ME6604 Gas dynamics and Jet Propulsions Question Bank

This post covers the unit wise 2 marks and 16 marks of the subject Gas Dynamics and Jet propulsion. Make it for your reference and get more marks in university examinations.

ANNA UNIVERSITY,CHENNAI

REGULATION 2013

ME6604-Gas Dynamics and Jet Propulsion

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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

Air is discharged from a reservoir at po = 6.91 bar and to = 325  C through a nozzle to an exit pressure of 0.98 bar. If the flow rate is 3600 Kg/hr, determine throat area, pressure and velocity at the throat, exit area, exit Mach number and maximum velocity. Consider flow is isentropic. (AU: May 2012, Dec 2009, May 2008)

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.

(AU: May 2012)

A supersonic nozzle expands air from Po = 25 bar and T0 = 1050 K to an exit pressure of 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.

(AU: May 2011, May 2010)

A conical diffuser has entry and exit diameters of 15 cm and 30 cm respectively. The pressure, temperature and velocity of air at entry are 0.69 bar, 340 K and 180 m/s respectively. Determine i) exit pressure ii) the exit velocity and iii) the force exerted on the diffuser walls assume isentropic flow, γ =1.4, Cp = 1.00 J/Kg K

(AU: May 2011, May 2010, May 2009 Dec 2008, Dec 2007)

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 a perfect gas (γ = 1.3, R = 0.469 kJ/Kg K). I) Stagnation temperature. ii) Temperature and velocity of gas at exit. Iii) the flow rate per square metre of the inlet cross-section. (AU: May 2010, May 2008)

Air (γ = 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 cross-section 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. (AU: May 2010, May 2008)

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.

(AU: Dec 2009)

Air is drawn isentropically from a standard atmosphere at sea level (101.3 KPa and 15 C) through a converging diverging nozzle. The static pressure at two different locations at 80 KPa and 40 KPa respectively. Determine the Mach number at each of these locations. Also determine the velocity at each of these locations.

(AU: May 2009)

Air (Cp = 1.05 KJ/Kg-K, γ = 1.38) at P1 = 3 × 105 N/m2 and T1 = 500 k flows with a velocity of 200 m/s in a 0.3 m diameter duct. Calculate: Mass flow rate, Stagnation temperature, Mach number and stagnation pressure values assuming the flow as compressible and incompressible respectively.

(AU: Dec 2008, Dec 2007)

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 = 287 J/Kg K. Determine: i) Stagnation pressure and temperature. ii) Velocity of sound in the dynamic and stagnation conditions. Iii) Stagnation pressure assuming constant density.

(AU: May 2008, Dec 2007)

What is the effect of Mach number on compressibility? Prove for γ=1.4, Po –P / ½ P c² = 1 +¼ M² + 1/40 M 4 + ……. (AU: May 2009, Dec 2007, Dec 2006)

Derive area ratio as a function of Mach number for one dimensional isentropic flow

(AU: Dec 2008)

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:

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

(AU: May 2004)

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. (AU: Dec 2004)

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 γ = 1.4 and Cp = 1.005 Kj/KgK. (AU: Dec 2005)

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.

(AU: May 2012, May 2010, May 2009, Dec 2007)

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?

(AU: May 2012)

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.

(AU: May 2011)

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 γ = 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.

(AU: May 2011, Dec 2007)

The stagnation temperature of air in a combustion chamber is increased to 3.5 times its initial value. If the air at entry is at 5 bar, 105 C and a mach number of 0.25 determine: i) the Mach number, pressure and temperature at exit. ii) Stagnation pressure loss and iii) the heat supplied per kg of air.

(AU: May 2010, May 2008)

Air enters a constant area duct at M1 = 3, P1 = 1 atm and T1 = 300 K. inside the duct the heat added per unit mass is q = 3 × 105 J/Kg. Calculate the flow properties M2, P2, T2, ρ2, To2 and Po2 at the exit. (AU: Dec 2009)

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.

(AU: Dec 2009)

Air flows with negligible friction in a constant are duct. At section one, the flow properties are T1 = 60.4 C, P1 = 135 kPa absolute and velocity 732 m/s. Heat is added to the flow between section one and section two, where the mach number is 1.2. Determine the flow

properties at section two, the heat transfer per unit mass and the entropy change. (AU: May 2009)

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.

(AU: Dec 2008, May 2008)

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.

(AU: Dec 2008)

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:

Derive the equation for Mach number in the downstream of the normal shock wave

(AU: May 2012)

The velocity of a normal shock wave moving into stagnant air (P = 1.0 bar, T = 17 C) is 500m/s. if the area of cross section of the duct is constant, determine pressure, temperature, velocity of air, stagnation temperature and Mach number imparted upstream of the wave front.

(AU: May 2012)

Air approaches a symmetrical wedge (angle of deflection δ= 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.

