Blade-stiffened, compression-loaded cover panels were designed, manufactured, analyzed, and tested. All panels were fabricated from IM6/1808I interleafed graphite-epoxy. An orthotropic blade stiffener and an orthotropic skin were selected to satisfy the design requirements for an advanced aircraft configuration. All specimens were impact damaged prior to testing. Experimental results were obtained for three- and five-stiffener panels. Analytical results described interlaminar forces caused by impact and predicted specimen residual strength. The analytical results compared reasonably with the experimental results for residual strength of the specimens.

Composite materials are sensitive to out-of-plane loads because they are weaker through the thickness than in the plane of lamination. Consequently, composite structure subjected to impact may suffer significant damage, which results in a loss in strength. The effect of impact energy on different thicknesses of laminate with respect to ultimate compressive strength was evaluated using the NASA ST-1 specimen test. Impact damage was measured using ultrasonic C-scans and quality was assessed by visual observations. A semiempirical model for impact damage was proposed to determine residual strength in terms of laminate properties, static influence coefficients, and impact energy. To determine residual compression strength, single- and multiple-impact damage was inflicted on bonded stiffened structures. The results were verified by the residual strength method, which uses laminate properties and measured damage size to predict the residual strength of composites. The correlation between predicted and actual strength was very encouraging. Although the damage was inflicted with different impactor masses, velocities, and impact energy, the data indicated that residual strength is a function only of damage present and does not reflect the manner in which the damage is inflicted. Finally, a semiempirical relationship was developed for damaged areas in stitched laminates.

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