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 8.0
Turbine inlet temperature 1200K
Of compressor ηc 0.87
Of turbine ηt 0.90
Of intake ηi 0.93
Of propelling nozzle ηj 0.95
Mechanical transmission efficiency ηm 0.99
Combustion efficiency ηb 0.98
Combustion chamber pressure loss ΔPb 4% of compressor outlet
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
6. A turbojet engine is traveling at 270 m/s at an altitude of 5000m. The compressor pressure ratio is 8:1 and maximum cycle temperature is 1200K. By assuming the following data,
Ram efficiency 93%
Isentropic efficiency of compressor 87%
Pressure loss in combustion chamber 4%of compressor delivery pressure
Calorific value of fuel 43,100 kj/kg
Combustion efficiency 98%
Mechanical transmission efficiency 99%
Isentropic efficiency of turbine 90%
Propelling nozzle efficiency 95%
Ambient conditions at 5000 m are 0.5405 bar and 255.7 K.
(i) Specific thrust and
(ii) TSFC 16
7. (i) Define thrust of an engine and derive the thrust equation for a general propulsion system. 8
(ii) Discuss the typical turbojet cycle performance with suitable sketches. 8
8. An ideal turbojet flies at sea level at a Mach number of 0.75. It ingests 74.83 kg/s of air, and the compressor operates with a total pressure ratio of 15. The fuel has a heating value of 41,000 kj/kg, and the burner exit total temperature is 1389 K. Find the thrust developed and the TSFC. Assume that the specific heat ratio is 1.4. 16
9. Air enters a turbojet engine at a rate of 12*104 kg/h at 150C &1.03 bar and is compressed adiabatically to 1820C & four times the pressure. Products of combustion enter the turbine at 8150C & leave it at 6500C to enter the nozzle. Calculate the isentropic efficiency of the compressor, the power required to drive the compressor, the exit speed of gasses & thrust developed when flying at 800 km/h. Assume the isentropic efficiency of the turbine is same as that of the compressor and the nozzle efficiency is 90%.Assume the data required suitably. 16
10. A jet propelled plane consuming air at the rate of 18.2 kg/s is to fly at Mach number of 0.6 at an altitude of 4500m (Pa = 0.55 bar, Ta = 255K ). The diffuser which has a pressure coefficient of 0.9, decreases the flow to a negligible velocity. The compressor pressure ratio is 5 & maximum temperature in the combustion chamber is 1273 K. After expanding in the turbine, the gases continue to expand in the nozzle to a pressure of 0.69 bar. The isentropic efficiency of compressor, turbine and nozzle are 0.81, 0.85 & 0.915 respectively. The heating value of the fuel is 45870 kj/kg. Assume Cp = 1.005 kj/kg-K, Cpg = 1.147 kj/kg-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. 16
1. What are the requirements of an aircraft intake? 2
2. Write notes on pressure recovery factor of the intake? 2
3. What are the starting problems in supersonic inlets? 2
4. What are the factors to be considered while designing a subsonic inlet? 2
5. What are the factors to be considered while designing a supersonic inlet? 2
6. What is meant by sub critical mode of inlet operation? State its advantages and disadvantages. 2
1. (i) Explain successive steps in the acceleration and over speeding of a one-
dimensional supersonic inlet with sketches. 8
(ii) Derive the relation between area ratio Amax/Ai and external deceleration ratio ui/ua. 8
2. A supersonic inlet is designed with a two-dimensional conical spike (with two half-cone angles 100 and 200 relative to the axial centerline, respectively). The inlet is to operate at a flight Mach number of 1.9.The two standing oblique shocks are attached to the spike and cowl, and a converging inlet section with a throat of area A* is used to decelerate the flow through internal compression. Assume γ = 1.4 and internal diffuser pressure recover factor Πr = 0.97. Estimate the overall recovery factor Πd on the assumption that the inlet starts (i.e., the normal shock is swallowed). Also, find the required A*/A1.
3. What are the different modes of inlet operation? Explain with suitable sketches. 16
4. Air enters a two-dimensional supersonic diffuser at a pressure of 14.102 kPa, a temperature of 217 K, and with a Mach number of 3.0. The two-dimensional oblique shock diffuser has an oblique shock angle of 27.80, which is followed by a normal shock. Determine, assuming constant specific heats.
(i) The velocity, total temperature and pressure of the air entering the oblique shock.
(ii) The Mach number, total pressure after the oblique shock.
(iii) The flow deflection angle.
(iv) The Mach number, total and static pressure and static temperature after the normal shock.
