Need help with this problem

Description Wing Horizontal Tail Vertical Tail b Span 50 ft 25 ft 15 ft C Chord Length 6 ft 4 ft 4 ft G Lift Slope ) ! i 2 2 g1 degree o degree "~ degree oL Zero-lift AoA -4 degrees 0 degrees 0 degrees Cm,p,airfoil | Airfoil moment coefficient | -0.05 0 0 about aerodynamic center i Incidence Angle +2 degrees +3 degrees 0 degrees Note also: (a) Fuselage diameter of the plane tapers flush with the tails (b) Aerodynamic centers of the wing, horizontal tail, and vertical tail are at their respective quarter-chords (c) Fuselage, flap, and propulsion contributions to pitching moment are negligible (d) Leading edge of the wing is located 20 ft aft of the aircraft's nose (e) Aircraft's center of gravity is located 23 ft aft of the aircraft's nose (f) Leading edge of the horizontal tail is located 38 ft aft of the aircraft's nose (g) Leading edge of the vertical tail is located 41 ft aft of the aircraft's nose is 7 ft where it intersects the wings, while the aft body (h) When calculating static pitch stability, j; =03,n,=10 (i) When calculating static yaw stability, By _ 2B Calculate the following aircraft properties: Moment coefficient Cr c; when the vehicle angle of attack a = 2 degrees. At what angle of attack would the vehicle achieve steady pitch (i.e. Cin g = 0)? Pitching moment derivative Cma 1. DUk wN Neutral point Xyp 1.0,n, = 1.0 I 1 1 I I I I I I I - oo I | T 1 w @ s e Static margin SM. State whether this aircraft has static longitudinal stability. Yaw moment derivative for only the vertical tail: Cnpy