Hi, I'm having problems with the equations used to calculate Cd, Cl, Cm and Cn for an airfoil analisys (2D the extruded to 3D using 1 cell thick). The equations I use are as follow:

Cd = Drag / (0.5*Density*Chord*Speed^2) Cl = Lift / (0.5*Density*Chord*Speed^2) Cm = Moment / (0.5*Density*Chord*Speed^2) Cn = Normal / (0.5*Density*Chord*Speed^2)

All the expressions used as CEL expressions in CFX-pre gives me a dimensional result (most of them in meters) and it must be dimensionless..I can't find equations for airfoils (2D) since all the equations are used in wings. I want to compare some results with X-Foil so I need tp use 2D expressions.

Anyone have any idea?

Andres Bernal Bogota - Colombia.

I think the problem is with the reference Area...I don't know how to calculate these value for an airfoil; Can anyone help me?.

Thanks in advance for your time,

Andres Bernal O.