FUND OF AERODYNAMICS(LLF) +CONNECT (1YR)
6th Edition
ISBN: 9781265141387
Author: Anderson
Publisher: MCG
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Chapter 10, Problem 10.15P
Return to Problem 9.18. where the average Mach number across the two-dimensional flow in a duct was calculated, and where
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Chapter 10 Solutions
FUND OF AERODYNAMICS(LLF) +CONNECT (1YR)
Ch. 10 - The reservoir pressure and temperature for a...Ch. 10 - A flow is isentropically expanded to supersonic...Ch. 10 - A Pitot tube inserted at the exit of a supersonic...Ch. 10 - For the nozzle flow given in Problem 10.1, the...Ch. 10 - A closed-form expression for the mass flow through...Ch. 10 - Prob. 10.6PCh. 10 - A convergent-divergent nozzle with an...Ch. 10 - For the flow in Problem 10.7, calculate the mass...Ch. 10 - Consider a convergent-divergent nozzle with an...Ch. 10 - A 20 half-angle wedge is mounted at 0 angle of...
Ch. 10 - The nozzle of a supersonic wind tunnel has an...Ch. 10 - We wish to design a supersonic wind tunnel that...Ch. 10 - Consider a rocket engine burning hydrogen and...Ch. 10 - For supersonic and hypersonic wind tunnels, a...Ch. 10 - Return to Problem 9.18. where the average Mach...Ch. 10 - Return to Problem 9.19, where the average Mach...Ch. 10 - A horizontal flow initially at Mach I flows over a...Ch. 10 - Consider a centered expansion wave where M1=1.0...
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- b) In a converging-diverging duct the flow conditions are such that a normal shock is formed at the exit of the diverging section. Upstream of the shock the flow conditions are as below. Mach number= '5.198', Pressure= '311886.9' Pa and Temperature= '228.72' K. Obtain the Mach number, pressure, and temperature at the downstream of the shock. Assume standard air properties (y=1.4).arrow_forwardUnder low-speed incompressible flow conditions, the pressure coefficientat a given point on an airfoil is −0.54. Calculate Cp at this point when thefreestream Mach number is 0.58, usinga. The Prandtl-Glauert ruleb. The Karman-Tsien rulec. Laitone’s rulearrow_forwardCalculate the Mach number of the exit if the diverging duct has an inlet area and velocity of air of 0.5m^2 and 60 m/s respectively. The exit area is 8 times the inlet, and the temperature on the exit is standard atmospheric value. (flow is incompressible)arrow_forward
- For Air Assume: y =1.4 and R = 287 J/kg K Question B1 a) Air is flowing at a Mach number of 0.6 in a two-dimensional duct at a location where the area is of 0.75 m². At this location the static pressure is 50 kPa and the static temperature is 300 K. 1- Calculate the mass flow rate through the duct; [2 marks] 2- What percentage in area change would be necessary to reach a Mach number of 0.75? [2 marks] 3- What percentage in area change would be necessary to reach a Mach number of 1.00? [2 marks] b) A normal shock wave occurs in a gas with an unknown specific heat ratio. The static pressure ratio across the normal shock wave is 10.6. The Mach number downstream of the shock wave is equal to 0.495, find the specific heat ratio of the gas and the velocity ratio across the shock wave. [8 marks] c) Explain why successive infinitesimal compression waves tend to reinforce and form a shock wave. Use sketches to illustrate your answer. [5 marks] d) Show that for a choked convergent nozzle further…arrow_forwardA Pitot tube is inserted into an airflow where the static pressure is 1 atm. Calculate the flow Mach number when the Pitot tube measures (a) 1.276 atm, (b) 2.714 atm, (c) 12.06 atm.arrow_forward02: The large compressed-air tank shown in the figure bellow exhausts from a nozzle at an exit velocity of V= 235 m/s. Assuming isentropic flow, compute: a) the exit Mach number. b) the pressure in the tank. (take; k = 1.396, C, = 917 J/Kg.k and R = 260 J/Kg.k) air at 30 "C tank conditions remain constant P- 101 kPaarrow_forward
- QUESTION 2 A compressor blade design tested in a cascade is found to choke with an inlet Mach number of 0.9 when the inlet flow angle is 52 deg. If the ratio of the throat area to the frontal area, A*/H1s, for the cascade is).625, calculate the loss of stagnation pressure between the far upstream and the throat and express this as a loss coefficient. Give your comments on what could cause this loss.arrow_forwardThe large compressed - air tank shown in the figure bellow exhausts from a nozzle at an exit velocity of Ve = 235 m / s. Assuming isentropic flow, compute: a) the exit Mach number. b) the pressure in the tank. (take; k = 1.396, Cp = 917 J / Kg.k and R = 260 J / Kg.k) air at 30 ° C tank conditions remain constant Pate = 101 kPaarrow_forwardA nozzle for a supersonic wind tunnel is designed to achieve a Mach number of 3.2, with a velocity of 2500 m/s, and a density of 1.0 kg/ m³ in the test section. Find the temperature and pressure in the test section and the upstream stagnation conditions. The fluid is helium. Te = i Pe= To= Po= Mi K kPa K kPaarrow_forward
- If the Mach number behind a normal shock is equal to 0.4, calculate the ratio of the readings of pitot tubes located before and after the shock.arrow_forwardUsing the Image below solve the following two questions, AND ONLY SOLVE FOR i and iii a.Determine The total pressure loss in the final state, after the 20 degree turn, for i and iii. (I'd also like to confirm for ii that the total pressure loss is 0?) b.Determine the Mach number in the final state. For which geometry is this highest? ( For ii it is 2.12)arrow_forward5. The upstream Mach number for a given Prandtl-Meyer expansion is 3.2 and the pressure ratio throughout the wave is P2/P1 = 0.4. Calculate the angles of the expansion fan's forward and rearward Mach lines in relation to the freestream direction.arrow_forward
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