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Research Article Open Access
Aircraft designer needs to ensure the structural integrity of the airframe without compromising on the safety of the structure. This requires several stress analysis of different components, which represents the features of the airframe. Also these analysis predictions have to be substantiated by structural testing during the developmental phase. In current study a representative stiffened panel from a centre fuselage segment of an aircraft will be considered for the evaluation. The stress analysis of the stiffened panel will be carried out by using FEM approach. Aluminium alloy 2024-T351 material is used for the stiffened panel. Fuselage structure experiences the hoop tension and longitudinal tension because of the internal pressurization. If there is a crack in the unstiffened fuselage, under the flight condition it could lead to catastrophic failure of the structure. The fatigue crack will initiate normally from the highest tensile stress locations. Therefore rivet locations are the most probable locations for fatigue crack initiation. Miner’s rule will be employed for the fatigue damage calculation and life estimation of the structure. S-N curve of the aluminium alloy 2024-T351 will be used for obtaining the number of cycles to failure at particular stress magnitudes.