Aircraft & Rocket Propulsion Question Paper 2014 is given below. This question was asked in the 2nd Semester M.Tech Mechanical Engineering (Heat Power Engg. ) Examination held in 2015 for 2013-15 admission batch. If you are going to appear the M.tech 3rd sem exam, than prepare this paper from the text book of “Gas turbine ” by V.Ganeshan, TMH Publication.
2nd Semester Regular/Back Examination 2014
Aircraft & Rocket Propulsion Question Paper 2014
BRANCH(S): MECHANICAL ENGINEERING (HEAT POWER ENGINEERING)
Time: 3 Hours
Max Marks: 70
Answer Question No.1 which is compulsory and any five from the rest.
The figures in the right hand margin indicate marks.
01. Answer the following questions (2×10)
a) What do you mean by air-breathing engine? Give four examples of air-breathing engine.
b) Define degree of reaction for an axial flow compressor
c) Why the liquid propellant rocket engine produces low thrusts as compared that of a solid propellant rocket engine.
d) What do you mean by homogenous solid propellant?
e) Distinguish among the sonic, supersonic and transonic flow.
f) Define the surging in a compressor.
g) Draw the air speed vs propulsion efficiency curve (free hand curve) for a pure turbojet, low bypass ratio turbojet, high bypass ratio turbojet and turboprop engine.
h) Draw a free hand curve for the thermal efficiency vs pressure ratio for an ideal Brayton cycle.
i) Define a shock in compressible flow. What happens to pressure and velocity
after the shock wave?
j) What do you mean by a level flight condition? In a level flight condition, how the drag is related to the vehicle weight.
Q2 .
a) The first stage of an axial flow compressor is designed on free-vortex principle, with no inlet guide vanes. The rotational speed is 6000 rpm and stagnation pressure rise is 20 K. The hub-tip ratio is 0.6, the work done factor is 0.93 and isentropic efficiency of stage is 0.89. Assuming an inlet velocity of 140 m/s
and ambient conditions of 1.01 bar and 288 K, Find the tip radius and corresponding rotor air angles, if the Mach number relative to tip is limited to 0.95. Also find the mass flow rate entering the stage.
b) For the above problem, find the stage stagnation pressure ratio and power required.
Q3 .
(a) A ramjet is to propel an aircraft at Mach 3 at high altitude where the ambient pressure is 8.5 k Pa and temperature Ta is 220 K. The turbine inlet temperature is 2540 K. If all the components of the engine are ideal, and assuming the specific heat ratio 1.4 and fuel air ratio f «1, determine the thermal efficiency.
(b) Also determine the propulsion and overall efficiencies for the above case.
Q4 .
a) Define compressor stall on basis of blade angle of attack.
b) An aircraft engine is fitted with a single sided centrifugal compressor The aircraft flies with a speed of 800 km/hr at an altitude where the pressure is 0.23 bar and temperature of 217 K. The inlet duct of impeller eye contains fixed varies which give the air pre-whirl of 25o at all radii. The inner and outer diameters of eye are 180 and 300 mm respectively, the diameter of the impeller tip is 540 mm and rotational speed 16,000 rpm. Estimate the stagnation pressure at the compressor outlet when the mass flow is 216 kg/min. Neglect losses in inlet and fixed vanes and assume isentropic efficiency of compressor is 0.8. Take slip factor as 0.9 and power input factor as 1.04
where T denotes the stagnation temperature and the subscripts 1 denotes the compressor and fan inlet, 2 is the compressor outlet, f is the fan outlet, 3 is the turbine inlet and β
is bypass ratio, respectively.
Q6.
a) A turbojet engine with zero bypass ratio has a pressure ratio of 30 and maximum temperature of 1700 K. The component efficiencies and ambient conditions are given as follows:
- Diffuser efficiency ƞd = 0.97
- Compressor efficiency ƞc = 0.85
- Burner efficiency ƞb = 1.00
- Turbine efficiency ƞt = 0.9
Nozzle efficiency ƞn = 0.98 - Fuel heating value 45,000 kJ/kg
The aircraft is flying at Mach 0.85 where ambient pressure and temperature are 18.75 k Pa and 216.7 K. Taking constant specific heat of 1.4, determine specific thrust and the thrust specific fuel consumption.
b) For the above problem, find also the engine thermal efficiency, propulsion efficiency and overall efficiency.
Q7. A simple gas turbine with heat exchanger has a compressor and turbine efficiencie ƞc and ƞt respectively. The pressure drop in gas side of the heat exchanger is ∆Phg and the exchanger is ∆P . The turbine inlet is T03 , compressor inlet pressure is P01 , compressor pressure ratio is rp ; the specific heat at constant pressure cp , and ratio of specific heats ұ are constant throughout the cycle. Show that the reduction in specific work output is
Q8. Write Short Notes (Any Two)
a) Liquid propellant rockets
b) Solid propellant rockets
c) Regenerative cooling in liquid propellant rockets