
#1
Sep2409, 06:32 PM

P: 3

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. 



#3
Nov1009, 02:11 PM

P: 261

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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