(AU: May 2012)

The ratio of the exit to entry area in a subsonic diffuser is 4.0. The Mach number of a jet of air approaching the diffuser at Po = 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, I) Mach number ii) Temperature and pressure iii) What is the stagnation pressure loss between the initial and final stages of the flow

(AU: May 2011, May 2010, Dec 2008, Dec 2007, May 2007)

Derive the equation for static pressure ratio across the shock waves (AU: May 2012)

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 i) Deflection angle ii) Final Mach number and iii) the temperature of

the gas (AU: May 2011, May 2008)

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 (γ = 1.4, R = 0.287kJ/kg-k) 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.

(AU: May 2010, May 2008)

A convergent-divergent 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.

(AU: Dec 2009)

A pilot tube kept in a supersonic wind tunnel forms a bow shock ahead 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. (AU: Dec 2009)

A convergent-divergent 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.

(AU: May 2009)

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 36o 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.

(AU: May 2009)

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.

(AU: Dec 2007)

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:

Differentiate turbojet and turboprop propulsion engines with suitable diagrams

(AU: May2012)

Write the equations to calculate propulsion efficiency and thermal efficiency of an aircraft.

(AU: May 2012)

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.

(AU: May 2012)

Explain the principle of operation of a turbojet engine and state its advantages and disadvantage

(AU: May 2011)

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

(AU: May 2011, May 2007)

Explain with a neat sketch the principle of operation of a ramjet engine and state its advantages and disadvantages. (AU: May 2010, May 2009).

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 air-fuel 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.(AU: Dec 2009)

Describe the working of supersonic ramjet engine with a neat sketch. List out its advantages and disadvantages. (AU: May 2009)

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. (AU: Dec 2008)

Explain with a neat sketch the principle of operation of a turbojet engine and state its advantages and disadvantages. (AU: May 2008)

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?

Part-B

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.

(AU: May 2012)

Explain briefly about the propellant feed system of a liquid propellant rocket engine with suitable schematic sketches. (AU: May 2012)

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.

(AU: May 2012, Dec 2009)

Explain with a neat sketch the working of a gas pressure feed system used in liquid propellant rocket engines (AU: May 2011)

Describe the important properties of liquid and solid propellants desired for rocket propulsion. (AU: May 2011, May 2010, May 2008)

Explain the working of a turbo-pump feed system used in a liquid propellant rocket

(AU: May 2010, Dec 2007)

Deduce expressions for propulsion efficiency specific impulse and overall efficiency of a rocket engine. (AU: Dec 2009)

Explain the principle of operation of liquid propellant and solid propellant engines with neat sketch. (AU: May 2009)

List down the advantages of liquid propellant rockets. (AU: May 2009)

The effective jet velocity from a rocket is 2700 m/s. The forwared flight velocity is 1350 m/s and the propellant consumption is 78.6 kg/s. Calculate: thrust, Thrust power and propulsion

efficiency.

(AU: Dec 2008)

Derive the thrust equation for rocket engines. (AU: Dec 2008)

ME6604 Gas Dynamics and Jet propulsions Question Bank

Anna university important questions for all mechanical engineering subjects go to

 : annaunivstudymaterials.blogspot.com

GAS DYNAMICS AND JET PROPULSION

UNIT-1

2 MARKS

1. Define: Stagnation enthalpy.                                                                       
[Nov/Dec-2005] & [Nov/Dec-2009]

2. Distinguish between Mach wave and normal shock.                              
[Nov/Dec-2005] & [Nov/Dec-2009]

3. Differentiate between compressible and incompressible flows.                          
 [May/June-2006]

4.  What is Mach cone?                                                                        
[May/June-2006] & [April/May-2008]

5. Define stagnation state of a system. 
[Nov/Dec-2006], [April/May-2010], [Nov/Dec-2010], [Nov/Dec-2012]

6. An air jet (ᵞ=1.4, R= 287 J/kg-K) at 400K has sonic velocity. Determine its velocity.                                                                                                                             [Nov/Dec-2006]        

7. Draw the Mach cone and identify its salient features.                                                        
  [Nov/Dec-2006]

8. What is the effect of Mach number on compressibility?                                       
   [Nov/Dec-2006]

9. Define compressible flow and Mach number.                                   
 [Nov/Dec-2007], [April/May-2010]

10. Define stagnation temperature and stagnation pressure.                                    
 [Nov/Dec-2007]

11. Define Mach number.                                                       
 [April/May-2008], [May/June-2012]

12. How will you classify the compressible flow based on Mach number range? 
   [April/May-2008]

13. What are the difference between nozzle and diffuser?     
 [April/May-2008],[May/June-2014]

14. What is the advantage of using M^* (second kind of Mach number) instead of M (Local Mach  number) in some cases?                                                                                                             
[Nov/Dec-2008]

15. The wave front caused by firing a bullet gave a Mach angle of 35⁰. Find the velocity of the bullet if the static temperature of atmosphere is 276K.                                                             [Nov/Dec-2008]

16. Draw the disturbances wave propagation in compressible flow for M=1 and M>1.                                        

      [May/June-2009]

 17.When M^* is used instead of M?                                                                       [May/June-2009]

18. Express the stagnation enthalpy in terms of static enthalpy and velocity of flow

                                                                                                            [Nov/Dec-2009], [May/June-2012]

19. When air is released adiabatically from a tyre, the temperature of air at the nozzle exit is 37⁰C below that of air inside the tyre. Neglecting irreversibility calculate the exit velocity of air.

