Fatigue life estimation for a circumferential joint of the aircraft fuselage structure through stress analysis

Aircraft is a mechanical structure which can fly in air. Civil Aircraft is used for transportation and fighter aircraft for defense purpose. It is a challenge in front of the aircraft structural designer to bring out a safest structure with minimum weight. Airframe is the load bearing structure of the aircraft. Fuselage and wing are the major primary structural components of the airframe. Fuselage is a cylindrical structure which houses passenger seats. Fuselage is a semi-monologue structure. It is stiffened orthogonally with stiffening members along longitudinal and circumferential directions. The skin of the fuselage is spliced both in longitudinal and circumferential directions. These joints will be critical due to bending and pressurization loads in the fuselage. In this project stress analysis of the circumferential splice joint in the fuselage structure is carried out. The internal pressurization load case is considered to be one of the critical load cases during the design and development of the aircraft. A panel representing the location of circumferential joint is identified for evaluation of the structural integrity of the fuselage. The panel consists of two skin plates, part of the bulkhead, tear strap and rivets. Longitudinal stress produced because of the internal pressurization and the tensile stress due to bending of the fuselage will make this joint critical from the fatigue crack initiation point of view. Stress analysis will be carried out on this panel to identify the maximum tensile stress location in the panel. Fatigue cracks will always get initiated from the location of maximum tensile stress.