Computation of supersonic inviscid flow around wings with a detached shock wave

The numerical method for computing inviscid supersonic flow around the front part of the arbitrary planform wings was presented in 1,2]. The aim of this paper, which is the extension of the previous ones, is to present a numerical method for computing the flow field above the remaining parts of the wing—the wingtip section and the central section of the wing. Here the problems are formulated for a three-dimensional steady gas-dynamic system of equations written in special curvilinear coordinates. Complicated physical domains of the solution are mapped on simple computational domains. For approximation of the differential equations the finite-difference second-order implicit schemes are used. The approximation of the wing surfaces is made with the help of the local cubic splines. According to the obtained algorithms calculations were made for the wing with elliptical planform and thick airfoil at M_∞ = 2 and M_∞ = 3.5 with the angle of attack α = 5°.