Consider a finite wing with an area and aspect ratio of 21.5 m2 and 5, respectively (this is comparable to the wing on a Gates Learjet, a twin-jet executive transport). Assume that the wing has a NACA 65–210 airfoil, a span efficiency factor of 0.9, and a profile drag coefficient of 0.004. If the wing is at a 6° angle of attack, calculate CL and CD.
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