Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 3, Problem 3.4P
Consider a low-speed open-circuit subsonic wind tunnel with an inlet-to-throat area ratio of 12. The tunnel is turned on. and the pressure difference between the inlet (the settling chamber) and the test section is read as a height difference of 10 cm on a U-tube mercury manometer. (The density of liquid mercury is
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Consider a low-speed subsonic wind tunnel with a 12/1 contraction area ratio for the nozzle. If
the flow in the test section is at a standard sea level conditions with a velocity of 50 m/s,
calculate the height difference in a U-tube mercury manometer with one side connected to the
nozzle inlet and the other to the test section. pHg = 13.6 x 103 kg/m3.
An airplane is flying at a standard altitude of 10,000 ft. A Pitot tube mounted at the nose measures a pressure of 2220 lb/ft2. The airplane is flying at a high subsonic speed, faster than 300 mph. The flow should be considered compressible. Calculate the velocity of the airplane.
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Consider a low-speed subsonic wind tunnel with a 12/1 contraction area ratio for the nozzle. If the flow in the test section is at a standard sea level conditions with a velocity of 50 m /s, calculate the height difference in a U-tube mercury manometer with one side connected to the nozzle inlet and the other to the test section. ρHg = 13.6 x 103 kg/ m^3.
Chapter 3 Solutions
Fundamentals of Aerodynamics
Ch. 3 - For an irrotational flow. show that Bernoullis...Ch. 3 - Consider a venturi with a throat-to-inlet area...Ch. 3 - Consider a venturi with a small hole drilled in...Ch. 3 - Consider a low-speed open-circuit subsonic wind...Ch. 3 - Assume that a Pitot tube is inserted into the...Ch. 3 - A Pilot tube on an airplane flying at standard sea...Ch. 3 - At a given point on the surface of the wing of the...Ch. 3 - Consider a uniform flow with velocity V. Show that...Ch. 3 - Show that a source flow is a physically possible...Ch. 3 - Prove that the velocity potential and the stream...
Ch. 3 - Prove that the velocity potential and the stream...Ch. 3 - Consider the flow over a semi-infinite body as...Ch. 3 - Derive Equation (3.81). Hint: Make use of the...Ch. 3 - Derive the velocity potential for a doublet; that...Ch. 3 - Consider the nonlifting flow over a circular...Ch. 3 - Consider the nonlifting flow over a circular...Ch. 3 - Consider the lifting flow over a circular cylinder...Ch. 3 - The lift on a spinning circular cylinder in a...Ch. 3 - A typical World War I biplane fighter (such as the...Ch. 3 - The Kutta-Joukowski theorem, Equation (3.140), was...Ch. 3 - Consider the streamlines over a circular cylinder...Ch. 3 - Consider the flow field over a circular cylinder...Ch. 3 - Prove that the flow field specified in Example 2.1...
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- In a low-speed subsonic wind tunnel with a closed test section, a static pressure tap on the wall of the tunnel test section measures 0.98 atm. The temperature of the air in the test section is 80 ˚F. A Pitot tube is inserted in the middle of the flow in the test section in order to measure the flow velocity. The pressure measured by the Pitot tube is 2,200 psf. Calculate the flow velocity (in KPH) in the test section.arrow_forwardConsider a wing mounted in the test-section of a subsonic wind tunnel. The velocity of the airflow is 160 ft/s. If the velocity at a point on the wing is 195 ft/s, what is the pressure coefficient at this point?arrow_forwardQ6/ Air is heated as it flows subsonically through a duct. When the amount of heat transfer reaches 60 kJ/kg, the flow is observed to be choked, and the velocity and the static pressure are measured to be 620 m/s and 270 kPa. Disregarding frictional losses, determine the velocity, static temperature, and static pressure at the duct inlet.arrow_forward
- An airplane is flying at a pressure altitude of 15 km with a velocity of 619 m/s. The outside air temperature is 220 K. What is the pressure measured by a Pitot tube mounted on the nose of the airplane?arrow_forwardAn engineer is designing a subsonic wind tunnel. The test section is to have a cross-sectional area of 4 m2 and an airspeed of 60 m/s. The air density is 1.2 kg/m3. The area of the tunnel exit is 10 m2. The head loss through the tunnel is given by hL=0.025VT2/2g, where VT is the airspeed in the test section. Calculate the power needed to operate the wind tunnel. Hint: Assume negligible energy loss for the flow approaching the tunnel in region A, and assume atmospheric pressure at the outlet section of the tunnel. Assume α = 1.0 at all locations.arrow_forwardSuppose an aircraft is flying at standard sea-level at M = 0.8 and using a pitot-static tube for airspeed measurement. Determine the actual difference in total and static pressures as would be measured by the pitot-static system. Compare the speed of the aircraft as determined using the actual pressure difference and incompressible flow versus that knowing the actual Mach number. Use k = 1.4.arrow_forward
- A nozzle for a supersonic wind tunnel is designed to achieve a Mach number of 3.2, with a velocity of 2500 m/s, and a density of 1.0 kg/ m³ in the test section. Find the temperature and pressure in the test section and the upstream stagnation conditions. The fluid is helium. Te = i Pe= To= Po= Mi K kPa K kPaarrow_forwardA Pitot tube inserted at the exit of a supersonic nozzle reads8.92 × 104 N/m2. If the reservoir pressure is 2.02 × 105 N/m2, calculatethe area ratio Ae/A∗ of the nozzle.arrow_forwardConsider a circular cylinder in a hypersonic flow, with its axisperpendicular to the flow. Let φ be the angle measured between radiidrawn to the leading edge (the stagnation point) and to any arbitrary pointon the cylinder. The pressure coefficient distribution along the cylindricalsurface is given by Cp = 2 cos2 φ for 0 ≤ φ ≤ π/2 and 3π/2 ≤ φ ≤ 2πand Cp = 0 for π/2 ≤ φ ≤ 3π/2. Calculate the drag coefficient for thecylinder, based on projected frontal area of the cylinder.arrow_forward
- How to Calibration of Air Velocity from Pressure Using Sub-Sonic Wind Tunnel and Write the components of Sub-Sonic Wind Tunnel?arrow_forwardThe nozzle of a supersonic wind tunnel has an exit-to-throat area ratioof 6.79. When the tunnel is running, a Pitot tube mounted in the testsection measures 1.448 atm. What is the reservoir pressure for thetunnel?arrow_forward85. Calculate the altitude of a supersonic airplane (v = 800 m-s1, T=-34 C°) when the time between seeing above observer and hearing is 9.9 sec.arrow_forward
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