                                                                                                                                [Nov/Dec-2009]

20. Define closed and open system.                                                                     [April/May-2010]

21. What is the difference between intensive and extensive property?                             [April/May-2010]

22. What is the velocity of sound at 293K in Hydrogen?                                                    [Nov/Dec-2010]

23. ‘Zone of silence’ is absent in subsonic flow. Why?                                         [May/June-2012]

24. What is the cross section of the nozzle required to increase the velocity of compressible fluid flow from

a)      Subsonic to supersonic,

b)      Subsonic to sonic                                                                     [May/June-2012]

25. Write the energy equation in differential form.                    [Nov/Dec-2012], [May/June-2013]

26. Draw the Mach cone and indicate various zones.                                            [May/June-2013]

27. Name the different regimes of compressible fluid flow.                                     [Nov/Dec-2013]

28. Define zone of action and zone of silence.                                                        [Nov/Dec-2013]

29. When does maximum flow occur for an isentropic flow with variable area duct?

      [May/June-2014]

30. What are the basic differences between compressible and incompressible flows?

      [May/June-2013]

31. Name the four reference velocities that are used in expressing the third velocities in non-dimensional form.                                                                                                                            [May/June-2013]

16 MARKS

1. Deduce expressions for   and   for a compressible fluid flow under isentropic conditions              (6)

                                                                                                                                    [Nov/Dec-2005]

2. Differentiate supersonic and subsonic flows.                                                                                         (4) 

         [May/June-2006]

3. A supersonic wind tunnel is designed for M = 3.0 at the test section. If the air supply in the reservoir is at 4 Bar and 26⁰C. Determine mass flow rate, the area of the test section, the temperature and the density at the throat and test section. The throat area is 0.09m² .γ= 1.4.                                                            (12)

                                                                                                                              [May/June-2006]

4. What is the effect of Mach number on the compressibility? Prove for γ =1.4.

           

Where p₀- stagnation pressure, p- static pressure, ϼ- density, c- velocity.                                                 (12)

         [Nov/Dec-2006]

5. Air at 200KPa flows at a velocity of 50m/sec. Find the Mach number at a point where its density is 2.9Kg/m³.                                                                                                                                   (4)

         [Nov/Dec-2006]

6. Speed of an aircraft 800Km/hr. The stagnation conditions are105KPa & 308K. Find static conditions and flight Mach number. (Take γ =1.4, c_p= 1.005KJ/Kg-K.                                                             (6)

         [Nov/Dec-2006]

7. Air flows from a reservoir at 550KPa and 343K. Assuming isentropic flow, calculate the velocity, temperature, pressure at a section where M=0.6.                                                                                   (6)

         [Nov/Dec-2006]

8. Velocity of an aircraft which has same Mach number at all altitudes, flying at an altitude of 11000m is 50m/sec lower than that of its velocity at mean sea level. Determine its Mach number.

                                                                                                                                                      (4)

         [Nov/Dec-2006]

9.  Deduce the expression for the change in area for isentropic flow and identify the geometric shapes for nozzles and diffusers.                                                                                                                         (10)

         [Nov/Dec-2006]

10. An aircraft is flying at an altitude of 12,000 meters (T=216.65K, p=0.193 bar) at a Mach number of 0.82. The cross-sectional area of the inlet diffuser before the L.P compressor stage is 0.5m². Determine the mass of air entering the compressor per second, the speed of the aircraft and the stagnation pressure and temperature of air the diffuser entry.                                                                                    (6)

         [Nov/Dec-2006]

11. Derive the expression for the area expansion ratio of nozzles.                                             (10)

        [Nov/Dec-2006]

12. A Freon-turbine has to use a maximum flow rate of 5Kg/s of Freon employing a ring of convergent nozzles of total exit area of cross-section of 100cm². The pressure in the nozzle entry space is 20X10⁵N/m². Taking C_p= 0.845KJ/Kg-K, γ =1.2 calculate the stagnation temperature, static pressure and the temperature at the nozzle exit, and Mach number at the nozzle exit.

     (6)

         [Nov/Dec-2006]

13. (i) What is the effect of Mach number on the compressibility? Prove for γ =1.4.             

        (ii) Derive the following relations:

                                                                                                                 (16)

         [Nov/Dec-2007]

14. Air (c_p=1.05KJ/Kg-K, γ =1.38) at p₁= 3X10⁵ N/m and T₁=500K flows with a velocity of 200m/s in a 30cm diameter duct. Calculate

                    i.            Mass flow rate

                  ii.             Stagnation temperature

                iii.            Mach number

                iv.            Stagnation pressure values assuming the flow is compressible.                                (16)

   [Nov/Dec-2007]

15. Air at P₁= 3bar and T₁=227⁰C is flowing with a velocity of 200m/s in a 0.3m diameter duct. If C_p=1050J/KgC and γ=1.38, determine the following

                    i.             Stagnation temperature and pressure                                                                     (4)

                  ii.            Mach flow rate of air                                                                                             (4)

                iii.            Mach number                                                                                                        (4)        

                iv.            Stagnation pressure assuming the flow is incompressible                                      (4)

[May/June-2013]

15.  The pressure, temperature and Mach number at the entry of a flow passage are 2.45bar, 26.5⁰C and 1.4 respectively. If the exit Mach number is 2.5 determine for adiabatic flow of a perfect gas (γ =1.3), R=0.469KJ/Kg-K.

