Engineering

Engineering, 11.07.2019 02:40, cia196785920

Alarge air reservoir supplies (k = 1.4 and rair = 287 n m / kg. k) a convergent nozzle that discharges at atmospheric pressure (patm = 101 kpa). assuming that the flow is reversible and adiabatic inside the nozzle, determine: if the pressure in the tank is 600 kpa (abs) and the temperature is 600k, what is the mass flow through the nozzle if the output area is 1.29x10^-3 m²?

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Physics, 05.07.2019 03:20, 2308046

a convergent-divergent nozzle with an exit-to-throat area ratio of 1.616 has exit and reservoir pressures equal to 0.947 an 1.0 atm, respectively. in addition, the reservoir temperature is 288 k and the throat area is 0.3 m2. assuming isentropic flow through the nozzle, calculate the mach number, pressure, and mass flow rate at the throat. 4- (20 pts) consider the flow through a convergent-divergent duct with an exit-to-throat area ratio of 2. the reservoir pressure is 1 atm, and the exit pressure is 0.95 atm calculate the mach numbers at the throat and at the exit. 5 - (20 pts) consider a convergent-divergent nozzle with an exit-to-throat area ratio of 3. a normal shock wave is inside the divergent portion at a location where the local area ratio is a/a, = 2, calculate the exit-to-reservoir pressure ratio.

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Engineering, 20.08.2019 20:30, trvptrav

Aconverging nozzle has an exit area of 0.007 m^2 . nitrogen stored in a reservoir is to be discharged through the nozzle to an ambient pressure of 101.3 kpa. determine the flow rate through the nozzle for absolute reservoir pressures of 138 kpa, and 700 kpa. assume isentropic nozzle flow. in each case, determine the increase in mass flow to be gained by reducing the back pressure from 101.3 kpa to 0 kpa. reservoir temperature =289 k.

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Engineering, 26.09.2019 21:30, Galaxie85111

Space-shuttle main engine is designed with the given criteria x design altitude = 60,000 feet (18,300 m), where atmospheric pressure is, patm=7171.64 pa= 1.04016 psi. at this altitude the exhaust jet is fully expanded. x nozzle exit mach number = 5.05 (calculated) x throat area = 93 square inches (600 cm2) x nozzle exit area = 50.265 square feet (4.6698 m2 ) x chamber total pressure = 2,747 pounds per square inch (18,940 kpa) at 100% power assuming you can use the 2-d shock/expansion theory: calculate the static pressure at the nozzle exit plane and the flow-turning angle right after the flow exits the nozzle at: x sea level, where patm = 101325 pa =14.696 psi x b) 100,000 feet, where patm = 1090.16 pa = 0.158 psi

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Alarge air reservoir supplies (k = 1.4 and rair = 287 n m / kg. k) a convergent nozzle that discharg...

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