1

Aerospace Structures and Materials:

Composite Failure

Dr. Tom Dragone

Orbital Sciences Corporation

2

Structure Design / Analysis Process

BOX BEAM ANALYSIS Component Loads (Cap Forces, Shear Flow)

BOX BEAM ANALYSIS Component Loads (Cap Forces, Shear Flow)

JOINT LOADS Weld , Braze Bond, Bolt

Metal Yield Rupture

Composite FPF LPF

Stability Buckling Crippling

Fracture Toughness Crack Size

Fatigue Crack Initiation Crack Growth

MS>0?MS>0?

SHEAR-MOMENTDIAGRAM Section Loads

GLOBAL LOADS Aerodynamics Inertial Applied

GEOMETRY Planform Skin Construction Spar/Rib Layout

SIZING Thickness Ply Orientation

MATERIALS Metal Composite

StructureIdealization

Stiffness Lamination Theory

Done

FAILURE ANALYSIS

Yes No

3

Motivation

• Composite failure is very different from metal failure

Discussion Questions:• How does a composite “yield”? Does Von Mises or Tresca hold?• How does a composite “fail” or “rupture”? What are some of the

mechanisms involved?• Are composites better or worse than metals under fatigue

loading?• How would a composite fracture? Does LEFM apply?• What additional failure modes are possible with composites?

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Failure Envelopes

• Metal Failure: Homogeneous and Isotropic

• Composite Failure: Inhomogeneous and Anisotropic

VON MISES: 1

2

2221

2

1

tytyty FFF

TRESCA: syF2,,max 2121

1

2 COMPOSITE:

5

Stress-Strain Behavior

METAL

BIDIRECTIONALLAMINATE

UNIDIRECTIONALLAMINATE

Yield

FPF

LPF

Ultimate

FPF, LPF

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Ply Failure

• First Ply Failure (FPF)– Similar to yield

– First indication of non-reversible deformation

– Change in slope of loading curve (non-linear)

– Laminate has residual load-bearing potential

• Last Ply Failure (LPF)– Similar to Ultimate

– No more load bearing potential

– Rupture

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Ply Failure Criteria

First Ply Failure Criteria• Maximum Stress• Maximum Strain• Hill (Maximum Distortion Energy)• Tsai-Wu (Quadratic)• Matrix Tension• Matrix Compression

Last Ply Failure Criteria• Fiber Tension• Fiber Compression

No Description ofFailure Mechanism

IndicatesFailure Mechanism

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Failure Analysis Implementation

• “Weakest Link” Analogy– Failure criteria apply at the ply level

– When one layer fails, the entire laminate fails

• Which Failure Criteria to Use?– Depends on the particular fiber/matrix combination

– Must test to determine most appropriate criteria

• Failure Envelopes for Composites are Rarely Used– Complex ply interactions make visualization difficult

– Sometimes can be helpful for a particular laminate

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Failure Criteria

111 SYXxyyx

Maximum Stress

Maximum Strain

122

2

2

SYXXxyyyxx

Hill (Max Energy)

111 S

G

Y

E

X

E xyxyyyxx

Tsai-Wu

12

1111

2

222

yxijxyyx

yx

FSYtYcXtXc

YcYtXcXt

X = LongitudinalStrength Y = Transverse

StrengthS = ShearStrength

Xt = TensileStrength

Xc = CompressiveStrength

Fij = EmpiricalFactor ~ -0.5

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Failure Criteria

122

SYxyy

Matrix Tension

1122

222

SYcS

Yc

Sxyyy

Matrix Compression

Fiber Tension 122

SXtxyx

1Xc

xFiber Compression

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Stress Space Failure Envelope

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

-300

-200

-100

0

100

200

300

400

Tra

nsv

ers

e S

tre

ss (

ksi)

Longitudinal Stress (ksi)

MaxStress MaxStrain Hill TsaiWu

-400 -300 -200 -100 0 100 200 300 400-15

-10

-5

0

5

10

15

Tra

nsv

erse

Str

ess

(ksi

)

Longitudinal Stress (ksi)

MaxStress MaxStrain Hill TsaiWu

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Strain Space Failure Envelope

-15 -10 -5 0 5 10 15-15

-10

-5

0

5

10

15

Tra

nsve

rse

Str

ain

(m)

Longitudinal Strain (m)

MaxStress MaxStrain Hill TsaiWu

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Progressive Damage Models

• FPF Usually Implies Transverse Failure of Matrix– Fiber can still continue to bear load

