard specimen preparation, complex test procedures and data processing

• Requires only the stress/strain curves for tensile specimens as input. The SS curve can be obtained from:

– Experiments– Handbooks– Other literature

82

Fatigue Crack Growth (FCG)

• Computes da/dN vs. ∆K curves• Predicts the fatigue life and performance of metallic structures (aluminum,

titanium, steel) with minimal experimental testing• Cost and time effective approach for sophisticated fracture testing--eliminates

high standard specimen preparation, complex test procedures and data processing

• Requires fracture toughness as input. The fracture toughness can be obtained from:

– Experiments– Handbooks– Other literature– GENOA/FTD module

83

Flow Chart: Fatigue Crack Metal Approach

• KC• KIC• Stress Ratio• Maximum Stress• Yield Stress

•ASTM E647OR

FTD

Tensile test

Part I

PFAPart II

Part III

FCG

Test

Outputs

FEA Model component

component

material

Ref: B. Farahmand, C. Saff, De Xie and F. Abdi, “Estimation of Fatigue and Fracture Allowables For Metallic Materials Under Cyclic Loading”. AIAA-2007-2381, Honolulu, Hawaii, April, 2007

84

Flow Chart: Application of FTD and FCG

InputStress-Strain Curve (From

a Reliable Source)

Fracture Toughness

DeterminationFTD

Fracture Toughness Versus Thickness (2219-T87)

01020304050607080

0 0.5 1 1.5 2 2.5

Thickness, inch

Kc

-ksi

(in.)^

0.5

Small Plate'W=10", 2a=3"NASGRO

KIc

NASGRO

FTD

Delta K, ksi-(in)^0.5

1.E-09

1.E-08

1.E-07

1.E-06

1.E-05

1.E-04

1.E-03

1.E-02

1.E-01

1.E+00

1 10 100

da/d

n, in

/cyc

le

10 100

FCG

PhysicalTesting -

NASGRO

2219-T87

85

Composite Over-wrappedPressure Vessels (COPVs)

86

Filament Winding (FW)

• Supports design and analysis of composites over-wrapped vessels (COPVs)• Generates mesh for circular cylinder and dome shapes (Spherical, Geodesic,

Elliptic, Toroidal)• Allows for importing external mesh• Allows to model linerless COPVs• Allows for modeling bonded, unbonded, and frictional behavior between the

linear and the composite over-wrap• Duplicates the manufacturing process by generating the correct tape schedule at

each location on the COPV FEM model and calculating residual stresses caused by the filament winding process

• Simplifies the complex Hoop/Helical wrapping definitions for the user and need to requires:

– Winding Angle– Tape overlap– Number of windings

• Generated model can be used with Progressive Failure Analysis for simulating the certification process

87

COPV Delamination Initiation / Progression and Fracture Simulation

Delamination Initiation in Tank 3(Pressure 3,200 psi)

Fracture Initiation in Tank 3(Pressure 5,040 psi)

Delamination Progression in Tank 3(Pressure 4,480 psi)

Fracture Test/Prediction Comparison

Test : 4,890 to 5,303 psiTest Average: 5,057 psi

GENOA: 5,040 psi

Prediction burst pressure is 0.33 % lowerthan the average test pressure

Ref: G. Abumeri, F. Abdi, M. Baker, M. Triplet and, J. Griffin “Reliability Based Design of Composite Over-Wrapped Tanks”. SAE World Congress, 2007, 07M-312, Detroit Mi, April 2007

88

Reference: G. Abumeri, F. Abdi, M. Baker, M. Triplet and, J. Griffin “Reliability Based Design of Composite Over-Wrapped Tanks”. SAE World Congress, 2007, 07M-312, Detroit Mi, April 2007

Probabilistic Evaluation

Test Fracture Internal Pressure4,890 to 5,303 psi corresponds to cumulative probability of 0.425 to 0.70;

Test Average: 5,057 psi GENOA 50% Probability Prediction: 4,950

Test

Test

Methodology is applicable to all types of materials and structures

cont’dFracture Initiation Load of Composite Tank

89

Probability & Reliability Analysis

90

Progressive Failure Optimization (PFO)

• Helps maximize the structural performance of a designed composite structure by directing the design towards minimization of its service life damage.

• Optimizes the durability and damage tolerance of the part in consideration and improves the overall robustness

• Helps reduce experimental efforts by computing the optimized design parameters that produce the best performance

• Minimizes objective function in the overall material damage• Graphically shows the comparison of the optimized results and initial design• Requires the selection of manufacturing design variables and constraints• Recalculates the stress distribution at the micro level based on the input loads

91

Probabilistic Analysis (PA)

• Permits the simulation of progressive failure in composite structures taking into consideration uncertainties in material properties, loading conditions and service and manufacturing environments

• Generates various probabilistic responses and sensitivities based on the user defined perturbed random variables

• Can predict the complicated response of composite structures using the following:

– Cumulative Distribution Function (CDF)– Probability Density Function (PDF)– Probabilistic Sensitivities– Random variables most probable design vectors

• Requires the identification of variables and definition corresponding mean values, standard deviation, and distribution types.

• Helps rationalize between competing designs• Facilitates the assessment of the reliability of advanced materials and structures

for a variety of applications

92

Probabilistic Progressive Failure Analysis (PPFA)

• Evaluates the reliability of a structure in presence of the following:– Uncertainties in the constituent properties of the composite (or properties of metal)– Fabrication variables and geometry– Service conditions

• Enables quantify the relative effect of the various random variables on the performance of the structure

• Calculates the following– Cumulative Distribution Function (CDF)– Probability Density Functions (PDF) of the response– Sensitivities of design parameters to the response

• Automatic linking to major FEA solvers (e.g., NASTRAN, ANSYS, and ABAQUS). PPFA perturbs the random variables and automatically extracts response from the structural solver

• Choice of using either high fidelity methods (Advance Mean Value and Monte Carlo

93

Fabrication Process is a Primary Source ofVariation (Scatter) in Composites

Other Sources of Scatter Include: Loading, Environment, and Service

94

Ref: G. Abumeri, F. Abdi, M. Baker, M. Triplet and, J. Griffin “Reliability Based Design of Composite Over-Wrapped Tanks”. SAE World Congress, 2007, 07M-312, Detroit Mi, April 2007

Probabilistic Evaluation of Fracture Initiation Load of Composite Tank

Test Fracture Internal Pressure4,890 to 5,303 psi corresponds to cumulative probability of 0.425 to 0.70;

Test Average: 5,057 psi GENOA 50% Probability Prediction: 4,950

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