The desired method for repair of composite airframe structures is by joining with an adhesive(-like) material. The available choices of structurally efficient joint configurations are limited. Most of the commonly used joint configurations would result in repairs that would be either impractical, because of the long adherend overlap lengths required to sustain the design load, or unsafe, because a shortened lap length would limit the load-carrying capability of the joint. Single lap joint configurations based on a variety of patch material thicknesses and stiffnesses are investigated for elastic response and strength capability. The section stiffness of the patch material is kept constant in the direction of the first principal stress, as the number of plies is decreased. The average modulus of the patch adherend is increased by reorienting the plies in the laminate. This is done to demonstrate the possibility of increasing the average joint failure shear stress and the accompanying load transfer efficiency by increasing the stiffness of the adherends. Stress analyses of these configurations are carried out and compared to the results from testing the described configurations.
Experimental data show that it may be necessary to incorporate the nonlinear adhesive behavior into the model to account for the extensional response of the adherends. Also, it may be necessary to model each composite ply to predict the extent of incompatibility of the adhesive bondline and the adjacent composite material. The data also demonstrate that the failure of a composite joint depends on the strength of bondline-contacting interfacial plies, as well as the usual factors considered when estimating joint strength.