Surface damage from discrete sources on pressurized aircraft fuselages can cause cracks to form. As the fuselage experiences pressurization during ascent and descent, these cracks bulge-out or protrude from the original contour. This change in the geometry of the "bulging effect" significantly increases the stress-intensity factor (SIF) at the cracks tips. The aim of this research is to determine the contribution of the bulging effect in crack development on single curved fiber-polymer composite (FPC) shell structures. In order to accurately determine the major dynamic failure mode involved in this process, the loading condition of the specimens is limited to cyclic pressurization/ depressurization. Although extensive studies on this subject have been performed on aluminum structures, the dissimilar failure evolution of FPCs requires a separate characterization of the crack development and its relationship to bulging factor. The recent work performed at Florida International University entailed the analysis of 304.8 mm x 304.8 mm (12" x 12") carbon-fiber curved specimens replicating a Boeing 727 access door, chosen to be representative of the curvature of a pressurized airliner. The study included the design and fabrication of a robust experimental set-up and specimens as well as the numerical analysis using the finite element method (FEM). In order to be consistent with FEM studies performed on the bulging phenomenon on aluminum structures by the Federal Aviation Administration (FAA), the Modified Crack Closure Integral (MCCI) method was used to calculate the SIF of the cracks. The study concluded that FPC parts with cracks on pressurized cylindrical structures preserve their structural integrity when designed to withstand planar shear stress. It was also concluded that for a 50.8 mm (2 inch) long crack, the bulging factor is independent of the cabin pressure of the airplane.
