Advances in the Bonded Composite Repair o f Metallic Aircraft Structure phần 9 pptx

Chapter 15. Graphite/epo.xy patching eficiency studies 439

Fig. 15.18. Different damage configurations of "equivalent" width

that the combined contribution of Kl, and KIIr to the effective stress intensity factor

was less than 8% for the configurations studied.

Fatigue testing confirmed that various forms of damage could be repaired

effectively with single patches. For example, the fatigue lives of panels containing

40mm diameter holes and either single cracks or four 45 kinked cracks were

improved by factors of 5.1 and > 15, respectively, by single-sided repairs with a

1 mm thick patch [80 mm x 80mml. The measured fatigue crack growth rates were

Table 15.5

Comparison of stress intensity factor ranges for four 45" kinked cracks

and two diametrically opposite cracks, with tips at x, =4Omm and hole

radius = 20 mm (R = 0.1, omrx = 65 MPa).

Crack configuration Patch thickness, mm AKP AKu AKp/AKL'

4 x 45" kinked 2.0 6.0 22.3 0.30

cracks 1 .o 8.8 22.3 0.39

2 x diam. opposite 2.0

cracks 1 .o

11.3 27.7 0.41

13.7 27.7 0.49

440 Advances in the bonded composite repair of metallic aircraft structure

in good general agreement with theoretical predictions. For example, the mean

crack growth rate of 9 x 10~9m/cycle measured for four kinked cracks at a half

crack length of 35mm during R=0.1, am,,=41.25MPa loading, was in good

agreement with predicted values of AKp= 5.7 MPam’I2 and Kmin/Kmax = 0.51, and

da/dN-AK data for 2024-T35 1 aluminium alloy sheet. Furthermore, the observed

crack paths indicated little effect due to Mode I1 loading. In the case of a 40mm

hole and two cracks, double sided patching resulted in crack arrest, in agreement

with theoretical predictions. Work is in progress to determine the effectiveness of

patch repairs for other damage configurations.

15.10. Future work

Although adhesively bonded gr/ep patch repair of cracked metallic structures has

been studied extensively and service experience with repairs has been good, it

appears that further work is required to address some remaining problems and to

assess the full potential of the repair technique. Specific research objectives include

the following:

(a) To investigate the effect of variable amplitude loading spectra on patch

debonding and hence on patch efficiency. There is a clear requirement for a

model to predict debonding, and for incorporation of this in a general model,

which will enable the effect of patching on fatigue crack growth to be predicted

for a wide range of loading spectra.

(b) To assess the influence of hot-wet fatigue test environments on patch efficiency,

and the effect of long-term pre-exposure to hot-wet environments on such

behaviour.

(c) To establish the advantages and limitations of repairs carried out by co-cure of

prepreg and adhesive.

(d) To develop and assess bonded patch repair schemes for applications involving

elevated service temperatures.

(e) To investigate the effectiveness of bonded patches for the repair of various

forms of corrosion damage and battle damage in aluminium alloy structures.

(f) To develop and assess patch repairs for applications involving bonding over

fasteners.

(g) To assess the potential of bonded patches for the repair of SPF/DB titanium

alloy structures, and to develop optimum repair schemes.

(h) To develop “Smart” patches for monitoring repair performance in service, and

improved NDE techniques for (i) inspecting pretreated surfaces prior to

bonding, and (ii) assessing the strength and durability bonded patch repairs.

15.11 Acknowledgements

0 British Crown Copyright 2001. Published with the permission of the Defence

Evaluation and Research Agency on behalf of the Controller of HMSO.

Chapter 15. Graphitelepoxy patching efficiency studies 44 1

References

I. Kemp, R.M.J., Murphy, D.J., Butt, R.I., et al. (1983). RAE Technical Report TR 83005.

2. Sutton, G.R., Stone, M.H., Poole, P. et a/. (1984). In: Repair and Reclamation, The Metals Society;

3. Poole, P., Stone, M.H.. Sutton, G.R., et al. (1986). In: The Repair of Aircraft Structures Involving

4. Sutton, G.R. and Stone, M.H. (1989). RAE Technical Report TR 89034.

5. Dowrick, G., Cartwright, D.J. and Rooke, D.P. (1980). RAE Technical Report TR 80098.

6. Young, A,, Cartwright, D.J. and Rooke, D.P. (1988). Aeronautical J., pp. 41&421.

7. Young, A,, Rooke, D.P. and Cartwright, D.J. (1989). Aeronautical J., pp. 327-332.

8. Ball, AS. (1993). MOD Contract SLS 41B/2093, Final Report BAe-KDD-FCP-0104.

9. Poole, P., Brown, K. and Young, A. (1990). RAE Technical Report TR 90055.

pp. 17.1-17.6.

Composite Materials, AGARD-CP-402, pp. 9.1-9.21.

10. Poole, P., Lock, D.S. and Young, A. (1991). In: Aircraft Damage Assessment and Repair. The

11. Poole, P. and Young, A. (1992). In: Theoretical concepts and Numerical Analysis of Fatigue [A.F.

12. Baker, A.A. (1988). In: Bonded Repair of Aircraft Structures (A.A. Baker and R. Jones, eds.),

13. Poole, P., Young, A. and Ball, AS. (1994). In: Composite Repair of Military Aircraft, AGARD￾14. Poole, P., Lock. D.S. and Young, A. (1997). In: Proc. of 1997 USAF Aircrufi Structural Infqrit.)’

15. Poole, P., Brown, K., Lock, D.S.. et a/. (1999). In: Proc. of IW9 USAF Aircruft Structural Infegrit?,

16. Poole, P., Stone, M.H., Sutton, G.R., el al. (2000). In: Proc. of 2000 USAF Aircraft Structural

Institution of Engineers, Australia, pp. 85-91.

Blom and C.J. Beevers, eds.], EMAS, pp. 421438.

Martinus Nijhoff, pp. 107-173.

CP-550, pp. 3.1-3.12.

Conf., USAF.

ConA, USAF.

Integrity Conf., USAF.