Anna University, Chennai
SRINIVASAN ENGINEERING COLLEGE, PERAMBALUR
DEPARTMENT OF MECHANICAL ENGINEERING
QUESTION BANK
Sub. Code/Name: ME1303 Gas Dynamics and Jet Propulsion Year/Sem: III/V
UNIT1
PART A (2 Marks)
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 (16 Marks)
1) Derive the energy equations
a² /γ 1 +½ c² = ½ c² max =ao² /γ 1 =ho

2.) The pressure, temperature and Mach number at the entry of a flow passage are 2.45 bar 26.5˚ C
and 1.4 respectively. If the exit Mach number is 2.5 determine for adiabatic flow
of perfect gas (γ =1.3, R=0.469 KJ/Kg K). (16)
3) Air (γ =1.4, R=287.43 J/Kg K) enters a straight axis symmetric duct at 300 K, 3.45 bar and
150 m/s and leaves it at 277 k, 500cm². Assuming adiabatic flow determines:
1.  stagnation temperature,  (4) 
2.  maximum velocity,  (4) 
3.  mass flow rate, and,  (4) 
4.  Area of crosssection at exit.  (4) 
4) An aircraft flies at 800 Km/hr at an altitude of 10,000 meters (T=223.15 K, P=0.264 bar). The air is reversibly compressed in an inlet diffuser. If the Mach number at the exit of the diffuser is
0.36 determine (a) entry Mach number and (b) velocity, pressure and temperature of air at
diffuser exit. (16)
5) Air (Cp =1.05 KJ/Kg K, γ =1.38) at p1 =3x 10 5 N/m² and T1 =500 k flows with a velocity of  
200 m/s in a 30 cm diameter duct. Calculate mass flow rate, stagnation temperature, Mach number, and Stagnation pressure values assuming the flow as compressible and  
incompressible.  (16) 
6) (a) What is the effect of Mach number on compressibility prove for
γ=1.4, Þo –þ / ½ þ c² = 1 +¼ M² + 1/40 M 4 + ……. (8) (b) Show that for sonic flow the deviation between the compressible and incompressible
flow values of the pressure coefficient of a percent gas (γ =1.4) is about 27.5 per cent. (8)
7) Air at stagnation condition has a temperature of 800 K. Determine the stagnation velocity of
Sound and the maximum possible fluid velocity. What is the velocity of the sound when
the flow velocity is at half the maximum velocity (16)
8) Air flow through a duct. The pressure and temperature at station one are pressure is0 .7 bar and temperature is 300C. At a second station the pressure is 0.5 bar. Calculate temperature
and density at the second station. Assume the flow is to be Isentropic (16)
Unit 2
PART A (2 Marks)
1) Differentiate Adiabatic and Isentropic process.
2) Differentiate nozzle and diffuser?
3) What is Impulse function?
4) Differentiate between adiabatic flow and adiabatic flow?
5) State the expression for dA/A as a function of Mach number?
6) Give the expression for T/To and T/T* for isentropic flow through variable area in terms of
Mach number?
7) Draw the variation of Mach number along the length of a convergent divergent duct when it acts as a (a) Nozzle (b) Diffuser (c) Venturi
8) What is chocked flow through a nozzle?
9) What type of nozzle used for sonic flow and supersonic flow?
10) When does the maximum mass flow occur for an isentropic flow with variable area?
Part  B (16 Marks)

2) A conical diffuser has entry and exit diameters of 15 cm and 30cm respectively.
.The pressure, temperature and velocity of air at entry are 0.69bar,340 k and
180 m/s respectively. Determine
1)  The exit pressure,  (4) 
2)  The exit velocity and  (6) 
3)  The force exerted on the diffuser walls.  (6) 
Assume isentropic flow, γ =1.4, Cp =1.00 KJ KgK.
3) A nozzle in a wind tunnel gives a test –section Mach number of 2.0 .Air enters the nozzle from a large reservoir at 0.69 bar and 310 k .The cross –sectional area of the throat is
1000cm².Determine the following quantities for the tunnel for one dimensional isentropic flow
1) Pressures, temperature and velocities at the throat and test sections, (4)
2) Area of cross sectional of the test section , (4)
3) Mass flow rate, (4)
4) Power rate required to drive the compressor.