                                i.            Stagnation temperature

                              ii.            Temperature and velocity of gas at exit, and

                            iii.            The flow rate per square meter of the inlet cross-section.                                                   (16)

          [April/May-2008], [April/May-2010]

16. Air (γ =1.4, R=287.43J/Kg-K) enters a straight axi-symmetric duct at 300K, 3.45bar and 150m/s and leaves it at 277K, 2.058bar and 260m/s. The area of cross-section at entry is 500cm². Assuming adiabatic flow determines:

                                i.            Stagnation temperature   

                              ii.            Maximum velocity

                            iii.            Mass flow rate and

                            iv.            Area of cross-section at exit.                                                                               (16)

                    [April/May-2008], [April/May-2010],  [May/June-2012], [Nov/Dec-2013]

17. Air flow through a nozzle which has inlet area of 10cm². If the air has a velocity of 80m/s, temperature of 28⁰C and pressure of 700KPa at the inlet section and a pressure of 250KPa at the exit, find the mass flow rate through the nozzle and assuming one dimensional isentropic flow, the velocity at the exit section of the nozzle.                                                                                                                    (16)

      [April/May-2008]

18. Air (c_p=1.05KJ/Kg-K, γ =1.38) at p₁=3X10⁵N/m² and T₁=500K flows with a velocity of 200m/s in a 0.3m diameter duct. Calculate: Mass flow rate, Stagnation temperature, Mach number and Stagnation pressure values assuming the flow as compressible and incompressible respectively.                                                                                                                                                                                         (13)

         [Nov/Dec-2008]

19. An air plane travels at Mach 1.2 at an elevation where the temperature is 233K. Determine the speed of the air plane in Km/hr. Assume γ =1.4.                                                                                          (3)

         [Nov/Dec-2008]

20. A jet fighter is flying at Mach number 2.5. It is observed directly overhead at a height of 10Km. How much distance it would cover before a sonic boom is heard on the ground?

       (7)

         [Nov/Dec-2008]

21. An air jet at 400K has sonic velocity. Determine: (a) Velocity of sound at 400K, (b) Velocity of sound at stagnation condition, (c) Maximum velocity of jet, (d) Stagnation enthalpy.(e) crocco number.                                                                                                                                                                 (12)

          [Nov/Dec-2012], [May/June-2013], [Nov/Dec-2008]

22. Air at rest at 90⁰C is accelerated isentropically (take γ =1.4).

                                i.            What is the air speed in m/s when the Mach number becomes 0.8?

                              ii.            What is the air speed when the flow becomes sonic?

                            iii.            What is the Mach number when the air speed becomes 600m/s.                                      (16)

      [May/June-2009]

23. What is the effect of Mach number on compressibility? Prove for γ =1.4                                                                                                                                                  (16)

      [May/June-2009]

24. Prove that                                                                                    (12)                                                                         [May/June-2013]

25. Explain the difference between flow and flow work.                                                                             (6)

      [May/June-2009]

26. Carbon-di-oxide expands isontropically through a nozzle from a pressure of 3.2bar to 1.0bar.If the initial temperature is 475⁰K determine the final temperature, the enthalpy drop and the change in internal energy.                                                                                                                                            (10)

      [May/June-2009]

27. What is meant by velocity of sound? Derive the expression for the velocity of sound.

    (16)

       [May/June-2009] & [April/May-2010]

28. Deduce an expression for sonic velocity in terms of the properties of air.                                            (6) 

         [Nov/Dec-2009]

29. Sketch the effect of disturbance in still air as it moves from rest to supersonic velocity for the following Mach number: M=0, M=0.5, M=1.0, M=2. Explain in detail the observed phenomena.

       (16)                                               [Nov/Dec-2009]

30. Ambient air at altitude of 6000m above sea level enters the engine of an aircraft flying at 500Km/hr. If the flow rate through the engine is 25Kg/sec, determine the diameter of the inlet to the engine.                                                                                                                                                                              (8)

      [April/May-2010]

31. The pressure, velocity and temperature of isentropic flow of air at the entry of a nozzle are 2bar, 150m/sec and 330K and the exit pressure is 1.5bar.

                                                        i.            Determine the shape of the nozzle

                                                      ii.            Determine the Mach number at the entry and exit of the nozzle

                                                    iii.            Mass flow rate through the nozzle.

