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Problem 2 (20 points) The measured lift slope for a new airfoil is 0.0975 degree, with...
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 +...
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...
P2. Consider two wings with a NACA 23012 airfoil section: (a) one with an aspect ratio of 4 and (...
P2. Consider two wings with a NACA 23012 airfoil section: (a) one with an aspect ratio of 4 and (b) the other wing with an spetratio of 10. The span efficiency factor is e-0.95 for both wings. Both wings Note: the lift slope for a NACA 23012 is 0.106 per degree, and are flying at an angle of attack of 2 1.5° alculate the change in lit coefficient for both wings if the angle of attack is perturbed by an...
Problem 4 (20 points) (a) An NACA 2412 airfoil with chord length of 6 ft is...
Problem 4 (20 points) (a) An NACA 2412 airfoil with chord length of 6 ft is at a 5° angle of attack during sea-level flight at V = 100 ft/s. Calculate the lift and moment (about the quarter chord) per unit span. (See Fig. 4.10). (b) Find the profile drag coefficient for the airfoil of Part (a), and drag force per unit span. (See Fig. 4.11).
Problem 4 (20 points) (a) An NACA 2412 airfoil with chord length of 6 ft is...
Problem 4 (20 points) (a) An NACA 2412 airfoil with chord length of 6 ft is at a 5° angle of attack during sea-level flight at V = 100 ft/s. Calculate the lift and moment (about the quarter chord) per unit span. (See Fig. 4.10). (b) Find the profile drag coefficient for the airfoil of Part (a), and drag force per unit span. (See Fig. 4.11).
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)
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...
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...
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 +...
P2. Consider two wings with a NACA 23012 airfoil section: (a) one with an aspect ratio of 4 and (b) the other wing with an spetratio of 10. The span efficiency factor is e-0.95 for both wings. Both wings Note: the lift slope for a NACA 23012 is 0.106 per degree, and are flying at an angle of attack of 2 1.5° alculate the change in lit coefficient for both wings if the angle of attack is perturbed by an...
Problem 4 (20 points) (a) An NACA 2412 airfoil with chord length of 6 ft is at a 5° angle of attack during sea-level flight at V = 100 ft/s. Calculate the lift and moment (about the quarter chord) per unit span. (See Fig. 4.10). (b) Find the profile drag coefficient for the airfoil of Part (a), and drag force per unit span. (See Fig. 4.11).
Problem 4 (20 points) (a) An NACA 2412 airfoil with chord length of 6 ft is at a 5° angle of attack during sea-level flight at V = 100 ft/s. Calculate the lift and moment (about the quarter chord) per unit span. (See Fig. 4.10). (b) Find the profile drag coefficient for the airfoil of Part (a), and drag force per unit span. (See Fig. 4.11).
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)
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...
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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