Drag coefficient error
 

Hi all,

I'm relatively new to CFD and could use a few pointers on my following problem.

I'm using Gambit 2.4.6 to generate a NACA0006 airfoil then Fluent 13 to calculate lift and drag coefficients at several angles of attack. Reynolds number is 3e6, laminar model, pressure based solver.

I have refined the mesh several times and have converged on a lift and drag value for each case. Comparing to Xfoil and 'Theory of Wing Sections' my lift is in the order of 2-7% accurate but for almost every case my drag is approximately double.

I've tried to refine the mesh further close to the airfoil surface but after a certain level of refinement my continuity residual no longer converges.

Thoughts?

Regards,
Josh

What turbulence model are you using?