Take ratio of specific heats (γ) for air as 1.4 and C_p=1.0KJ/Kg-K.                                                 (8)

      [April/May-2010]

32. An aircraft is driven by propellers with a diameter of 4m. At what engine speed will the tips of the propellers reach sonic velocity if the air temperature is 288K?                                                          (6)

         [Nov/Dec-2010]

33. Derive the energy equation.                                                                                                     (10)       

                                                                                        [Nov/Dec-2010]

34. Explain the effect of Mach number on compressibility. Calculate the percentage deviation due to the assumption of incompressibility when Mach number is equal to 0.5 and specific heat ratio is 1.4.                                                                                                                                                                    (16)

         [Nov/Dec-2010]

35. Air is discharged from a reservoir at P₀=6.91bar and t₀=325⁰C through a nozzle to an exit pressure of 0.98bar. If the flow rate is 3600Kg/hr, determine throat area, pressure and velocity at the throat, exit area, exit Mach number and maximum velocity. Consider the flow is isentropic.

                                                                                                                                                    (16)

      [May/June-2012]

36.  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 are at inlet are 70KPa and -7⁰C. If the mass flow rate of air is 125Kg/s, determine the stagnation conditions, areas at throat and exit, static conditions (pressure, temperature, velocity) of air at exit.                                                                                                                     (16)

      [May/June-2012]

37. An air jet (γ=1.4, R=287J/Kg-K has sonic velocity. Determine

                                i.            Velocity of sound at 400K,

                              ii.            Velocity of sound at the stagnation conditions,

                            iii.            Maximum velocity of the jet,

                            iv.            Stagnation enthalpy,

                              v.            Crocco number.                                                                                                    (16)                                                                                                                   [May/June-2012]

38. Show that T₀/T= (1 + (γ -  M²)                                                                                                 (4)

      [May/June-2012]

39. Derive the energy equation.

                                                                                             (16)

         [Nov/Dec-2012]

40. Write short notes on effect of Mach number on compressibility.                                              (4)

         [Nov/Dec-2012]

41. Explain What is meant by stagnation properties of fluid and supersonic flow.                                    (4)

      [May/June-2013]

42. What is the effect of Mach number on compressibility? Derive the relation between pressure coefficient and Mach number.                                                                                          [May/June-2013]

43. In settling chamber air is maintained at a temperature of 400K and a pressure of 6bar. Calculate the following (1) stagnation enthalpy (2) stagnation velocity of sound (3) maximum velocity (4) critical velocity of fluid (5) critical velocity of sound.                                                                                     (10)                

         [Nov/Dec-2013]

44. Show that P₀/P = (1+ (γ-1/2) M²) ^γ/γ-1.                                                                                      (6)

         [Nov/Dec-2013]

45. Air flows down a variable area duct. Measurements indicate that the temperature is 278K and velocity is 150m/s at a certain section of the duct. Measurements at a second section indicate that the temperature has decreased to 253K. Assuming that the flow is adiabatic and one dimensional, find the velocity at this second section.                                                                                                                                  (6)

      [May/June-2014]

46. Typical cruising speeds and altitudes for three commercial aircraft are:

            Dash 8: Cruising speed 500Km/hr at an altitude of 4500m

            Boeing 747: Cruising speed: 978 Km/hr at an altitude of 9500m

Find the Mach number of the aircraft when flying at these cruise conditions.                                            (10) 

      [May/June-2014]

47. Air at bar pressure flows with a velocity of 180Km/hr. Find the Mach number if the density of air is 3.0Kg/m³.                                                                                                                                           (4)

      [May/June-2013]

UNIT-2

2 MARKS

1. Give any two assumptions regarding Fanno flow. [Nov/Dec-2005] & [Nov/Dec-2009]

2. State the two governing equations used in plotting Rayleigh line. [Nov/Dec-2005]

3. What is meant by stagnation properties? [May/June-2006]

4. Define Normal shock. [May/June-2006]

5. Define one dimensional steady flow. [Nov/Dec-2006]

6. Show an adiabatic expansion process through a nozzle on T-s coordinates. [Nov/Dec-2006]

7. What is Fanno line? [Nov/Dec-2006]

8. How do the flow properties change in Rayleigh flow? [Nov/Dec-2006]

9. What is subsonic, sonic and supersonic flow with respect to Mach number? [Nov/Dec-2007]

10.  How the area and velocity vary in supersonic flow of nozzle and diffuser? [Nov/Dec-2007], [April/May-2010]

11. Plot the variation of Area ratio with Mach number. [April/May-2008]

12. How the velocity vary along the axis of flow in supersonic nozzle and diffuser? [April/May-2008]

13. Sketch the Fanno flow on the T-S plane and explain the significance of it. [April/May-2008]

14. Write down the ratio of velocities between any two sections in terms of their Mac h number in Fanno flow. [April/May-2008]

15. Draw h-s (Enthalpy-Entropy) diagram for the flow through a nozzle showing stagnation states. [Nov/Dec-2008]

16. Show graphically the variation of Mach number across a convergent-divergent nozzle. [Nov/Dec-2008]

17. Draw the curve between A/A* and Mach number M. [May/June-2009]

18. List the condition of choking in CD nozzle. [May/June-2009]

19. Why is expansion shock impossible? [Nov/Dec-2009]

20. Sketch the isentropic and actual expansion through a nozzle and give the expression for nozzle efficiency. [Nov/Dec-2009]

