
SRI VIDYA COLLEGE OF ENGINEERING
& TECHNOLOGY
Question Bank


DEPARTMENT
OFMECHANICAL ENGINEERING
Academic
year : 20162017 (Even) Year/section : III/A&B
Subject code/Name: ME6604/Gas
Dynamics and Jet Propulsion
Unit I – Basic concepts and isentropic flows
Part A
1. State the difference between compressible fluid and
incompressible fluid?
2. Define stagnation pressure?
3. Express the stagnation enthalpy in terms of static enthalpy
and velocity of flow?
4. Explain Mach cone and Mach angle?
5. Define adiabatic process?
6. Define Mach number?
7. Define zone of action and zone of silence?
8. Define closed and open system?
9. What is the difference between intensive and extensive
properties?
10. Distinguish between Mach wave and normal shock?
Part B
1. What is the effect of Mach
number on compressibility?
Prove for Î³=1.4, Po –P / ½ P c² = 1 +. M2 +
1/40 M4 + …….
2.Air at P1=3 bar and T1=227 C is flowing with a
velocity of 200 m/s in a 0.3m diameter duct. If Cp=1050 J/KgK and Î½=1.38.
Determine the following (i) Stagnation temperature and pressure(ii) Mass flow
rate of air (iii) Mach number (iv) Stagnation pressure.

3. Derive the expression for an energy equation.
4.Air flowing in a duct has a
velocity of 300 m/s, pressure 1.0 bar and temperature 290 k. Taking Î³ = 1.4 and
R = 287 J/Kg K. Determine: i)
Stagnation pressure and temperature. ii) Velocity of sound in the dynamic and
stagnation conditions. ii) Stagnation pressure assuming constant density.
5. Air (Î³ = 1.4, R = 287.43 J/Kg K)
enters a straight axisymmetric duct at 300 K, 3.45 bar and 150 m/s and leaves
it at 277 k, 2.058 bar and 260 m/s. The area of crosssection at entry is
500cm2. Assuming adiabatic flow determine i) Stagnation temperature ii) maximum
velocity iii) Mass flow rate iv) Area of cross section at exit.
OR
1.
A supersonic diffuser diffuses air in an
isentropic flow from a mach number of 3 to a mach
number of 1.5. The static conditions
of air at inlet are 70 kpa and 7 C. If the mass flow rate of air is
125 kg/s, determine the stagnation
conditions, areas at throat and exit, static conditions (pressure,
temperature, velocity) of air at exit.

2.
A supersonic nozzle expands air from Po = 25 bar
and T0 = 1050 K to an exit pressure
4.35 bar the exit
are of the nozzle is 100 cm2. Determine i) throat area ii) pressure and
temperature at the throat iii)
temperature at exit iv) Exit velocity as fraction of the maximum
attainable velocity v) mass flow rate.

3.
A conical diffuser has entry and exit area of 0.11
m^{2 }and 0.44 m^{2} respectively. The pressure,
temperature
and velocity of air at entry are 0.18 Mpa, 37 ‘C and 267 m/s
respectively. Determine
i)
mass flow rate
ii) the mach number, static temperature and static pressure of the air
leaving
diffuser and iii) the net thrust acting on the
diffuser
4.
In an isentropic flow diffuser the inlet area is 0.15 m2. At the
inlet velocity 240m/s, static
temperature = 300 k and static
pressure 0.7 bar. Air leaves he diffuser with a velocity of
120 m/s. Calculate at the exit
the mass flow rate, stagnation pressure, stagnation temperature
area and entropy change across the diffuser.

Unit II Flow through ducts
Part A
1. What are the consumption made for fanno
flow?
2. Differentiate Fanno flow and Rayleigh
flow?
3. Explain chocking in Fanno flow?
4. Explain the difference between Fanno
flow and Isothermal flow?
5. Write down the ratio of velocities
between any two sections in terms of their Mach number in a fanno flow ?
6. Write down the ratio of density between
any two section in terms of their Mach number in a fanno flow?
7. What are the three equation governing
Fanno flow?
8. Give the expression to find increase in
entropy for Fanno flow?
9. Give two practical examples where the
Fanno flow occurs?
