A preliminary study into the matrix dominated non-linear behaviour of graphite/epoxy laminates

The potential of composites for weight-saving is now widely recognized and they are used extensively in both control surfaces and doors. However, if major weight-saving is to be realized it is essential that composites be used in 'primary' components i.e. wing and fuselage skins. Before this can be achieved it must be shown that composites are an adequate substitute for thin-skinned metallic components, particularly in regions where they undergo large strains and/or post-buckling deformations. For stiffened composite panels one potential failure mechanism is the separation of the skin from the stiffeners; resulting from excessive 'through the thickness' stresses. To address this potential interlaminar mechanism, the present paper discusses the non-linear, time-dependent, response of a T300/914C, [ 0 90]_s graphite/epoxy laminate. Initially a series of ASTM standard, D4255-83, shear rail tests were performed and the equivalent σ_eq-ε{lunate}_eq curve for the laminate was derived using combined experimental and computational techniques. In general, the shear behaviour of the laminate was very similar to that observed for epoxy adhesives with extensive strain rate effects, stress relaxation and creep even at room temperature. To further substantiate these results a series of ASTM standard, D3518-76, tests were then performed on ± 45 ° laminates. These tests revealed that the σ_eq-ε{lunate}_eq curve was indeed rate dependent, often with a significant inelastic behaviour. Furthermore, the failure stresses and strains were found to be strongly dependent on the loading rate.