Damage tolerance and failure analysis of a composite geodesically stiffened compression panel

A geodesically stiffened continuous-filament composite structural concept has been designed for a transport aircraft fuselage application and fabricated using an automated manufacturing process. Both large panels and element specimens derived from these panels have been experimentally and analytically investigated when subjected to axial compression to understand their buckling, postbuckling, and failure responses. The primary failure mode for this structural concept is skin-stiffene r separation in the skin postbuckling load range. The large panels are subjected to low-speed impact damage and tested to failure in axial compression to evaluate the damage tolerance of this structural concept. These results suggest that damage to the stiffener and a stiffener intersection point from the skin side do not influence the failure load or failure mode of this structural concept. Nonlinear finite element analysis using a detailed element specimen model indicates that failure of this specimen may have initiated at the skin-stiffener flange region close to the stiffener intersection. When the skin is in the postbuckling range at four times its initial buckling load, the interlaminar shear stress concentrations at a region where the stiffener makes a 20-deg turn away from the stiffener intersection seem to initiate panel failure.