Cd and Cl coefficients
 

Good day,

I am trying to get Cl and Cd coefficients in post result expression tab, by writting formulas:

Fy=force_y()@Airfoil
Fx=force_x()@Airfoil
Lift =cos(AOA)*Fy-sin(AOA)*Fx
Drag =cos(AOA)*Fx +sin(AOA)*Fy
Denom=0.5*massFlowAve(Density)@Inlet*Uinf^2*Area
cL=Lift/Denom
cD=Drag/Denom

Im struggling with the Area part, because I dont know what formula to use for calculating airfoil area. Im researching NACA0012 airfoil with angle of attack of 0.

Any help would be really appreciated because I want to validate my CFX results with XFOIL calculations of Cl and Cd

You need to read the papers you are comparing against because CL and CD definitions vary depending on what you are doing.

For most bodies CL and CD are based on frontal area. But for airfoils the area is the planform area. If you are doing NACA0012 work it is almost certainly planform area, but check first.

Hi,

Thank you for the reply. Could you tell me how can I calculate the planform area? Is there a formula for it regarding the airfoil?

Refer to: https://lmgtfy.com/?q=planform+area