Airfoil-values of cd/cl are too high-please help!

matthias

Hi, i try to recompute the polare of an airfoil in 2d (a laminar profile), that was testet in a windtunnel (Reynoldsnumber approximately 1400000). i tried it with different turbulence modells (spalart-allmaras,k-w standart,k-w sst/transition) but every time my values for cl and cd where too high (in the range of alpha 0°-6° cl was correct, for higher angles it was about 1.1-1.2 times higher then at the experiment, cd was about 1.8-2 times higher for all angles of attack (e.g for 0° cd(exp.):0.006/cd(fluent):0.011). is there perhaps anybody who can help me to advance my values?

my boundary conditions are (SI-values)-> turb intensity:0.01, turb lengthskale:0.01, roughness:0.000005, velocity inlet, pressure outlet, incompressible flow, y+max:1.25, steady-state, the length of the control volume is about 100xchordline, the reference values for cl and cd are computed from the velocity inlet, about 195000 nodes at the grid, hybrid grid.

thanks in advance matthias