4) Air is discharged from a reservoir at Po =6.91bar and To =325˚c through a nozzle to an exit pressure of 0.98 bar .If the flow rate is 3600Kg/hr determine for isentropic flow:
(4)
5) A supersonic wind tunnel settling chamber expands air or Freon21 through a nozzle from a nozzle from a pressure of 10 bar to 4bar in the test section. Calculate the stagnation temperature to the maintained in the setting chamber to obtain a velocity of
500 m/s in the test section for Air, Cp =1.025 KJ/Kg K, Cv =0.735 KJ/Kg K, Freon 21, Cp =0.785 KJ/Kg K, Cv= 0.675 KJ/Kg K.
What is the test section Mach number in each case? (16)
6) Derive the following relations for one dimensional isentropic flow:
dA/A =dP/þ c²(1 M²) (8)
p*/p =(2/γ+1 +γ1 /γ+1M²) (8)
7) Air flowing in a duct has a velocity of 300 m/s, pressure 1.0 bar and temperature
290 k.
Taking γ=1.4 and R =287J/Kg K determines:
1)Stagnation pressure and temperature,  (4) 
2)Velocity of sound in the dynamic and stagnation conditions  (6) 
3) Stagnation pressure assuming constant density.  (6) 
8) A conical diffuser has entry and exit diameters of 15 cm and 30cm respectively
The pressure ,temperature and velocity of air at entry are 0.69bar,340 k and
180 m/s respectively. Determine
1) The exit pressure, (4)
2)The exit velocity and (6)
3) The force exerted on the diffuser walls. Assume isentropic flow, γ =1.4,Cp =1.00 KJ KgK.
(6)
9) A nozzle in a wind tunnel gives a test –section Mach number of 2.0 .Air enters the nozzle
from a large reservoir at 0.69 bar and 310 k .The cross –sectional area of the throat is 1000cm².
Determine the following quantities for the tunnel for one dimensional isentropic flow:
1)Pressures, temperature and velocities at the throat and test sections,  (4) 
2)Area of cross sectional of the test section ,  (4) 
3)Mass flow rate,  (4) 
4) Power rate required to drive the compressor.  (4) 
10) Air is discharged from a reservoir at Po =6.91bar and To =325˚c through a nozzle to an exit pressure of 0.98 bar .If the flow rate is 3600Kg/hr determine for isentropic flow:
11) A supersonic wind tunnel settling chamber expands air or Freon21 through a nozzle
from a nozzle from a pressure of 10 bar to 4bar in the test section. Calculate the stagnation
temperature to the maintained in the setting chamber to obtain a velocity of 500 m/s in the test section for, Air, Cp =1.025 KJ/Kg K, Cv =0.735 KJ/Kg K, Freon 21, Cp =0.785 KJ/Kg K Cv= 0.675 KJ/Kg K.
What is the test section Mach number is each case?  (16) 
12) Derive the following relations for one dimensional isentropic flow: dA/A =dP/þ c²(1 M²)  (8) 
p*/p =(2/γ+1 +γ1 /γ+1M²)  (8) 
Unit 3
PARTA (2 Marks)
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 sections 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 (16MARKS)
1) A circular duct passes 8.25Kg/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 0.005.If the Mach number at entry is 0.15,
Determine:
I.  The diameter of the duct ,  (2) 
II.  Length of the duct,  (4) 
III.  Pressure and temperature at the exit,  (4) 
IV.  Stagnation pressure loss, and  (4) 
V.  Verify the exit Mach number through exit velocity and temperature.  (2) 
2) A gas (γ =1.3, R=0.287 KJ/KgK) at p1 =1bar, T1 =400 k enters a 30cm diameter duct at

3) Air enters a long circular duct (d =12.5cm,f=0.0045) at a Mach number 0.5, pressure 3.0 bar
and temperature 312 K. If the flow is isothermal throughout the duct determine (a) the  
length of the duct required to change the Mach number to 0.7,(b) pressure and temperature of air at M =0.7 (c) the length of the duct required to attain limiting Mach number, and (d) State of air at the limiting Mach number. Compare these values with those obtained in  
adiabatic flow.  (16) 
4) A convergent –divergent nozzle is provided with a pipe of constant crosssection at its exit the exit diameter of the nozzle and that of the pipe is 40cm. The mean coefficient of friction for t h e pipe is 0.0025. Stagnation pressure and temperature of air at the nozzle entry are 12 bar and 600k. The flow is isentropic in the nozzle and adiabatic in the pipe. The Mach numbers at the entry and exit of the pipe are 1.8 and 1.0 respectively .
