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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An air tank with a nozzle, has a pressure of twice the standard sea level pressure and
density of 2.45 kg/m³. Outside the converging-diverging nozzle, the pressure corresponds
to an altitude of 3km and designed to have a mach number of 1.0 and 1.8 at the throat
and exit, respectively. The area at the throat is 0.15 square meters.
Calculate:
a) the temperature and speed of sound in the tank.
b) pressure and density at the throat
c) mass flow rate at the exit.
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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- The Mach no. of a fixed geometry inlet at overspeeded for starting mode (Ai= 0.13 m2, At= 0.095 m2arrow_forward(please solve within 20 minutes i will give multiple thumbs up)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_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_forward43.4.5. If y = 1.2 and the fluid is a perfect gas, what Mach number will give a temperature ratio of T/T= 0.909? What will the ratio of p/p.be for this flow?arrow_forward
- Please helparrow_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_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_forward
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