[krvm]

Hi everyone, I am performing a simulation in ANSYS Fluent where I want to plot lift versus drag coefficient, at several angles of attack, over a NACA 0012 airfoil.

As a reference, I am using Abbott's data (1959), where it says that for alpha = 10 degrees, Reynolds = 3e06 and assuming K-omega SST turbulence hypotheses, the values should be: Cl = 1.08, Cd = 0.015. I'm getting the right value for Cl, but my Cd lays at around 0.04.

I asked my teacher about it and he said it could be due to an error in the wing surface I am using. Thing is, I don't know how to check or change this value, and I'd appreciate any help that you could give me. Thanks in advance!