DAMAGE TOLERANCE AND FAILURE ANALYSIS OF A COMPOSITE GEODESICALLY STIFFENED COMPRESSION PANEL

A geodesically stiffened continuous-filament composite structural conc ept has been designed for a transport aircraft fuselage application an d fabricated using an automated manufacturing process, Both large pane ls and element specimens derived from these panels have been experimen tally and analytically investigated when subjected to axial compressio n to understand their buckling, postbuckling, and failure responses, T he primary failure mode for this structural concept Is skin-stiffener separation in the skin postbuckling load range. The large panels are s ubjected to low-speed impact damage and tested to failure in axial com pression to evaluate the damage tolerance of this structural concept. These results suggest that damage to the stiffener and a stiffener int ersection point from the skin side do not influence the failure load o r failure mode of this structural concept. Nonlinear finite element an alysis using a detailed element specimen model indicates that failure of this specimen may have initiated at the skin-stiffener flange regio n close to the stiffener intersection, When the skin is in the postbuc kling 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.