**AE 2304- PROPULSION - II**

**PART –A**

1. Differentiate between impulse stage and reaction stage turbines.

2. Define match point.

3. Write down the merits and demerits of integral ram-rocket.

4. What do you mean by supercritical mode of operation of ramjet?

5. Name any two oxidizer-fuel combinations used for hybrid rockets.

6. Compare air breathing engine and rocket engine.

7. Define Specific impulse.

8. Define temperature sensitivity coefficient of a solid propellant.

9. Define Characteristic velocity.

10. What is the basic concept in using advanced propulsion technique?

11. Define (a) Impulse stage (b) Reaction stage.

12. Define total-to-total efficiency and state when it is appropriate to use this efficiency.

13. An ideal ramjet engine operates at M = 1.5 at an altitude of 6500 m. Find its cycle efficiency.

14. How do you classify ramjets based on combustion process?

15. What are the limitations of hybrid rockets?

16. Define discharge correction factor. Can it be more than one? Justify your answer.

17. Define characteristic exhaust velocity.

18. Define specific impulse.

19. Why electrical rockets are called essentially power limited?

20. What is the basic principle of operation of a solid propellant rocket?

**PART B - (5**

**x 16 = 80 marks)**

11. (a) (i) Describe the working of an axial flow turbine stage with a neat sketch. Draw the T-S diagram and velocity triangles.

(ii) Discuss the limiting factors in turbine design.

(b) A mean-diameter design of a turbine stage having equal inlet and outlet velocities leads to the following data.

Mass flow

*m*Inlet temperature

*TOI*Inlet pressure POI

Axial velocity (constant through stage) Ca

Blade sped

*U*Nozzle effiux angle

*a*_{2}Stage exit swirl

*a*_{3}20 kg/s 1000 K

4.0 bar

260 nfs

360 nfs

65 degrees

10 degrees

Determine the rotor blade ~as angles, degree of reaction, temperature drop coefficient

*(2cpD.Tos/U*^{2}*)*and power output. Assuming a nozzle loss coefficient*AW*of 0.05, calculate the nozzle throat area required (ignoring the effect of friction on the critical conditions).12. (a) . (i) Describe the working of a ramjet engine. Depict the various thermodynamic processes occurring in it on h-s diagram.

(ii) Discuss the performances of Supersonic Combustion Ramjet.

Compare Subsonic and Supersonic combustion Ramjets.

(b) A ramjet is traveling at Mach 3 at an altitude of 4572 m, the external static temperature is 258.4K, and the external static pressure is 57.1 kPa. The heating value of the fuel is 46,520 kJlkg. Air flows through the engine at 45.35 kg/so The burner exit total temperature is 1944 K Find the thrust, fuel ratio, and TSFC. The specific heat ratio can be assumed to be 14.

13. (a) A chemical rocket is used for launch into earth orbit. At the end of the combustion chamber the stagnation temperature is 3000 K, The molecular weight of the combustion products is 26. The gases expand isentropically as an ideal gas mixture with specific heat ratio' 1.2, The

area ratio

*A*/_{e}*A'*of the nozzle is 20, and the throat is 0.1 m. At sea level determine:(i) The stagnation pressure if the expansion is correct,

(ii) The rocket thnist.

(b) (i) Explain the working of liquid propellant rocket engine with a gas

pressure feed system. Write down its merits and demerits.

(ii) What are the important factors in selecting a liquid propellant?

Justify those points.

14. (a) (i) What are the important factors that influence the burning rate of a

solid propellant? Explain them with appropriate sketches.

(ii) How do you classify solid propellant rockets? Name any four solid propellant ingredients function with two examples for each function.

(b) A rocket is to be designed to produce 5 MN of thrust at sea level. The pressure in the combustion chamber is 7 MPa and the temperature is 2800 K. If the working fluid is assumed to be a perfect gas with the properties of air at room temperature, determine the following:

(i) Specific impulse,

(ii) Mass flow rate,

(iii) Throat diameter,

(iv) Exit diameter and

(v) Thrust at 30 km altitude .

