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Civil-Comp Proceedings
ISSN 1759-3433
CCP: 83
Edited by: B.H.V. Topping, G. Montero and R. Montenegro
Paper 102

The Effects of Thermal Residual Stresses on the Fatigue Life of Aluminum Panels Repaired with Various Bonded Composite Materials

H. Hosseini-Toudeshky, B. Mohammadi and M. Musivand-Arzanfudi

Department of Aerospace Engineering, Amirkabir University of Technology (Tehran Polytechnic), Tehran, Iran

Full Bibliographic Reference for this paper
H. Hosseini-Toudeshky, B. Mohammadi, M. Musivand-Arzanfudi, "The Effects of Thermal Residual Stresses on the Fatigue Life of Aluminum Panels Repaired with Various Bonded Composite Materials", in B.H.V. Topping, G. Montero, R. Montenegro, (Editors), "Proceedings of the Eighth International Conference on Computational Structures Technology", Civil-Comp Press, Stirlingshire, UK, Paper 102, 2006. doi:10.4203/ccp.83.102
Keywords: repair, aluminium panels, composite, thermal, residual stresses.

Patch techniques or using stiffeners on the cracked surfaces, are viable methods for providing life extension of cracked aluminum panels. Bonded reinforcement or patches provide a stiff load path; however, mechanically fastened patches provide a relatively compliant load path. Adhesively bonded composite patches are more structurally efficient with much less damage on the structure than the repairs based on mechanical fasteners with metallic patches. A single-side repair can reduce the effective stress intensity factor at the crack tip. However, it produces out-of-plane bending, which occurs due to the shift in the neutral axis of the panel under tension loading. This causes stress variations over the thickness of the cracked panel and therefore produces non-uniform crack propagation along the panel's thickness.

One of the main disadvantages of the composite patches such as boron-epoxy and graphite-epoxy results from their relatively low coefficient of thermal expansion compared to the parent material, which results in the residual stresses in the repaired components [1]. Because of the coefficient of thermal expansion (CTE) mismatch between the metal structure and the repair material, tensile and, or compression strains are introduced into the metallic component, thus affecting the opening of the crack. This occurs during the cool-down phase of the cure cycle that bonds the patch to the structure. During the heat-up phase, both materials are allowed to expand or contract freely because the adhesive has not yet cured and locked in its structure. When the adhesive is sufficiently cured, the three-dimensional molecular cross-linked structure limits its ability to viscoelstically respond to stress. When the repair has been cooled to the ambient temperature, the CTE mismatch becomes clearly visible.

Numerical studies have been carried out by Chow and Atluri [2] to examine the effect of the thermal cycling on the fatigue response of the aluminum panels repaired with the boron-epoxy patch. They found that due to the strong difference in the thermal expansion coefficient of the boron-epoxy patch and the aluminum panel, the fatigue life of the specimens, which undergoes cycles of high stress at low-temperature and low stress at high-temperature is dramatically reduced.

Crooker [3] investigated the thermal residual strain occurred in the 7075-T6 aluminum panel repaired with 15-ply graphite-epoxy patch. He concluded that reducing the cure cycle temperature could decrease thermal residual strains by 26.5% between the patch and the aluminum panels when FM73 adhesive is used to bond them together, and 7.4% when EA 9696 is used.

This study investigates the thermal residual stresses that occur as a direct result of the coefficient of thermal expansion (CTE) mismatch between the repair patch and the underlying cracked metallic structure to which the patch is bonded. In this investigation the numerical (FEM) fatigue crack-growth behaviour of centrally cracked aluminum panels (2024-T3) in mode-I condition repaired with various single-side 8-ply glass-epoxy, graphite-epoxy and boron epoxy composite patches are curried out. The crack growth behavior obtained, crack-front configuration and crack growth life of the repaired panels are discussed for the repaired panels with various patch materials. The sensitivity of the residual stresses to the curing temperatures of the repaired panels with various composite patches is also investigated. The obtained numerical fatigue life results are compared with the available experimental results in some cases.

It is shown that using a low curing temperature (less than 60oC) with a long curing cycles do not cause a considerable effect of residual stresses on the fatigue crack growth life for glass-epoxy, graphite-epoxy and boron-epoxy composite patches. It is also shown that considering the thermal residual stresses the FEM fatigue crack-front shapes for the repaired panels using glass-epoxy patches become very close to those obtained from the experiments.

Baker A.A. "Crack Patching", In: Baker A.A. and Jones R., (Eds), "Bonded Repair of Aircraft Structure", (Chapter 6: Experimental Studies, Practical Application), Dordrecht: Martinus Nijhoff, The Netherlands (1988).
W.T. Chow and S.N. Atluri, "Composite Patch Repairs of Metal Structure: Adhesive Nonlinearity, Thermal Cycling, and Debounding", AIAA Journal, Vol. 35, No. 9, pp 1528-1535, 1997. doi:10.2514/2.7481
H.R. Crooks, "Reduction of Thermal Residual Strains in Adhesively Bonded Composite Repairs", MSc Thesis, USAF, Air Force Institute of Technology, 2003.

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