21. Explain the difference between Fanno flow and isothermal flow. [Nov/Dec-2009]

22. What is meant by Fanno flow? [April/May-2010]

23. Is it possible to get subsonic to supersonic flow in a constant area duct? Give reasons. [April/May-2010]

24. Where are the convergent nozzles and convergent- divergent nozzles used? [April/May-2010]

25. When is the flow of a fluid said to be one dimensional? [Nov/Dec-2010]

26. What is choked flow through a nozzle? [Nov/Dec-2010]

27. List some flow properties. [May/June-2012]

28. What are the assumptions made in the analysis of Rayleigh flow? [May/June-2012]

29. Differentiate nozzle and diffuser. [May/June-2012]

30. Draw the variation of Mach number along the length of a convergent divergent duct when it act as a

a)      Nozzle

b)      Diffuser. [May/June-2012]

31. Represent the diffuser process in h-s diagram. [Nov/Dec-2012]

32. When will divergent passage act as nozzle? [Nov/Dec-2012]

33. Draw the p/p₀ along the length of a convergent divergent device when it functions as nozzle. [May/June-2013]

34. Define impulse function and its significance. [May/June-2013]

35. Differentiate between adiabatic flow and diabatic flow. [Nov/Dec-2013]

36. What is chocked flow through a nozzle? [Nov/Dec-2013]

37. Give assumptions made on Rayleigh flow. [May/June-2014]

38. Define critical condition in Fanno flow. [May/June-2014]

39. What is impulse function and give its uses? [May/June-2013]

40. Give the expression for T₀/T and T^*/T for isentropic flow through variable area in terms of Mach number. [May/June-2013]

16 MARKS

1. Sketch the Fanno line as h-s and h-ϼ diagrams and explain how these lines are constructed? (6) [Nov/Dec-2005]

2. The friction factor for a 50mm diameter steel pipe is 0.005. At the inlet to the pipe the velocity is 70m/s temperature is 80⁰C and pressure is 10bar. Find the temperature, pressure and Mach number at exit if the pipe is 25m long. Also determine the maximum possible length. (10) [Nov/Dec-2005]

3. Distinguish between Rayleigh flow and Fanno flow. (4) [Nov/Dec-2005]

4. Air enters a combustion chamber with a certain Mach number. Sufficient heat is added to obtain a stagnation temperature ratio of 3 and a final Mach number is 0.8. Determine the Mach number at entry and the percentage loss in static pressure. Take r = 1.4 and C_p = 1.005KJ/Kg K for air. (12) [Nov/Dec-2005]

5. For isentropic flow prove  = (1+ m²) (6) [MayJune-2006]

6. Air flows through a duct. The pressure and temperature at station 1 are p₁=0.7 Bar and T₁=30⁰C. At a second station the pressure p is 0.5 Bar. Calculate temperature and density at the second station. Assume the flow to be isentropic. (10) [MayJune-2006]

7. Air is allowed to expand from initial state A (where P_A=2.068X10⁵ N/m² and T_A=333⁰K) to state B (where P_B=1.034X10⁵ N/m² and T_B=305⁰K). Calculate change in specific entropy of the air and show that the change in entropy is the same for

                                            i.            an isobaric process from A to some intermediate state C followed by an isovolumetric change from C to B and

                                          ii.            an isothermal change from A to some intermediate state D followed by an isoentropic change from D to B. (16) [MayJune-2006]

8. A reservoir whose temperature can be varied in a wide range of temperature receives air at a constant pressure of 150KPa. The air is expanded isentropically in a nozzle to an exit pressure of 101.5KPa. Determine (without using Gas tables) the values of the temperature to be maintained to the reservoir to produce the following velocities at the nozzle exit. (i) 100m/sec (ii) 250m/sec (16) [Nov/Dec-2006]

9. A subsonic diffuser operating under isentropic conditions has inlet area of 0.15m². The inlet conditions are c₁= 240m/sec, T₁= 300K, p₁= 70KPa. The velocity leaving the diffuser is 120m/sec. Calculate for air (i) mass flow rate (ii) stagnation pressure at exit (iii) stagnation temperature at exit (iv) static pressure at exit (v) change in entropy (vi) exit area. (16) [Nov/Dec-2006]

10.  Describe how the property changes in Fanno flow with suitable diagram and justifications. (6) [Nov/Dec-2006]

11. Outline the assumptions made in Rayleigh flow and explain their implications. (6) [Nov/Dec-2006]

12. A conical diffuser has entry and exit diameters of 15cm and 30cm respectively. The pressure, temperature and velocity of air at entry are 0.69bar, 340K and 180m/s respectively. Determine:

                    i.            the exit pressure

                  ii.            the exit velocity and

                iii.            the force exerted on the diffuser walls.

