Problem 3. A Pitot tube is placed in a supersonic wind tunnel where the temperature is T = 300 K and theMach number is Max = 2.0. A normal shock stands in front of the probe. Behind the shock, the pressure jumpsto py = 40 kPa. (a) Find the pressure, velocity, and density upstream of the shock. (b) Find the stagnationtemperature and stagnation pressure before and after the shock. (c) Show static and stagnation points andprocess path on a T-s diagram and briefly describe the process.

Answered: Problem 3. A Pitot tube is placed in a…

Elements Of Electromagnetics

7th Edition

ISBN:9780190698614

Author:Sadiku, Matthew N. O.

Publisher:Sadiku, Matthew N. O.

ChapterMA: Math Assessment

Section: Chapter Questions

Problem 1.1MA

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Q.5 A supersonic wind tunnel, which is shown in the figure, is fed from a large tank. The air is discharged to the atmosphere. The Mach number in the test section is 2.65 and a normal shock wave stands at the exit of the wind tunnel. The pressure at section 3 immediately after the normal shock wave is 100 kPa. The flow is isentropic except through the normal shock wave. Find: 1. Po, P1, P2 and M3. 2. If a Pitot tube is placed in the exit jet, then calculate the pressure measured by the Pitot tube. Po Test section Normal shock wave Exit Jet Pitot tube
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.
i need the answer quickly
Q2) Uniform airflow at Mach 3 passes into a concave corner of angle 15°, as shown in Figure. The pressure and temperature in the supersonic flow are, respectively, 72 kPa and 290K. Determine the tangential and normal components of velocity and Mach number upstream and downstream of the wave for the weak shock solution. Also find the static and stagnation pressure ratios across the wave. M1 = 3
Please answer the second question related to the diffuser in detail. Dont solve the problem. Thanks!~
Qs: A two-dimensional supersonic inlet is to be designed to operate at Mach 2.4. Deceleration is to occur through a series of oblique shocks followed by a normal shock, as shown in Figure. Determine the loss of stagnation pressure for the cases of two, three, and four oblique shocks. Assume the wedge turning angles are each 6°. M. = 2.4 Vi, Pai P.- 50 kPa 12° Vi, Pai T.-0° C
When a pitot-static tube is placed in a supersonic flow, the probe reads P = 190 kPa and P = 150 kPa. If the stagnation temperature is 400 K, estimate the freestream Mach number and velocity.
A nozzle for a supersonic wind tunnel is designed to achieve a Mach number of 3.2, with a velocity of 2500 m/s, and a density of 1.0 kg/ m³ in the test section. Find the temperature and pressure in the test section and the upstream stagnation conditions. The fluid is helium. Te = i Pe= To= Po= Mi K kPa K kPa
1.AB
Air at V_1 = 800 m/s, p_1 = 100 kPa, and T_1 = 300 K passes through a normal shock. Calculate the velocity V_2, temperature T_2, and pressure p_2 after the shock. What would be the values of T_2 and p_2 if the same velocity change were accomplished isentropically?
(Q) A converging-diverging nozzle discharges an over-expanded jet into a room. Calculate the deflection angle in degrees caused by the oblique shock forming at the nozzle exit? Ratio between the room pressure and the pressure at the nozzle exit is 5.4. Mach number at the nozzle exit is 1.5. Assume the fluid to be air as a perfect gas with k=1.4
Nozzle is assuming steady one-dimensional flow. M = 2.731. This is the the supersonic flow of air through a convergent-divergent nozzle. The stagnation temperature = 300K, stagnation pressure at the inlet = 107500Pa, static pressure at the exit=4400Pa, C1 is a constant = 0.1097 for calculating circular cross-sectional area of a convergent-divergent nozzle: A = C1 + x^2 and x (axial distance from the throat) =1m. γ = 1.4 and R=287. Calculate the mass flow rate of air through the nozzle. Thank You.

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