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P2. Consider two wings with a NACA 23012 airfoil section: (a) one with an aspect ratio of 4 and (...
The large airplane has a span of 50 meters, an elliptic wing of aspect ratio 6,...
The large airplane has a span of 50 meters, an elliptic wing of aspect ratio 6, a mass of 60,000 kg and is flying at sea level density (1.2 kg/m^3) at 100 m/s. Assume that it has a NACA 0009 (symmetric) airfoil, i. Calculate the actual lift coefficient for the wing ii. Calculate the geometric angle of attack. iii. What is the downwash angle? iv. What is the downwash velocity at the wing? v. What is the strength (circulation) of...
A finite wing area of 1.5 ft2 and aspect ratio of 6 is tested in a...
A finite wing area of 1.5 ft2 and aspect ratio of 6 is tested in a subsonic wind tunnel at a velocity of 180 ft/s at standard sea-level conditions. At an angle of attack of ?1°, the measured lift and drag are 0 and 0.181 lb, respectively. At an angle of attack of 2°, the lift and drag are measured as 5.0 and 0.23 lb, respectively. Calculate the span efficiency factor and the infinite-wing lift slope. (Round the final answers...
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:...
3. The NACA 4412 airfoil has a mean camber line given by 0.25 0.8-- for 0...
3. The NACA 4412 airfoil has a mean camber line given by 0.25 0.8-- for 0 s -0.4 en for 0.4 -?1 Using thin airfoil theory, calculate ?1:0 and Cl when ? = 30. m,c/4 and xcp/c for a -3 Compare the results of part (a) and (b) with experimental data of NACA 4412 airfoil (see plots below) Lift per unit length of span and circulation for an airfoil with chord length of 2 m flying at a standard altitude...
use a Reynolds number of 2.6x10^5 to find using graphs 2. Consider an NACA 23015 airfoil...
use a Reynolds number of 2.6x10^5 to find using
graphs
2. Consider an NACA 23015 airfoil (Fig 5.2a and 5.2b in text) with a chord of 0.64 m in an airstream 1000m above sea level conditions. The freestream velocity is 70 m/s. The lift per unit span is 1200 N/m. Calculate the angle of attack and the drag per unit span. (See example 4.1) CHAPTER S incompressible Flow over Finite Wings Section lift coefficient - Moment coefficient, -20 -32 -24...
Problem 2 (20 points) The measured lift slope for a new airfoil is 0.0975 degree, with...
Problem 2 (20 points) The measured lift slope for a new airfoil is 0.0975 degree, with ac o =-1.15º. Consider a finite wing using this airfoil, with AR of 10 and taper ratio of 0.6. Calculate the lift and induced drag coefficients for this wing at a geometric angle of attack of 9º. Also, find the span efficiency factor. Hint: See Eq. (5.70) and Fig. 5.20. Are the values of lift and induced drag coefficients reasonable? Use 8 =1.
The measured lift slope for a new airfoil is 0.0975 degree-1, with al=0 =-1.15º. Consider a...
The measured lift slope for a new airfoil is 0.0975 degree-1, with al=0 =-1.15º. Consider a finite wing using this airfoil, with AR of 10 and taper ratio of 0.6. Calculate the lift and induced drag coefficients for this wing at a geometric angle of attack of 9º. Also, find the span efficiency factor. Hint: See Eq. (5.70) and Fig. 5.20. Are the values of lift and induced drag coefficients reasonable? do a = (5.70) 1 + (ao/A AR)(1 +...
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...
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)...
The large airplane has a span of 50 meters, an elliptic wing of aspect ratio 6, a mass of 60,000 kg and is flying at sea level density (1.2 kg/m^3) at 100 m/s. Assume that it has a NACA 0009 (symmetric) airfoil, i. Calculate the actual lift coefficient for the wing ii. Calculate the geometric angle of attack. iii. What is the downwash angle? iv. What is the downwash velocity at the wing? v. What is the strength (circulation) of...
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:...
3. The NACA 4412 airfoil has a mean camber line given by 0.25 0.8-- for 0 s -0.4 en for 0.4 -?1 Using thin airfoil theory, calculate ?1:0 and Cl when ? = 30. m,c/4 and xcp/c for a -3 Compare the results of part (a) and (b) with experimental data of NACA 4412 airfoil (see plots below) Lift per unit length of span and circulation for an airfoil with chord length of 2 m flying at a standard altitude...
use a Reynolds number of 2.6x10^5 to find using
graphs
2. Consider an NACA 23015 airfoil (Fig 5.2a and 5.2b in text) with a chord of 0.64 m in an airstream 1000m above sea level conditions. The freestream velocity is 70 m/s. The lift per unit span is 1200 N/m. Calculate the angle of attack and the drag per unit span. (See example 4.1) CHAPTER S incompressible Flow over Finite Wings Section lift coefficient - Moment coefficient, -20 -32 -24...
Problem 2 (20 points) The measured lift slope for a new airfoil is 0.0975 degree, with ac o =-1.15º. Consider a finite wing using this airfoil, with AR of 10 and taper ratio of 0.6. Calculate the lift and induced drag coefficients for this wing at a geometric angle of attack of 9º. Also, find the span efficiency factor. Hint: See Eq. (5.70) and Fig. 5.20. Are the values of lift and induced drag coefficients reasonable? Use 8 =1.
The measured lift slope for a new airfoil is 0.0975 degree-1, with al=0 =-1.15º. Consider a finite wing using this airfoil, with AR of 10 and taper ratio of 0.6. Calculate the lift and induced drag coefficients for this wing at a geometric angle of attack of 9º. Also, find the span efficiency factor. Hint: See Eq. (5.70) and Fig. 5.20. Are the values of lift and induced drag coefficients reasonable? do a = (5.70) 1 + (ao/A AR)(1 +...
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...
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)...
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