[PPT] - Damage Tolerance Philosophy for Bonded Aircraft Structures Towards PowerPoint Presentation

SLIDE 1

Challenge the future

Delft University of Technology

6-7-2009

Dr. C.D. Rans

27-05-2009

Damage Tolerance Philosophy for Bonded Aircraft Structures

Towards a generic assessment approach

Aerospace Materials

Faculty of Aerospace Engineering

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2 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Possible Damage Scenarios

Generic Bonded Structure

Adherent damage Bond line damage Adherent and bond line damage

Damage tolerance approach to deal with all three scenarios

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3 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

I Have Seen This Before!

Coupled Adherent and Bond Line Damage

Fatigue crack growth metals Delamination at metal-fiber interfaces

Crack opening constraint

Fatigue crack growth metals Delamination at metal-fiber interfaces

Crack opening constraint

FMLs

Can this be translated to generic bonded structures?

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4 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

1. Prediction Approach

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5 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

What Do We Need?

Bond line delamination growth
Adherent damage growth
Damage interaction

Required Prediction Approaches

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6 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Adherent Crack Growth

Fracture Mechanics description using Paris type relation
Mode I crack growth
Empirical relation for R-ratio effects

Prediction Approach

( )(

)

2 max

0.55 0.33 0.12 1

eff

K R R R K Δ = + + −

( )

cg

n cg eff

da C K dN = Δ

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7 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Bond Line Delamination Growth

Fracture Mechanics description using Paris type relation
Mode II delamination growth
Reformulated strain energy release rate range

Prediction Approach

( )

max min max min

2 II II II II II

G G G G G Δ ≠ − = − ( ) d

n d II

db C G dN = Δ

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8 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Mode II Growth Assumption

Adhesive joints designed to

transfer load through shear

Mode II assumption supported by

test data

FML delamination characterization
Composite laminates
Damage growth, not initiation

Bond Line Delamination Growth

√Gmax - √Gmin [MPa mm] db/dN [mm/cycle]

0.2 0.4 0.6 0.8 1 10

9

10

8

10

7

10

6

10

5

10

4

10

3

10

2

Data obtained from Alderliesten et al. (2006)

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9 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Strain Energy Release Rate Range

Linear elastic fracture mechanics description
Principle of Similitude
Principle of Superposition

2 (1) (2) (3) 2 (1) (2) (3) 2 (1) (2) (3) T I II III I I I I II II II II III III III III

G G G G G G G G G G G G G G G G = + + ⎡ ⎤ = + + + ⎣ ⎦ ⎡ ⎤ = + + + ⎣ ⎦ ⎡ ⎤ = + + + ⎣ ⎦

Bond Line Delamination Growth

( )

max min max min

2 II II II II II

G G G G G Δ ≠ − = −

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10 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Strain Energy Release Rate Range

Illustrate with Mode I DCB specimen

Bond Line Delamination Growth

2

I

P dC G b da =

For identical crack and specimen geometries

2 I

G const P = ⋅

Compare applied loading to obtain the same ΔG for 2 different R-ratios

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11 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Strain Energy Release Rate Range

Bond Line Delamination Growth

2 I

G const P = ⋅

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12 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

(√Gmax - √Gmin)

2 [(J/m 2)]

db/dN [m/cycle]

10

1

10

2

10

3

10

10

10

9

10

8

10

7

10

6

10

5

Post-stretched (C2) Post-stretched (C4) Post-stretched (C6) As cured (C2) As cured (C4) As cured (C6)

Gmax - Gmin [J/m

2]

db/dN [m/cycle]

10

1

10

2

10

3

10

10

10

9

10

8

10

7

10

6

10

5

Post-stretched (C2) Post-stretched (C4) Post-stretched (C6) As cured (C2) As cured (C4) As cured (C6)

Strain Energy Release Rate Range

Advantages of new formulation
Removal of residual stress influence on SERR range
Permits use of superposition in analysis
R-ratio effects can be studied

Bond Line Delamination Growth

Data obtained from Lin and Kao (1996)

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13 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Damage Interaction

Discretize damage area
Enforce displacement compatibility between adherents
Determine load redistribution resulting from damages
Superposition of load redistribution on damage growth

Prediction Approach

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Displacement Compatibility

Damage Interaction

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15 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Displacement Compatibility

Cracked adherent
Undamaged adherent

1 br

P P

v v v = −

Damage Interaction

2 br

P P ad

δ δ δ δ = + +

br

v P δ = →

Over delamination length, b

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16 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Influence on Damage Growth

Superposition of bridging load
Implementation
Analytical solutions
FEM

Damage Interaction

1

1 2

( , , )

br

P P br

K K K G f P P P = − =

Discretization into 1-D damage interaction zones

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17 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Case Studies

2.

