FUND OF AERODYNAMICS(LLF) +CONNECT (1YR)
6th Edition
ISBN: 9781265141387
Author: Anderson
Publisher: MCG
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Chapter 10, Problem 10.14P
For supersonic and hypersonic wind tunnels, a diffuser efficiency,
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Turbomachines 4
A supersonic diffuser for air (y%3D1.4) has an arca ratio of 0.416 with an inlet Mach number
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Mach number (2.7) a normal shock occurs inside the diffuser. Determine the upstream
Mach number and area ratio at the section where the shock occurs and pressure ratio
across the diffuser.
A long pipe
Chapter 10 Solutions
FUND OF AERODYNAMICS(LLF) +CONNECT (1YR)
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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- 1:07 GAS CH 4 HW 1.pptx > 1. Air is flowing isentropically through a converging duct which is fed from a large reservoir where the temperature and 250 kPa, pressure are 350 K and respectively. At a certain point along the duct, where the cross-sectional area is 0.005 m, the pressure is 150 kPa. Determine the Mach number, temperature and velocity at that point and also calculate the mass flow rate. (Ans: 330.1 K, 0.549, 162.8 kPa, 1.718 kg-s"). 2. Air is supplied to a converging nozzle from a large reservoir where the temperature and pressure are 400 K and 100 kPa, respectively. At a certain cross-section, the temperature and pressure are measured to be 383.8 K and 63 kPa, respectively. Assuming isentropic flow, find the Mach number at this cross- section and the mass flow rate per unit area. (Ans: 0.46, 103.3 kg-s'-m²). 3. A converging nozzle is fed with air from a large reservoir where the temperature and pressure are 400 K and 170 kPa, respectively. The nozzle has an exit…arrow_forwardThe flow inside the intake of a scramjet is found to be supersonic with Mach number of 2.5. From the instrumentations located inside the intake section, the static pressure and temperature were measured as 1 atm and 290 K, respectively. Inside the intake, the flow encounters the first compression ramp at an angle of 22º. As an intake engineer, investigatethe flow properties (Mach number, static pressure and temperature, stagnation pressure and temperature) after the resulting first oblique shock wave, so that the 2ndcompression ramp can be designed according to the requirements. Use the relevant tables and charts to assist your investigation. Commentif more shock waves are required to ensure that subsonic combustion is possible.arrow_forwardInlet stagnation air pressure of a two-dimensional compressor cascade is 1.0 bar and inlet stagnation temperature of 295K. For an inlet Mach number of 0.75 and an inlet flow angle of 480, the exit flow angle is measured as 16.50. Assuming the flow is isentropic, determine the mass flow rate per unit frontal area, the exit Mach number and the static pressure ratio across the cascade.arrow_forward
- 3. Consider 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 2.8, and (b) M1 = 4.4. Compare these two results and comment on their implication.arrow_forwardA long pipe of 0.0254 m diameter has a mean coefficient of friction of 0.003. Air enters the pipe at a mach number of 2.5, stagnation temperature 310 K and static pressure 0.507 bar. Determine for a section at which the mach number reaches 1.2: i) Static pressure and temperature, ii) Stagnation pressure and temperature, iii) Velocity of air, iv) Distance of this section from the inlet and v) mass flow rate of air.arrow_forwardA supersonic nozzle is used for discharging air from a reservoir. The reservoir pressure and temperature are 0.5 MPa and 500 K. The design Mach number is 2. With back pressure, a normal shock appears at the exit of the nozzle. Determine the Mach number, pressure and temperature of air after the shock. [Ans. M2 = 0.577, p2 = 0.288 MPa, T, = 468.14 K] %3D ||arrow_forward
- A perfect gas (gamma = 1.9) enters a converging-diverging nozzle with a Mach number of 0.98 and local pressure and temperature values of 490 kPa and 570 K, respectively. The nozzle throat area is 6.5 * 10^-4 m^2 and the nozzle exit area is 26*10^-4 m^2. The nozzle exit pressure is 170 КРа. Verify that the back pressure is such that a shock occurs. a) What are the values of the mach number and stream temperature at the exit? b) At what area does the shock occur?arrow_forwardThe stagnation pressure and temperature at the entry of a nozzle are 5 bar and 500 K respectively. The exit Mach number is 2 where a normal shock wave occurs. Calculate the following quantities before and after the shock. Static and stagnation pressures and temperatures, air velocities and Mach numbers. What are the values of stagnation pressure loss and increase in entropy across the shock?arrow_forwardAir flows isentropically at a rate of 1.3 kg/s from a large chamber through a convergent- divergent duct and leave to the outlet at Mach number 2.72. The air velocity, pressure OUTM EXAMINATION SESSION 2020/2021 (a) Sketch the system and label all components with subsonic/supersonic and UTM N 2020/2021 TM FINAL EXAMIN 2020/2021 answer. 2020/2021 st TION STAL EXAMINATION SEMESTAR I, SESSJO , SESSIOVON 3, SERIONON Esto ON b/202 2020/202 and temperature at a location somewhere along the system were found to be 900 m/s, OUTM 150 kPa and 60°C, respectively. FINAL EXAMINATIC SEMESTER IL SESSION 2025/202 RATION 2020/2021 FINAL EXAMINA BEMENTER I SESSION 2020/202 diffuser/nozzle according to the effect of area change. Justify your ALE eSTER R, SESSION 202b/202 (b) Determine the pressure and temperature of the air in the large chamber, the area at throat, and the velocity at outlet. zb/202arrow_forward
- I3arrow_forward01: A convergent-divergent nozzle has an exit area to throat area ratio of 2. Air enters the nozzles with a stagnation pressure of 6.5 bar and a stagnation temperature of 93°C. The throat area is 6.25 cm. If there is a normal shock wave standing at a point where M= 1.5, determine the pressure, temperature on either side of the plane of shock and the mach number on the downstream side of the plane. Find also the exit mach number of the nozzle. M<1arrow_forwardThe pressure upstream of a normal shock wave is 1 atm. The pressure and temperature downstream of the wave are 10.33 atm and 1,390 °R, respectively. Calculate the Mach number and temperature upstream of the wave.arrow_forward
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