Weight Effect of Buckled Composite Stiffened Panels Under Pure Shear Loading

Abstract

Composite hat-stiffened skin panels are of great interest to modern aircraft design because of their high strength and stiffness. This work aims to computationally predict damage within hatstiffened composite panels under pure shear loading. As the panels enter the post-buckling regime, matrix cracking, fiber rupture and delamination are crucial. The Enhanced Schapery Theory can capture the intraply damage while the Cohesive Zone Model delamination between the skin plies and debonding between skin-hat interfaces. For hat-stiffened panels, the four, three, and two stringers configurations experienced damage initiation at 166.6%, 295.0%, and 739.6% of the buckling load, respectively. Matrix cracking initiated in the panels least stiff to transverse loading, while panels more resistant to shear see debonding at the skin-hat interface initiated first. All configurations predicted delamination at the skin ply 1-ply 2 interface as the last damage mode before penetrating deeper into the skin. Debonding between the stringers and the skin was predicted as the primary driver for ultimate failure.