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.
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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
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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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- A 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_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_forwardA 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_forward
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