A supersonic flow at
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Fundamentals of Aerodynamics
- An aircraft is flying at supersonic speed. At a component of an aircraft where the flow is perpendicular, the density ratio is 5. Solve for: a.Mach Number Downstream b. Pressure Ratio c. Temperature Ratio d. Mach Number upstreamarrow_forwardAir 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?arrow_forwardAir enters a converging–diverging nozzle of a supersonic wind tunnel at 150 psia and 100°F with a low velocity. The flow area of the test section is equal to the exit area of the nozzle, which is 5 ft2. Calculate the pressure, temperature, velocity, and mass flow rate in the test section for a Mach number Ma = 2. Explain why the air must be very dry for this application.arrow_forward
- Engineers call the supersonic combustion in a scramjetengine almost miraculous, “like lighting a match in a hurricane.”Figure C9.8 is a crude idealization of the engine.Air enters, burns fuel in the narrow section, then exits, all at supersonic speeds. There are no shock waves. Assume areasof 1 m2 at sections 1 and 4 and 0.2 m2 at sections 2 and3. Let the entrance conditions be Ma1 = 6, at 10,000 mstandard altitude. Assume isentropic fl ow from 1 to 2,frictionless heat transfer from 2 to 3 with Q = 500 kJ/kg,and isentropic fl ow from 3 to 4. Calculate the exit conditionsand the thrust produced.arrow_forwardConsider 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 implicationarrow_forward1. Consider a low-speed subsonic wind tunnel designed with a reservoir cross-sectional area A, = 2 m2 and a test-section cross-sectional area A2 = 0.5 m2. The pressure in the test section is P2 = 1 atm. Assume constant density equal to standard sea level density, calculate the pressure (in kPa) required in the reservoir, P1, necessary to achieve a flow velocity V2: 40 m/s in the test section. a. From item no. 1, calculate the mass flow rate (in kg/s) through the wind tunnel. b. Calculate the Mach number of the vehicle in air. c. Calculate the Mach number of the vehicle in hydrogen.arrow_forward
- In compressible flow, velocity measurements with a Pitot probe can be grossly in error if relations developed for incompressible flow are used. Therefore, it is essential that compressible flow relations be used when evaluating flow velocity from Pitot probe measurements. Consider supersonic flow of air through a channel. A probe inserted into the flow causes a shock wave to occur upstream of the probe, and it measures the stagnation pressure and temperature to be 620 kPa and 340 K, respectively. If the static pressure upstream is 110 kPa, determine the flow velocity.arrow_forwardConsider a low-speed subsonic wind tunnel designed with a reservoir cross-sectional area of 2 m2 and a test-section cross-sectional area of 0.5 m2. The pressure in the test section is 1 atm.arrow_forwardA normal shock produced by an explosion propagate at a constant velocity of 450 m/s into still air of 100 kPa and 23 °C. The ratio of stagnation pressure upstream the shock to stagnation pressure of the gas flow behind the wave isarrow_forward
- 1. a. Consider a low-speed subsonic wind tunnel designed with a reservoir cross-sectional area A1 = 2 m2 and a test-section cross-sectional area A2 = 0.5 m2. The pressure in the test section is P2 = 1 atm. Assume constant density equal to standard sea level density, calculate the pressure (in kPa) required in the reservoir, P1, necessary to achieve a flow velocity V2 = 40 m/s in the test section. b. calculate the mass flow rate (in kg/s) through the wind tunnel.arrow_forwardNozzle 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.arrow_forwardA supersonic airflow at Ma1 = 3.2 and p1 = 50 kPa undergoesa compression shock followed by an isentropic expansionturn. The flow deflection is 30° for each turn. ComputeMa2 and p2 if (a) the shock is followed by the expansionand (b) the expansion is followed by the shock.arrow_forward
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