Aging metallic aircraft, manufactured in the early 70’s, are still being operated around the world, and are subjected to new types of damage, appearing in new locations on the aircraft. In past years, the use of composites has increased for Aluminum aircraft components. Various studies substantiated these repairs to be effective from durability point of view and evaluated the effect of plies application sequence, orientation sequence, effect of stress intensity factor and more. However, all those studies focused on simple tensile loading.
Forward and center sections of a combat aircraft fuselage contain a skin, which plays a major role in load transferring, usually due to torsion originated by fuselage bending. The present research aims at studying the effect of the above mentioned parameters on composite repair patch for a panel, subjected to shear loads. The study is based on a new approach of finite element modeling, consisting of 2D and 3D merged elements. Correspondingly, the model is parametrically analyzed for strength and buckling, as a function of patch size, overlaps, and sequences of plies. The analysis is then substantiated by a dedicated set of experiments on typical representative elements. Based on the test results, the accuracy of the finite element model was evaluated and thus validating the method of patch repair.