15. (a) (i) Mention the various methods of cooling of thrust chamber

assemblies and briefly explain anyone cooling method.

(ii) With the aid of neat sketches explain various techniques for thrust

vector control.

(b) (i) Draw a neat sketch and explain the working of ion propulsion

rocket.

(ii) How does the shape of the nozzle affect performance? How do you

overcome the thrust loss associated with over expansion?

**AE 2304- PROPULSION - II**

**PART B - (5**

**x**

**16**

**=**

**80 marks)**

11. (a) (i) Draw neat sketch and explain the general working principle of a

nuclear rocket.

(ii) Draw a neat sketch and briefly explain about electric rocket

propulsion technique.

(b) (i) Explain the working of electric plasma rocket with a neat sketch.

(ii) Describe the concept of Nozzleless propulsion with their merits and

demerits.

12. (a) The following data apply to a single-stage turbine designed on free-vortex theory.

Mass flow | 36 kg/s | |

Inlet temperature | TOI | 1200 K |

Inlet pressure | POI | 8.0 bar |

Temperature | !:J.T_{o} | 150 K |

Isentropic efficiency | TJi | 0.9 |

Mean blade speed | U_{m} | 320 rnfs |

Rotational speed | N | 250 rev/s |

Outlet velocity | C3 | 400 rnfs |

The outlet velocity is axial. Calculate the blade height and radius ratio of the anriulus from the outlet conditions. The turbine is designed with a constant annulus area through the stage, i.e. with no flare.

(b) Draw T-S diagram for a reaction stage turbine. Define the terms nozzle loss coefficient and rotor blade loss coefficient and prove that

*A*= 0.86 Y for even when the Mach number at the blade exit is unity.13. (a) A jet engine is to propel an aircraft at Mach 3 at high altitude where ambient pressure is 8.5 kPa and the ambient temperature is 220 K. The turbine inlet temperature is 2540 K. If all components of the engine are frictionless determine

(i) The thermal efficiency

(ii) The propulsion efficiency

(iii) The overall efficiency

Let the specific heat ratio be

*r*= 1.4 and make the approximation of*f«*1. (b) (i) With a neat sketch explain the concept of integral ram-rocket and mention its advantages and disadvantages.

(ii) Briefly discuss performance of supersonic combustion ramjet and compare subsonic combustion Ramjet with supersonic combustion Ramjet engine.

14. (a) (i) What are the important factors that influence the burning rate of a

solid propellant? Explain them with appropriate sketches.

(ii) How do you classify solid propellant rockets? Name any four solid

propellant ingredients with two examples for each.

(b) A chemical rocket is used for launch into earth orbit. At the end of the combustion chamber the stagnation temperature is 3000 K and the stagnation pressure is 2 MPa. The molecular weight of the combustion products is 26. The gases expand entropically as an ideal gas mixture with specific heat ratio 1.2. The area ratio

*A*/_{e}*A·*of the nozzle is 20, and the throat diameter is 0.1 m. At sea level, determine the rocket thrust.,

15. (a) How long would it take for a thrust of a rocket to diminish to 10% of its

steady value if the fuel and oxidant flows into the chamber were suddenly stopped? Consider, for example, the following conditions:

Initial combustion chamber pressure Po | = | 10 MPa |

Initial combustion chamber temperature To | = | 3000 K |

Combustion chamber volume V | = | 0.15 m ^{3} |

Throat area A· | = | 0.1 m ^{2} |

Molecular weight of propellant M | = | 10 |

Ratio of specific heats r | = | 1.2 |

Ambient pressure P_{a} | = | 0 |

(b) (i) How does the shape of the nozzle affect performance? How do you

overcome the thrust loss associated with over expansion?

(ii) Explain various methods of thrust vector control with sketches.

