Abstract
In this article, the elastic buckling behaviour of cylindrical GLARE (GLAss REinforced) panels with classically simply supported boundary conditions under uniaxial compression is investigated using the finite element method (FEM) and eigenvalue buckling analysis. The buckling coefficient-curvature parameter diagrams of five GLARE grades are obtained and studied along with the diagrams of two glass-fiber composites and monolithic 2024-T3 aluminum, using validated FEM models. It is found that aluminium has a stronger impact on the buckling behaviour of the GLARE panels than the composite layers. From the constructed buckling coefficient - curvature parameter diagrams in double logarithmic scale it is found that there is an approximately linear relation between the buckling coefficient and the curvature parameter of the panels. Based on this finding, appropriate regressions are implemented in order to derive approximate analytical formulas of the buckling coefficient as a function of the curvature parameter for the considered materials.
Highlights
Fiber Metal Laminates (FMLs) are hybrid composite materials built up from thin metal alloy sheets bonded into one laminate with intermediate fibre/epoxy layers
The buckling coefficient k will be determined from the finite element method (FEM) results for variable values of the panel's curvature parameter Z and (k, Z) curves of panels consisting of different materials will be constructed
This article deals with the elastic buckling of supported cylindrical GLARE panels subjected to uniaxial compression
Summary
Fiber Metal Laminates (FMLs) are hybrid composite materials built up from thin metal alloy sheets bonded into one laminate with intermediate fibre/epoxy layers. In 1990, high strength glass fibers were used creating GLARE (GLAss REinforced), which is the most successful FML up to now. GLARE laminates have been selected for many aerospace applications such as the upper fuselage skin material of the Airbus A380 and the cargo floor of Boeing 777 [2]. The elastic buckling of thin panels is a classical problem of the strength of materials with great practical importance and it must always be considered during the design of many engineering applications where stiffened thin-walled structures are employed. The fuselage and wing skin of aerospace structures consist of stringer stiffened curved plates. FMLs are mainly used for the construction of stiffened thin-walled fuselage structures and, as a result, the investigation of the buckling strength of FML panels is very important. The classical supported boundary conditions [3] are considered for the analyzed FMLs
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