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AE2255 PROPULSION-1 QUESTION BANK UNIT-1 1. Define Froude efficiency, what is its effect on thrust? 2 2.Compare air breathing engine and rocket engine. 2 3. Define SFC.Write down its significance. 2 4. Mention the factors affecting thrust. 2 5. Find the propulsive efficiency of a jet engine moving with 300 m/s at 7000m altitude and its exhaust gas velocity is 600 m/s. 2 6. Define by pass ratio. 2 7. Why rate of thrust for an air breathing engine decreases with altitude and increases for non air breathing engine? 2 8. Differentiate between Scramjet & Ramjet engine. 2 9. Why is the ‘reverse diffuser’ impractical? 2 10. What are the advantages and disadvantages of cooling gas turbine blades? 2 11. Mention relative merits of jet engines over piston engines. 2 1. An advanced fighter engine operating at Mach 0.8 and 10Km altitude where, Ta=223.297K & Pa=0.2649 bar has the following uninstalled performance data and uses a fuel with C.V= 42,800KJ/Kg: Thrust = 50 KN Mass flow of air = 45Kg/s Mass flow of fuel = 2.65 Kg/s Determine the specific thrust, thrust specific fuel consumption; exit velocity, thermal efficiency, propulsion efficiency, and overall efficiency (assume exit pressure equal to ambient pressure). 16 2. Find specific thrust and SFC of a simple turbojet engine, having the following component performance at which the cruise speed and altitude are M 0.8 and 10000m. Select ambient condition from the gas table. Compressor pressure ratio Turbine inlet temperature Isentropic efficiency: Of compressor ηc Of turbine ηt Of intake ηi Of propelling nozzle ηj Mechanical transmission efficiency ηm Combustion efficiency ηb Combustion chamber pressure loss ΔPb 8.0 1200K 0.87 0.90 0.93 0.95 0.99 0.98 4% of compressor outlet pressure. C.V of fuel is 43,000 KJ/Kg, assume data if necessary, Cpa ≠ Cpg 16 3. (a) Explain with neat sketch operating principles of turbofan engine 8 (b) What is thrust augmentation? Explain any two methods of thrust augmentation with sketches. 8 4. Compare the characteristics, advantages & disadvantages of turbojet, turbofan and turboprop engine. 5. (i)Discuss the different methods of thrust augmentation. Draw T-S diagram for turbojet engine with thrust augmentation. 8 (ii) Discuss the typical turbojet cycle performance with suitable sketches. 8 turbine and nozzle are 0. 16 UNIT-II 1.6 at an altitude of 4500m (Pa = 0. 16 10. The heating value of the fuel is 45870 kj/kg.5405 bar and 255.000 kj/kg. By assuming the following data. 2 1.4. The fuel has a heating value of 41. Calculate the (i) Specific thrust and (ii) TSFC 16 7.75. After expanding in the turbine. 8 (ii) Derive the relation between area ratio Amax/Ai and external deceleration ratio . What are the starting problems in supersonic inlets? 2 4. Assume Cp = 1. The compressor pressure ratio is 5 & maximum temperature in the combustion chamber is 1273 K. The isentropic efficiency of compressor. Air enters a turbojet engine at a rate of 12*10 adiabatically to 1820C & four times the pressure.147 kj/kg-K. decreases the flow to a negligible velocity.81.Assume the data required suitably. (i) Define thrust of an engine and derive the thrust equation for a general propulsion system.85 & 0. 8 8. A turbojet engine is traveling at 270 m/s at an altitude of 5000m. Assume the isentropic efficiency of the turbine is same as that of the compressor and the nozzle efficiency is 90%. the gases continue to expand in the nozzle to a pressure of 0. 0.2 kg/s is to fly at Mach number of 0.9. An ideal turbojet flies at sea level at a Mach number of 0.005 kj/kg-K.7 K. Calculate the isentropic efficiency of the compressor.55 bar. Write notes on pressure recovery factor of the intake? 2 3. the exit speed of gasses & thrust developed when flying at 800 km/h.69 bar. What are the factors to be considered while designing a subsonic inlet? 2 5. (i) Explain successive steps in the acceleration and over speeding of a onedimensional supersonic inlet with sketches. What are the requirements of an aircraft intake? 2 2. 16 4 kg/h at 150C &1.915 respectively. and the compressor operates with a total pressure ratio of 15. What are the factors to be considered while designing a supersonic inlet? 