ANNA UNIVESITY MODEL QUESTION PAPER, B.E. Aeronautical Engineering,

Semester V- PROPULSION I

**PART – A** (10 x 2 = 20 Marks)

- 1.Bring out any two essential difference between aircraft piston engines and jet engines from operating principle point of view.
- 2.Mention any two advantages and two limitations of a turbojet powered airplane?
- 3.What is the effect of increase in air temperature during compression process on the overall jet engine performance?
- 4.Sketch typical flow pattern over a supersonic inlet of a jet engine?
- 5.Distinguish between primary zone and secondary zone of a gas turbine combustion chamber?
- 6.What is the function of a swirler in a can type gas turbine combustion chamber?
- 7.Sketch a typical flame stability curve of a gas turbine combustion chamber?
- 8.What is the function of impeller in a centrifugal air compressor?
- 9.Suggest any two methods to improve compression efficiency of an axial flow compressor.

10.Explain the phenomenon ‘surging’ of compressor.

**PART – B (5 x 16 = 80 Marks)**

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11.i)Sketch thrust vs air speed, TSFC vs air speed curves for turboprop, turbofan and turbojet engines. Also compare propulsive efficiency vs flight speed curves of the above three engines?

ii) Sketch the performance characteristic curves of centrifugal compressor and axial flow compressor. What conclusions can be made from these curves?

12.a)i) Classify gas turbine combustion chambers.

Bring out the advantages and limitations of any three types of combustion chambers from structural design and combustion performance point of view?

ii)What is the need for flame stabilization in a gas turbine combustion chamber? How is flame stability achieved in a jet turbine combustion chamber?

12.b)i) Sketch the typical flow pattern in the flame tube of a typical gas turbine combustion chamber and mark all the regions in the flow pattern.Explain the salient features of these regions.

ii) List the important operating variables and performance variables of a gas turbine combustion chamber. What is the effect of operating variables on the performance of a combustion chamber.

13.a)i) Draw pressure-specific volume, pressure–temperature and temperature-entropy diagrams for a gas turbine engine cycle. Marks all the events in the diagrams

ii) The overall air to fuel ratio for a turbojet is 70:1 and the engine flies at a mach number of 1.78 at a certain altitude where the free stream temperature is 245 k. The exhaust gas temperature is 623 k and the exit mach number of the jet nozzle is 2.34. Estimate the thrust specific fuel consumption. Assuming a total airflow rate of 67 kg/s, estimate the thrust developed by the engine.

OR

** **13.b)i) What is thrust augmentation? Explain how temperature vs entropy diagram of a turbojet gets modified by thrust augmentation.

ii) A turboprop engine operates at its design point (I.S.A, sea level conditions). Gas enters the turbine of the engine at a temperature and pressure of 950 k and 2.5 x 10^{5} p_{a} respectively. The gas flow rate is 55 kg/s. Calculate shaft horse power and jet thrust assuming that isentropic efficiencies of turbine and nozzle are 0.85 and 0.9 respectively. The jet velocity is 310 m/s and reduction gear efficiency is 0.97.

14.a)i) How do mass flux and heat flux of a convergent–divergent nozzle flow vary along the longitudinal axis?

ii) What are the important assumptions that are made in the isentropic theory of jet engine nozzle flows?

iii) The throat area a convergent–divergent jet nozzle is 0.0124m^{2 }and the nozzle entry conditions for the flow are 2.1 Mpa and 1518 k respectively.

Assuming stagnation conditions at the entry, estimate the flow rate through the nozzle. The ratio of specific heats may be assumed as 1.37.

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14.b)i) Explain the following types of expression in the nozzle flows

- Under expansion
- Over expansion
- Optimal expansion.

ii) Assuming stagnation conditions of nozzle inlet, determine the following

1.Temperature at the exit of the nozzl

2.exit velocity as a fraction of maximum attainable velocity.

The nozzle exit area = 0.01m^{2},inlet pressure and temperature are 2.5 Mpa and 1075 k respectively. Exit pressure is 113745 p_{a}. The ratio of specific heats may be assumed as 1.35.

15.a)i) Explain the starting problem associated with supersonic inlets. What is the remedy for this problem?

ii) A single sided centrifugal compressor is fitted to an aircraft which is flying at a certain altitude. The impeller eye reaches air at a velocity of 230 m/s. The free stream pressure and temperature of the air are 0.23×10^{5 }p_{a} and 217 k respectively. The impeller eye has a provision of air pre whirl of 25°at all radii. The inner and outer diameters of the eye are 0.18 m and 0.33 m respectively. The diameter of the impeller periphery is 0.54 m and the impeller rotates at a speed of 16,200 rpm. The air mass flow rate is 3.6 kg/s. Estimate the stagnation pressure at the compressor outlet.

OR

15.b)i) Derive an expression for minimum frontal area ratio as a function of deceleration ratio and maximum allowable velocity ratio for a subsonic inlet.

ii) A symmetrical blading axial flow compressor has an airflow with axial velocity of 145 m/s. The blading is designed for 50% reaction at mean diameter. Pressure ratio is 1.5 and isentropic efficiency is 86 %. Assuming that the flow is of Vortex type, estimate the degree of reaction at the root and tip of the blade, if the ratio of inside diameter to outside diameter is 0.75. The inlet conditions to compressor correspond to standard state sea level

i want aeronautical department 4th semester propulsion 1 may/june2013 question paper