You need to apply C = CL(a a=0) to both the wing and the horizontal...

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You need to apply C = CL(a – a=0) to both the wing and the horizontal tail. You can assume that the dynamic pressures at the wing and the tail are the same. Problem 5.2 refers to the previous problem. For the aircraft in Problem 5.2, with the additional information ARW =7.65 aL-Owing = -2 deg e =1 aL-Oail = 0 deg AR = 5 Sw = 240 ft? S = 40 ft? a) What is the trimmed angle of attack of the tail? b) What is the trimmed angle of attack of the wing? You need to apply C = CL(a – a=0) to both the wing and the horizontal tail. You can assume that the dynamic pressures at the wing and the tail are the same. Problem 5.2 refers to the previous problem. For the aircraft in Problem 5.2, with the additional information ARW =7.65 aL-Owing = -2 deg e =1 aL-Oail = 0 deg AR = 5 Sw = 240 ft? S = 40 ft? a) What is the trimmed angle of attack of the tail? b) What is the trimmed angle of attack of the wing?

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