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Aircraft Stability and Control Submit your answers to all questions below. Show your working steps insufficient details 1. (25pts.) From wind-tunnel test, it is found that flying-wing aircraft model has zero-lif angle of attack of 1.5° and lift-curve ope of 4.6/rad. The center of gravity of the aircraft is located at 35% of the mean aerodynamic chord (c) from the wing leading edge. The pitching moment coefficients about the aircraft center of gravity are measured as -0.01 when the angle of attack is 1.0° and 0.05 when the angle of attack is .88°. a. Calculate Cm and determine whether the aircraft is statically stable longitudinally b. Determine the location of the aerodynamic center of the flying wing. A horizontal tail with tail-volume ratio of 0 34 and lift-curve slope of 5.7 /rad added to the flying-wing to modify its stability characteristics. It is estimated that effect of the downwash on the tail delda 0.35. The horizontal tail efficiency is assumed to be 100% Neglect the lift contribution of the tail to the whole aircraft aerodynamics. Also assume that aircraft 's center of gravity (CG) is not affected by the installation of the horizontal tail c. Determine the stick-fixed static margin of the new configuration and its longitudinal static stability d. What the most aft allowable center of gravity location for static longitudinal stability? e. Calculate the angle of attack to trim the modified aircraft if the pitching moment coefficient value about c is 0.06 at zero angle of attack. 2. (30pts.) Pitching moment coefficient vs lift coefficient data of an aircraft at various fixed elevator deflections are given in the figure below a. Determine the stick-fixed neutral point of the aircraft and its static margin Is the aircraft statically stable longitudinally? b. If the aircraft : tobe flown steady and level at 38.4 m/s, determine the trim lift coefficient and the elevator angle to trim. c. Estimate the elevator control power of the aircraft from the figure. Hint: Determine the change in Cm values with the change of for fixed Cr value. d. If the elevator control power of the aircraft is as calculated above and the maximum +10 elevator deflection i S, determine the most forward CG location so that the -20° aircraft can be trimmed at its maximum lift coefficient< 1.4. e. What the range of CGlocation to satisfy longitudinal static stability and trim requirement of the aircraft? 1 0.15 Weight 12,237 N 0.10 Wing area m² xcg=25% MAC 0.05 Cmco CL 0.0 0.4 0.6 0.8 8,=-15° -0.05 -0.10 80=-5° 80=0° -0.15 8,=5° 3. (20pts.) A conventional airplane has the following wing and tail lift slope characteristics: C1. -0.09/° CL =0.08/° Its horizontal tail volume ratic is 0.4 with tail efficiency of The downwash effect on the tail is given by delda 0.4. The effect of elevator on lift coefficient is given by CLE =0.015/°. The hinge moment of this airplane has the following characteristics: Ca =0 Co. =-0.003/° =-0.005/° a. Determine the ratio of the free- -floating angle of the elevator to the angle of attack of the tail. What does the sign of this atio mean? b. If the stick -fixed neutral point is located at 0.45c , determine the location of the stick-free neutral point. c. In stick- free condition is the airplane statica ally more stable or less stable (longitudinally) than in stick-fixed condition? Briefly explain your answer. 4. (25pts.) Wing fuselage data of an airplane is given in the figure below. a. Estimate wing fuselage contribution to static directional stability of the airplane for flight at 150 m/s at sea level. b. If : vertical tail with lift slope of 10.08/° is be added on the airplane at distance< of 4 m from the CG, determine the area of this vertical tail so that the directional stability of the airplane has a value of Cno =0 the flight conditions above. Assume 2 S-21.3m³ b-10.4m z_-0.4m d-1.6m 7=13.7m w,= 1.6 m n-1.6m h,=1.6m h e h2 : Appendix Characteristics of the International Standard Atmosphere, SI Units Altitude. Temperature, T Pressure, P Density,A Speed of Viscosity,A km K N/m² kg/m Sound. kg/ms m/s 288.16 101325 1.225 340.3 1.79E-05 0.5 284.91 95461 1.1673 338.4 1.77E-05 281.66 89876 1.1117 336.4 1.76E-05 1.5 278.41 84560 1.0581 334.5 1.74E-05 2 275.16 79501 1.0066 332.5 1.73E-05 2.5 271.92 74692 0.95696 330.6 1.71E-05 3 268.67 70121 0.90926 328.6 1.69E-05 3.5 265.42 65780 0.86341 326.6 1.68E-05 262.18 61660 0.81935 324.6 1.66E-05 4.5 258.93 57752 0.77704 322.6 1.65E-05 5 255.69 54048 0.73643 320.5 1.63E-05

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Aircraft Stability and Control
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