Abstract
A numerical method for computing inviscid transonic flow around an arbitrary aircraft configuration is described in this paper. Basic equations are simultaneous equations for velocity potential and Mach number. These equations are equivalent to the full potential equation. The simultaneous equations are numerically solved satisfying given boundary conditions. Shock waves are captured as discontinuous surfaces where shock waves are expected to occur.
The calculated flow field is a rectangular box, where grid points and boundary points are generated in Cartesian coordinates. The grid points are equidistantly distributed, but their intervals can be varied with the location. External forms of the aircraft and lifting surfaces are shaped by use of numerous small triangular planes. The majority of the boundary points are on these external forms. The basic equations are expressed as difference equations using the grid points and the boundary points. The difference equations are solved in the numerical procedure. Several calculated results of this method are compared with the experimental and other numerical results.
