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
