A nozzle for a supersonic wind tunnel is designed to achieve a Mach number of 3.0, with a velocity of 2000 m/s, and a density of 1.0 kg/m in the test section. Find the temperature and pressure in the test section and the upstream stagnation conditions. The fluid is helium.

QUESTION 2 A compressor blade design tested in a cascade is found to choke with an inlet Mach number of 0.9 when the inlet flow angle is 52 deg. If the ratio of the throat area to the frontal area, A*/H1s, for the cascade is).625, calculate the loss of stagnation pressure between the far upstream and the throat and express this as a loss coefficient. Give your comments on what could cause this loss.

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In the test section of a supersonic wind tunnel, a Pitot tube in the flowreads a pressure of 1.13 atm. A static pressure measurement (from apressure tap on the sidewall of the test section) yields 0.1 atm. Calculatethe Mach number of the flow in the test section.

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.

Q2: Air (y = 1.4) enters a converging-diverging diffuser with a Mach number of 2.8, static pressure pi of 100 kPa, and a static temperature of 20°C. For the flow situation shown in Figure, find the exit velocity, exit static pressure, and exit stagnation pressure. Ans: Ve = 55.093 m/s; Pe = 2236.678 kPa; Poe = 2252.44 kPa i M₁ = 2.8 A₁ = 0.10 m² A₂ = 0.50 m² A₁ = 0.25 m²

43.4.5. 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?

Please show all work of a,b,&c. A supersonic converging-diverging nozzle is near the exit of a turbojet engine. Its cross-section is circular. The gas has the properties of air, and the flow is isentropic. At the throat location (the minimum area) the air flow is choked (is Mach one) and the throat has diameter of 20 cm.

he reservoir pressure and temperature for a Laval nozzle are 5 atm [abs] and 520 K,respectively. The flow is isentropically expanded to supersonic speed at the nozzle exitwith area Ae = 4in2. If the exit-to-throat area ratio is 2.295, calculate the followingexit properties: d) ρe; e) p0,e; f) T0,e; gamma = 1.4, R = 287 Please show all work!!

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