Consider an aircraft with a finite wing area of 60.0 m2 and an aspect ratio of 8. Assume the wing is composed of a NACA 65-210 airfoil with

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Consider an aircraft with a finite wing area of 60.0 m2 and an aspect ratio of 8. Assume the wing is composed of a NACA 65-210 airfoil with a span efficiency factor of 0.9. The measured profile drag coefficient is 0.012 for the entire aircraft. If the wing is at a 4 degree angle of attack, calculate CL and CD.

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  1. Answer:

    CL = 0.57

    CD = 0.027

    Explanation:

    Thinking process:

    Let the parameters be:

    wing area = 21.5 m²

    aspect ratio = 5

    span efficiency factor = 0.9

    CD₀ = 0.004

    Angle AOA = 6°

    Therefore,

    [tex]CD = CD_{0} + \frac{(CL)^{2} }{ITEAR}[/tex]

    For the NACA 65210

    α = 9° ; CL = 1.05

    α = -1.5 ; Cl = 0

    Therefore, [tex]a_{0} = \frac{1.05- 0}{9-(-1.5)} \\ = 0.11[/tex]

    Lift slope for finite wing is given by:

    [tex]a = \frac{a_{0} }{1+\frac{a_{0} }{ITEAR} } \\ = \frac{0.11}{1+\frac{57.3(0.11)}{\pi }(0.9)(5) }\\ = 0.076 deg[/tex]

    at α = 6°, CL is given by:

    a ([tex]\alpha -\alpha _{L=0}) = 0.076(6-(-1.5)\\ CL= 0.57[/tex]

    CD = [tex]0.004 + \frac{(0.57)^{2} }{\pi*0.9*6 }[/tex]

         = 0.027

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