SUMMARY

Carbon Fiber / Epoxy Laminates are increasingly being used in the primary structure of aircraft. To make effective use these materials, it is necessary to consider the ability of a laminate to resist damage, as well as material strength and stiffness. A possible means for improving damage tolerance is the use of non-traditional composite laminates, in which the longitudinal 0° plies are replaced with ±5° or ±10° plies. The main objectives of this collaborative Georgia Tech / Boeing research was the characterization of these non-traditional laminates, and the determination of appropriate lamina-level analytical techniques that are capable of predicting the changes caused by the use of slightly off-axis longitudinal plies. A quasi-isotropic [45/90/-45/q/45/90/-45/-q]s and “hard” [45/±q/-45/±q/90/±45]s lay-up, where q =5° or 10°, were tested in open hole tension, filled hole tension, open hole compression, single shear bearing, and unnotched tension. These coupon level tests illustrated the effects of lay-up, notch constraint, and load type on traditional and non-traditional laminates. Die penetrant enhanced in-situ radiography was performed to determine the extent of damage suppression. The use of non-traditional laminates was found to reduce longitudinal ply cracking and delamination, with significant effect on the stress distribution around the notch. The use of non-traditional laminates also resulted in a 15%-20% improvement in bearing strength of the traditional laminates. Several predictive techniques were implemented to evaluate their ability to predict the effect of slight changes in ply orientations. A progressive damage model was written to compare Tsai-Wu, Hashin, and Maximum Stress unnotched strength criterion. Additionally, several semi-empirical failure theories for notched strength prediction were compared with linear and bi-linear cohesive zone models to determine applicability to non-traditional laminates.