In Problem 11.8, the critical Mach number for a circular cylinder is given as
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Fundamentals of Aerodynamics
- Consider a rocket engine burning hydrogen and oxygen; the combustion chamber temperature and pressure are 3517 K and 26 atm, respectively. The molecular weight of the chemically reacting gas in the combustion chamber is 16, and y = 1.22. The pressure at the exit of the convergent-divergent rocket nozzle is 1.174 x10-2 atm. The area of the throat is 0.4 m2. Assuming a calorically perfect gas and isentropic flow, calculate: (a) the exit Mach number, (b) the exit velocity, (c) the mass flow through the nozzle, and (d) the area of the exitarrow_forwardan airstream with a velocity of 650 m/s static pressure of 222 kpa and a static temperature of 452 k undergoes a normal shock find 1- mach number and the velocity after the normal shock wave 2-the static conditions after the normal shock wave 3-the stagnation conditions after the normal shock wave 4-the entropy change across the normal shock wavearrow_forwardFor an adiabatic flow of Argon (k=1.667) as a perfect gas, the maximum to critical velocity ratio will be ______.arrow_forward
- For a projectile traveling at 800 mph through air at 50 °F and standard atmospheric pressure, what is the value of the Mach number?arrow_forwardIn compressible flow, velocity measurements with a Pitot probe can be grossly in error if relations developed for incompressible flow are used. Therefore, it is essential that compressible flow relations be used when evaluating flow velocity from Pitot probe measurements. Consider supersonic flow of air through a channel. A probe inserted into the flow causes a shock wave to occur upstream of the probe, and it measures the stagnation pressure and temperature to be 620 kPa and 340 K, respectively. If the static pressure upstream is 110 kPa, determine the flow velocity.arrow_forwardGas flows through a converging-diverging duct. At point "A'' the cross-sectional area is 50 [cm2]and the Mach number was measured to be 0.4. At point B in the duct the cross-sectional area is 40[cm2]. Find the Mach number at point B. Assume that the flow is isentropic and the gas specificheat ratio is 1.4.arrow_forward
- An 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_forwardFor an ideal gas obtain an expression for the ratio of the speed of sound where Ma = 1 to the speed of sound based on the stagnation temperature, c*/c0.arrow_forwardMach 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 byarrow_forward
- Carbon dioxide enters an adiabatic nozzle at 1200 K with a velocity of 50 m/s and leaves at 400 K. Assuming constant specific heats at room temperature, determine the Mach number (a) at the inlet and (b) at the exit of the nozzle. Assess the accuracy of the constant specific heat approximation.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_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
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