Prediction of the postbuckling response of composite airframe panels including ply failure

Future design scenarios aim to allow buckling in composite airframe panels. Reliable simulation procedures should be able to capture the postbuckling elastic as well as the inelastic response associated with damage. Damage in composite laminates in terms of ply failure may primarily occur as fiber fracture or matrix cracking. This paper presents a model which is able to capture both geometrical and material nonlinearity. It bases on the finite element formulation of a layered, iso-parametric, quadrilateral shell element which allows for an arbitrary reference surface as well as an arbitrary stacking sequence. Geometrical nonlinearity is accounted for by utilizing Green strains and second Piola–Kirchhoff stresses. Material nonlinearity is considered via a layerwise ideally brittle damage model. The model is applied to a buckling test of a stringer-stiffened composite airframe panel. The numerical results are compared with an experiment proving the applicability of the proposed concept.