Determine
a)  The length of the pipe ,  (4) 
b)  Diameter of the nozzle throat, and  (6) 
c)  Pressure and temperature at the pipe exit.  (6) 
5) Show that the upper and lower branches of a Fanno curve represent subsonic and supersonic flows respectively. Prove that at the maximum entropy point Mach number is unity and all processes approach this point .How would the state of a gas in a flow change from
the supersonic to subsonic branch ? (16)
Flow in constant area ducts with heat transfer(Rayleigh flow)
6) The Mach number at the exit of a combustion chamber is 0.9. The ratio of stagnation t e m p e r ature at exit and entry is 3.74. If the pressure and temperature of the gas at exit are 2.5 bar and 1000˚C respectively determine (a) Mach number, pressure and temperature of the gas at entry, (b) the heat supplied per kg of the gas and (c) the maximum heat that can be supplied. Take γ= 1.3,
Cp= 1.218 KJ/KgK (16)
7) The conditions of a gas in a combuster at entry are: P1=0.343bar, T1 = 310K, C1= 60m/s.
Determine the Mach number, pressure ,temperature and velocity at the exit if the increase in stagnation enthalpy of the gas between entry and exit is 1172.5KJ/Kg.
Take Cp=1.005KJ/KgK, γ =1.4 (16)
8) A combustion chamber in a gas turbine plant receives air at 350 K ,0.55bar and 75 m/s .The air – fuel ratio is 29 and the calorific value of the fuel is 41.87 MJ/Kg .Taking γ=1.4 and R =0.287 KJ/kg K for the gas determine.
a)  The initial and final Mach numbers,  (4) 
b)  Final pressure ,temperature and velocity of the gas,  (4) 
c)  Percent stagnation pressure loss in the combustion chamber , and  (4) 
d)  The maximum stagnation temperature attainable.  (4) 
9) Obtain an equation representing the Rayleigh line . Draw Rayleigh lines on the hs and
pv planes for two different values of the mass flux. Show that the slope of the Rayleigh line
on the pv plane is {dp/dv} = þ² c² (16)
Unit4
PARTA (2 Marks)
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 (16 Marks)
Flow with normal shock
1) The state of a gas (γ=1.3, R =0.469 KJ/Kg K) upstream of a normal shock is given by the following data:
Mx =2.5, px= 2bar ,Tx =275K calculate the Mach number ,pressure, temperature and velocity of the gas downstream of the shock; check the calculated values with those give in the gas tables. (16)
2) The ratio of th exit to entry area in a subsonic diffuser is 4.0 .The Mach number of a jet of air approaching the diffuser at p0=1.013 bar, T =290 K is 2.2 .There is a standing normal
Shock wave just outside the diffuser entry. The flow in the diffuser is isentropic. Determine
at the exit of the diffuser.
1.  Mach number ,  (4) 
2.  Temperature, and  (4) 
3.  Pressure  (4) 
4.  What is the stagnation pressure loss between the initial and final states of the flow ?  (4) 
3) The velocity of a normal shock wave moving into stagnant air (p=1.0 bar, t=17˚C) is 500 m/s .
