Experimental and finite element studies of bolted, bonded and hybrid step lap joints of thick carbon fibre/epoxy panels used in aircraft structures

A prior study conducted by the authors had investigated the behaviour of thin composite double lap joint repairs. This investigation was an extension to the previous study and focused on analysing the behaviour of thick step lap joint repairs which are more complicated in nature. Finite element analysis (FEA) was performed to verify the static and fatigue strength of bolted, bonded and hybrid joint configurations. Thick carbon fibre/epoxy laminates and aerospace grade film adhesive and fasteners were selected. Several configurations were considered in this case. These include varying the number of fasteners in the joint region as well as inclusion of bondline defects. Adhesive non-linear material properties, fastener contacts and frictional forces were all included in the three dimensional (3D) finite element (FE) models. The progressive failure process was simulated and the Multicontinuum Theory (MCT) was used to determine the stress states in all specimens considered. The strain energy release rate (SERR) as a function of crack length was also utilised to gain further understanding into the fatigue behaviour. Overall the FE models are able to accurately predict the bonded, bolted and hybrid joint strengths. SERR results suggest it is vital to place fasteners closer to the ends of the overlap to suppress the peak peeling stresses and to delay the effects of early crack initiation. This overall improves the longevity of a conventional bonded joint design.