1. What is need for supersonic combustion? 2
2. Define equivalence ratio and stochiometric fuel air ratio. 2
3. Define efficiency of the combustor. 2
4. What is the purpose of primary air in combustion chamber? 2
5. What is the purpose of secondary air in combustion chamber? 2
6. What is the purpose of dilution air in combustion chamber? 2
7. Define combustion intensity? 2
8. State the advantages and disadvantages of annular combustor. 2
1. (a) What are the important factors affecting combustor design? 8
(b)Write down the methods of flame stabilization and explain with sketch. 8
2. (a)What are the three types of combustion chamber? Compare its advantages and disadvantages. 8
(b) Name the material used for combustion chamber and discuss the special qualities of the material used for combustion chamber? 8
3. (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. 8
4. (i) With a neat sketch explain the working of a combustion chamber. 8
(ii) Consider n-decane fuel, balance the chemical equation for the stoichiometric combustion of this fuel in air and find the stoichiometric fuel-to-air ratio. 8
1. What is choked nozzle? 2
2. Is it possible to have over expanded jets in convergent nozzle? Justify your
3. Give any four functions of an exhaust nozzle. 2
1. (a) Plot Mach number, static temperature, static pressure and static density variations along the longitudinal axis of a convergent-divergent nozzle, when it flows full. Explain the variations. 8
(b)A De Laval nozzle has to be designed for an exit Mach number of 1.5 with exit diameter of 200 mm. Find the ratio of throat area/exit area necessary. The reservoir conditions are given as Po = 106 Pa, To = 200C. Find also the maximum mass flow rate through the nozzle. What will be the exit pressure and temperature? 8
2. A converging-diverging is designed to operate with an exit Mach number of 1.75. The nozzle is supplied from an air reservoir at 68bar (abs.). Assuming 1-d flow, calculate:
(i) Maximum backpressure to choke the nozzle. 4
(ii) Range of backpressure over which a normal shock will appear in the nozzle. 4
(iii) Back pressure for the nozzle to be perfectly expanded to design M. 4
(iv) Range of back pressure for supersonic flow at the nozzle exit plane. 4
3. (i) What are the types of nozzle? Explain various operating conditions of a C-D nozzle with suitable sketch. 8
(ii) Write short notes on the following:
(a) Ejector and variable area nozzles 4
(b) Thrust reversing 4
4. An exhaust air stream at Mach 2.9, pressure 68.95kPa, and temperature 777.8 K enters a frictionless diverging nozzle with a ratio of exit area to inlet area of 3.0. Determine the back pressure necessary to produce a normal shock in the channel at an area equal to twice the inlet area. Assume one-dimensional steady flow with the air behaving as a perfect gas with constant specific heats and a specific heat ratio of 1.36; assume isentropic flow except for the normal shock. 16
1. Write down the difference between centrifugal and axial flow compressors. 2
2. Define degree of reaction for an axial flow compressor. 2
3. Define rotating stall for compressors. 2
4. What are the causes for stalling in axial flow compressors? 2
5. Define slip factor. 2
6. Write down the conditions for free and forced vortex flows. 2
7. Distinguish between surging and stalling. 2
1. An axial compressor stage has a mean diameter of 60cm and runs at 15000rpm. If the actual temperature rise and pressure ratio developed are 300C and 1.4 respectively.
(i) The power required to drive the compressor while delivering 57 Kg/s of air; assume mechanical efficiency of 86 % and an initial temperature of 350C.
(ii) The stage loading coefficient.
(iii) The stage efficiency and
(iv) The degree of reaction if the temperature at the rotor exits is 550C.
2. (i) Explain the working of a centrifugal compressor and draw the velocity
(ii) A centrifugal compressor has an impeller tip speed of 366 m/s. 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 30.5 m/s and the slip factor is 0.9. Given that the flow area at impeller exit is 0.1m2 and the total-to-total efficiency of the impeller is 90%, determine the mass flow rate. 8
3. (i) A sixteen-stage axial flow compressor is to have a pressure ratio of 6.3. Tests have shown that a stage total-to-total efficiency of 0.9 can be obtained for each of the first six stages and 0.89 for each of the remaining ten stages. Assuming constant work done in each stage and similar stages fine the compressor overall total-to –total efficiency. For a mass flow rate of 40 kg/s determine the power required by the compressor. Assume an inlet total temperature of 288 K. 8
(ii) Discuss the factors affecting stage pressure rise of an axial flow compressor with suitable sketches. 8
4. A stage of a radial compressor is to be analyzed. It rotates at 12,300 rpm and compresses 31.75 kg/s of air. The inlet pressure and temperature are 241.325 kPa and 306K respectively. The hub and tip radii of the blades at the inlet are 7.62 and 13.97cm respectively. The exit radius is 27.94cm and the exit blade height is 2.54cm. The slip factor is unity. Flow enters the inducer with no prewhirl and the impeller has straight radial blades. The efficiency of the stage is 88%. The value of Cp and γ are 1.005 kj/kg-K and 1.397 respectively.
Find the following:
(i) Mean relative flow angle at the inlet.
(ii) The static pressure at the impeller exit.
(iii) The total pressure ratio for the stage,
(iv) The Mach numbers at the impeller inlet and exit.
(v) The required power for the stage. 16
5. An axial flow compressor stage is designed to give free-vortex tangential velocity distributions for all radii before and after the rotor blade row. The tip diameter is constant and 1.0m; the hub diameter is 0.9m and constant for the stage. at the rotor tip the flow angles are as follows: 16
Absolute inlet angle, α1 = 300
Relative inlet angle, β1 = 600
Absolute outlet angle, α2 = 600
Relative outlet angle, β2 = 300
(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 1.5 kg/m3 which can be assumed constant for the stage. It can be further assumed that stagnation enthalpy and entropy are constant and after the rotor row.
6. The mass flow rate of flow at 288 K and 101.3 KPa at the inlet to the impeller of the centrifugal-flow compressor is 1.814 kg/s. The inlet flow is in the axial direction. The impeller eye has the minimum diameter of 3.81cm and a maximum diameter of 12.7cm and rotates at 35,000rpm. Assuming no blockage due to the blade, calculate the ideal angle at the hub and tip at the inlet to the impeller. Draw velocity diagram at the hub and at the tip. 16
AE2255 Propulsion - I Anna University Question bank, question paper pervious year question paper for Unit 1 unit 2 unit3 unit 4 unit 5,important 2 marks and 16 marks questions Reviewed by Rejin Paul on 4:05 AM Rating: