Question Bank
Unit-1
1. Define Froude efficiency, what is
its effect on
thrust?
2
2. Compare air breathing engine and
rocket
engine.
2
3. Define SFC.Write down its
significance.
2
4. Mention the factors affecting
thrust.
2
5.
Find the propulsive efficiency of a jet engine moving with 300 m/s at 7000m
altitude and its exhaust gas velocity is 600
m/s.
2
6. Define by pass
ratio.
2
7. Why rate of thrust for an air
breathing engine decreases with altitude and increases
for non air breathing
engine?
2
8. Differentiate between Scramjet
& Ramjet
engine.
2
9. Why is the ‘reverse diffuser’
impractical?
2
10. What are the advantages and
disadvantages of cooling gas turbine
blades?
2
11. Mention relative merits of jet
engines over piston
engines.
2
1.
An advanced fighter engine operating at Mach 0.8 and 10Km altitude where,
Ta=223.297K & Pa=0.2649 bar has the following uninstalled performance data
and uses a fuel with C.V= 42,800KJ/Kg:
Thrust
= 50 KN
Mass flow of air = 45Kg/s
Mass flow of fuel = 2.65 Kg/s
Determine the specific thrust,
thrust specific fuel consumption; exit velocity, thermal efficiency, propulsion
efficiency, and overall efficiency (assume exit pressure equal to ambient
pressure).
16
2.
Find specific thrust and SFC of a simple turbojet engine, having the following
component performance at which the cruise speed and altitude are M 0.8 and
10000m. Select ambient condition from the gas table.
Compressor pressure
ratio
8.0
Turbine inlet
temperature
1200K
Isentropic efficiency:
Of compressor ηc
0.87
Of turbine ηt
0.90
Of intake ηi
0.93
Of propelling nozzle
ηj
0.95
Mechanical transmission efficiency
ηm 0.99
Combustion efficiency
ηb
0.98
Combustion chamber pressure loss
ΔPb 4% of compressor outlet
pressure.
C.V of fuel is 43,000 KJ/Kg, assume
data if necessary, Cpa ≠ Cpg
16
3. (a) Explain with neat sketch
operating principles of turbofan
engine
8
(b) What is thrust augmentation? Explain any two methods of thrust augmentation
with
sketches.
8
4.
Compare the characteristics, advantages & disadvantages of turbojet,
turbofan and turboprop engine.
5.
(i)Discuss the different methods of thrust augmentation. Draw T-S diagram for
turbojet engine with thrust
augmentation.
8
(ii) Discuss the
typical turbojet cycle performance with suitable
sketches.
8
6. A turbojet engine is traveling at
270 m/s at an altitude of 5000m. The compressor pressure ratio is 8:1 and
maximum cycle temperature is 1200K. By assuming the following data,
Ram
efficiency
93%
Isentropic efficiency of
compressor
87%
Pressure loss in combustion
chamber
4%of compressor delivery
pressure
Calorific value of
fuel
43,100 kj/kg
Combustion
efficiency
98%
Mechanical transmission
efficiency
99%
Isentropic efficiency of
turbine
90%
Propelling nozzle
efficiency
95%
Ambient conditions at 5000 m are
0.5405 bar and 255.7 K.
Calculate the
(i)
Specific thrust and
(ii)
TSFC
16
7. (i) Define thrust of an
engine and derive the thrust equation for a general propulsion
system.
8
(ii)
Discuss the typical turbojet cycle performance with suitable
sketches.
8
8. An ideal turbojet flies at sea
level at a Mach number of 0.75. It ingests 74.83 kg/s of air, and the
compressor operates with a total pressure ratio of 15. The fuel has a heating
value of 41,000 kj/kg, and the burner exit total temperature is 1389 K. Find
the thrust developed and the TSFC. Assume that the specific heat ratio is 1.4.
16
9. Air enters a turbojet engine at a
rate of 12*104 kg/h at 150C &1.03 bar and is
compressed adiabatically to 1820C & four times the pressure.
Products of combustion enter the turbine at 8150C & leave it at
6500C to enter the nozzle. Calculate the isentropic efficiency of
the compressor, the power required to drive the compressor, the exit speed of
gasses & thrust developed when flying at 800 km/h. Assume the isentropic
efficiency of the turbine is same as that of the compressor and the nozzle
efficiency is 90%.Assume the data required
suitably.
