Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 10, Problem 10.7P
A convergent-divergent nozzle with an exit-to-throat area ratio of 1.616 has exit and reservoir pressures equal to 0.947 and 1.0 atm, respectively. Assuming isentropic flow through the nozzle, calculate the Mach number and pressure at the throat.
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A convergent-divergent nozzle with an exit-to-throat area ratio of 1.616has exit and reservoir pressures equal to 0.947 and 1.0 atm, respectively.Assuming isentropic flow through the nozzle, calculate the Mach numberand pressure at the throat.
D5.
A converging-diverging nozzle operates with air and subsonic flow at
the inlet and supersonic flow at the outlet with no shacks inside the
nozzle. Determine the mach number and static pressure at the
outlet if the area of the outlet is 0.900 m2, the area of the throat is
0.490m2, and the stagnation pressure is 215 kPa.
Air flowing steadily in a nozzle experiences a normal shock at a Mach number of Ma = 2.6. If the pressure and temperature of air are 58 kPa and 270 K, respectively, upstream of the shock, calculate the pressure, temperature velocity, Mach number, and stagnation pressure downstream of the shock. Calculate the entropy changes of air and helium across the normal shock wave
Chapter 10 Solutions
Fundamentals of Aerodynamics
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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
- The discharge air from a convergent-divergent nozzle at a Mach number of 1.72. The exit pressure is 15 psia, and the temperature is 155 degree F. For isentropic expansion and a negligible inlet velocity, calculate the stagnation pressure, psia.A. 72.31 B. 74.31 C. 76.31 D. 78.31arrow_forwardA convergent-divergent nozzle with an exit-to-throat area ratio of 1.616has exit and reservoir pressures equal to 0.947 and 1.0 atm, respectively.Assuming isentropic flow through the nozzle,calculate the mass flow through the nozzle, assuming that the reservoir temperature is 288 K and the throat area is 0.3 m2.arrow_forwardAir enters a converging–diverging nozzle at 1.2 MPa with a negligible velocity. Approximating the flow as isentropic, determine the back pressure that would result in an exit Mach number of 1.8.arrow_forward
- Consider the isentropic supersonic flow through a convergent-divergent nozzle with an exit-to-throat area ratio of 10.25. The reservoir pressure and temperature are 5 atm and 600◦R, respectively. Calculate M, p, and T at the nozzle exit.arrow_forwardConsider a normal shock wave in a supersonic airstream where the pressure upstream of the shock is 1 atm. Calculate the loss of total pressure across the shock wave when the upstream Mach number is (a) M1 = 2, and (b) M1 = 4. Compare these two results and comment on their implication.arrow_forwardConsider the isentropic flow through a convergent-divergent nozzle with an exit-to-throat area ratio of 2.5. The reservoir pressure and temperature are 1 atm and 288 K, respectively. Calculate the Mach number, pressure, and temperature at both the throat and the exit for the cases where (a) the flow is supersonic at the exit and (b) the flow is subsonic throughout the entire nozzle except at the throat, where M = 1.arrow_forward
- Air enters a converging–diverging nozzle at a pressure of 1200 kPa with negligible velocity. What is the lowest pressure that can be obtained at the throat of the nozzle?arrow_forwardNitrogen enters a converging–diverging nozzle at 620 kPa and 310 K with a negligible velocity, and it experiences a normal shock at a location where the Mach number is Ma = 3.0. Calculate the pressure, temperature, velocity, Mach number, and stagnation pressure downstream of the shock. Compare these results to those of air undergoing a normal shock at the same conditions.arrow_forwardConsider the isentropic flow through a convergent-divergent nozzle with an exit-to-throat area ratio of 2. The reservoir pressure and temperature are 1 atm and 288 K, respectively.Calculate the Mach number, pressure, and temperature at both the throat and the exit for the cases where (a) the flow is supersonic at the exit and (b) the flow is subsonic throughout the entire nozzle except at the throat, where M = 1.arrow_forward
- Consider the isentropic flow through a supersonic nozzle. If thetest-section conditions are given by p = 1 atm, T = 230 K, and M = 2,calculate the reservoir pressure and temperature.arrow_forwardAir flowing steadily in a nozzle experiences a normal shock at a Mach number of Ma = 2.6. If the pressure and temperature of air are 58 kPa and 270 K, respectively, upstream of the shock, calculate the pressure, temperature, velocity, Mach number, and stagnation pressure downstream of the shock. Compare these results to those for helium undergoing a normal shock under the same conditions.arrow_forwardAn ideal gas flows through a passage that first converges and then diverges during an adiabatic, reversible, steady-flow process. For supersonic flow at the inlet, sketch the variation of pressure, velocity, and Mach number along the length of the nozzle when the Mach number at the minimum flow area is equal to unity.arrow_forward
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