If the area of cross section of the duct is constant determine (a) pressure (b) temperature (c) velocity of air (d) stagnation temperature and (e) the Mach number imparted upstream of the wave front. (16)
4) The following data refers to a supersonic wind tunnel: Nozzle throat area =200cm²
Test section cross section =337.5cm²
Working fluid ;air (γ =1.4, Cp =0.287 KJ/Kg K)
Determine the test section Mach number and the diffuser throat area if a normal
shock is located in the test section. (16)
5) A supersonic diffuser for air (γ =1.4) has an area ratio of 0.416 with an inlet Mach number of
2.4 (design value). Determine the exit Mach number and the design value of the pressure ratio across the diffuser for isentropic flow. At an off design value of the inlet Mach number
(2.7) a normal shock occurs inside the diffuser .Determine the upstream Mach number and area ratio at the section where the shock occurs, diffuser efficiency and the pressure ratio
across the diffuser. Depict graphically the static pressure distribution at off design. (16)
6) Starting from the energy equation for flow through a normal shock obtain the following relations (or) Prandtl – Meyer relation Cx Cy =a* ² M*x M*y =1 (16)
Flow with oblique shock waves:
7) Air approaches a symmetrical wedge (δ =15˚) at a Mach number of 2.0.Determine for the strong and weak waves (a) wave angle (b) pressure ratio (c) density ratio,
(d) Temperature ratio and (e) downstream Mach number Verify these values using
Gas tables for normal shocks. (16)
8) A gas (γ =1.3) at p1 =345 Mbar, T1= 350 K and M1=1.5 is to be isentropically expanded
to 138 Mbar. Determine (a) the deflection angle, (b) final Mach number and (c) the temperature
of the gas. (16)
9) A jet of air at Mach number of 2.5 is deflected inwards at the corner of a curved wall.The wave angle at the corner is 60˚.Determine the deflection angle of the wall, pressure
and temperature ratios and final Mach number. (16)
10) Derive the Rankine –Hugoniot relation for an oblique shock
Þ2 /þ 1 = γ + 1 p2 γ + 1 p2
  +1  + 
γ 1 p1 γ 1 p1
Compare graphically the variation of density ratio with the initial Mach number in isentropic flow
and flow with oblique shock. (16)
11) The Mach number at the exit of a combustion chamber is 0.9. The ratio of stagnation
temperature at exit and entry is 3.74.If the pressure and temperature of a gas at exit are  
2.5 bar and 1000˚C respectively determine (a) Mach number ,pressure and temperature of the gas at entry,(b) the heat supplied per Kg of the gas and (c) the maximum heat that  
can be supplied.  
Take γ =1.3 and Cp =1.218 KJ/Kg K  (16) 
12) The conditions of a gas in a combuster at entry are: P1=0.343 bar,T1= 310K ,C1=60m/s Determine the Mach number ,pressure, temperature and velocity at the exit if the increase in stagnation enthalpy of the gas between entry and exit is 1172.5KJ/Kg.
Take Cp=1.005KJ/kg, γ =1.4. (16)
13) 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 41.87 MJ/Kg. Taking γ =1.4 and R =0.287 KJ/Kg K for the gas determine:
a) The initial and final Mach number, (4)
b)  Final pressure, temperature and velocity of the gas,  (4) 
c)  Percent stagnation pressure loss in the combustion chamber and  (4), 
d)  The maximum stagnation temperature attainable.  (4) 
14) Obtain an equation representing the Rayleigh line. Draw Rayleigh lines on the hs and pv planes for two different values of the mass flux.
Show that the slope of the Rayleigh line on the pv plane is {dP/dV} r = þ² c² (16)
Unit 5
PART A (2 Marks)
1) Differentiate jet propulsion and rocket propulsion (or) differentiate between air breathing and rocket propulsion?
2) What is monopropellant? Give one example for that?
3) What is bipropellant?
4) Classify the rocket engines based on sources of energy employed?
5) What is specify impulse of rocket?
6) Define specific consumption?
7) What is weight flow coefficient?
8) What is IWR?
9) What is thrust coefficient?
10) Define propulsive efficiency?
