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Coefficient of lift for a wing is CL

  1. Nov 6, 2005 #1
    OK, so the coefficient of lift for a wing is CL.

    For an airfoil, it is cl.

    How do I go from Lift for an airfoil to Lift for the whole wing? Is it just multiplying by the wingspan?

  2. jcsd
  3. Nov 6, 2005 #2


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    You use the CL to calculate the force of lift. The lift would be over the entire span if there is no variation in the cord along the entire span.
  4. Nov 6, 2005 #3
    I understand that part, but my book has L for the wing, and L' for the airfoil.

    So, assuming that chord is constant, would L'*Span = L?
  5. Nov 7, 2005 #4
    Show the definition of [tex]L[/tex] and [tex]L'[/tex]..?
  6. Nov 7, 2005 #5
    I think I figured it out.

    I found an equation that relates the L' to the position on the wing. By integrating that over the span, I ought to have the total lift of the wing. I can then use that to find my CL variable.

    Now is there a way to solve for the limits if I know what the integral needs to equal? Other than guess and check?

    Obviously the lower limit would be the negative of the upper limit.

    BTW, the wing is continuous over the whole craft, the fuselage does not interfere with the wing at all.
  7. Nov 8, 2005 #6


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    Not that I can think of off the top of my head. I would probably do trial and error. Then again, integration was always my bane. You might want to pose that question in the math forums.
  8. Nov 8, 2005 #7
    Sounds good.

    Oh, the equation I found applies only to elliptical wings. What would I do if the wing was a trapezoid?

    L' = L'o*(.5*(b1+b2)*h)?

    That equation being lift for an airfoil section any distance from the root.

    If no one knows, not a big deal, i have a meeting with profs today.
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