Fluid Mechanics: Fundamentals and Applications
4th Edition
ISBN: 9781259696534
Author: Yunus A. Cengel Dr., John M. Cimbala
Publisher: McGraw-Hill Education
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Chapter 12, Problem 117P
A subsonic airplane is flying at a 5000-m altitude where the atmospheric conditions are 54 kPa and 256 K. A Pitot static probe measures the difference between the static and stagnation pressures to be 16 kPa. Calculate the speed of the airplane and the flight Mach number.
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A subsonic airplane is flying at a 5000-m altitude where the atmospheric conditions are 54 kPa and 256 K. A Pitot static probe measures the difference between the static and stagnation pressures to be 16 kPa. Calculate the speed of the airplane and the flight Mach number.
An ideal isentropic nozzle is attached to an infinite reservoir that has stagnation conditions 3 MPa and 2250 K, and a constant specific heat of 1.2. If the nozzle's static exit pressure is 38.871 kPa, what is the exit static temperature? Also determine the nozzle's exit Mach number, stagnation pressure, and stagnation temperature.
Air flowing steadily in a nozzle experiences a normal shock at a Mach number of Ma = 2.5. If the pressure and temperature of air are 10.0 psia and 440.5 R, 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.
Chapter 12 Solutions
Fluid Mechanics: Fundamentals and Applications
Ch. 12 - What is dynamic temperature?Ch. 12 - Calculate the stagnation temperature and pressure...Ch. 12 - Prob. 6PCh. 12 - Prob. 7PCh. 12 - Prob. 8EPCh. 12 - Prob. 9PCh. 12 - Products of combustion enter a gas turbine with a...Ch. 12 - Is it possible to accelerate a gas to a supersonic...Ch. 12 - Prob. 72EPCh. 12 - Prob. 73P
Ch. 12 - Prob. 74PCh. 12 - Prob. 75PCh. 12 - For an ideal gas flowing through a normal shock,...Ch. 12 - Prob. 77CPCh. 12 - On a T-s diagram of Raleigh flow, what do the...Ch. 12 - What is the effect of heat gain and heat toss on...Ch. 12 - Prob. 80CPCh. 12 - Prob. 81CPCh. 12 - Prob. 82CPCh. 12 - Argon gas enters a constant cross-sectional area...Ch. 12 - Prob. 84EPCh. 12 - Prob. 85PCh. 12 - Prob. 86PCh. 12 - Prob. 87EPCh. 12 - Prob. 88PCh. 12 - Prob. 89PCh. 12 - Prob. 90PCh. 12 - Prob. 91PCh. 12 - Prob. 93CPCh. 12 - Prob. 94CPCh. 12 - Prob. 95CPCh. 12 - Prob. 96CPCh. 12 - Prob. 97CPCh. 12 - Prob. 98CPCh. 12 - Prob. 99CPCh. 12 - Prob. 100CPCh. 12 - Prob. 101PCh. 12 - Air enters a 5-cm-diameter, 4-m-long adiabatic...Ch. 12 - Helium gas with k=1.667 enters a 6-in-diameter...Ch. 12 - Air enters a 12-cm-diameter adiabatic duct at...Ch. 12 - Prob. 105PCh. 12 - Air flows through a 6-in-diameter, 50-ft-long...Ch. 12 - Air in a room at T0=300k and P0=100kPa is drawn...Ch. 12 - Prob. 110PCh. 12 - Prob. 112PCh. 12 - Prob. 113PCh. 12 - Prob. 114PCh. 12 - Prob. 115PCh. 12 - Prob. 116EPCh. 12 - A subsonic airplane is flying at a 5000-m altitude...Ch. 12 - Prob. 118PCh. 12 - Prob. 119PCh. 12 - Prob. 120PCh. 12 - Prob. 121PCh. 12 - Prob. 122PCh. 12 - Prob. 123PCh. 12 - An aircraft flies with a Mach number Ma1=0.9 at an...Ch. 12 - Prob. 125PCh. 12 - Helium expands in a nozzle from 220 psia, 740 R,...Ch. 12 - Prob. 127PCh. 12 - Prob. 128PCh. 12 - Prob. 129PCh. 12 - Prob. 130PCh. 12 - Prob. 131PCh. 12 - Prob. 132PCh. 12 - Prob. 133PCh. 12 - Prob. 134PCh. 12 - Prob. 135PCh. 12 - Prob. 136PCh. 12 - Prob. 137PCh. 12 - Prob. 138PCh. 12 - Air is cooled as it flows through a 30-cm-diameter...Ch. 12 - Prob. 140PCh. 12 - Prob. 141PCh. 12 - Prob. 142PCh. 12 - Prob. 145PCh. 12 - Prob. 148PCh. 12 - Prob. 149PCh. 12 - Prob. 150PCh. 12 - Prob. 151PCh. 12 - Prob. 153PCh. 12 - Prob. 154PCh. 12 - Prob. 155PCh. 12 - Prob. 156PCh. 12 - Prob. 157PCh. 12 - Prob. 158PCh. 12 - Prob. 159PCh. 12 - Prob. 160PCh. 12 - Prob. 161PCh. 12 - Prob. 162PCh. 12 - Assuming you have a thermometer and a device to...
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- 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.arrow_forwardHow does the Mach number affect the behavior of compressible flow in a supersonic nozzle?arrow_forwardIs it possible to accelerate a fluid to supersonic velocities with a velocity other than the sonic velocity at the throat? Explainarrow_forward
- 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.arrow_forwardIs it possible to accelerate a gas to a supersonic velocity in a converging nozzle? Explainarrow_forwardConsider subsonic Fanno flow of air with an inlet Mach number of 0.70. If the Mach number increases to 0.90 at the duct exit as a result of friction, will the (a) stagnation temperature T0, (b) stagnation pressure P0, and (c) entropy s of the fluid increase, decrease, or remain constant during this process?arrow_forward
- 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.arrow_forwardNitrogen enters a duct with varying flow area at 400 K, 100 kPa, and a Mach number of 0.5. Assuming a steady, isentropic flow, determine the temperature, pressure, and Mach number at a location where the flow area has been reduced by 20 percent.arrow_forwardAir flowing at 8 psia, 480 R, and Ma1 = 2.0 is forced to undergo a compression turn of 15°. Determine the Mach number, pressure, and temperature of the air after the compression.arrow_forward
- vortex generators on the upper surface of a wing will a. decrease the spanwise flow at high Mach numbers b. increase the critical Mach number c. decrease the intensity of the shockwave effects d. increase intensity of the shockwave effectsarrow_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_forwardNitrogen enters a converging–diverging nozzle at 800 kPa and 400 K with a negligible velocity. Determine the critical velocity, pressure, temperature, and density in the nozzle. Use data from the tables. The properties of nitrogen are k = 1.4 and R = 0.2968 kJ/kg·Karrow_forward
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