piper cherokee wing area is 16m², span is 9.75m. total weight is 1100kg. airfoil is NACA...

piper cherokee wing area is 16m², span is 9.75m. total weight is 1100kg. airfoil is NACA 65-415 and lift coefficient slope is 0.1033 per degree. αL=0 is -1.5degree. Assume τ = 0.12. when air density is 1.23Kg/m³ and speed is 193Km/h. Determine the wing's angle of attack

Solutions

Expert Solution

Let the angle of attack be degrees.

The lift coefficient is given by

The lift force is given by

V= velocity=193x1000/3600=53.61 m/s

1.23 kg/m³ (density)

S= wing area= 16 m²

Hence

The lift force needs to be balance the weight of aircraft

Hence

samet mamat answered 1 year ago

Next > < Previous

Related Solutions

A rectangular wing with a NACA 65-210 airfoil section and a chord of 6 in is...

A rectangular wing with a NACA 65-210 airfoil section and a chord of 6 in is mounted in the Auburn University 2 ft × 2 ft wind tunnel, completely spanning the test-section. The tunnel is operating with test section velocity, pressure, temperature, and velocity of 1 atm, 520°R, and 75 ft/s, respectively. (a) Calculate the lift, drag, and moment about the quarter-chord when the angle of attack is 5°. (b) What angle of attack would be required to generate a...

8. A Piper Cherokee is flying at 4000 ft standard altitude at 105 mph. The wing...

8. A Piper Cherokee is flying at 4000 ft standard altitude at 105 mph. The wing has a rectangular planform (shape) with a span of 30 ft and a chord of 5.33 ft. Assuming that the turbulent boundary layer grows as on a smooth plate and a completely turbulent boundary layer flow, determine: (a) The boundary layer thickness at the downstream edge of the wing. (b) The total skin friction drag coefficient of the wing. (c) The total skin friction...

Delta Wing Aircraft Specifications (do the necessary assumptions if needed) Wing Area: 585 m2 Wing Span:...

Delta Wing Aircraft Specifications (do the necessary assumptions if needed) Wing Area: 585 m2 Wing Span: 32 m Taper ratio: 0.019 Leading Edge Sweep angle: 65.57° Wing loading: 4070 N/m2 Cruise Mach Number: 3 Cruise altitude: 22.5 km (L/D) cruise = (L/D) max = 9 Calculate the cruising angle of attack? Write down the cruising drag polar equation? Using the drag polar found in (b), calculate (CL1/2/CD) max and corresponding Mach number maximizing (CL1/2/CD).

5. Consider an airplane with wingspan 40 ft, wing area 250 ft2 , and weight 12,000...

5. Consider an airplane with wingspan 40 ft, wing area 250 ft2 , and weight 12,000 lb. Assume that the Oswald efficiency factor is 0.9. The airplane is flying at a velocity of 250 mph at a standard altitude of 5000 ft. Calculate the airplane drag due to lift.

Subjects