I have a big problem to calculate my drag and lift coefficients. On the one hand, I use the command AERO which calculate these coefficients given the velocity, the density and the area. In fact my problem is the area to use. I have a 2d airfoil, 1 meter long and 0.12 meter high (Naca 0012). As with StarCD we have to use 3d airfoil, I have a span from 0.1 meter long. So the area I use is 0.12. This funtion gives drag coefficient twice much bigger as the experimental one, and lift coefficient really really too small ! On the other hand, I calculate myself the coefficients. I get the wall forces x and y data ; with excel I divide them by 0.5*rho*v^2*S, and I make the sum around all the airfoil. The drag coefficient is exactly the same as the one calculated directly by the AERO command, but the lift coefficient is very bigger, and closer with the experimental data but not still in agreeement. I am completely lost ! Who can help me ? Thanks for your answer.

Sorry, the command I use is called ACOEF

I am dealing with exactly the same problem. What I do is to calculate myself the aerodyanmic coefficients(as you said using excel). What exactly do you mean by saying that you make the sum around all the airfoil? Waht are the values you are getting for the aerodynamic coefficients?

I mean I take each value of all the forces which apply on each cell on the airfoil´s surface (with the command GETW FTXY) and I make the sum.

Ok, I found the problem. You have to select only the airfoil´s cell and this between the GETW and the ACOEF command. That means the command are the following : LOAD CASE.PST GETW FTOT CSET NEWS "airfoil" ACOEF DRAG UREF DREF AREF

Normally you will find the same results as the one calculated yourself with excel (I do).

Which area do you use in the ACOEF command ?

You don't really have to do that. If you open the file with extension erd the last three values are you lift and drag forces. Then, you only have to apply the formula in order to get the aerodynamic coefficients