: +♡العنوانX-Ra>۱/۱Air flows through an isentropic convergence-divergence nozzlewith stagnation pressure and temperature of 250 kPa and 320K.respectively. The throat area is 0.1 m² and the exit area is 0.4 m²(b) Is the.The velocity v-180-65.5A-500A2: (a) Is dA/dVthroat chocked? Explain by sketch (c) Mach number pressureand temperature at the exit section.Q=5+Q=3+R9TΤM

The problem involves isentropic flow through a converging-diverging nozzle. We are given the…

Answered: : +♡ العنوان X-Ra > ۱/۱ Air flows…

Elements Of Electromagnetics

7th Edition

ISBN:9780190698614

Author:Sadiku, Matthew N. O.

Publisher:Sadiku, Matthew N. O.

ChapterMA: Math Assessment

Section: Chapter Questions

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A large tank at 500 K and 165 kPa feeds air to a convergingnozzle. The back pressure outside the nozzle exit is sealevelstandard. What is the appropriate exit diameter if thedesired mass fl ow is 72 kg/h?
A convergent divergent diffuser is exposed to air at pressure of 80kPa and temperature of 350K and Mach number Mi=2.5 as shown in fig. The inlet area Ai = 20 cm2, throat area = 10 cm2, and exit area = 25 cm2. 1- If the back pressure is 650kPa the shock wave is take place at the divergent part find: Throat Mach number , mass flow rate, Shock wave location ,exit Mach number. 2- If the incoming Mach number is decrease to 1.8 State the flow condition through the duct and find, the mass flow rate , exit pressure and Mach number M=2.5, P=80, T=350 Flow condition through the duct for CASE 1)
-Converging nozzle is supplied with Po=10 bars and To=D1000 K and discharges to atmosphere with pressure of 9.8 bars and the nozzle exit area is 10 cm. The Mach number at nozzle exit will be: Select one: O a. M-0.46 O b. M=0.55 Oc. M-0.17 O d. M-0.27
A compressed Air enters a convergent-divergent nozzle, at 4.0 atm, 300 K and at Mach 2.0. The nozzle is fitted at the exit of a 0.3 m constant diameter pipe and 2.0 m long that supplies the air. Assuming that the friction factor for the pipe is 0.003, and that a normal shock occurs in the pipe at Mach number 1.8, a) Find the length of the pipe between the location of the normal shock and the inlet of the nozzle. b) Claculate the throat and the exit plane diameters of the nozzle if the mach number at the exit plane is 2.0 c) Calculate the exit pressure. d) On a graph, skecth the variation of static and stagnation pressure along the pipe and the nozzle.
Air flows steadily from a tank through the pipe inFig. . There is a converging nozzle on the end. Ifthe mass flow is 3 kg/s and the nozzle is choked, estimate(a) the Mach number at section 1 and (b) the pressureinside the tank.
A converging–diverging nozzle exits smoothly to sealevelstandard atmosphere. It is supplied by a 40-m3 tankinitially at 800 kPa and 100 °C. Assuming isentropicflow in the nozzle, estimate (a) the throat area and(b) the tank pressure after 10 s of operation. The exitarea is 10 cm2.
a) A convergent nozzle with a given area ratio is supplied with air at a constant stagnation pressure Poand a constant stagnation temperature To. With the aid of a sketch plot the nozzle mass flow rate as a function of pressure ratio () across P the nozzle. In your answer indicate the critical value of the pressure ratio. For air you may assume y = 1.4 and R = 287J/kg K. b) The engine of a supersonic aircraft is equipped with a convergent-divergent nozzle. A A, A -4.0).Determine the design subsonic and supersonic Mach numbers for this area ratio. You may assume the combustion products have the same properties as air and the expansion process through the nozzle is isentropic. The nozzle's area ratio is ( c) Determine at what pressure ratios ( ) a normal shock wave appears in the P divergent part of the nozzle at a location where -3.0) Aheck A₁
Air flows at the rate of 0.8 kg/s through a convergent-divergent nozzle. The entrance area is 0.004 m? and the inlet temperature and pressure are 300 K and 400 kPa. If the exit pressure is 140 kPa and the expansion is isentropic, find: The velocity at entrance, The stagnation temperature and stagnation pressure, The throat and exit areas, The exit velocity.
Thermodynamics/fluid mechanics question is in the image.
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Qs: A two-dimensional supersonic inlet is to be designed to operate at Mach 2.4. Deceleration is to occur through a series of oblique shocks followed by a normal shock, as shown in Figure. Determine the loss of stagnation pressure for the cases of two, three, and four oblique shocks. Assume the wedge turning angles are each 6°. M. = 2.4 Vi, Pai P.- 50 kPa 12° Vi, Pai T.-0° C

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