Trouble with airfoil drag prediction.
 

Hello.
I know quite a few questions of this sort have been posted here, so I looked through some pages but found nothing useful for my case.

I try to calculate Cl and Cd for an airfoil at Re=2e6. I get more or less accurate results for Cl but Cd is nit correct no matter what I'm doing to the mesh. I tried over 20 different mesh sizes, Inflation settings and so on but it doesn't seem to work...
Could you please suggest some "example" settings or something like that for CFX-mesh and CFX-Pre so I have a starting point?

Hi,

Can you describe what you are trying to model? Is the wing stalled? Does it have small separations or fully attached flow? Is turbulence transition important? 2D or 3D? Compressible or incompressible? What Mach number?

Glenn Horrocks

Sorry :)

It is a NACA 23020 airfoil with 1.2m long chord. It is at 0degree angle of attack. The simulation is 2D, the velocity of flow is 20m\s, normal pressure, density = 1.184.
Using k-omega turbulence model, Eddy Length Scale in Global Initialisastion is set to 0.1m. Y-plus is somewhere around 1 in dense meshes but not more than 12.
I need only accurate Cd and Cl that are close to the experimental data. Right now I'm getting Cl = 0.04-0.06 and Cd = 0.01, but it should be: Cl = 0.06 and Cd = 0.006.

Yesterday I somehow achieved almost correct results but then I decreased the mesh scale (0.5 which led to 700 000+ element mesh) and the results were incorrect (similar to above).

Try SST turbulece model

I did... Same results.