Hi there,

i really need help in solving a 2D NACA0012 for inviscid flow. In my code, i've employed the following methods:
: 1st-order Steger-Warming FVS
: Finite difference
: Structured grid
: 1st-order explicit time integration
: Local time stepping

When i run my code for the problem at M=0.8(transonic) and 0 degree angle of attack, the shock location is positioned at the trailing edge of the airfoil instead at the middle.

Can anyone advise me on this, i have tried my code with other flow problems (i.e. channel flow, blunt body, regular shock & etc.) and it seems to work out fine.

Thanks in advance.

If you are looking for a second opinion, download my shareware MicroTunnel from www.microcfd.com which you can run for three days. Here a Mach 0.7 flow over NACA0012 at 10 deg angle of attack,

www.microcfd.com/gallery/airfoil-trans-m.htm

Hi; I don't wonder the result since you are using =>first<= order =>FVS=< and =>first<= order time integration scheme. The schock wave will move in the transient stage to end up at a position of about 60% of the chord, this does not occur because of the high diffusion associated to (1)flux vector splitting(2)truncation error of the first order schemes. Try to switch to second order accuracy. Good luck.