Composite cylindrical shells and panels are widely used in aerospace structures. Delaminations within the composite structure reduce the compressive strength of laminates, and often result because of damage incurred during manufacturing and in-service use. This paper investigates the buckling behaviour of laminated cylindrical panels loaded in axial compression using the finite element method. The use of three-dimensional finite elements for predicting the delamination buckling of these structures is computationally expensive. Here the analysis has been carried out using a layerwise shell finite element based on the first-order, shear deformation theory. Contact elements were placed between the delaminated regions to avoid physical interpenetration of the elements. It is shown that through-the-thickness delamination can be modelled and analysed effectively without requiring a great deal of computing time and memory. Delamination shapes considered in mis study were square and rectangular - extended longitudinally over the entire length or extended along the entire circumference of the panel. Some of the results were compared with the corresponding analytical results which were in good agreement. The most influential parameters for a given laminated panel were the size of the delamination and its through-the-thickness position. The effect of the curvature on the global buckling strength of a delaminated panel was also studied. Depending on the size and through the thickness position of delaminations, three different modes of buckling behaviour occur. The local mode occurs when the delamination is near the free surface of the laminate and the area of the delamination is large. The global mode occurs when the delamination is deeper within the laminate and has a small area. The mixed mode is a combination of global and local modes.
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