theoretical lift slope for thin airfoils

[heinekenisnic]

hi,

I am required to search the internet to find out what the theoretical value of the lift slope (dcl/dalpha) is for thin airfoils.

Cl is the lift coefficient and alpha is the angle of attack of the airfoil. Does anyone have any ideas? Thanks for your time.

[Cyrus]

The answer is 2pi.

[Brian_C]

Ignoring thickness effects, the slope is simply 2*Pi as stated above. For a Joukowski airfoil with a small but finite thickness, the slope is 2Pi(1+.77t/l), where t is the maximum thickness and l is the chord. The effective angle of attack is alpha+2h/l, where h is the maximum camber of the centerline.