JAWAHARLAL NEHRU TECHNOLOGICAL UNIVERSITY HYDERABAD
JAWAHARLAL NEHRU TECHNOLOGICAL UNIVERSITY HYDERABAD B.Tech III Year AIRCRAFT PROPULSION Previous Question Paper (February - 2022 - R18)
Answer any five questions
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Main
A turbojet engine is traveling at 920 km/h at standard sea-level conditions. The ram efficiency is 87%, the compressor pressure ratio is 4.3:1, the compressor efficiency is 82%, the burner pressure loss is 2%, the fuel-air ratio is 0.0119, the turbine inlet total temperature is 688 oC. The turbine efficiency is 83.5%. Calculate
the specific gross thrust
Answer
Given: V = 920 km/h = 255.56 m/s. Standard sea level: T0 = 288.15 K, P0 = 101.325 kPa, a0 = 340.3 m/s, Cp = 1.005 kJ/kgK, gamma = 1.4. Mach number M = 255.56 / 340.3 = 0.751. Total pressure at intake P02 = P0 * (1 + eta_ram * ((gamma-1)/2) * M^2)^(gamma/(gamma-1)) = 101.325 * (1 + 0.87 * 0.2 * 0.751^2)^3.5 = 134.7 kPa. Compressor exit pressure P03 = 4.3 * 134.7 = 579.2 kPa. Burner exit pressure P04 = 0.98 * 579.2 = 567.6 kPa. Compressor exit temperature T03 = T0 * (1 + (1/eta_c) * (PR^((gamma-1)/gamma) - 1)) = 288.15 * (1 + (1/0.82) * (4.3^0.286 - 1)) = 466.8 K. Turbine inlet T04 = 688 + 273.15 = 961.15 K. Turbine work equals compressor work: Cp * (T04 - T05) = Cp * (T03 - T02). T05 = T04 - (T03 - T02) = 961.15 - (466.8 - 288.15) = 782.5 K. Turbine expansion: T05/T04 = 1 - eta_t * (1 - (P05/P04)^((gamma-1)/gamma)). Solving for P05 gives 182.4 kPa. Exit velocity V9 = sqrt(2 * Cp * T05 * (1 - (P0/P05)^((gamma-1)/gamma))) = 512.3 m/s. Specific gross thrust = V9 - V = 512.3 - 255.56 = 256.74 Ns/kg.
the thrust specific fuel consumption
Answer
Thrust specific fuel consumption (TSFC) is defined as the ratio of fuel-air ratio to the net thrust produced per unit mass flow. Net thrust = V9 - V = 256.74 Ns/kg. Given fuel-air ratio f = 0.0119. TSFC = f / (V9 - V) = 0.0119 / 256.74 = 4.635 * 10^-5 kg/Ns. Converting to standard units (kg/kN-hr), TSFC = 4.635 * 10^-5 * 1000 * 3600 = 166.86 kg/kN-hr.
Evaluate the working of turbojet with afterburner with a neat diagram.
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Elaborate the working conditions of subsonic and supersonic inlets with appropriate diagrams.
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Elaborate the factors affecting combustion chamber design.
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Explain the following performance coefficients with respect to thrust generated: (i) Nozzle coefficient (ii) Cross thrust coefficient (iii) Flow coefficient
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Which type of nozzle do you use in a fighter aircraft? Support your answer using ejector and variable area nozzles?
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A turbojet engine powering an aircraft flying at an altitude of 11,000m where Ta = 216.7 K and Pa = 24.444 kPa. The flight Mach number is 0.9. The inlet conditions to the nozzle are 1000 K and 60 kPa. The specific heat ratio of air and gases at nozzle are 1.4 and 4/3. The nozzle efficiency is 0.98. Determine the thrust per inlet frontal area and ratio between the thrust forces in C-D nozzle.
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At inlet of a centrifugal compressor eye, the relative Mach number is to be less than 0.97 to eliminate the formation of shock waves and its attendant losses. The hub to tip radius ratio of the inducer is 0.4. There is no IGV and the inlet velocity is axial and uniform over the impeller eye. The eye tip diameter is 20 cm. Determine:
The maximum mass flow rate throughout for a given rotational speed of 486 rps
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The blade angle at the inducer tip corresponding to this mass flow rate (Take air inlet conditions to be 101.3 kPa, 288 K.)
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An axial flow gas turbine stage has a flow coefficient of 0.65, a constant axial velocity, and the gas leaves the stator blades at an angle of 65o to axial direction. Calculate
The blade-loading coefficient
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The rotor blade relative flow angle
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The degree of reaction
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