2 6. the power required to drive the compressor. A jet propelled plane consuming air at the rate of 18. and the burner exit total temperature is 1389 K. Calculate (i) Power input to the compressor (ii) Power output of the turbine (iii) The fuel air ratio (iv) The thrust provided by the engine (v) The thrust power developed.83 kg/s of air.6. Cpg = 1. Ta = 255K ). Products of combustion enter the turbine at 8150C & leave it at 6500C to enter the nozzle. Find the thrust developed and the TSFC. Assume that the specific heat ratio is 1. The diffuser which has a pressure coefficient of 0. It ingests 74.100 kj/kg Combustion efficiency 98% Mechanical transmission efficiency 99% Isentropic efficiency of turbine 90% Propelling nozzle efficiency 95% Ambient conditions at 5000 m are 0. The compressor pressure ratio is 8:1 and maximum cycle temperature is 1200K. 8 (ii) Discuss the typical turbojet cycle performance with suitable sketches. Ram efficiency 93% Isentropic efficiency of compressor 87% Pressure loss in combustion chamber 4%of compressor delivery pressure Calorific value of fuel 43.03 bar and is compressed 9. What is meant by sub critical mode of inlet operation? State its advantages and disadvantages. 8 4. Define equivalence ratio and stochiometric fuel air ratio. What is the purpose of secondary air in combustion chamber? 2 6. and with a Mach number of 3. The inlet is to operate at a flight Mach number of 1. 2 1. assuming constant specific heats.97. static temperature. UNIT-III 1. State the advantages and disadvantages of annular combustor. A supersonic inlet is designed with a two-dimensional conical spike (with two half-cone angles 100 and 200 relative to the axial centerline. What are the different modes of inlet operation? Explain with suitable sketches.The two standing oblique shocks are attached to the spike and cowl. (iv) The Mach number.4 and internal diffuser pressure recover factor Πr = 0. 2 3. 2 1.102 kPa. Give any four functions of an exhaust nozzle. (a) What are the important factors affecting combustor design? 8 (b)Write down the methods of flame stabilization and explain with sketch. 8 UNIT-IV 1. A converging-diverging is designed to operate with an exit Mach number of 1. What will be the exit pressure and temperature? 8 2. (a)What are the three types of combustion chamber? Compare its advantages and disadvantages.5 with exit diameter of 200 mm. which is followed by a normal shock. Find the ratio of throat area/exit area necessary. Determine. balance the chemical equation for the stoichiometric combustion of this fuel in air and find the stoichiometric fuel-to-air ratio. 8 (b) Name the material used for combustion chamber and discuss the special qualities of the material used for combustion chamber? 8 3. Find also the maximum mass flow rate through the nozzle. Assume γ = 1. The reservoir conditions are given as Po = 106 Pa. the normal shock is swallowed). 2 3. Is it possible to have over expanded jets in convergent nozzle? Justify your answer. Estimate the overall recovery factor Πd on the assumption that the inlet starts (i.e. 2 4.ui/ua. Also.0. To = 200C. What is choked nozzle? 2 2. (ii) The Mach number.9.80. (i) With a neat sketch explain the working of a combustion chamber. find the required A*/A1. when it flows full. (i) The velocity. total and static pressure and static temperature after the normal shock. 8 (b)A De Laval nozzle has to be designed for an exit Mach number of 1. Explain the variations. (iii) The flow deflection angle. and a converging inlet section with a throat of area A* is used to decelerate the flow through internal compression. Air enters a two-dimensional supersonic diffuser at a pressure of 14. respectively).. 3. static pressure and static density variations along the longitudinal axis of a convergent-divergent nozzle. What is the purpose of primary air in combustion chamber? 2 5. 8 (ii) Consider n-decane fuel. total pressure after the oblique shock. total temperature and pressure of the air entering the oblique shock. 8 2. Define combustion intensity? 2 8. The nozzle is . a temperature of 217 K. The two-dimensional oblique shock diffuser has an oblique shock angle of 27. (a)What are the factors affecting combustion chamber? Explain briefly? 8 (b) With the aid of a simplified picture explain the operation of a flame holder. Define efficiency of the combustor. (a) Plot Mach number. What is need for supersonic combustion? 2 2. What is the purpose of dilution air in combustion chamber? 2 7.75. 16 4. 8 2. (ii) The stage loading coefficient.36. (i) The power required to drive the compressor while delivering 57 Kg/s of air. 8 (ii) A centrifugal compressor has an impeller tip speed of 366 m/s. 16 UNIT-V 1. Define degree of reaction for an axial flow compressor.supplied from an air reservoir at 68bar (abs. 2 1. (i) A sixteen-stage axial flow compressor is to have a pressure ratio of 6. If the actual temperature rise and pressure ratio developed are 300C and 1. (i) Explain the working of a centrifugal compressor and draw the velocity triangles. and temperature 777.3. pressure 68.300 rpm and compresses 31. Tests have shown that a stage total-to-total efficiency of 0. The value of Cp and γ .75 kg/s of air. An exhaust air stream at Mach 2. What are the causes for stalling in axial flow compressors? 