10. What is Rayleigh line and Fanno line?
Part B
1.Air having mach number 3 with
total temperature 295 C and static pressure 0.5 bar flows through a constant
are duct adiabatically to another section where the mach number is 1.5.
Determine the amount of heat transfer and the change in stagnation pressure
2. Air flow through a constant area
duct with inlet temperature of 20 C and inlet Mach number of 0.5. What is the
possible exit stagnation temperature? It is desired to transfer heat such that at
exit of the duct the stagnation temperature is 1180 K. For this condition what
must be the limiting inlet Mach number? Neglect
friction.
3.Air enters a combustion chamber
with certain Mach number. Sufficient heat is added to obtain a stagnation
temperature ratio of 3 and a final Mach number of 0.8. Determine the Mach number
at entry and the percentage loss in static pressure. Take Î³ = 1.4 and Cp =
1.005 Kj/KgK
4.The mach number at inlet and exit
for a Rayleigh flow are 3 and 1.5 respectively. At inlet static pressure is 50
kPa and stagnation temperature is 295 K. Consider the fluid is air. Find i) the
static pressure, temperature and velocity at exit, ii) stagnation pressure at
inlet and exit, iii) heat transferred, iv) maximum possible heat
transfer, v) change in entropy
between the two sections, vi) is it a cooling or heating process?
5.The mach number at the exit of a
combustion chamber is 0.9. the ratio of stagnation temperatures at exit and
entry is 3.74. If the pressure and temperature of the gas at exit are 2.5 bar
and 1273 K respectively, determine: i) Mach number, pressure and temperature of
the gas at entry ii) the heat supplied per Kg of the gas and iii) the maximum
heat that can be supplied.
6.A combustion chamber in a gas
turbine plant receives air at 350 k, 0.55 bar and 75m/s. The air fuel ratio is
29 and the calorific value of the fuel is 41.87MJ/Kg. Taking Î³ = 1.4 and R=
0.287 KJ/Kg K for the gas determine: i) the initial and final mach numbers ii)
final pressure, temperature and velocity of the gas. iii) percent stagnation
pressure loss in the combustion chamber and iv) the maximum stagnation
temperature attainable.
OR
1.Air at Po = 10 bar, To = 400 K is
supplied to a 50 mm diameter pipe. The friction factor for the pipe surface is
0.002. If the Mach number changes from 3.0 at the entry to 1.0 at the exit
determine i) the length of the pipe and ii) the mass flow rate.
2.A circular duct passes 8.25 kg/s
of air at an exit Mach number of 0.5. The entry pressure and temperature are
3.45 bar and 38 C respectively and the coefficient of friction is 0.005. If the
Mach number at entry is 0.15, determine the diameter of the duct, length of the
duct, pressure and temperature at the exit, and stagnation pressure loss.
3.Air at an inlet temperature of 60
C flows with subsonic velocity through an insulated pipe having inside diameter
of 50 mm and a length of 5 m. The pressure at the exit of the pipe is 101 kPa and
the flow is choked at the end of the pipe. If the friction factor 4f = 0.005.
determine the inlet Mach number, the mass flow rate and the exit temperature.
4.A long pipe of 0.0254 m diameter
has a mean coefficient of friction of 0.003. Air enters the pipe at a mach
number of 2.5, stagnation temperature 310 K and static pressure 0.507 bar.
Determine for a section at which the mach number reaches 1.2: i) Static
pressure and temperature, ii) Stagnation pressure and temperature, iii)
Velocity of air, iv) Distance of this section from the inlet and v) mass flow
rate of air.
5. A long pipe of 0.0254 m diameter has a
mean coefficient of friction of 0.003. Air enters the pipe at a mach number of
2.5, stagnation temperature 310 K and static pressure 0.507 bar. Determine for
a section at which the mach number reaches 1.2: i) Static pressure and
temperature, ii) Stagnation pressure and temperature, iii) Velocity of air, iv)
Distance of this section from the inlet and v) mass flow rate of air.
6.Air at an inlet temperature of 60
C flows with subsonic velocity through an insulated pipe having inside diameter
of 50 mm and a length of 5 m. The pressure at the exit of the pipe is 101 kPa and
the flow is choked at the end of the pipe. If the friction factor 4f = 0.005.
determine the inlet Mach number, the mass
flow rate and the exit temperature.