Assume isentropic flow, γ =1.4, c_p=1.00KJ/Kg-K. (16) [Nov/Dec-2007], [April/May-2010]

13. Air flowing in a duct has a velocity of 300m/s, pressure 1.0bar and temperature 290K. Taking γ =1.4 and R=287J/Kg=K. Determine

                    i.            Stagnation pressure and temperature

                  ii.            Velocity of sound in the dynamic and stagnation conditions

                iii.            Stagnation pressure assuming constant density. (16) [Nov/Dec-2007] & [April/May-2008]

14. Derive area ratio as a function of Mach number for one dimensional isentropic flow. (10) [Nov/Dec-2008]

15. Explain for a convergent nozzle the variation of pressure and Mach number when the back pressure is gradually lowered from stagnation pressure. (6) [Nov/Dec-2008]

16. A conical diffuser has entry and exit diameters as 0.15m and 0.3m respectively. The pressure, temperature and velocity of air at entry are 0.96bar, 340K and 185m/s respectively. Determine: Exit pressure, Exit velocity and Force exerted on the diffuser walls. Assume γ =1.4 and c_p=1.005KJ/Kg-K. (16) [Nov/Dec-2008]

17. Air is drawn isentropically from a standard atmosphere at sea level (101.3KPa and 15⁰C) through a converging diverging nozzle. The static pressure at two different locations is 80KPa and 40KPa, respectively. Determine the Mach number at each of these locations. Also determine the velocity at each of these locations. (16) [May/June-2009]

18. A conical diffuser has entry and exit diameters as 150mm and 3mm respectively. The pressure, temperature and air velocity at entry are 69KPa, 67⁰C and 180m/s respectively. Determine: Exit pressure, Exit velocity and Force exerted on the diffuser walls. Assume isentropic flow. Take γ =1.4 and c_p=1KJ/Kg-K. (16) [May/June-2009]

19. Starting from the continuity equation derive the expression for the area variation in terms of Mach number and velocity variation and hence obtain the shape (geometry) for both subsonic and supersonic nozzles and diffusers. (16) [Nov/Dec-2009]

20. In an isentropic flow diffuser the inlet area is 0.15m². At the inlet velocity 240m/s, static temperature=300K, and static pressure 0.7bar. Air leaves the diffuser with a velocity of 120m/s. Calculate at the exit the mass flow rate, stagnation pressure, stagnation temperature, area and entropy change across the diffuser. (16) [Nov/Dec-2009]

21. A supersonic nozzle expands air from P₀=25bar and T_o=1050K to an exit pressure of 4.35bar; the exit area of the nozzle is 100cm². Determine

                                i.            Throat area;

                              ii.            Pressure and temperature at the throat;;

                            iii.            Exit velocity as fraction of the maximum attainable velocity;

                            iv.            Mass flow rate. (16) [April/May-2010]

23. Derive the following relation for one dimensional isentropic flow. (6) [Nov/Dec-2010], (4) [May/June-2013]

           

24. Starting from adiabatic energy equation derive the following for a one-dimensional isentropic flow in an axisymmetric duct. (12) [May/June-2013]

           

25. The Mach number at inlet and exit for a Rayleigh flow are 3 and 1.5 respectively. At inlet static pressure is 50KPa and stagnation temperature is 295K. Calculate the fluid is air. Find [May/June-2012]

                                                        i.            The static pressure, static temperature and velocity at exit,   (3)

                                                      ii.            Stagnation pressure at inlet and exit,                                      (3)

                                                    iii.            Feat transferred,                                                                      (3)

                                                    iv.            Maximum possible heat transfer,                                            (2)

                                                      v.            Change in entropy between the two sections,                                    (3)

                                                    vi.            Is it a cooling or heating process?                                           (2)

26. An air nozzle is to be designed for an exit Mach number of 3.5. The stagnation conditions for the isentropic flow are 800KPa and 240⁰C. Estimate pressure, temperature, velocity and area at throat and exit for a mass flow rate of 3.5Kg/s. (16) [May/June-2012], [Nov/Dec-2013]

27. Derive an expression for mass flow rate through varying cross sectional passage for isentropic flow in terms of pressure ratio. (16) [Nov/Dec-2012]

28. A supersonic wind tunnel is designed for M=2 with a throat section 890cm². The air at 1.2bar and 25⁰C is supplied with negligible velocity. Find the mass flow rate, the area of test section and fluid property at the throat section. (16) [Nov/Dec-2012]

29. A gas is isentropically expanded from 10bar, 525⁰C in a nozzle to a pressure of 7.0bar. If the rate flow of gas is 1.5 Kg/sec.  Determine

                                i.            Pressure velocity and temperature at the nozzle throat and exit

                              ii.            Maximum possible velocity attainable by the gas and

                            iii.            Type of nozzle and its throat area.

Take r=1.5 and R=0.464KJ/Kg-K. (12) [May/June-2013]

30. Distinguish between the nozzle and diffusers. (4) [May/June-2013]

31. Explain the chocking condition with example. (4) [May/June-2013]

32. A conical air diffuser has an inlet diameter of 40cm and an exit diameter of 80cm. Air enters the diffuser with a static pressure of 200KPa, static temperature of 37⁰C and velocity of 265m/s. Calculate (i) mass flow rate (ii) properties at exit. (16) [Nov/Dec-2013]

33. The conditions of a gas in a combustor at entry is p₁=0.343baar, T₁=310K, c₁=60m/sec. 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 c_p=1.005KJ/Kg-K, γ=1.4. (16) [May/June-2014]