– Does not cause rupture

– Causes change in failed ply stiffness

• Set Ply Transverse Modulus and Shear Modulus = 0• Load is Shifted to Other Layers• Other Plies MAY Fail Leading to FPF = LPF

or

• Stable Equilibrium Reached Such That Laminate Can Take More Applied Load

• Process Continues Until Fiber Failure Occurs in Weakest Ply• Progressive Damage Models Typically Used in Failure

Investigations, Not in Design Because They are Cumbersome

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COMPFAIL Process

• Apply Loads• Return Strains and Curvatures• Return Equivalent Moduli (For Symmetric Laminates ONLY)• Return Ply Strains and Ply Stresses

– 1, 2, 6, 1, 2, 6 for Global (Laminate) Coordinate System

– x, y, s, x, y, s for Local (Material) Coordinate System

Two Values:Top and Bottom

of Ply

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COMPFAIL Failure Analysis Process

• Calculate Failure Criteria for Each Ply

22

2

2

SYXXxyyyxx

22

SXtxyx

Xcx

yxijxyyx

yx

FSYtYcXtXc

YcYtXcXt

2

1111

2

222

22

SYxyy

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COMPFAIL Failure Analysis Process

• Calculate Failure Criteria for Each Ply• Calculate R Value for Each Ply

– R = Factor x Applied Load That Gives Failure Index = 1

– R ~ 1/(Failure Index)^2

122

S

R

Y

R xyy

22

SYxyy

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COMPFAIL Failure Analysis Process

• Calculate Failure Criteria for Each Ply• Calculate R Value for Each Ply• Search for Minimum R Value Through Thickness

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COMPFAIL Failure Analysis Process

• Calculate Failure Criteria for Each Ply• Calculate R Value for Each Ply• Search for Minimum R Value Through Thickness• Summarize Values

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COMPFAIL Failure Analysis Process

• Calculate Failure Criteria for Each Ply• Calculate R Value for Each Ply• Search for Minimum R Value Through Thickness• Summarize Values

Color Code:Green = FI > 1.5Yellow = 1.25 < FI < 1.5Red = FI<1.25

Color Code:Green = FI > 1.5Yellow = 1.25 < FI < 1.5Red = FI<1.25

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Other Failure Mechanisms

Failure

Mechanism

Characteristics

HygroscopicSwelling

Organic polymer matrices tend to absorb moisture Absorbed moisture causes the polymer to swell, resulting

in stress if the volume is constrained Composite swelling described by Moisture Expansion

Coefficient, analogous to Thermal Expansion Coefficient Hot/Wet properties can be 30% less than RT properties

Delamination Separation between plies in a laminate or between thecore and the skin of a sandwich structure

Very difficult to predict Usually requires fracture mechanics approach to

determine stable or unstable energy release rates

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Delamination

045

-459090

-45450

Crack Initiation

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Delamination

Between Plies

Interface

Delamination Growth

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Other Failure Mechanisms

Failure

Mechanism

Characteristics

ImpactDamage

Impact may be caused by dropped tools (low velocity),Foreign Object Damage (FOD) kicked up from runway,hail, bird strikes, ballistic impact, hypervelocity impact ofmicrometeoroid or orbital debris (high velocity)

Impact may cause damage that is undetectable (matrixcracking within laminate), visible (usually on the rear sideof a laminate) or complete penetration

Impact damage may be matrix cracking, delamination,skin debond, or fiber breakage

Greater impactor energy => greater damage Tougher matrix => less damage Impact damage may cause ultimate failure immediately

(rupture of a tank), or may be the site of crackpropagation for subsequent failure

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Impact Damage

Impact Visible Damage

Ultrasonic Image

Internal Damage

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Impact Damage

Internal Rib Damage

Core Damage

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Other Failure Mechanisms

Failure

Mechanism

Characteristics

Fatigue Fatigue in composites is generally better than metalsbecause the fibers act to deflect the crack and stop crackgrowth

Exact mechanisms are complex, but follow same generalpattern as for metals:

LCF: Failure set by ultimate strain of material MCF: Allowable strain decreases with number of cycles HCF: Below minimum strain threshold, composites have

infinite fatigue life because matrix does not crack, so nocracks can grow

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Other Failure Modes

• th ~ 6000 for many resins

• Design Below This to Eliminate Microcracking and Fatigue Damage

High Cycle Fatigue

FATIGUE

th

Strain

Cycles

Low Cycle Fatigue

c

Matrix CrackingInterface Shear

Fatigue Limitfor Matrix

Fiber BreakageInterface Debonding

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Delamination

045

-4590

90-45

450