## 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

Question

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 year 2021-08-31T21:32:44+00:00 1 Answers 0 views 0

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,

$$CD = CD_{0} + \frac{(CL)^{2} }{ITEAR}$$

For the NACA 65210

α = 9° ; CL = 1.05

α = -1.5 ; Cl = 0

Therefore, $$a_{0} = \frac{1.05- 0}{9-(-1.5)} \\ = 0.11$$

Lift slope for finite wing is given by:

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

at α = 6°, CL is given by:

a ($$\alpha -\alpha _{L=0}) = 0.076(6-(-1.5)\\ CL= 0.57$$

CD = $$0.004 + \frac{(0.57)^{2} }{\pi*0.9*6 }$$

= 0.027