Investigation of Fatigue Crack Growth Rate in Fuselage of Large Transport Aircraft using FEA Approach
Transport aircraft is a highly complex airframe structure. The aircraft fuselage shell is composed of stressed skin, circumferential frames and longitudinal stringers. The skin is connected to frames and stringers mostly by rivets. Fuselage has a number of riveted joints and is subjected to a major loading of differential internal pressurization. When the fuselage is pressurized and depressurized during each takeoff and landing cycle of aircraft, the metal skin of fuselage expands and contracts resulting in metal fatigue. Fatigue damage accumulates during every cycle of loading in the airframe structure during its operation. The accumulated damage reaches a critical value, a fatigue cracks initiate from riveted holes and propagate to critical sizes leading to catastrophic failure of the structure. The large transport aircraft are designed to tolerate large fatigue cracks. This paper focuses its attention on damage tolerance design of a fuselage structure of transport aircraft. The objective of this paper is to investigate crack initiation, and crack growth rate in the flat stiffened panel of fuselage structure. The longitudinal crack is initiated from the rivet hole location and stress intensity factor is calculated using modified virtual crack closure integral (MVCCI) method during each stage of crack propagation. Fatigue crack growth rate can be estimated by using Paris law under spectrum loading analysis in the structure. In this paper for modeling CATIA V5 software is used and MSC PATRAN is used for meshing the stiffened panel and static linear stress analysis is carried out using MSC NASTRAN.
