|July 12, 2011, 02:01||
Drag coefficient error
Join Date: Jul 2011
Posts: 3Rep Power: 5
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.
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