16
10. A jet propelled plane consuming
air at the rate of 18.2 kg/s is to fly at Mach number of 0.6 at an altitude of
4500m (Pa = 0.55 bar, Ta = 255K ). The diffuser which has a pressure
coefficient of 0.9, decreases the flow to a negligible velocity. The compressor
pressure ratio is 5 & maximum temperature in the combustion chamber is 1273
K. After expanding in the turbine, the gases continue to expand in the nozzle
to a pressure of 0.69 bar. The isentropic efficiency of compressor, turbine and
nozzle are 0.81, 0.85 & 0.915 respectively. The heating value of the fuel
is 45870 kj/kg. Assume Cp = 1.005 kj/kg-K, Cpg = 1.147 kj/kg-K. Calculate
(i)
Power input to the compressor
(ii)
Power output of the turbine
(iii)
The fuel air ratio
(iv)
The thrust provided by the engine
(v)
The thrust power
developed.
16
unit-ii
1. What are the requirements of an aircraft
intake?
2
2. Write notes on pressure recovery
factor of the
intake?
2
3. What are the starting problems in
supersonic
inlets?
2
4. What are the factors to be
considered while designing a subsonic
inlet?
2
5. What are the factors to be
considered while designing a supersonic
inlet?
2
6. What is meant by sub critical
mode of inlet operation? State its advantages and
disadvantages.
2
1. (i) Explain successive
steps in the acceleration and over speeding of a one-
dimensional
supersonic inlet with sketches.
8
(ii)
Derive the relation between area ratio Amax/Ai and external deceleration ratio
ui/ua.
8
2. A supersonic inlet is designed
with a two-dimensional conical spike (with two half-cone angles 100
and 200 relative to the axial centerline, respectively). The inlet
is to operate at a flight Mach number of 1.9.The two standing oblique shocks
are attached to the spike and cowl, and a converging inlet section with a
throat of area A* is used to decelerate the flow through internal compression.
Assume γ = 1.4 and internal diffuser pressure recover factor Πr = 0.97.
Estimate the overall recovery factor Πd on the assumption that the
inlet starts (i.e., the normal shock is swallowed). Also, find the required
A*/A1.
3. What are the different modes of
inlet operation? Explain with suitable sketches. 16
4. Air enters a two-dimensional
supersonic diffuser at a pressure of 14.102 kPa, a temperature of 217 K, and
with a Mach number of 3.0. The two-dimensional oblique shock diffuser has an
oblique shock angle of 27.80, which is followed by a normal shock.
Determine, assuming constant specific heats.
(i)
The velocity, total temperature and
pressure of the air entering the oblique shock.
(ii)
The Mach number, total pressure
after the oblique shock.
(iii)
The flow deflection angle.
(iv)
The Mach number, total and static
pressure and static temperature after the normal shock.
unit-iii
1. What is need for supersonic
combustion?
2
2. Define equivalence ratio and
stochiometric fuel air
ratio.
2
3. Define efficiency of the
combustor.
2
4. What is the purpose of primary
air in combustion
chamber?
2
5. What is the purpose of secondary
air in combustion
chamber?
2
6. What is the purpose of dilution
air in combustion
chamber?
2
7. Define combustion
intensity?
2
8. State the advantages and
disadvantages of annular
combustor.
2
1. (a) What are the important
factors affecting combustor
design?
8
(b)Write
down the methods of flame stabilization and explain with
sketch.
8
2. (a)What are the three types of combustion
chamber? Compare its advantages and
disadvantages.
8
(b) Name the
material used for combustion chamber and discuss the special qualities of the
material used for combustion
chamber?
8
3. (a)What are the factors affecting
combustion chamber? Explain
briefly?
8
(b) With the aid of a simplified
picture explain the operation of a flame holder.
8
4. (i) With a neat sketch explain
the working of a combustion
chamber.
8
(ii) Consider n-decane fuel,
balance the chemical equation for the stoichiometric combustion of this fuel in
air and find the stoichiometric fuel-to-air
ratio.
8
unit-iv
1. What is choked
nozzle?
2
2. Is it possible to have over
expanded jets in convergent nozzle? Justify your
answer.
2
3. Give any four functions of an
exhaust
nozzle.
2
1. (a) Plot Mach number, static
temperature, static pressure and static density variations along the
longitudinal axis of a convergent-divergent nozzle, when it flows full. Explain
the
variations.
8
(b)A De Laval nozzle has to
be designed for an exit Mach number of 1.5 with exit diameter of 200 mm. Find
the ratio of throat area/exit area necessary. The reservoir conditions are
given as Po = 106 Pa, To = 200C. Find also the maximum mass flow
rate through the nozzle. What will be the exit pressure and
temperature?
8
2. A converging-diverging is
designed to operate with an exit Mach number of 1.75. The nozzle is supplied
from an air reservoir at 68bar (abs.). Assuming 1-d flow, calculate:
(i) Maximum backpressure to choke
the
nozzle.
4
(ii) Range of backpressure over
which a normal shock will appear in the nozzle.
4
(iii) Back pressure for the nozzle
to be perfectly expanded to design
M.
4
(iv) Range of back pressure for
supersonic flow at the nozzle exit
plane.
