A supersonic converging-diverging nozzle has an exit-to-throat area ratio of 1.555. The flow at the outside of the nozzle is observed to make an 8° deflection inward. The back pressure is known to be 100 kPa. (a) Determine the Mach number and the static pressure at the exit of the nozzle. (b) Determine the angle of the first oblique shock and the Mach number downstream from the first oblique shock. (c) A second oblique shock turns the flow back in parallel with the jet axis; determine the angle of the second oblique shock. (d) Determine the Mach number and the static pressure in region C. A B C

A supersonic converging-diverging nozzle has an exit-to-throat area ratio of 1.555. The flow at the outside of the nozzle is observed to make an 8° deflection inward. The back pressure is known to be 100 kPa. (a) Determine the Mach number and the static pressure at the exit of the nozzle. (b) Determine the angle of the first oblique shock and the Mach number downstream from the first oblique shock. (c) A second oblique shock turns the flow back in parallel with the jet axis; determine the angle of the second oblique shock. (d) Determine the Mach number and the static pressure in region C. A B C

Elements Of Electromagnetics

7th Edition

ISBN:9780190698614

Author:Sadiku, Matthew N. O.

Publisher:Sadiku, Matthew N. O.

ChapterMA: Math Assessment

Section: Chapter Questions

Problem 1.1MA

See similar textbooks

Similar questions

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
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. Compare these results to those for helium undergoing a normal shock under the same conditions.
Nitrogen 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.
Consider air entering a heated duct at p1 = 1 atm and T1 = 288 K. Ignore the effect of friction. Calculate the amount of heat per unit mass (in joules per kilogram) necessary to choke the flow at the exit of the duct for an inlet Mach number of M1 = 2.2.
Consider a hypersonic vehicle with a spherical nose flying at Mach 20at a standard altitude of 150,000 ft, where the ambient temperature andpressure are 500◦R and 3.06 lb/ft2, respectively. At the point on thesurface of the nose located 20◦ away from the stagnation point, estimatethe: (a) pressure, (b) temperature, (c) Mach number, and (d) velocity ofthe flow.
Find the expression for the ratio of the stagnation pressure after a shock wave to the static pressure before the shock wave as a function of k and the Mach number upstream of the shock wave Ma1
A flow of air with Mach number M1 = 2, pressure p1 = 0.7 atm, and temperature 630 degR is turned away from itself through an angle of 26.38 deg. Determine the Mach number, the staticpressure, the static temperature, and the stagnation pressure after the turn (all pressures in atm).Also determine the Mach angles at the beginning and end of the expansion fan.
Mach number at a certain section in a diverging part of a c-d nozzle is 0.7. At farther section (towards the exit), Mach number reduces to 0.4. Hence, the cross-sectional area of the farther section is greater by
For a given Prandtl-Meyer expansion, the upstream Mach number is 3 and the pressure ratio across the wave is p,/p, = 0.4. Calculate the angles of the forward and rearward Mach lines of the expansion fan relative to the free-stream direction.
Air flows isentropically through a duct. At section 1, thepressure and temperature are 250 kPa and 125°C, andthe velocity is 200 m/s. At section 2, the area is 0.25 m2 andthe Mach number is 2.0. Determine (a) Ma1; (b) T2; (c) V2;and (d ) the mass fl ow.
Consider 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.
Consider 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.