XFOIL coordinates

[Paul_BB]

Hello everybody,

Can someone tell me in which reference frame are computed the lift and drag coefficients CL & CD?
Is it in the wind frame (x = velocity) or the airfoil frame (x = chord)?

Thanks

Regards,
Paul

[kcaldwel]

L = 1/2*ro*V^2*S*Cl
D = 1/2*ro*V^2*S*Cd

ro = mass density of the air
V = velocity
S = wing area

Cl and Cd are non-dimensional, and do not include the air velocity or the chord. They are both highly dependent on the Reynolds Number (Re), which is a factor of the wing chord and velocity.

Kevin

[Montag DP]

Lift is always defined normal to the upstream velocity, and drag is always parallel to it, if that's what you're asking. So no matter what the angle of attack of the airfoil, Cl and Cd refer to forces relative to the direction of the upstream airflow.

If you wanted to get the components normal to the airfoil (N) and parallel to the airfoil (A), you'd use the following equations:

N = Lcos(alpha) + Dsin(alpha)
A = Dcos(alpha) - Lsin(alpha)

[Paul_BB]

Lift is always defined normal to the upstream velocity, and drag is always parallel to it, if that's what you're asking.

Thanks, that is what I wanted to know. :)

Best regards,
Paul

[bwalt822]

Lift is always defined normal to the upstream velocity, and drag is always parallel to it, if that's what you're asking. So no matter what the angle of attack of the airfoil, Cl and Cd refer to forces relative to the direction of the upstream airflow.

If you wanted to get the components normal to the airfoil (N) and parallel to the airfoil (A), you'd use the following equations:

N = Lcos(alpha) + Dsin(alpha)
A = Dcos(alpha) - Lsin(alpha)


I second Montags correctness.