Hall Ticket No Question Paper Code: AAE007
INSTITUTE OF AERONAUTICAL ENGINEERING
(Autonomous)
B.Tech V Semester End Examinations (Supplementary) January, 2019
Regulation: IARE R16
AIRCRAFT PROPULSION
Time: 3 Hours Max Marks: 70
Answer ONE Question from each Unit
All Questions Carry Equal Marks
All parts of the question must be answered in one place only
UNIT I
1. Construct a simple line diagram of gas turbine and explain with T-s diagram how it functions?

A gas turbine operating at a pressure ratio of 11.314 produces zero net work output when 473.35
kJ of heat is added per kg of air. If the inlet air temperature is 300 K and the turbine efficiency
if find the compressor efficiency?
2. Derive the thrust equation for gas turbine engine.
The effective jet exit velocity from a jet engine is 2700 m/s. the forward flight velocity is 1350
m/s and the air flow rate is78.6 kg/s. Calculate thrust, thrust power and propulsive efficiency.

UNIT II
3. Discuss the following:
Supersonic inlets
ii) Factors affecting diffuser performance.
Differentiate between internal compression and external compression in a supersonic inlet.
4. State the main factors which are effecting combustion chamber performance? and explain in
detail?
Differentiate between Can-type and Cannular-type combustor in gas turbine with a neat sketch?

UNIT III
5. Explain in detail about different operating conditions in CD nozzle.
Write brief notes on thrust vectoring and various methods of thrust vectoring.
Page 1 of 2
6. Derive the equation for nozzle efficiency and explain the losses in nozzle.
A turbojet engine powering an aircraft flying at an altitude of 11,000m where Ta 216.7 K and
Pa 24.444 kPa. The flight Mach number is 0.9. The inlet conditions to the nozzle are 1000
K and 60 kPa. The specific heat ratio of air and gases at nozzle are 1.4 and 4/3. The nozzle
efficiency is 0.98. Determine the thrust per inlet frontal area for C-D nozzle
UNIT IV
7. Explain the various components of typical centrifugal compressors with the help of a schematic
diagram. Discuss the actual pressure and velocity variations of flow across the impeller and
diffuser
A centrifugal compressor compresses 30kg of air per second at a rotational speed of 15000 rpm.
The air enters the compressor axially, and the conditions at the exit sections are radius =0.3m,
relative velocity of air at tip=100m/s at an angle of 800 with respect to the plane of rotation
take p01=1 bar and T01=300K. Find the torque and power required to drive the compressor and
also the ideal head developed.
8. Write short notes on performance characteristics of axial compressors
Explain the operating principle of centrifugal compressor with neat diagram.
UNIT V
9. Write short notes on work done and pressure rise by radial flow turbine and derive the equations.

Combustion gases enter the first stage of a gas turbine at a stagnation temperature and pressure
of 1200 K and 4.0 bar. The rotor blade tip diameter is 0.75m, the blade height is 0.12 m and the
shaft speed is 10,500 rpm. At the mean radius the stage operates with a reaction of a flow
coefficient of 0.7and a stage loading coefficient of 2.5. Determine the relative and absolute
flow angles for the stage; the velocity at nozzle exit; the static temperature and pressure
at nozzle exit assuming a nozzle efficiency of 0.96 and the mass flow
10. What is axial and radial flow turbine? What are the limitations of axial and radial flow turbine?

A single stage gas turbine operates at its design condition with an axial absolute flow at entry
and exit from the stage. The absolute flow angle at the nozzle exit is 70 deg. At stage entry, the
total pressure and temperature are 311 kPa and 8500C respectively. The exhaust static pressure
is 100 kPa, the total to static efficiency is 0.87 and mean blade speed is 500 m/s. Assuming
constant axial velocity through the stage, determine
the specific work done
(ii)the Mach number leaving the nozzle
iii)the axial velocity
total to total efficiency
stage reaction.