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18 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Bonded Patch Repair

Predict edge delamination growth behaviour
Adherents undamaged
Delamination growth model only
Initiation assumed at 1st load cycle
Approximate as a 1-D problem

Case Studies

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19 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Bonded Patch Repair

Good agreement in damage

growth rates

Shift in data due to damage

initiation behaviour

Case Studies

N [kcycles] b [mm]

20 40 60 80 100 20 40 60 80 100 Smax = 106 MPa Smax = 120 MPa mode II prediction

experimental measurement

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20 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

FML Panel with a Bonded Strap

Cracked Glare panel with an intact

bonded titanium strap

Damages
Cracked metal layers of FML
Delamination between metallic and

fibre FML layers

Delamination between strap and

FML

Superposition of multiple bridging

effects

Case Studies

C-scan

, ,

,

farfield br FML br stiffner FML stiffener

K K K K G G = − −

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21 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

FML Panel with a Bonded Strap

Prediction
Intact central strap

0.00002 0.00004 0.00006 0.00008 0.0001 0.00012 0.00014 0.00016 0.00018 0.0002 12.5 17.5 22.5 27.5 32.5 37.5 42.5 47.5 52.5 57.5 Crack length [mm] da/dN [mm/cycle] 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 12.5 17.5 22.5 27.5 32.5 37.5 42.5 47.5 52.5 x [mm] Crack opening V(x) [mm]

2 4 6 8 10 12 14 16 18 20 12.5 17.5 22.5 27.5 32.5 37.5 42.5 47.5 52.5 x [mm] b [mm] front left front right Rear Left Rear Right Prediction

Case Studies da dN

delamination shape crack opening displacement

Data obtained from Rodi (2007)

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22 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

FML Panel with a Bonded Strap

Prediction
Broken central strap

0.0005 0.001 0.0015 0.002 0.0025 0.003 12.5 22.5 32.5 42.5 52.5 62.5 72.5 82.5 Crack length [mm] da/dN [mm/cycle] 0.05 0.1 0.15 0.2 0.25 12.5 22.5 32.5 42.5 52.5 62.5 72.5 82.5 x [mm] Crack opening V(x) [mm]

Case Studies

, , farfield br FML br stiffner

K K K K = − + da dN

crack opening displacement

Data obtained from Rodi (2007)

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23 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Metallic Skin with Bonded Stiffeners

Predict crack growth in stiffened

panel

Aluminum skin
7 bonded aluminum stringers
7 bonded aluminum straps
Central stringer initially broken
Stiffener failure assumption
Superimpose bridging effects of all

stringers

Case Studies

1,

,

farfield stiffeners stiffner stiffnerN

K K K G G = +∑

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24 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Metallic Skin with Bonded Stiffeners

Case Studies

a [mm] da/dN [mm/cycle] N [cycles]

50 100 150 200 250 10

3

10

2

10

1

10 5000 10000 15000

Broken Stringer 1

st Doubler

1

st Stringer

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25 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Summary

Damage tolerance analysis philosophy
Simultaneous analysis of adherent and bond line damage
Linear elastic fracture mechanics description of damage growth
Damage interaction through displacement compatibility and

superposition

Philosophy demonstrated to work for bonded metallic and hybrid

structures

Cracked metal adherents
Bond line delamination growth
Crack opening displacement

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26 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Summary

Potential for composite structures
Stiffness reduction in damaged composite
Determination of load redistribution through displacement

compatibility (superposition of effects)

Requires further work and understanding of composite damage

growth

Potential power of superposition and linear elastic fracture

mechanics for delamination growth prediction

Proper formulation of SERR range

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27 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

Questions?

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28 Damage Tolerant Philosophy for Bonded Aircraft Structures | 27

References

R.C. Alderliesten, J. Schijve, and S. van der Zwaag, Application of the SERR

approach for delamination growth in Glare. Eng. Fract. Mech., 73(6), 2006, pp. 697-709.

C.T. Lin and P.W. Kao, Delamination growth and its effects on crack propagation

in carbon fibre reinforced aluminum laminates under fatigue loading. Acta Mater., 44(3), 1996, pp. 1181-88.

R. Rodi, The effect of external stiffening elements on the fatigue crack growth in

Fibre Metal Laminates. Masters thesis, 2007, Delft University of Technology, Delft, the Netherlands