[shahrooz88]

Dear All, I have been trying to conduct numerical simulation on NACA43012 airfoil, but unfortunately i was not able to achieve verified results for cl, while the answers i get for cd is correct. I am using C-mesh topology with 7 million cells, i prepared the mesh Using POINTWISE software and imported into FLUENT, I achieved cl of 0.04 while the published data is 0.2, and for Cd i achieved 0.0069 which is the same as the published data, i would appreciate if anyone could guide me or give me some hints regarding the uncorrect cl.

Inlet Velocity : 23.9
Re : 500,000
Incompressible flow
Chord: 0.3048
Span: 0.85407