Two bay crack arrest capability evaluation for metallic fuselage

Abstract

Abstract Transport aircraft is a highly complex structure. The aircraft fuselage shell is composed of stressed skin, longitudinal stringers, and circumferential frames. The skin is connected to the stringers and frames mostly by rivets. The 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 take-off and landing cycle of aircraft, the metal skin of fuselage expands and contracts to result in metal fatigue. Due to the presence of a large number of rivet holes, the fuselage skin has a large number of high-stress locations and these are locations of potential crack initiation. The wide-bodied transport aircraft are designed to tolerate large fatigue cracks. This project focuses attention on damage tolerance design of a fuselage structure of transport aircraft. In this project, the stress intensity factor, quantifying the intensity of the stress field around a crack tip for a longitudinal crack under the pressurization load is studied. The objective of this project is to investigate crack initiation, crack growth, fast fracture and crack arrest features in the stiffened panel. The longitudinal crack is initiated from the rivet hole and stress intensity factor is calculated using modified virtual crack closure integral (MVCCI) method at each stage of crack propagation. In order to arrest crack propagation which is the capital importance of tear straps are used, which prevent the further crack propagation. In this project, the linear static stress analysis of the stiffened panel of a fuselage is performed using MSC NASTRAN solver. The pre-processing of the model is done by using MSC PATRAN software.