Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 10, Problem 10.11P
The nozzle of a supersonic wind tunnel has an exit-to-throat area ratio of 6.79. When the tunnel is running, a Pitot tube mounted in the test section measures 1.448 atm. What is the reservoir pressure for the tunnel?
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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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- A perfect gas enters a frictionless, insulated passage with a supersonic speed. It is desired to increase the static pressure of the gas as it flow through the passage. The passage should bearrow_forwardIn 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.arrow_forwardAir flows at 200 m/s through the pipe. The temperature is 500 K and the absolute stagnation pressure is 200 kPa. Assume isentropic flow. For air R = 286.9 J/kg · K] and k = 1.40. (Figure 1) 0.3 m Determine the Mach number of the flow. Express your answer using three significant figures. M = Determine the mass flow. Express your answer using three significant figures. m = kg/sarrow_forward
- How do I find reservoir pressure given the exit-to-throat area ratio of a supersonic nozzle and a Pitot tube measurement from the test section.arrow_forward100% High pressure water and compressed air are delivered to some operating rooms. Bored one day, with no patients around, you take the cap off the nozzle to see what would happen. Water flows through the vertical pipe and nozzle and then exhausts into the air as shown in the diagram. The water pressure in the pipe at position 1 is 3 atm. The pressure when the water discharges to the air is 1 atm. Position 2 identifies the nozzle, 0.75 m above position 1, where the water flow has been constricted from an 8-cm diameter pipe to a 3-cm diameter nozzle. Assume that the free jet remains a stream of water with a 3-cm diameter. Assume no friction losses. Free jet -3 cm diameter 0.75 m a) Determine the mass flow rate of water through pipe at position 1 in kg/s. 3 atm 8 cm diameter b) Determine the height, h, the water jet can rise. c) Calculate the friction factor inside the 8-cm pipe and determine if the assumptions of negligible friction losses was reasonable. Water AR d 128 PM 10/3/2021…arrow_forwardThe velocity ratio (v1/v2) of an isentropic flow through a supersonic wind tunnel is 0.615. if the mach number at the tunnels entry section is 0.95 find the value of mach number at the exitarrow_forward
- A jet-propelled airliner flies at an altitude of 15,000 m and a velocity of 265 m/s. What is the stagnation pressure and stagnation temperature of the air entering the engine? If the airliner is stationary on the ground, the engines are running with the same Mach number flow into the engines, and the air flow to each engine is 35 kg/s, what is the airflow to each engine at the above flight conditions? Po = Pa To = i mf = K kg/sarrow_forwardA shock occurs in the nozzle when pres=p0,res = 3 atm and the Pitot tube measurement at the nozzle exit is P 0,Pitot at the exit = 1.085 atm. Calculate the location of the shock.arrow_forwardProblem 04.020 - Calculate pressure using a Pitot tube A Pitot tube is mounted in the test section of a low-speed subsonic wind tunnel. The flow in the test section has a velocity, static pressure, and temperature of 148 mi/h, 1 atm, and 68°F, respectively. Calculate the pressure measured by the Pitot tube. The pressure measured by the Pitot tube is | lb/ft2.arrow_forward
- Suppose 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_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_forwardWhat is the Mach of airflow with 150 m/s at a pressure of 0.95 atm and a Density of 1.15 kg/m^3?arrow_forward
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