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 a2
Stage exit swirl a3
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/U2) 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 Ae / 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.To
|
150
K
|
Isentropic
efficiency
|
TJi
|
0.9
|
Mean blade
speed
|
Um
|
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
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
|
(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.
RAJALAKSHMI ENGINEERING COLLEGE
AE2304 PROPULSION-II
PART
A – (10 X 2 = 20 marks)
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.
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