The substantiation of the optimum design decisions providing the minimal weight of a structure, is one of the basic preconditions of successful realization of the project ultra light UAV. Therefore at definition of key structural parameters of a wing there was a task in view of optimization which in connection with the big complexity was solved at enough rigid restrictions and in some stages. At the first stage the choice of the basic structural scheme, potentially capable to provide the requirement of static strength in all limiting cases loading, and also requirements of dynamic strength was carried out. As a result the thin-walled structure with a twice-closed contour of the cross-section section, made of composite materials, mainly type a sandwich has been chosen. Together with the theoretical geometrical sizes of a wing which are growing out previous aerodynamic researches, the choice of the structural scheme defined the basic restrictions in a problem of optimization. At the second stage parametrical researches of dependence of the intense condition and bearing ability of simple elements of a design of a wing (type of flat panels or cylindrical shell of small curvature) from a kind of external loading were carried out. Optimum parameters of panels (thickness and structure of face layers, type and the sizes of a core) were defined at various variants of external loading: tension, compression, bending, shift and their various combinations. The problem of buckling of panels and post-buckling behavior as advantages of the thin-walled composite panel are substantially lost in conditions of compression has special value. The problem of an optimum choice was put as follows. Let overall dimensions of the rectangular panel fixed on all four parties (Figure1) are given. Rigidity of a material of an internal layer of the panel is fixed, and external running load for single length 2T is set. It is required to define an optimum combination of thickness of face layers and a core. For definition of critical loading 2Tc the formula from [1] was used. where B is tensile rigidity of a face layer and δ its thickness, ar is reduced length of a panel, H is its thickness, G is shear rigidity of a core. On the third stage a settlement schematization of a wing in the form of a thin-walled cylindrical shell of twice-closed cross-section section at the bending and torsion is used for stress and strain state analysis. The post-buckling behavior of compressed panel is performed by energy approach. Optimization of parameters of a design (thickness of the walls and layers) is performed by a method of iteration.