Unit III – Normal and oblique shocks
Part A
1. What is mean by shock wave?
2. What is mean by Normal shock?
3. What is oblique shock?
4. Define strength of shock wave?
5. What are applications of moving shock
wave?
6. Shock waves cannot develop in subsonic
flow? Why?
7. Define compression and rarefaction
shock? Is the latter possible?
8. State the necessary conditions for a
normal shock to occur in compressible flow?
9. Give the difference between normal and
oblique shock?
10. what are the properties change across a
normal shock ?
Part B
1. Derive the equation for static pressure
ratio across the shock waves.
2. Derive PrandtlMeyer relation.
3. A supersonic nozzle is provided with a
constant diameter circular duct at its exit. The duct diameter is same as the
nozzle exit diameter. Nozzle exit cross section is three times that of its throat.
The entry conditions of the gas (Î³ = 1.4, R = 0.287kJ/kgk) are Po = 10 bar, To
= 600 K. Calculate the static pressure, Mach number and the velocity of the gas
in the duct: i) when the nozzle operates at this design condition ii) when a
normal shock
occurs at this design condition.
ii) when a normal shock occurs at its exit.
4. A convergentdivergent nozzle is
designed to expand air from a reservoir in which the pressure is 800 kpa and
temperature is 40 C to give a mach number at exit of 2.5. the throat area is 25=
cm2. Find i) mass flow rate, ii) exit area and iii) when a normal shock appears
at a section where the area is 40 cm2 determine
the pressure and temperature at
exit.
5. A pilot tube kept in a supersonic wind
tunnel forms a bow shock a head of it. The static pressure upstream of the
shock is 16 kPa and the pressure at the mouth is 70 kPa. Estimate the mach number
of the tunnel. If the stagnation temperature is 300 C, calculate the static
temperature and total pressure upstream and downstream of the tube.
OR
1.A gas (Î³ = 1.3) at P1 = 345 mbar, T1 =
350 K and M1 = 1.5 is to be isentropically expanded to 138 mbar. Determine i)
Deflection angle ii) Final Mach number and iii) the temperature of the gas
2. Air approaches a symmetrical wedge (angle
of deflection Î´= 15') at a Mach number of 2. Consider strong waves conditions. Determine
the wave angle, pressure ratio, density ratio, temperature ratio and downstream
Mach number.
3. A jet of air at a mach number of 2.5 is
deflected inwards at the corner of a curved wall. The wave angle at the corner
is 60o. Determine the deflection angle on the wall, pressure and temperature
ratios and final Mach number.
4. A convergent divergent nozzle operates at
off design condition while conducting air from a high pressure tank to a large container.
A normal shock occurs in the divergent part of the nozzle at a section where
the cross section area is 18.75 cm2. The stagnation pressure and stagnation
temperature at the inlet of the nozzle are 0.21 Mpa and 360 C respectively. The
throat area is 12.5 cm2 and the exit area is 25 cm2. Estimate the exit mach
number, exit pressure, loss in stagnation pressure and entropy increase during
the flow between the tanks.
5. A convergentdivergent nozzle has an exit
area to throat area ratio of 2. Air enters this nozzle with a stagnation
pressure of 1000 kPa and a stagnation temperature of 360 K. the throat area is
500 mm2. The divergent section of the nozzle acts as a supersonic nozzle.
Assume that a normal shock stands at a point
M = 1.5. Determine the exit plane
of the nozzle, the static pressure and temperature and Mach number.
Unit IV Jet propulsion
Part A
1. What is thrust (or) drag?
2. What is Thrust Specific Fuel Consumption
(TSFC)?
3. Define Specific impulse
4. What are the various types of air
breathing engine?
5. What is scram jet?
6. How is turbofan engine different from
turbo prop engine?
7. What is thrust augmentation?
8. Give the difference between Ramjet and
Turbojet engine
9. What is the difference between turboprop
and turbojet engine
10. What type of compressor used in
turbojet? Why?
Part B
1.Explain with a neat sketch the
principle of operation of a turbojet engine and state its advantages and
disadvantages.
