Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 10, Problem 10.12P
We wish to design a supersonic wind tunnel that produces a Mach 2.8 flow at standard sea level conditions in the test section and has a mass flow of air equal to 1 slug/s. Calculate the necessary reservoir pressure and temperature, the nozzle throat and exit areas, and the diffuser throat area.
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We wish to design a supersonic wind tunnel that produces a Mach 2.8flow at standard sea level conditions in the test section and has a massflow of air equal to 1 slug/s. Calculate the necessary reservoir pressureand temperature, the nozzle throat and exit areas, and the diffuser throatarea.
In the test section of a supersonic wind tunnel, a Pitot tube in the flowreads a pressure of 1.13 atm. A static pressure measurement (from apressure tap on the sidewall of the test section) yields 0.1 atm. Calculatethe Mach number of the flow in the test section.
If the Mach number behind a normal shock is equal to 0.4, calculate the ratio of the readings of pitot tubes located before and after the shock.
Chapter 10 Solutions
Fundamentals of Aerodynamics
Ch. 10 - The reservoir pressure and temperature for a...Ch. 10 - A flow is isentropically expanded to supersonic...Ch. 10 - A Pitot tube inserted at the exit of a supersonic...Ch. 10 - For the nozzle flow given in Problem 10.1, the...Ch. 10 - A closed-form expression for the mass flow through...Ch. 10 - Prob. 10.6PCh. 10 - A convergent-divergent nozzle with an...Ch. 10 - For the flow in Problem 10.7, calculate the mass...Ch. 10 - Consider a convergent-divergent nozzle with an...Ch. 10 - A 20 half-angle wedge is mounted at 0 angle of...
Ch. 10 - The nozzle of a supersonic wind tunnel has an...Ch. 10 - We wish to design a supersonic wind tunnel that...Ch. 10 - Consider a rocket engine burning hydrogen and...Ch. 10 - For supersonic and hypersonic wind tunnels, a...Ch. 10 - Return to Problem 9.18. where the average Mach...Ch. 10 - Return to Problem 9.19, where the average Mach...Ch. 10 - A horizontal flow initially at Mach I flows over a...Ch. 10 - Consider a centered expansion wave where M1=1.0...
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Need a deep-dive on the concept behind this application? Look no further. Learn more about this topic, mechanical-engineering and related others by exploring similar questions and additional content below.Similar questions
- An air tank with a nozzle, has a pressure of twice the standard sea level pressure and density of 2.45 kg/m³. Outside the converging-diverging nozzle, the pressure corresponds to an altitude of 3km and designed to have a mach number of 1.0 and 1.8 at the throat and exit, respectively. The area at the throat is 0.15 square meters. Calculate: a) the temperature and speed of sound in the tank. b) pressure and density at the throat c) mass flow rate at the exit.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(Q) A converging-diverging nozzle discharges an over-expanded jet into a room. Calculate the deflection angle in degrees caused by the oblique shock forming at the nozzle exit? Ratio between the room pressure and the pressure at the nozzle exit is 5.4. Mach number at the nozzle exit is 1.5. Assume the fluid to be air as a perfect gas with k=1.4arrow_forward
- 2. An air tank with a nozzle has a pressure of 196.32 KPa and density of 1.9 Kg/m³. Outside the converging- diverging nozzle, the pressure is atmospheric and designed to have a Mach No. of 1.0 and 1.5 at the throat and exit respectively. The area at the throat is 0.11m2. Calculate the following: (a) Temperature and speed of sound at the tank. (b) Pressure, density, temperature and speed of sound at the throat. (c) Mass flow at the exit.arrow_forwardNozzle is assuming steady one-dimensional flow. M = 2.731. This is the the supersonic flow of air through a convergent-divergent nozzle. The stagnation temperature = 300K, stagnation pressure at the inlet = 107500Pa, static pressure at the exit=4400Pa, C1 is a constant = 0.1097 for calculating circular cross-sectional area of a convergent-divergent nozzle: A = C1 + x^2 and x (axial distance from the throat) =1m. γ = 1.4 and R=287. Calculate the mass flow rate of air through the nozzle. Thank You.arrow_forwardConsider 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.arrow_forward
- The pressure ratio across a normal shock wave that occurs in air is 1.25. Ahead of the shock wave, the pressure is 100 kPa and the temperature is 15 C. Find the velocity, pressure, and temperature of the air behind the shock wave.arrow_forwardA jet transport is flying at a standard altitude of 30,000 feet with a velocity of 550 miles per hour. What is the Mach number?arrow_forwardAir is flowing in a convergent nozzle. The stagnation pressure is 350 KPa , the stagnation temperature is 400 K and Mach number is 0.7. If the cross sectional area at this location is 10 × 10-3 m² , find 1) The static temperature and static pressure 2) The velocity at this location. 3) The area and static temperature and static pressure at the exit where Mach number =1arrow_forward
- An industrial nozzle as shown in Figure 3 works with an inert gas rush through the nozzle where the speed of the gas at the nozzle throat is recorded equal to the speed of sound. The inlet velocity, temperature, and pressure are 98 m/s, 600 oC, and 88 psi, respectively.Take the property of inert gas for k = 1.66 and cp = 5.192 kJ/kg. K. Calculate:a) Stagnation pressureb) Mach number at the nozzle inletc) Temperature of the gas at the throatd) Pressure of the gas at the throatarrow_forwardQ1. A turbojet aircraft flying at speed of 1645 km/h and altitude of 9000 m. A supersonic intake was designed with a conical center body, which generated one oblique shock wave and normal shock wave at the inlet of the aircraft. The deflection of the conical center body is a 10°. Calculate the fluid flow properties and velocity at the aircraft inlet of the intake. Assume a weak solution. Take: R= 0.287 kJ/kg K, and Cp=0.9 kJ/kg K.arrow_forwardGood day. Here is my question (10) 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.arrow_forward
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