139. Explore the theoretical foundations of aerospace propulsion systems and their applications in aircraft and spacecraft, considering concepts such as jet engine dynamics, rocket propulsion, and aerodynamic efficiency. ????
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A propulsion system uses a turbojet engine. The combustion gas drives the turbine and is then accelerated to high speeds through a converging-diverging nozzle. The temperature at the exit of the combustion chamber (inlet of the turbine) is 1313 K. It is assumed that the combustion process takes place at very low Mach numbers ( < 0.2), and therefore is a constant pressure process (no pressure loss). The thermodynamic processes in the compressor and in the turbine are considered isentropic ( do not confuse this isentropic process with that in variable area flows) and the Mach numbers at their inlets and outlets are low. Consequently, in the calculations the thermodynamic properties in the combustion chamber, compressor, and the turbine can be considered as stagnation properties , (as we described in the class). The flow in the nozzle is also assumed to be isentropic and the flow at the nozzle exit is perfectly expanded. Assume the mass flow rate of the fuel is negligible. Determine the optimal compressor pressure ratio to achieve the maximum specific thrust on the ground (temperature=293K, pressure=IOOkPa, and negligible flight Mach number, e.g., during take off). Determine the air mass flow rate needed to achieve a thrust of 55kN. If an afterburner is used, determine the afterburner temperature in order to achieve a thrust of 70kN.
Supreeta N.
4.4 a A simple no bypass turbojet engine flies at 256.5 m/s. The compressor and turbine are as described in Exercises 4.1, 4.2, and 4.3. Find the jet velocity which would be produced with the temperatures given as appropriate for initial cruise (31000 feet, M = 0.85 in part d of Exercise 4.3). Assume that all the net work is used to increase the kinetic energy of the flow. Explain why the propulsive efficiency and overall efficiency are low. Indicate ways in which no could be raised at this flight speed. (Ans: Vj = 832 m/s, np = 0.471, no = 0.225) b Recalculate the efficiencies if the flight speed were 600 m/s, M = 1.99 at 31000 feet. (Ans: Vi = 993 m/s, np = 0.753, no = 0.360)
A turbojet aircraft flies with a velocity of $1100 \mathrm{~km} / \mathrm{h}$ at an altitude where the air temperature and pressure are $-35^{\circ} \mathrm{C}$ and $40 \mathrm{kPa}$. Air leaves the diffuser at $50 \mathrm{kPa}$ with a velocity of $15 \mathrm{~m} / \mathrm{s},$ and combustion gases enter the turbine at $450 \mathrm{kPa}$ and $950^{\circ} \mathrm{C}$. The turbine produces $800 \mathrm{~kW}$ of power, all of which is used to drive the compressor. Assuming an isentropic efficiency of 83 percent for the compressor, turbine, and nozzle, and using variable specific heats, determine $(a)$ the pressure of combustion gases at the turbine exit, $(b)$ the mass flow rate of air through the compressor, ( $c$ ) the velocity of the gases at the nozzle exit, and $(d)$ the propulsive power and the propulsive efficiency for this engine. This problem is solved using appropriate software.
Madhur L.
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