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.

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heinekenisnic	theoretical lift slope for thin airfoils Tweet 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.

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.