Stress Intensity Factors in Un-repaired and Repaired Cracked Aluminum Panels by CFRP Composite Patch-FE Analysis

Abstract

Aging aircraft may contain many cracks resulting from their long service. Therefore, it is necessary to repair the cracked aircraft structural components to extend the service life of the aircraft. Adhesively bonded composite repair is an efficient method to extend the fatigue life of cracked components in advanced aerospace structures. In this study, a 2-D finite element model of an aluminum panel with a crack perpendicular to the load direction has been developed which is also used to model and analyze the repaired thick panel.. The panel is reinforced with singlesided composite patch made of Carbon/Epoxy composite material. Different lay-ups of 0°, +45°, and 90° have been used to achieve the minimum stress intensity factor at the crack tip. At +45°, due to generation of maximum in-plane shear stress in composite patch and adhesive the effect of patch on crack suppression is minimized. The configuration consists of aluminum panel, composite patch and adhesive layer denoted here as PCA. The uniqueness of this configuration is due to its stacking and lay-up of the adhesive layer. Instead of spring elements commonly used in previous studies, this analysis modeled the adhesive layer as an elastic continuum. The stress intensity factor is calculated using both the displacement and the JIntegral methods. The importance of using J-Integral method here is the selection of path around the crack tip, where, one path is chosen in the plastic zone at crack tip and the other is selected in the elastic region. For patched cracked aluminum panel it has been found that both counters have resulted in close values of stress intensity factors due to effectiveness of the PCA configuration. Finally, methods to measure the stress intensity factor and the fiber angle for achieving the minimum stress intensity at crack tip have been discussed.