Transport aircraft is a highly complexairframestructure. The aircraft fuselage shell is composed of stressedskin, circumferential frames and longitudinal stringers. The skinisconnected toframes and stringers mostly by rivets. Fuselagehas a number of riveted joints and is subjected to a majorloading of differential internal pressurization. When the fuselageis pressurized and depressurized during each takeoff and landingcycle of aircraft, the metal skin of fuselage expands and contractsresulting in metal fatigue.Fatigue damage accumulates duringevery cycle of loading in the airframestructure during itsoperation. The accumulated damage reaches a critical value, afatiguecracks initiate from riveted holes and propagate tocritical sizes leading to catastrophic failure of the structure. Thelargetransport aircraft are designed to tolerate large fatiguecracks. This paper focusesitsattention on damage tolerancedesign of a fuselage structure of transport aircraft. The objectiveof this paper is to investigate crack initiation, and crack growthratein theflatstiffened panel of fuselagestructure. Thelongitudinal crack is initiated from the rivet holelocationandstress intensity factor is calculated using modified virtual crackclosure integral (MVCCI) method duringeach stage of crackpropagation.Fatigue crack growth rate can be estimated byusing Paris law under spectrum loading analysis in the structure.In this paper for modeling CATIA V5 software is used and MSCPATRAN is used for meshing the stiffened panel and static linearstress analysis is carried out using MSC NASTRAN.