AERO PROPULSION - II Question Paper

PART B - (5 x 16 = 80 marks)               Download as pdf

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.To 150 K Isentropic efficiency TJi 0.9 Mean blade speed Um 320 rnfs Rotational speed 250 rev/s Outlet velocity
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 Ae / 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 m3
Throat area A·                                              Download previous year question paper
=
0.1 m2
Molecular weight of propellant M
 =
10
Ratio of specific heats r
=
1.2
Ambient pressure Pa
=
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                                         - 650
      Stage exit swirl                                               - 100

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 Ae / 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 m3
Throat area A·
=
0.1 m2
Molecular weight of propellant M
 =
10
Ratio of specific heats r
=
1.2
Ambient pressure Pa
=
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.

TERIMA KASIH ATAS KUNJUNGAN SAUDARA
Semoga artikel ini bermanfaat bagi saudara. Jika ingin mengutip, baik itu sebagian atau keseluruhan dari isi artikel ini harap menyertakan link dofollow ke https://androidjones7.blogspot.com/2012/03/aero-propulsion-ii-question-paper.html. Terima kasih sudah singgah membaca artikel ini.