Abstract
Aircraft structural fatigue is a serious issue. Untreated, it could lead to failure. Several aircraft accidents were caused by widespread fatigue damage. For cracks in the fuselage, the stress state is due to the bulging around the crack caused by applied internal pressure. This condition is characterized by a parameter called the bulging factor; it compares the stress intensity factor of a crack in curved shell to its counterpart in a plate. Bulging factor expressions are available for longitudinal and circumferential cracks but less for slanted cracks.
The underlying structure that stiffens the skin of the fuselage has a direct impact on the stress intensity factor and the bulging factor of the crack. A hexagonal grid stiffening pattern has been shown to provide sufficient stiffening to the fuselage skin while using a lesser amount of material compared to the traditional orthogonal grid design. However, the response of this grid pattern in the presence of a cracked fuselage has not been studied.
The current paper aims to estimate the effect of the hexagonal grid pattern on the bulging factor and stress intensity factor of cracks of various lengths and orientations in the skin of the fuselage. Results obtained are compared to the conventional orthogonal grid stiffening pattern of a fuselage structure. Several patterns were considered.
Results
show that cracks of different lengths and orientations in hexagonal grid stiffened panels had stress intensity factors and bulging factors that are comparable to the base case within a margin of 2%–8%.
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