Fundamentals of Aerodynamics
6th Edition
ISBN: 9781259129919
Author: John D. Anderson Jr.
Publisher: McGraw-Hill Education
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Chapter 8, Problem 8.10P
The how just upstream of a normal shock wave is given by
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The flow just upstream of a normal shock wave is given by p1 = 1 atm,T1 = 288 K, and M1 = 2.6. Calculate the following properties justdownstream of the shock: p2, T2, ρ2, M2, p0,2, T0,2, and the change inentropy across the shock.
The entropy change across a normal shock wave (Tox= 505 K, vx= 490 ms-1)
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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- d*, 2. Consider a normal shock wave in air where the upstream flow properties are u1 = 690 m/s, T1 = 288 K, and p1=1 atm. Calculate the velocity, temperature, and pressure downstream of the shock.arrow_forwardThe flow just upstream of a normal shock wave is given byp1 = 1800 lb/ft2, T1 = 480 ◦R, and M1 = 3.1. Calculate the velocity andM∗ behind the shock.arrow_forwardPlease answer fastarrow_forward
- The driver and driven gases in a shock tube are both air at 300 K. Assume p4/p1 = 8.006. After the diaphragm is broken, an incident shock wave is propagating into the driven section and an incident expansion wave is propagating into the driver section. If the strength of the incident shock (p2/p1) is 2.6, a. Calculate the strength of the incident expansion wave (p3/p4). b. Calculate the induced mass motion velocity behind the expansion wave. c. Find the propagation velocity of the head expansion wave. d. Find the propagation velocity of the tail expansion wave.arrow_forward2arrow_forwardConsider an oblique shock wave with a wave angle of 36.87◦. Theupstream flow is given by M1 = 3 and p1 = 1 atm. Calculate the totalpressure behind the shock usinga. p0,2/p0,1 b. p0,2/p1 Compare the results.arrow_forward
- The pressure ratio across a normal shock wave that occurs in air is 1.25. Ahead of the shock wave, the pressure is 100 kPa and the temperature is 15 C. Find the velocity, pressure, and temperature of the air behind the shock wave.arrow_forwardA normal shock wave propagates into atmosphere where the atmospheric temperature is equal 336K and pressure is 0.8 atm. For a given pressure ratio of p2/p₁ 41 to T₁ = a) Assuming stationary atmosphere, calculate the shock wave velocity and the velocity induced behind the shock wave. Also calculate the temperature, total pressure and total temperature of the fluid particles behind the shock wave using equations for moving normal shock. =arrow_forwardThe velocity of flow after a normal shock wave when (Tx= 315 K, vx= 555 ms-1)arrow_forward
- 5. Airflow at Mach 2 passes through an oblique shock as shown and deflects 10°. A second oblique shock reflects from the solid wall. What is the pressure ratio (p3/p₁) across the two-shock system? Assume that there is no boundary layer near the wall, so the flow is uniform in each of the regions bounded by the shocks and that y = 1.4. Reflected oblique shock 10° M=2arrow_forwardQ1. A turbojet aircraft flying at speed of 1645 km/h and altitude of 9000 m. A supersonic intake was designed with a conical center body, which generated one oblique shock wave and normal shock wave at the inlet of the aircraft. The deflection of the conical center body is a 10°. Calculate the fluid flow properties and velocity at the aircraft inlet of the intake. Assume a weak solution. Take: R= 0.287 kJ/kg K, and Cp=0.9 kJ/kg K.arrow_forward1. Calculate the density of air entering the nozzle, if the air leaves a nozzle at a P=1.5atm and T=600K at a velocity of 400m/s. The air enters at 150 m/s and the inlet to exit diameter ratio is 3:1.5.arrow_forward
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