Hello, I'm doing simulation on airfoil. Rn: 400 000 I'm not able to obtain a good lift or drag coeficient. I'm getting lift coefficient in the 500 to 2500 range and sometime negative drag! (wow that a revolution!) When I reduce the speed by 2 the Cl reduce by approx. 4!

I can't understand why. I also tryed the Cornell tutorial and i'm getting crazy result, I tryed the fluent tutorial about 0.8 mach airfoil... calculation error.. and that even if i'm using my mesh or the one from the tutorial!

Someone had this problem?

Set the reference values ....

Monitor the forces on the wing to see if the forces are constant for each iteration.

hi are you peforming tutorial from Fluent 6.2.16/6.3?

Hello, The reference value was a big part of the problem... Long time i did any simulation and i forgot that setting. I was monitoring the force and in fact it was looking fine: no fluctuation and was looking like converging but in fact i was not. I changed the solver to coupled implicit and now thats fine!

A last question: The Cl sem very good but de Cd is not that good. I wanna be sure that the Cd given by fluent is relative to the frontal aera of the airfoil. And that in practice the Cd given in airfoil spec is relative to the surface of the wing...

Thanks Martin