Keywords: lateral wing upper cover, stiffened panel, carbon fibre composite, finite element method, non-linear buckling analysis, optimisation.

Due to increasing application of advanced composites in aircraft structures, significant progress has been achieved recently in the buckling and post-buckling analyses, and optimisation of stiffened laminated composite panels, but the majority of such analyses are devoted to the design of primary fibre composite fuselage structures under axial compression loads. This is why the present investigations are devoted to the methodology development based on the planning of experiments and the response surface technique for optimal design of stiffened laminated composite panels with special emphasize on more close conformity of the developed non-linear finite element analysis and the operational requirements for aircraft lateral wing upper covers under combined axial compression, shear and internal pressure loads.

A stiffened composite panel with the best weight/design performance obtained from the linear buckling analysis and optimisation is used in the present study. This panel is examined as a part of the aircraft lateral wing upper cover and consists of skin, ribs and HAT-stiffeners made from unidirectional intermediate modulus carbon fibre pre-preg tapes (T800/M21 or IMS/977-2). The stiffened panel is attached to the wing box structure by ribs perpendicular to stiffeners with a rib pitch of 800 mm. The non-linear buckling finite element analysis with a set of four load cases (100, 500, 1000 and 1500 kN per stiffener bay) is carried out using ANSYS for the panels with a HAT-stiffener pitch of 240 mm to study an effect of shear and fuel pressure on the performance of stiffened composite panels, and to investigate their behaviour under skin post-buckling.

An optimisation problem is formulated as the minimum weight design problem in term of cross-section area for one stiffener pitch. The combined influence of skin and stiffener lay-ups, stiffener height, stiffener top and root width on the buckling behaviour of composite panels under compression and shear loads is investigated. The constraints on manufacture, repair and exploitation concerning the minimum fibre percentage in each direction, the Poisson's ratio mismatch between skin and stiffener flange, and damage tolerance are taken into consideration. Due to the large scale of the numerical problem to be solved, an optimisation methodology is developed employing the method of experimental design and response surface technique. The basic idea of this approach is that simple mathematical models (response surfaces) are determined only by the finite element solutions in the reference points of the experimental design. A significant reduction in calculations is achieved in this case in comparison with the conventional optimisation methods.

Minimal weight optimisation problems have been solved for the given load levels, taking into account manufacturing, repairability and damage tolerance requirements. Re-optimisation has been carried out with the purpose of obtaining stiffened panels with permitted skin post-buckling behaviour. Recommendations for a utilisation of stiffened composite panels as parts of the lateral wing upper covers are given for designers from the non-linear buckling analysis and optimisation.

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