2 5. assume isentropic flow except for the normal shock. determine the mass flow rate. 8 (ii) Write short notes on the following: (a) Ejector and variable area nozzles 4 (b) Thrust reversing 4 4.8 K enters a frictionless diverging nozzle with a ratio of exit area to inlet area of 3. Determine the absolute Mach number of the flow leaving the radial vanes of the impeller when the radial component of velocity at impeller exit is 30. Assume one-dimensional steady flow with the air behaving as a perfect gas with constant specific heats and a specific heat ratio of 1. The inlet pressure and temperature are 241. Assume an inlet total temperature of 288 K. The efficiency of the stage is 88%. Assuming 1-d flow. calculate: (i) Maximum backpressure to choke the nozzle. 2 3. (i) What are the types of nozzle? Explain various operating conditions of a C-D nozzle with suitable sketch.95kPa.62 and 13. Write down the difference between centrifugal and axial flow compressors. Assuming constant work done in each stage and similar stages fine the compressor overall total-to –total efficiency. 2 2. The slip factor is unity. An axial compressor stage has a mean diameter of 60cm and runs at 15000rpm. A stage of a radial compressor is to be analyzed. 2. 4 (ii) Range of backpressure over which a normal shock will appear in the nozzle. The hub and tip radii of the blades at the inlet are 7.94cm and the exit blade height is 2.9 can be obtained for each of the first six stages and 0. 2 4. For a mass flow rate of 40 kg/s determine the power required by the compressor.5 m/s and the slip factor is 0. 8 4. The exit radius is 27. 2 6. It rotates at 12. assume mechanical efficiency of 86 % and an initial temperature of 350C.).97cm respectively.325 kPa and 306K respectively. Determine the back pressure necessary to produce a normal shock in the channel at an area equal to twice the inlet area.4 respectively.9.0.1m2 and the total-to-total efficiency of the impeller is 90%. 2 7. Given that the flow area at impeller exit is 0. 4 (iv) Range of back pressure for supersonic flow at the nozzle exit plane.54cm. 4 (iii) Back pressure for the nozzle to be perfectly expanded to design M.89 for each of the remaining ten stages. 8 3. Define slip factor. Write down the conditions for free and forced vortex flows. Define rotating stall for compressors. 8 (ii) Discuss the factors affecting stage pressure rise of an axial flow compressor with suitable sketches. Distinguish between surging and stalling. (iii) The stage efficiency and (iv) The degree of reaction if the temperature at the rotor exits is 550C.9. 4 3. Flow enters the inducer with no prewhirl and the impeller has straight radial blades. It can be further assumed that stagnation enthalpy and entropy are constant and after the rotor row. 6.Determine. (iv) The Mach numbers at the impeller inlet and exit. 16 http://feeds. 16 5.0m.3 KPa at the inlet to the impeller of the centrifugal-flow compressor is 1. The impeller eye has the minimum diameter of 3. (iii) The total pressure ratio for the stage. (v) The required power for the stage. The inlet flow is in the axial direction. The mass flow rate of flow at 288 K and 101.005 kj/kg-K and 1.feedburner.7cm and rotates at 35. β2 = 300 .com/rejinpaul kg/m3 which . calculate the ideal angle at the hub and tip at the inlet to the impeller. The tip diameter is constant and 1.000rpm. An axial flow compressor stage is designed to give free-vortex tangential velocity distributions for all radii before and after the rotor blade row. β1 Absolute outlet angle. α2 = 600 = 600 Relative outlet angle. at the rotor tip the flow angles are as follows: 16 0 Absolute inlet angle.397 respectively.5 can be assumed constant for the stage. (ii) The static pressure at the impeller exit.81cm and a maximum diameter of 12. Find the following: (i) Mean relative flow angle at the inlet. α1 = 30 Relative inlet angle. the hub diameter is 0. (i) the axial velocity (ii) the mass flow rate (iii) the power absorbed by the stage (iv) the flow angles at the hub (v) the reaction ratio of the state at the hub Given that the rotational speed of the rotor is 6000 rpm and the gas density is 1. Draw velocity diagram at the hub and at the tip. Assuming no blockage due to the blade.are 1.814 kg/s.9m and constant for the stage.