Loose Leaf for Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259683992
Author: Anderson, John
Publisher: McGraw-Hill Education
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Textbook Question
Chapter 8, Problem 8.5P
Consider the isentropic flow through a supersonic nozzle. If the test-section conditions are given by
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Consider the isentropic flow through a supersonic nozzle. If thetest-section conditions are given by p = 1 atm, T = 230 K, and M = 2,calculate the reservoir pressure and temperature.
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Chapter 8 Solutions
Loose Leaf for Fundamentals of Aerodynamics
Ch. 8 - Consider air at a temperature of 230 K. Calculate...Ch. 8 - The temperature in the reservoir of a supersonic...Ch. 8 - At a given point in a flow, T=300K,p=1.2atm, and...Ch. 8 - At a given point in a flow, T=700R,p=1.6atm, and...Ch. 8 - Consider the isentropic flow through a supersonic...Ch. 8 - Consider the isentropic flow over an airfoil. The...Ch. 8 - The flow just upstream of a normal shock wave is...Ch. 8 - The pressure upstream of a normal shock wave is 1...Ch. 8 - The entropy increase across a normal shock wave is...Ch. 8 - The how just upstream of a normal shock wave is...
Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Consider a flow with a pressure and temperature of...Ch. 8 - Repeat Problems 8.11 and 8.12 using (incorrectly)...Ch. 8 - Derive the Rayleigh Pitot tube formula, Equation...Ch. 8 - On March 16, 1990, an Air Force SR-71 set a new...Ch. 8 - In the test section of a supersonic wind tunnel, a...Ch. 8 - When the Apollo command module returned to earth...Ch. 8 - The stagnation temperature on the Apollo vehicle...Ch. 8 - Prove that the total pressure is constant...
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- A flow of air with Mach number M1 = 2, pressure p1 = 0.7 atm, and temperature 630 degR is turned away from itself through an angle of 26.38 deg. Determine the Mach number, the staticpressure, the static temperature, and the stagnation pressure after the turn (all pressures in atm).Also determine the Mach angles at the beginning and end of the expansion fan.arrow_forwardAir flows through a long, isentropic nozzle. The temp. and pressure at the reservoir are 1000 K and 20 atm respectively. If the Mach number at the entrance is 0.2, what is the gas velocity at the entrancearrow_forwardIn a wind tunnel air enters with a velocity of 200kmph. The static pressure and temperature of the air at the inlet of the tunnel is 110kPa and 27°C respectively. Determine Mach number, stagnation temperature, stagnation pressure and the stagnation density on a test model installed in the wind tunnel. The size of the tunnel is given as 1m x1m square cross-section. Determine the mass flow rate of the air. For air assume R=287J/kgK ; γ=1.4.arrow_forward
- Q3/( ) Air is flowing isentropically through a converging duct which is fed from a large reservoir where the temperature and pressure are 300K and 400kpa, respectively. The maximum mass flow rate through the duct is 4kg/s. Determine the Mach number, pressure, temperature and velocity at a point where the cross-sectional area is 0.01m².arrow_forwardAt some section in the convergent-divergent nozzle, in which air is flowing, pressure, velocity, temperature and cross-sectional area are 200 kN/m², 170 m/s, 200°C and 1000 mm² respectively. If the flow conditions are isentropic, determine: (i) Stagnation temperature and stagnation pressure, (ii) Sonic velocity and Mach number at this section, (iii) Velocity, Mach number and flow area at outlet section where pressure is 110 kN/m², (iv) Pressure and temperature at throat of the nozzle. Take for air: R=287 J/kg K, cp = 1.0 kJ/kg K and k = 1.4arrow_forwardAir is expanded from 200 kPa and 500°K through a throat to an exit Mach number of 2.5. If the desired mass flow is 3 kg/s, calculate the throat area, exit pressure, exit temperature, exit velocity, and the discharge area assuming isentropic flow.arrow_forward
- Q1: Air enters a converging-diverging nozzle of a supersonic wind tunnel at 1 MPa and 300 K with a low velocity. If a normal shock wave occurs at the exit plane of the nozzle at Ma = 2.4, determine the pressure, temperature, Mach number, velocity, and stagnation pressure after the shock wave.arrow_forwardThe ratio of stagnation temperature at the exit and entry of a combustion chamber is 3.75. If the pressure, temperature and flow Mach number at the exit are 2.5 bar, 1000°C and 0.9 respectively, determine (i) Mach number, pressure and temperature of the gas at entry, (ii) total heat supplied per kg of gas, and (iii) the maximum heat that can be supplied. Take y= 1.4 and C, = 1.2 kJ/kg K. [Ans. M1 = 0.255, p1 1.9 bar, T, = 391.4 K, Q = 1301.7 kJ/kg, Qmax 1315.82 kJ/kg]arrow_forwardAir flows through a long, isentropic nozzle. The temperature and pressure at the * reservoir are 1000K and 20 atm, respectively. If the Mach number at the entrance is 0.2, determine the gas velocity at the entrance. 634 m/s 127 m/s 478 m/s 254 m/s For a large centrifugal pump, the required net positive suction head is typically around 5 ft 2 ft 15 ftarrow_forward
- The velocity ratio (v1/v2) of an isentropic flow through a supersonic wind tunnel is 0.615. if the mach number at the tunnels entry section is 0.95 find the value of mach number at the exitarrow_forwardConsider the isentropic flow through a convergent-divergent nozzle with an exit-to-throat area ratio of 2.5. The reservoir pressure and temperature are 1 atm and 288 K, respectively. Calculate the Mach number, pressure, and temperature at both the throat and the exit for the cases where (a) the flow is supersonic at the exit and (b) the flow is subsonic throughout the entire nozzle except at the throat, where M = 1.arrow_forwardHeat is rejected from the air flowing in a constant area duct. At the duct entrance, the air is moving at 200 m/s and possesses static conditions of 300 K and 100 kPa. If 50000 J/kg is rejected along the duct, find the exit Mach number, stagnation temperature change and the stagnation pressure change. [Ans. M2 = 0.48, To1 – To2 = 49.75 K, p02 – poi 13.02 kPa]arrow_forward
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