Damage Tolerance and Damage Growth in Composite Aerostructures

P E IrvingCranfield University

Workshop on NDT/SHM requirements for Aerospace compositesNational Composites Centre, 9/10 February 2016

Plan of Presentation

• Damage processes in metal aircraft

• The role of NDT and inspection in Damage

tolerance

• Damage processes in Polymer composites

• Damage Tolerant aircraft in polymer composites

Damage growth aluminium aircraft

structuresD

am

age

or

crack

siz

e

Crack length at

failure

Limit of NDT detection

Manufacture defect size

Service

life

1st

inspection

Intermediate

Inspections

Inspection

Threshold

Crack growth development determined by

material constants (including defects), by

structure design and by service stress spectrum

3

Visual inspection for cracks JAL Boeing

757 aircraft

0.0

0.2

0.4

0.6

0.8

1.0

0 20 40 60 80 100 120 140 160

Crack length mm

Pro

ba

bil

ity

of

det

ecti

on

With prior

information

without prior

information

4

90% probability

of detection

5

High Strength Metallic Materials; Difficulties with

traditional approaches to damage tolerance

Cra

ck l

ength

Detectable crack

size

Crack length at

failure

Service life

Initiation period

prolonged

First

inspection

Period available

for inspection reduced

Impact damage in polymer composites

Surface dent

Residual Compression strength after impact

Impact damage area

Boeing test

150

mm

100 mm

Composite

material

performance

ranking test

7

Threshold impact damage & compression strength

- Small samples in Laboratory

0

100

200

300

400

500

600

700

0 5 10 15

Impact Energy (J)

Com

pre

ssio

n s

tren

gth

aft

er

imp

act

(M

Pa)

0

5

10

15

20

25

30

Dam

age

wid

th (

mm

)

CAI MPa

Damage width

From Mitrovic et al. Comp Sci Tech;1999, 59, pp 2059-2078

BVID damage

threshold

Threshold for damage

initiation

Largest

possible CAI

strength

Damage at

BVID during

inspection

8

Inspection for impact damage• Difficult to see

damage in the first place

• No one knows if what they can see is actually damage

• Not the most comfortable place to be, even on a sunny day

• Can’t spend all day up there!

Photograph courtesy of Tobias Rose, Airliners.net

9

In-plane compression fatigue of pristine and impact damaged cfrpQI laminate

Adapted From Isa et al. Comp Struct.2011, 93, 2269-2276

Compression fatigue of pristine and damaged cfrp, normalised wrt CAI static strength

Adapted from Uda et al. Comp Scie Tech. 2009, 69, 2309-2314.

Compression fatigue pristine and impact damaged

QI cfrp normalised wrt pristine compression

strength

0

0.2

0.4

0.6

0.8

1

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

Str

ess/

UC

S

BVID damage

pristine

Constant

amplitude

loading

Variable

amplitude w.

compressive

overloads

Constant amplitude loading

From Mitrovic et al. Comp Sci Tech;1999, 59, pp 2059-2078 12

2024 T3 aluminium in tension -fatigue lives with

manufacturing damage

0

0.2

0.4

0.6

0.8

1

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

Str

ess/

UT

S

Pristine

1.25 mm crack

13

Delamination crack growth; UD cfrp Mode I, DCB

samples; 2024 aluminium compact tension

1.E-11

1.E-10

1.E-09

1.E-08

1.E-07

1.E-06

1.0E+01 1.0E+02 1.0E+03 1.0E+04

DeltaG N/m

da

/dN

m/c

ycl

e

T300/914

924

PEEK

Al 2024

Using ΔG=ΔK2 /E to

represent aluminium

data from metals

database

Cfrp data from Hojo et al Eng Frac Mech 1994, 49, 35-47

mGCdNda )(/

14

Fatigue delamination growth from impact

damage; max compression stress 85% static CAI

Large impact & large initial damageSmall impact small initial damage

From Isa et al (2011); Comp. struct. 93 pp 2269-2276.

15

Calculated G values at tip of idealised delamination in compression strain

From Mitrovic et al. Comp Sci Tech;1999, 59, pp 2059-2078

Design load levels and damage severity- EASA

AMC 20-29 Quantification?

Ultimate

Limit1.5 factor

of safety

Max load

per lifetime

Continued

safe flight

Allowable

damage limit

Critical damage

threshold

Category 1

damage,

BVD,Allowed

M/f damage

Category 2

damage VID

damage

requiring

repair- normal

inspection

process

Category 3

damage obvious

damage needing

repair- found

within a few

flightsCategory 4 Damage

Discrete source damage-

obvious to flight crew

needing repair after flight

Category 5 damage

Anomalous damage

17

Conclusions

• Current approaches to satisfy regulatory airworthiness requirements for design against fatigue were developed in a way suited to fatigue crack development in metals- particularly aluminium.

• Concept of slow crack growth with a high probability of detection is central to continuing airworthiness in metallic structures.

• Current generation of polymer composite laminates not behaving in a way suited to this approach• High thresholds; rapid crack growth

NDT and damage growth Questions

• How are inspection intervals set for zero damage

growth?

• Can inspection detect when a defect is beginning

to grow?

• Is there an alternative to visual inspection for

damage?

• What service environment factors determine when

a defect begins to grow? Can this be predicted?

• Could structural health monitoring play a role?

QUESTIONS?