Aircraft & Rocket Propulsion Question Paper 2011 for M.Tech in Thermal Engineering (BPUT) is given below. This question was asked in BPUT M.Tech 2nd Sem. examination. Questions in this subject are asked from the “Gas Turbine” Book By V.Ganesan of TMH Publication. It is a problematic paper.
Second Semester Examination
Aircraft & Rocket Propulsion Question Paper 2011
Time: 3 Hour:
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.
1. Answer the following questions:
(a) Distinguish between fan, blower and Compressor. Give two examples from each from application view point.
(b) Draw Fanno curve and explain its salient points.
(c) How the forward motion of an air craft is achieved by propeller action? Explain with the help of illustrative sketches.
(d) What are the main components of a gas turbine engine used for turbojet aircrafts? Show the various process occurring in the engine on a T-S Diagram.
(e) Represent the processes of a ramjet engihe on a h -s diagram. What is the effect of flight Mach number on its efficiency?
(f) Draw the variation curves for propulsive efficiency with flight speed for various propulsive devices and analyse the result.
(g) Why is a rocket called non-air breathing ‘engine?
(h) Draw the schematic diagram of a multistage rocket vehicle system and with salient features.
(i) How does the development of thrust in a rocket engine differ from that in a turbojet engine?
(j) What is cryogenic propellant? Give name and importani properties of two cryogenic propellant?
2. (a) How the static and total efficiencies of fans defined?
(b) An axial ducted fan without any guide vanes has a pressure coefficient of 0.38 and delivers 3 kg/s of air at 750 rpm. Its hub and tip diameters are 25 cm and 75 cm respectively. If the pressure and temperature at the entry are 1bar and 37 Degree C respectively. Determine:
- (i) air angles at the entry and exit.
- (ii) pressure developed in mm of W.G
- (iii) Fan efficiency
- (iv) Power required to drive the fan if the overall efficiency of the drive is 85%.
3. An air craft flies at 960 kmph. One of its turbojet engines take in 40 Kg of air and expands the gases to the ambient pressure. The air fuel ratio is 50 and the lower calorific value of the fuel is 43 MJ/kg. For maximum thrust powers determine:
- (i) Jet Velocity
- (ii) Thrust
- (iii) Specific Thrust
- (iv) Thrust Power
- (v) Propulsive, thermal and overall efficiencies
- (vi) TSFC
Also derive the condition for maximum thrust power.
5. A gas turbine set draws in atmospheric air at 1.01325bar and 27 °C. there are two pressure stages with perfect intercooler and the total pressure ratio is 25.1. The maximum temperature of the cycle is 1300 degree celcius as there expansion. A regenerator is used and recovers 705 of the available heat in the exhaust. The turbine and compressor efficiencies are 0.87 arid 0.85 respectively. Further considering the overall mechanical efficiency and generator efficiency as 0.92 and 0.95 respectively, Determine:
- (i) Draw the schematic and T -S diagram
- (ii) The efficiency of the plant
- (iii) Work ratio
6. (a) Define characteristic velocity and derive the relation from fundamentals
V= ao / [y( 2/y+1)”.1r2″.1
Where go, y and v are standard notations.
(b) A rocket flies at 10,080kmph with an effective exhaust jet velocity of 1400m/s and propellant flow rate of 5.0 kg/s. If the heat of reaction of the propellant is 6500kJ/kg of the propellant mixture, Determine
- (i) Propulsion efficiency end propulsion- power
- (ii) Engine output and therrnal efficiency
- (iii) Overall efficiency
7. A rocket has the following data:
- Propellant flow rate = 5.25 kg/s
- Nozzle exit diameter = 9.5 cm
- Nozzle exit pressure = 1.018 bar
- Ambient pressure = 1.0325 bar
- Thrust chamber pressure = 21 bar
- Thrust = 71 kN
(i) Determine the effective jet velocity, actual jet velocity, specific impulse and the specific, propellent consumption. •
(ii) Recalculate the value of thrust and specific impulse for an altitude where the ambient pressure is 12mbar,
(iii) If the exhaust gases „have specific heat ratio -value of 1.3 for exit pressure 1.018. Determine exit Mach number. nozzle area ratio, thrust and propellant weight flow coefficients.
8. Write short note
(a) What is the purpose of injectors in rocket engines? Describe an injector with add of a sketch,
(b) Describe the events leading to pressure oscillation In a rocket Combustor.
(c ) How are regressive, neutral and-progressive burning of the solid propellant grain achieved? Explain With the aid Of diagram
(d) Expand and explain briefly-the followings
(i) UDMII and (ii) RFNA