Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 8, Problem 8.8P
The pressure upstream of a normal shock wave is 1 atm. The pressure and temperature downstream of the wave are 10.33 atm and
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Consider 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 = 2.5, and (b) M1 = 4.5. Compare these two results and comment on their implication
The pressure upstream of a normal shock wave is 1 atm. The pressureand temperature downstream of the wave are 10.33 atm and 1390 ◦R,respectively. Calculate the Mach number and temperature upstream of thewave and the total temperature and total pressure downstream of the wave.
4. Determine the upstream Mach number,
considering an oblique shock wave with ew =
32° and a pressure ratio, P2/P1 = 3.0.
Chapter 8 Solutions
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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- The entropy increase across a normal shock wave is 199.5 J/(kg · K). What is the upstream Mach number?PLease show step by step solns for better understanding thank you!arrow_forwardCan the Mach number of a fluid be greater than 1 after a normal shock wave? Explain.arrow_forwardThe entropy increase across a normal shock wave is 199.5 J/(kg · K). Whatis the upstream Mach number?arrow_forward
- The 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_forwardA 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_forwardA 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.arrow_forward
- 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_forwardASAParrow_forwardAhead of the normal shock wave, the upstream pressure , temperature, and Mach number are 0.53 atm, 255 K, and 2.8, respectively. Determine the pressure downstream of the shock wave.arrow_forward
- In 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_forwardThe pressure ratio across a normal shock wave in air is 4.5. What are the Mach numbers in front of and behind the wave? What are the density and temperature ratios across the wave?arrow_forwardQ.1. A uniform supersonic air flow travelling at a Mach number of 2.5 passes over a wedge which deflects the flow by 15°, as shown in the figure. If the pressure and temperature in the uniform flow are 40 kPa and 263 K, respectively, determine the shock wave angle, Mach number, pressure and temperature downstream of the oblique shock wave. Oblique Shock Wave M = 2.5 15° T= 263 K P = 40 KPaarrow_forward
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