5) A wedge airfoil is placed in a supersonic flow at M = 2. G. Calculate the lift and drag coefficients at 4° angle of attack. Hint: Consider the free-stream pressure at the base of the airfoil. 4°--
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- The NACA 64(l)–412 airfoil has a lift-to-drag ratio of 50 at 0° angle of attack. At what angle of attack does this ratio increase to 80?a)What is the impact of increasing Reynolds number on skin friction and pressure drags over an airfoil? What can be happened for separation in this case? b) What is an adverse pressure gradient and where does it occur on an airfoil (show that on a sketch)? c) Why lift-to-drag ratio is an important parameter for an aircraft? d)How can changing in altitude affect the aircraft power required, PR? Show thatmathematically and graphically?The Boeing 747 has a wingspan of 65 meters and AR of 7.7. (a)Assuming take off weight of 910,000 pounds and a take off velocity 160 knots, calculate the lift coefficient at take off for standard sea-level conditions.(b) Compare the above result with the lift coefficient for cruise at Mach number of 92% of the speed of sound.
- Air flows at Ma = 2.5 past a half-wedge airfoil whoseangles are 4°, as in Fig. Compute the lift and dragcoefficient at α equal to (a) 0° and (b) 6°.Consider the same Lockheed F-104 supersonic fighter shown, with the same flight conditions of Mach 2 at an altitude of 11 km. For these conditions the wing angle of attack is α = 0.035 rad = 1.98◦. Assume the chord length of the airfoil is 2.2 m, which is approximately the mean chord length for the wing. Also, assume fully turbulent flow over the airfoil. Calculate: (a) the airfoil skin friction drag coefficient, and (b) the airfoil wave-drag coefficient. Compare the two values of drag.Air is flowing past a symmetrical airfoil at an angle of attack of 5°. Is the (a) lift and (b) drag acting on the airfoil zero or nonzero?
- Consider an airfoil at 12◦ angle of attack. The normal and axial forcecoefficients are 1.2 and 0.03, respectively. Calculate the lift and dragcoefficients.Consider a NACA 2412 airfoil at an angle of attack of 4 degrees. If the freestream Mach number is 0.72 determine the lift coefficient.Consider a flat plate at an angle of attack in an inviscid supersonic flow.From linear theory, what is the value of the maximum lift-to-drag ratio,and at what angle of attack does it occur?
- Consider a flat plate at zero angle of attack in an airflow at standard sea level conditions (p∞ = 1.01 × 105 N/m2 and T∞ = 288 K). The chord length of the plate (distance from the leading edge to the trailing edge) is 2 m. The planform area of the plate is 40 m2. At standard sea level conditions, μ∞ = 1.7894 × 10−5 kg/(m)(s). Assume the wall temperature is the adiabatic wall temperature Taw. Calculate the friction drag on the plate assuming a turbulent boundary layer for a freestream velocity of (a) 100 m/s, and (b) 1000 m/s.An airfoil generates a 1200 N/m sectional lift when traveling at 75 m/s at sea level. What is the circulation generated by this airfoil?Read the question carefully and give me right solutions with clear calculations. The air flows past a 1m diameter disk oriented normal to upstream flow at 6.8m/s. The pressure distribution on front of the disk is function of the radius, where p(r) = (25 − 100 r2 ) N/m2 , and the pressure acts uniformly on back side with suction pressure of pback = −22 P a. a. Determine the total drag force on the disk. b. Calculate the drag coefficient for the disk . (Density = 1.25 kg/m3 ).