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For an aircraft shown in Fig. 1, assume that the wing span is b- 60 ft...
3. An airplane (weight 3,000 lb, wing span 35.5 ft, wing area 135 ft2, Oswald efficiency...
3. An airplane (weight 3,000 lb, wing span 35.5 ft, wing area 135 ft2, Oswald efficiency factor 0.85, zero-lift drag coefficient 0.05) flies with 90 kts at 8,000 ft. The asymmetrical airfoil has a lift-curve slope of 0.1 per degree, a zero-lift angle of attack of -3.0°, and a stall angle of attack of 15°. Assume a linear lift curve all the way to the stall. a. Calculate the density at the flight altitude in slug/ft? (6 points)
Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear supersonic theor...
Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear supersonic theory. Ignore 3D effects on lift coefficient and lift force associated with the tail and the fuselage. Useful data: • Total Wing Area: 18.22 m2 • Aircraft Mass: 8,000 kg -Is supersonic linear theory valid for this flight regime? Explain your answer.
3. For this problem, use the following aircraft properties: Gross weight, W-20,000 lbs Wing area, S...
3. For this problem, use the following aircraft properties: Gross weight, W-20,000 lbs Wing area, S 300 ft Mean aerodynamic chord of wing. -6f CG range: Max forward: 25% MAC, Max aft: 55% MAC Aerodynamic center of wing/body is at 25% MAC Tail lift slope, a-4.5 rad Elevator effectiveness, ae 2.0 rad Wing/body lift coefficient, CLwb 5.50ab Wing/body pitching moment coefficient about aerodynamic center, Cmac wb 0.1 · · b, or σε %α-0.3 radi Tail downwash angle, ε-0.3 Aircraft drag...
Remaining Time: 1 hour, 59 minutes, 29 seconds. Question Completion Status: 1. Consider the following airfoil data: 0 80 Stedaro 224 12 Secrion irt cofficent, NACA 2412 Wing Beetion NACA 2412 Wi...
Remaining Time: 1 hour, 59 minutes, 29 seconds. Question Completion Status: 1. Consider the following airfoil data: 0 80 Stedaro 224 12 Secrion irt cofficent, NACA 2412 Wing Beetion NACA 2412 Wing Section (centisund Notice that the arifoil data is shown at Reynold numbers from 3.1 to 8.9 million. The airfoil has a chord of 4 feet and is flying at 83.1 mph at sea level. 3. What are the (2-D) Ch Ca C .. at4 dearees AOA? Remaining Time:...
question 8-Supersonic Flight of F-104 (10 points): Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat pl...
question 8-Supersonic Flight of F-104 (10 points): Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear a) supersonic theory. Ignore 3D effects on lift coefficient and lift force associated with the tail and the fuselage. Useful data Total Wing Area: 18.22 m Aircraft Mass: 8,000 kg Is supersonic linear theory valid for this flight regime? Explain...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
1. For the airfoil Cp data shown below; • What is the maximum airspeed in flow...
1. For the airfoil Cp data shown below; • What is the maximum airspeed in flow just outside the boundary layer if the freestream speed Vo = 120 m/s? • What is the local Mach number at this point if the altitude is 12 km? • What is the approximate value of C, for these conditions? Ans: V = 204.35 m/s, M = 0.693, CL ~ 1.15. -1.00 .. .... -0.50 .... ..... 0 0.1 0.2 0.3 0.4 0.5 0.6...
3. An airplane (weight 3,000 lb, wing span 35.5 ft, wing area 135 ft2, Oswald efficiency factor 0.85, zero-lift drag coefficient 0.05) flies with 90 kts at 8,000 ft. The asymmetrical airfoil has a lift-curve slope of 0.1 per degree, a zero-lift angle of attack of -3.0°, and a stall angle of attack of 15°. Assume a linear lift curve all the way to the stall. a. Calculate the density at the flight altitude in slug/ft? (6 points)
3. For this problem, use the following aircraft properties: Gross weight, W-20,000 lbs Wing area, S 300 ft Mean aerodynamic chord of wing. -6f CG range: Max forward: 25% MAC, Max aft: 55% MAC Aerodynamic center of wing/body is at 25% MAC Tail lift slope, a-4.5 rad Elevator effectiveness, ae 2.0 rad Wing/body lift coefficient, CLwb 5.50ab Wing/body pitching moment coefficient about aerodynamic center, Cmac wb 0.1 · · b, or σε %α-0.3 radi Tail downwash angle, ε-0.3 Aircraft drag...
Remaining Time: 1 hour, 59 minutes, 29 seconds. Question Completion Status: 1. Consider the following airfoil data: 0 80 Stedaro 224 12 Secrion irt cofficent, NACA 2412 Wing Beetion NACA 2412 Wing Section (centisund Notice that the arifoil data is shown at Reynold numbers from 3.1 to 8.9 million. The airfoil has a chord of 4 feet and is flying at 83.1 mph at sea level. 3. What are the (2-D) Ch Ca C .. at4 dearees AOA? Remaining Time:...
question 8-Supersonic Flight of F-104 (10 points): Estimate the angle of attack for the F-104 aircraft flying level (and unaccelerated) at Mach 2.0 and at an altitude of 20,000 ft. Assume flat plate for the wing airfoil and use linear a) supersonic theory. Ignore 3D effects on lift coefficient and lift force associated with the tail and the fuselage. Useful data Total Wing Area: 18.22 m Aircraft Mass: 8,000 kg Is supersonic linear theory valid for this flight regime? Explain...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
80 m 845 m2 7.53 Wing Span Wing Planform Area Wing Aspect Ratio Sweep angle at quarter chord (0.25c) Taper ratio 33.5 0.3 A380 1500 m² Component Wetted Area Fuselage Wing 1700 m Horizontal tail Vertical tail Engines (4) 290 m² (each) 400 m² 280 m² Cruise Mach Number 0.85 Cruising Altitude 12.5 km Wing Loading 16500 N/m2) Temperature Geo potential Altitude above Sea Level -- (m) Acceleration of Gravity (m/s) Absolute Pressure -p- (104 N/m2) Density .p. (101 kg/m3)...
1. For the airfoil Cp data shown below; • What is the maximum airspeed in flow just outside the boundary layer if the freestream speed Vo = 120 m/s? • What is the local Mach number at this point if the altitude is 12 km? • What is the approximate value of C, for these conditions? Ans: V = 204.35 m/s, M = 0.693, CL ~ 1.15. -1.00 .. .... -0.50 .... ..... 0 0.1 0.2 0.3 0.4 0.5 0.6...
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