2.Differentiate turbojet and
turboprop propulsion engines with suitable diagrams
3. A turbojet engine operating at a
Mach number of 0.8 and the altitude is 10Km has the following data. Calorific
value of the fuel is 42,899 kJ/Kg. thrust force is 50 kN, mass flow rate of air
is 45 kg/s, mass flow rate of fuel is 2.65 kg/s. determine the specific thrust,
thrust specific fuel consumption, jet velocity, thermal efficiency, propulsion
efficiency and overall efficiency. Assuming the exit pressure is equal to
ambient pressure.
4. A turbojet aircraft flies at 875
Kmph at an attitude of 10,000 m above mean sea level. Calculate i) air flow
rate through the engine, ii) thrust, iii) specific thrust, iv) specific impulse
v) thrust power and TSFC from the following data: Diameter of the air at inlet
section = 0.75m Diameter of jet pipe at exit = 0.5m Velocity of the gases at
the exit of the jet pipe = 500m/s Pressure at the exit of the jet pipe = 0.30
bar Air to fuel ratio = 40
5.A turbojet propels an aircraft at
a speed of 900 km/hr, while taking 3000 kg of air per minute. The isentropic
enthalpy drop in the nozzle is 200 kJ/kg and the nozzle efficiency is 90%. The
airfuel ratio is 85 and the combustion efficiency is 95%. The calorific value
of the fuel is 42,000 kJ/Kg. Calculate: i) The propulsion power, ii) Thrust
power, iii) Thermal efficiency and iv) Propulsion efficiency
OR
1.Explain with a neat sketch the
principle of operation of a turboprop engine and state its advantages and
disadvantages.
2.Write the equations to calculate
propulsion efficiency and thermal efficiency of an aircraft.
3.Explain with a neat sketch the
principle of operation of a ramjet engine and state its advantages and
disadvantages.
4.Describe the working of turbofan
engine with a neat sketch. List out its advantages and disadvantages.
5.The diameter of the propeller of
an aircraft is 2.5m; it flies at a speed of 500 km/hr at an altitude of 8000 m.
For a flight to jet speed ratio of 0.75, determine: the flow rate of air
through the propeller, thrust produced, specific thrust, specific impulse and thrust
power.
Unit V Space propulsion
Part A
1. Differentiate jet propulsion and
Rocket propulsion.
2. What is mono propellant
3. What is bi propellant
4. Classify the rocket engines
based on source of energy employed
5. What is specific impulse of a
rocket?
6. Define thrust
7. What is IWR?
8. What is thrust coefficient?
9. Define propulsion efficiency
10. What is weight flow
coefficient?
PartB
1.Explain the principle of
operation of liquid propellant and solid propellant engines with neat sketch.
2.Explain briefly about the
propellant feed system of a liquid propellant rocket engine with suitable
schematic sketches.
3. Explain with a neat sketch the
working of a gas pressure feed system used in liquid propellant rocket engines
4.Explain the working of a
turbopump feed system used in a liquid propellant rocket
5.Deduce expressions for propulsion
efficiency specific impulse and overall efficiency of a rocket engine.
OR
1.A rocket has the following data: propellant
flow rate = 5 Kg/s, Nozzle exit diameter = 10 cm, Nozzle exit pressure = 1.02
bar, Ambient pressure = 1.013 bar, Thrust chamber pressure = 20 bar, Thrust = 7
KN. Determine the effective jet velocity, actual jet velocity, specific impulse
and the specific propellant consumption. Recalculate the values of thrust and
specific impulse for an altitude where the ambient pressure is 10 m bar.
2.Describe the important properties
of liquid and solid propellants desired for rocket propulsion.
3.The effective jet velocity from a
rocket is 2700 m/s. The forward flight velocity is 1350 m/s and the propellant consumption
is 78.6 kg/s. Calculate: thrust, Thrust power and propulsion efficiency.
4.A rocket engine has the following
data. Combustion chamber pressure is 38 bar, combustion chamber temperature is
3500 K, oxidizer flow rate is 41.67 Kg/s, mixture ratio is 5, and the
properties of exhaust gases are Cp/Cv = 1.3 and R = 0.287 kJ/KgK. The expansion
takes place to the ambient pressure of 0.0582 bar. Calculate the nozzle throat
area, thrust, thrust coefficient, exit velocity of the exhaust and maximum
possible exhaust velocity.
5. Write short notes on space
flights.