Q3/A combustion chamber consists of tubular combustors of 15-cm diameter. Compressed air enters the tubes at 550 K, 480 kPa, and 80 m/s. Fuel with a heating value of 42,000 kJ/kg is injected into the air and is burned with an air-fuel mass ratio of 40. Approximating combustion as a heat transfer process to air, determine the temperature, pressure, velocity, and Mach number at the exit of the combustion chamber.
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- Air Is flowing isentropically through a converging duct wl)lcJ,,is fed from a. large reservoir where )he temperature and pressure are 300 1<" and 400 kPa, respectively. The mass flow rate through the duel Is 4 kg s-'. Determine the Mach number, pressure, temperature and velocity at a point where the cross-sectional area is 0.01 m2A combustion chamber consists of tubular combustors of 15-cm diameter. Compressed air enters the tubes at 550 K, 480 kPa, and 80 m/s. Fuel with a heating value of 42,000 kJ/kg is injected into the air and is burned with an air–fuel mass ratio of 40. Approximating combustion as a heat transfer process to air, determine the temperature, pressure, velocity, and Mach number at the exit of the combustion chamber.Ex: A combustion chamber consists of tubular combustors of 15-cm diameter. Compressed air enters the tubes at 550 K, 480 kPa, and 80 m/s. Fuel with a heating value of 42,000 kJ/kg is injected into the air and is burned with an air–fuel mass ratio of 40. Approximating combustion as a heat transfer process to air, determine the temperature, pressure, velocity, and Mach number at the exit of the combustion chamber
- Fuel with heating value (HV) of 39,000 kJ/kg is burned by spraying in to 16-cm-diameter tabular combustionchamber. The air enters at 450 K, 380 kPa, and 55 m/s as shown in Figure Q3. If the exit Mach number is 0.8,determine (a) the exit temperature and,(b) the rate at which the fuel is burned.Properties of air: Specific heat capacity ratio k = 1.4, Specific heat at constant pressure c, = 1.005 kJ/kg.K,and Gas constant R = 0.287 kJ/kg.K.Use Appendix 2 for Rayleigh flow functions for an ideal gas with k = 1.4 Assume:1. Rayleigh flow (i.e., steady one-dimensional flow of an ideal gas with constant properties through aconstant cross-sectional area duct with negligible frictional effects) are valid. 2. combustion is compleso, and it is treated as a heat addition process, with no change in the chemical composition of flow 3.The increase in mass flow rate due to fuel in ection is disregardeo.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.Consider the isentropic flow through a convergent-divergent nozzle with an exit-to-throat area ratio of 2.5. The reservoir pressure and temperature are 1 atm and 288 K, respectively. Calculate the Mach number, pressure, and temperature at both the throat and the exit for the cases where (a) the flow is supersonic at the exit and (b) the flow is subsonic throughout the entire nozzle except at the throat, where M = 1.
- Example: Carbon dioxide flows steadily at a mass flow rate of 3 kg/s and pressure of 1400 kPa and 200°C with a low velocity. It expands after the nozzle to a pressure of 200 kPa. The duct is designed so that the flow can be approximated as isentropic. Determine the density, velocity, flow area, and Mach number at each location along the duct that corresponds to a pressure drop of 200 kPaConsider a rocket engine burning hydrogen and oxygen; the combustion chamber temperature and pressure are 3517 K and 26 atm, respectively. The molecular weight of the chemically reacting gas in the combustion chamber is 16, and y = 1.22. The pressure at the exit of the convergent-divergent rocket nozzle is 1.174 x10-2 atm. The area of the throat is 0.4 m2. Assuming a calorically perfect gas and isentropic flow, calculate: (a) the exit Mach number, (b) the exit velocity, (c) the mass flow through the nozzle, and (d) the area of the exitAir flows through a long, isentropic nozzle. The temp. and pressure at the reservoir are 1000 K and 20 atm respectively. If the Mach number at the entrance is 0.2, what is the gas velocity at the entrance
- The converging-diverging nozzle of the engine of a new supersonic unmanned aerial vehicle (UAV) has an exit area of 50 cm^2 . At the design operating condition, the mass flow rate of exhaust gases (? ≈ 1.4, molecular weight = 26 g/mol) is 1.75 kg/s. The exhaust gases enter the nozzle with stagnation pressure and temperature equal to Po = 300 kPa and To = 500 K, respectively a) What is the optimal cruising altitude for the operation of the nozzle of this UAV? (b) What is the lowest altitude at which the plane can fly such that the flow inside the diverging part of the nozzle is kept supersonic? Note: - Assume that the stagnation conditions of the flow of exhaust gases entering the nozzle are kept constant with varying altitudeIn 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.Air at a total pressure and temperature of 8 atm and 450 K enters a frictionless constant cross-section duct. A heat addition of 850 kJ/kg makes the flow to choke at the duct exit, determine the inlet Mach number and the total pressure and total temperature at the exit.