21. Differentiate between impulse stage and reaction stage turbines.

22. Define match point.

23. Write down the merits and demerits of integral ram-rocket.

24. What do you mean by supercritical mode of operation of ramjet?

25. Name any two oxidizer-fuel combinations used for hybrid rockets.

26. Define total-to-total efficiency and state when it is appropriate to use this efficiency.

27. An ideal ramjet engine operates at M = 1.5 at an altitude of 6500 m. Find its cycle efficiency.

28. How do you classify ramjets based on combustion process?

29. What are the limitations of hybrid rockets?

30. Define discharge correction factor. Can it be more than one? Justify your answer

**PART B – (5 X 16 = 80 marks)**

11. (a) (i) Describe the working of an axial flow turbine stage with a neat sketch and Draw the T-S diagram and velocity triangles.

(ii) Discuss the limiting factors in turbine design.

(OR)

(b) A mean-diameter design of a turbine stage having equal inlet and outlet velocities leads to the following data.

Mass flow m - 20 kg/s

Inlet temperature - 1000 k

Inlet pressure - 4.0 bar

Axial velocity (constant through stage) - 260 rps

Blade speed U - 360 rps

Nozzle efflux angle - 65

^{0} Stage exit swirl - 10

^{0}Determine the rotor blade angles, degree of reaction, temperature drop coefficient

*and*power output. Assuming a nozzle loss coefficient*A*of 0.05, calculate the nozzle throat area required (ignoring the effect of friction on the critical conditions).12. (a) . (i) Describe the working of a ramjet engine. Depict the Various

Thermodynamic processes occurring in it on h-s diagram.

(ii) Discuss the performances of Supersonic Combustion Ramjet.

Compare Subsonic and Supersonic combustion Ramjets.

(OR)

(b) A ramjet is traveling at Mach 3 at an altitude of 4572 m, the external static temperature is 258.4K, and the external static pressure is 57.1 kPa. The heating value of the fuel is 46,520 kJ/kg. Air flows through the engine at 45.35 kg/s. The burner exit total temperature is 1944K.Find the thrust, fuel ratio, and TSFC. The specific heat ratio can be assumed to be 1.4.

13. (a) A jet pressure is 8.5 kPa and the ambient temperature is 220 K. The turbine inlet temperature is 2540 engine is to propel an aircraft at Mach 3 at high altitude where ambient K. If all components of the engine are frictionless determine

(i) The thermal efficiency

(ii) The propulsion efficiency

(iii) The overall efficiency

Let the specific heat ratio be

*r*= 1.4 and make the approximation of*f«*1. (OR)

(b) (i) With a neat sketch explain the concept of integral ram-rocket and mention its advantages and disadvantages.

(ii) Briefly discuss performance of supersonic combustion ramjet and compare subsonic combustion Ramjet with supersonic combustion Ramjet engine.

14. (a) (i) What are the important factors that influence the burning rate of a

solid propellant? Explain them with appropriate sketches.

(ii) How do you classify solid propellant rockets? Name any four solid

propellant ingredients with two examples for each.

(OR)

(b) A chemical rocket is used for launch into earth orbit. At the end of the combustion chamber the stagnation temperature is 3000 K and the stagnation pressure is 2 MPa. The molecular weight of the combustion products is 26. The gases expand entropically as an ideal gas mixture with specific heat ratio 1.2. The area ratio

*A*/_{e}*A·*of the nozzle is 20, and the throat diameter is 0.1 m. At sea level, determine the rocket thrust.15. (a) How long would it take for a thrust of a rocket to diminish to 10% of its

steady value if the fuel and oxidant flows into the chamber were suddenly stopped? Consider, for example, the following conditions:

Initial combustion chamber pressure Po | = | 10 MPa |

Initial combustion chamber temperature To | = | 3000 K |

Combustion chamber volume V | = | 0.15 m ^{3} |

Throat area A· | = | 0.1 m ^{2} |

Molecular weight of propellant M | = | 10 |

Ratio of specific heats r | = | 1.2 |

Ambient pressure P_{a} | = | 0 |

(OR)

(b) (i) How does the shape of the nozzle affect performance? How do you

overcome the thrust loss associated with over expansion?

(ii) Explain various methods of thrust vector control with sketches.

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