![]() 'Dent le batie thermodyasmici confune you about this P V diagram and these P, V, T, S variables? (iv) Find a formula for the minimum power required to fly with that W, S, C and general 3D wing. (iii) Find the formula for the speed at which minimum power P is required to maintain ntralght and level flight for those W, S, p and general 3D wing. (ii) Shetch this P vs V curve f with V on the horizontal ( x ) axis and P on the vertical ( g ) axis. ![]() Use formula for a general wing with general aspect ratio A R. (i) Derive the relation between the power P supplied by the engine and the airplane speed V, for a given weight W, wing area S and air density ρ. T times speed V, P = T V (where V is the airplane speed, that is V = V ∞ ). The power P supplied by the engine equals thirust. For straight and Jevel flight, the thrust T provided by the engine must equal the total drag D, and the lift L must balance the weight W. Assume that the turbulent drag coefficient C t is constant, independent of speed V ∞0 . (e) The total drag coefficient C D is the sum of the induced drag coefficient C D 8 and the turbulent (or 'parasitic') drag coefficient C t (written c d in Anderson eqn. Derive a symbolic formula for the stall speed V n ( = V ∞ ) in terms of the known parameters ( W, S, etc.) clearly showing your logic, then obtain numerical values for α S = 15 degrees, with W = 2500 lbs and S = 16 m 2 at sea level with ρ = 1.225 kg / m 3 for both airplanes. (d) Assume both airplanes have the same total weight W and that they stall at the same angle of attack α n . Clearly display/state which formula or data you are using (from book or lecture notes), then obtain the numerical values. (c) Use Prandtl's lifting line theory to extimate the total lift slope a = d C L / da and induced drag coefficients C D for both (3D) wings. (b) Find the wingspan b 1 and root chord c 1 for the elliptical wing, and the wingspan b 2 and chord c 2 for the rectangular wing, all in terms of the common wing ares S. Find the numerical values in both radians and degrees. Clearly write which formula you are using and show your work. (a) Use thin airfol theory to calculate the angle of sero lift for both types of wing camber lines. Wing 2 has the same total area S as wing 1, but a rectangular planform with aspect ratio b 2 / S = 4, and a symmetric NACA 0012 airfoil with a straight camber line, no twist. ![]() The section camber line is z = θ ⋅ 2 x ( 1 − c x ), 0 ≤ x ≤ c with the same shape at every y and no twist (no geometric or aerodynamic twist). Consider two wing designs with the same total wing area S - Wing I his an elliptical planform with aspect ratio b 2 / S = 8 with chord c ( y ) = c 1 1 − 4 y 2 / b 2 where b is the total wingspan and − b /2 ≤ y ≤ b /2 is the coordinate along the span.
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