4
3. (i) What are the types of nozzle?
Explain various operating conditions of a C-D nozzle with suitable
sketch.
8
(ii) Write short notes
on the following:
(a) Ejector and variable area
nozzles
4
(b) Thrust reversing
4
4. An exhaust air stream at Mach
2.9, pressure 68.95kPa, and temperature 777.8 K enters a frictionless diverging
nozzle with a ratio of exit area to inlet area of 3.0. Determine the back
pressure necessary to produce a normal shock in the channel at an area equal to
twice the inlet area. Assume one-dimensional steady flow with the air behaving
as a perfect gas with constant specific heats and a specific heat ratio of
1.36; assume isentropic flow except for the normal
shock.
16
unit-v
1. Write down the difference between centrifugal and axial flow
compressors.
2
2. Define degree of reaction for an
axial flow
compressor.
2
3. Define rotating stall for
compressors.
2
4. What are the causes for stalling
in axial flow
compressors?
2
5. Define slip
factor.
2
6. Write down the conditions for
free and forced vortex
flows.
2
7. Distinguish between surging and
stalling.
2
1. An axial compressor stage has a
mean diameter of 60cm and runs at 15000rpm. If the actual temperature rise and
pressure ratio developed are 300C and 1.4 respectively.
(i)
The power required to drive the
compressor while delivering 57 Kg/s of air; assume mechanical efficiency of 86
% and an initial temperature of 350C.
(ii)
The stage loading coefficient.
(iii)
The stage efficiency and
(iv)
The degree of reaction if the
temperature at the rotor exits is 550C.
2. (i) Explain the working of a
centrifugal compressor and draw the velocity
triangles.
8
(ii) A
centrifugal compressor has an impeller tip speed of 366 m/s. Determine the
absolute Mach number of the flow leaving the radial vanes of the impeller when
the radial component of velocity at impeller exit is 30.5 m/s and the slip
factor is 0.9. Given that the flow area at impeller exit is 0.1m2
and the total-to-total efficiency of the impeller is 90%, determine the mass
flow
rate.
8
3. (i) A sixteen-stage axial flow
compressor is to have a pressure ratio of 6.3. Tests have shown that a stage
total-to-total efficiency of 0.9 can be obtained for each of the first six
stages and 0.89 for each of the remaining ten stages. Assuming constant work
done in each stage and similar stages fine the compressor overall total-to
–total efficiency. For a mass flow rate of 40 kg/s determine the power required
by the compressor. Assume an inlet total temperature of 288
K.
8
(ii) Discuss the factors
affecting stage pressure rise of an axial flow compressor with suitable
sketches.
8
4. A stage of a radial compressor is
to be analyzed. It rotates at 12,300 rpm and compresses 31.75 kg/s of air. The
inlet pressure and temperature are 241.325 kPa and 306K respectively. The hub
and tip radii of the blades at the inlet are 7.62 and 13.97cm respectively. The
exit radius is 27.94cm and the exit blade height is 2.54cm. The slip factor is
unity. Flow enters the inducer with no prewhirl and the impeller has straight
radial blades. The efficiency of the stage is 88%. The value of Cp and γ are
1.005 kj/kg-K and 1.397 respectively.
Find the following:
(i)
Mean relative flow angle at the
inlet.
(ii)
The static pressure at the impeller
exit.
(iii)
The total pressure ratio for the
stage,
(iv)
The Mach numbers at the impeller
inlet and exit.
(v)
The required power for the
stage.
16
5. An axial flow compressor stage is
designed to give free-vortex tangential velocity distributions for all radii
before and after the rotor blade row. The tip diameter is constant and 1.0m;
the hub diameter is 0.9m and constant for the stage. at the rotor tip the flow
angles are as
follows:
16
Absolute inlet angle, α1
= 300
Relative inlet angle, β1
= 600
Absolute outlet angle, α2
= 600
Relative outlet angle, β2
= 300
.Determine,
(i)
the axial velocity
(ii)
the mass flow rate
(iii)
the power absorbed by the stage
(iv)
the flow angles at the hub
(v)
the reaction ratio of the state at
the hub
Given that the rotational speed of the rotor is 6000 rpm and the gas density is
1.5 kg/m3 which can be assumed constant for the stage. It can be
further assumed that stagnation enthalpy and entropy are constant and after the
rotor row.
6. The mass flow rate of flow at 288
K and 101.3 KPa at the inlet to the impeller of the centrifugal-flow compressor
is 1.814 kg/s. The inlet flow is in the axial direction. The impeller eye has
the minimum diameter of 3.81cm and a maximum diameter of 12.7cm and rotates at
35,000rpm. Assuming no blockage due to the blade, calculate the ideal angle at
the hub and tip at the inlet to the impeller. Draw velocity diagram at the hub
and at the
tip.
16
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