Part  B (16 Marks)
1) A turboprop engine operates at an altitude of 3000 meters above mean sea level and an aircraft speed of 525 Kmph. The data for the engine is given below
Inlet diffuser efficiency =0.875
Compressor efficiency =0.790
Velocity of air at compressor entry =90m/s
Properties of air: γ =1.4, Cp =1.005 KJ/kg K (16)
2) The diameter of the propeller of an aircraft is 2.5m; It flies at a speed of 500Kmph at an altitude of 8000m. For a flight to jet speed ratio of 0.75 determine (a) the flow rate of air
through the propeller, (b) thrust produced (c) specific thrust, (d) specific impulse and
(e) The thrust power. (16)
3) An aircraft flies at 960Kmph. One of its turbojet engines takes in 40 kg/s of air and expands the gases to the ambient pressure .The air –fuel ratio is 50 and the lower calorific value of the fuel is 43 MJ/Kg .For maximum thrust power determine (a)jet velocity (b) thrust (c) specific thrust
(d) Thrust power (e) propulsive, thermal and overall efficiencies and (f) TSFC (16)
3) A turbo jet engine propels an aircraft at a Mach number of 0.8 in level flight at an altitude of 10 km
The data for the engine is given below: Stagnation temperature at the turbine inlet =1200K  
Stagnation temperature rise through the compressor =175 K  
Calorific value of the fuel =43 MJ/Kg  
Compressor efficiency =0.75  
Combustion chamber efficiency =0.975  
Turbine efficiency =0.81  
Mechanical efficiency of the power transmission between turbine and compressor =0.98  
Exhaust nozzle efficiency=0.97  
Specific impulse =25 seconds  
Assuming the same properties for air and combustion gases calculate  
Fuel –air ratio,  (2) 
Compressor pressure ratio,  (4) 
Turbine pressure ratio,  (4) 
Exhaust nozzles pressure ratio ,and  (4) 
Mach number of exhaust jet  (2) 
5) A ramjet engine operates at M=1.5 at an altitude of 6500m.The diameter of the inlet diffuser at entry is 50cm and the stagnation temperature at the nozzle entry is 1600K.The calorific value
of the fuel used is 40MJ/Kg .The properties of the combustion gases are same as those of
air (γ =1.4, R=287J/Kg K ). The velocity of air at the diffuser exit is negligible
Calculate  
(a) the efficiency of the ideal cycle, (b) flight speed (c) air flow rate (d) diffuser pressure ratio (e) fuel –ratio (f)nozzle pressure ratio (g) nozzle jet Mach number (h) propulsive efficiency  
(i) and thrust. Assume the following values: 0D =0.90, 0B =0.98, 0 j= 0.96.  
Stagnation pressure loss in the combustion chamber =0.002Po2.  (16) 
7) A rocket flies at 10,080 Kmph with an effective exhaust jet velocity of 1400m/s and propellant flow rate of 5.0Kg/s .If the heat of reaction of the propellants is 6500KJ/Kg of the
propel at mixture determine;
a)  Propulsion efficiency and propulsion power,  (6) 
b)  Engine output and thermal efficiency ,and  (6) 
c)  Overall efficiency.  (4) 
7) Determine the maximum velocity of a rocket and the altitude attained from the following data: Mass ratio =0.15
Burn out time =75s
Effective jet velocity =2500m/s
What are the values of the velocity and altitude losses due to gravity? Ignore drag and
Assume vertical trajectory. (16)
8) A missile has a maximum flight speed to jet speed ratio of 0.2105 and specific impulse equal to 203.88 seconds .Determine for a burn out time of 8 seconds
a)  Effective jet velocity  (4) 
b)  Mass ratio and propellant mass functions  (4) 
c)  Maximum flight speed, and  (4) 
d)  Altitude gain during powered and coasting flights  (4) 
9) Calculate the orbital and escape velocities of a rocket at mean sea level and an altitude of 3 0 0 k m from the following data:

13 ) What are the advantages and disadvantages of liquid propellants compared to solid propellants. (16)
14) Discuss in detail the various propellants used in solid fuel rockets and liquid fuel system
.Also sketch the propellant feedsystem for a liquid propellant rocket motor. (16)
15) Briefly explain the construction and working of:
A. Rocket engine (6) B. Ramjet engine (6) C. Pulsejet engine (4)
16) With the help of a neat sketch describe the working of a ramjet engine. Depict the
various thermodynamic process occurring in it on hs diagram. What is the effect of
flight Mach number on its efficiency? (16)
17) Explain with a neat sketch the working of a turbopump feed system used in a liquid pr o p e l l a n t rocket? (16)