On the analysis of composite structures with material and geometric non-linearities

If major weight saving is to be realised it is essential that composites be used in "primary" structural components, i.e., wing and fuselage skins. To this end it is essential that analytical tools be developed to ensure that composite structures meet the FAA damage tolerance certification requirements. For stiffened composite panels one potential failure mechanism is the separation of the skin from the stiffeners; resulting from excessive "through the thickness" stresses. This failure mechanism is also present in bonded composite joints and composite repairs. Currently failure prediction due to in-plane loading appears to be relatively well handled. Unfortunately, this is not yet true for matrix-dominated failures. Consequently, it is essential that a valid analysis methodology capable of addressing all of the possible failure mechanisms, including failure due to interlaminar failure, be developed. To aid in achieving this objective the present paper outlines the results of a series of experimental, analytical and numerical studies into the matrix-dominated failures of rib stiffened structures.