ABSTRACT
Fuselage panels are commonly fabricated as skin-stringer constructions, which are permitted to locally buckle under normal flight loads. Current aerospace analysis techniques applied to determine the buckling and post buckling behaviour of stiffened panels typically employ design formulae based on empirical / semi-empirical data. These empirical methods possess an inherent conservatism, which can lead to over-designed structures. This paper contributes to the development of a robust Finite Element analysis methodology capable of accurately predicting the buckling and post buckling response of conventional aircraft fuselage panels. The work presented focuses on curved panels subjected to uniform axial compression. Two panel configurations have been examined, one utilising bulbed-t section stringers, the other utilising z-section stringers. The Finite Element predictions have been correlated to experimental data in both cases, with post-buckling results proving favourable. The research undertaken has demonstrated that using a commercial implicit code, the Finite Element method can be used successfully to model the behaviour of curved panels in the post-buckling range. The work undertaken has emphasised the need for Finite Element analysis models to accurately represent the initial geometric and residual stress imperfections.
