B.E-B.E Aeronautical Engineering Aerospace Propulsion II(Sathyabama University, Chennai, Tamil Nadu-2010)
Friday, 23 August 2013 01:54Duraimani
(Established under section 3 of UGC Act,1956)
Course & Branch :B.E - AERO
Title of the Paper :Aerospace Propulsion – II Max. Marks :80
Sub. Code :526604/626604 Time : 3 Hours
Date :29/04/2010 Session :FN
PART - A (10 x 2 = 20)
Answer ALL the Questions
1. State the advantage and disadvantages of cooling a gas turbine blade.
2. Define blade – loading coefficient.
3. Mention the advantages and disadvantages of ramjet engine.
4. Distinguish between ramjet and scramjet.
5. Define discharge correction factor. Can it be more than one? Justify your answer.
6. Compare air breathing engine and rocket engine.
7. List any two fuel-oxidizer combination for hybrid propellant rocket.
8. Define specific impulse.
9. Why are electrical rockets called essentially power limited?
10. What do you mean by resist jet?
PART – B (5 x 12 = 60)
Answer All the Questions
11. Briefly discuss the methods of turbine blade cooling and mention its advantages and disadvantages. What is the need for matching of compressor and turbine?
(or)
12. An axial flow gas turbine has the following data. Turbine stage inlet temperature = 1300K, Turbine nozzle outlet temperature = 340k, Turbine stage outlet temperature = 82k, Blade speed = 318m/s, Axial velocity = 184m/s (constant). Turbine inlet velocity is eequal to turbine outlet velocity. Calculate the following:
(a) Degree of reaction, (b) blade loading coefficient, (c) flow coefficient, (d) work done, (e) Draw the combined velocity diagram. Assume suitable data if required.
13. For a ramjet the flight Mach number is 2.5, at a constant combustion chamber temperature of 1945k. The heat of reaction QR of the fuel is 42.8 MJ/kg and the inlet area is 0.0929 m2. Determine the gross thrust, fuel to air ratio, specific impulse, TSFC, and overall efficiency of the ideal engine when it is operated at sea level.
(or)
14. A ramjet engine is intended to be designed for operation at a Mach number of 3.1 and at an altitude of 15.8 km. The temperature and pressure at this altitude are 216.8k and 0.1289 x 105 Pa respectively. The maximum combustor chamber temperature is 2287K. The total pressure ratio across the air intake is 0.687. The diffuser exit Mach number may be assumed as 0.31. The heat of combustion of fuel is 434 KJ/kg. Calculate specific thrust and specific fuel consumption of the engine assuming that flow is frictionless and fully expand. Assume suitable design data if required.
15. With the help of schematic diagram explain the working of a liquid propellant rocket engine employing a gas pressure feed system. Mention the advantages and disadvantages of the liquid propellant.
(or)
16. A liquid oxygen-liquid hydrogen rocket thrust chamber of 44.5 KN thrust operates at a chamber pressure of 6.9 Mpa, a mixture ratio of 3.4 exhaust products with a mean molecular weight 4.04 kg/mole a combustion temperature of 2415°C, and a specific heat ratio of 1.26, determine the nozzle throat area, exit area for optimum at an altitude where P2 = 0.10907 bar, the propellant weight flow rates. Assume the actual specific impulse is 97% of the theoretical value.
17. What are the important factors that influence the burning rate of a solid propellant? Explain them with appropriate sketches. How do you classify solid propellant rockets and name any four solid propellants?
(or)
18. What are the various liquid fuel – oxidizer combinations used and explain their use. Explain the various injectors used with neat sketches. Explain the hybrid rocket with neat sketch.
19. With a neat sketch explain ion propulsion rocket. Explain various methods of thrust vector control with sketches.
(or)
20. Mention the various methods of cooling of thrust chamber assemblies and briefly explain any one cooling method. Explain the electric rocket propulsion technique with neat sketch.
| Earning: ₹ 3.15/- |
