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Aerospace Theoretical lift slope for thin airfoils

  1. Sep 24, 2009 #1

    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.
  2. jcsd
  3. Nov 5, 2009 #2
    The answer is 2pi.
  4. Nov 10, 2009 #3
    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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