Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 10, Problem 10.16P
Return to Problem 9.19, where the average Mach number across the two-dimensional flow in a duct was calculated, and where
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Chapter 10 Solutions
Fundamentals of Aerodynamics
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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- 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_forwardb) 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_forwardAn impact tube is used to measure the Mach number of a certain airflowin a wind tunnel. The static pressure is 3 psia, and the impact pressure at theprobe is 116 kPa. What is the Mach number of the flow (k = 1.4 for air)? If thefree-stream air temperature is −40 ◦C, what is the flow velocity?arrow_forward
- Under 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_forwardQUESTION 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_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
- A 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_forward6.6. Air flows through a 0.25 m diameter duct. At the inlet the velocity is 300 m/s, and the stagnation temperature is 360 K. If the Mach number at the exit is 0.3, determine the direction and the rate of heat transfer. With the same conditions at the inlet, determine the amount of heat that must be transferred to the system if the flow is sonic at the exit.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
- Example 2. A high-speed AC 130 gunship is flying at a pressure altitude of 10 km. A Pitot tube on the wingtip measures a pressure of 4.24 x 10ª N/m2. Calculate the Mach number at which the aircraft is flying. Solution: Solving for P1 at an altitude of 10000 m, we get 2.65 x 104 N/m2 k-1 1.4-1 Po k 4.24 x 104) 1.4 M? k – 1 - 1 - 1 1.4 – 1 2.65 x 104 M? = 0.719 M1 = 0.848arrow_forward3. Assume the turbine as an incompressible flow machine. The average meridional speed is given as 150m/s. The flow coefficient is 0.5. Determine the blade speed that satisfies the above conditions.4. In a two-dimensional compressor cascade, inlet stagnation air pressure is 1.5 bar and inlet stagnation temperature is 330K. For an inlet Mach number of 0.75 and an inlet flow angle of 500, the exit flow angle is measured as 15.80. Determine the mass flow rate per unit frontal area, the exit Mach number and the static pressure ratio across the cascade. Assume the flow is isentropic.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_forward
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