34. Air at P₀=10bar, T₀=400K is supplied to a 5cm 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, the length of the pipe and the mass flow rate. (6+6) [May/June-2013]

35. Air is supplied to a combustion chamber is a gas turbine plant at 350K, 0.55bar and 75m/s. The air-fuel ratio is 29 and the calorific value of the fuel is 42mJ/Kg. Assuming γ=1.4 and R=287J/Kg-K for the gas, determine

                                i.            The initial and final Mach numbers                            (4)

                              ii.            Final pressure, temperature and velocity of the gas    (4)

                            iii.            The maximum stagnation temperature attainable                   (4)

                            iv.            Stagnation pressure loss in the combustion chamber  (4) [May/June-2013]

UNIT-3

2 MARKS

1. Define the term: Strength of a shock wave. [Nov/Dec-2005], [April/May-2008], [Nov/Dec-2009], [Nov/Dec-2009], [April/May-2010]

2. Sketch an oblique shock and show the angles associated with flow through it. [Nov/Dec-2005]

3. What is Prandtl-Mayer relation? [May/June-2006]

4. Define propulsive efficiency. [May/June-2006]

5. Give the assumptions that are used in the analysis of Rayleigh flow process. [Nov/Dec-2006] & [Nov/Dec-2007]

6.  Give two examples of Fanno flow in thermal systems. [Nov/Dec-2006]

7. What is meant by Mach reflection? [Nov/Dec-2006]

8. Differentiate between impulse and specific impulse. [Nov/Dec-2006]

9. Give two practical examples for Fanno flow and Rayleigh flow analysis. [Nov/Dec-2007]

10. What is Rayleigh flow? Give two practical examples. [April/May-2008]

11. Draw Fanno curve and represent subsonic and supersonic flows. [April/May-2008]

12. What do you understand by oblique shock? [April/May-2008]

13. What are the assumptions made on Rayleigh flow? [Nov/Dec-2008], [May/June-2009], [Nov/Dec-2010], [May/June-2012], [Nov/Dec-2013]

14. What is the limiting Mach number in isothermal flow? [Nov/Dec-2008], [Nov/Dec-2010]

15. Give four examples of Fanno flow in thermal systems. [May/June-2009]

16. What is shock polar? [Nov/Dec-2009]

17. Sketch the Rayleigh line on the T-s plane and explain the significance of it. [Nov/Dec-2009]

18. Explain how the pitot tube could be used to measure the Mach number in supersonic flow. [April/May-2010]

19. Give two practical examples for Fanno flow and Rayleigh flow. [April/May-2010]

20. What are the assumptions made in the analysis of Rayleigh process? [April/May-2010]

21. A normal shock occurs at a point in air flow where the pressure is 530KPa and the temperature is -30⁰C. If the pressure ratio across this shock wave is 2.6, find the Mach number and static temperature at the downstream of the shock waves. [May/June-2012]

22. What are the beneficial and adverse effects of shock waves? [May/June-2012]

23. Explain the difference between Fanno flow and isothermal flow. [May/June-2012]

24. Define Fanno flow. [Nov/Dec-2012]

25. Give examples for Rayleigh flow. [Nov/Dec-2012]

26. Differentiate between Fanno and Rayleigh flow. [May/June-2013], [May/June-2013]

27. For constant area Fanno flow how limiting length for the pipe is determined? [May/June-2013]

28. Give the Fanno flow in h-s diagram show various Mach number regions and write the Fanno flow equation. [Nov/Dec-2013]

29. Why the efficiency of a machine, experiencing shock wave is considerably low? [May/June-2014]

30. What is the use of pitot tube in supersonic flow? [May/June-2014]

31. State the assumptions made to derive the isothermal flow equations. [May/June-2013]

16 MARKS

1. Derive the Rankine-Hergonoit reactions. (6) [Nov/Dec-2005]

2. When a converging divergine nozzle is operated at off-design condition a normal shock occurs at a section where the cross sectional area is 18.75cm² in the diverging portion. At inlet to the nozzle the stagnation state is given as 0.21 MPa and 36⁰C. The throat area is 12.5cm² and exit area is 25cm². Estimate the exit Mach number, exit pressure and loss in stagnation pressure for flow through nozzle. (10) [Nov/Dec-2005], [April/May-2010]

3. For flow through a normal shock deduce the relation M_y² =  (8) [Nov/Dec-2005], [April/May-2010]

4. A bow shock occurs infront of a pitot tube when it is used in a supersonic flow field. It measures 16KPa and 70KPa for static pressure upstream of the shock and the pressure at the mouth of the tube respectively. Estimate the Mach number of the supersonic flow. If the stagnation temperature is 300⁰C. Calculate the static temperature and the total (stagnation) pressure upstream and the downstream of the pitot tube. (8) [Nov/Dec-2005], [April/May-2010]

5. The pressure across an orifice in a duct delivering air drops from 10 bar to 6 bar. If the upstream Mach number is 0.60 determine final Mach number, the temperature ratio across the orifice and irreversibility of the process γ =1.4. (8) [MayJune-2006]

6. Air enters a constant area duct with a Mach number 0.4. The duct length is 2.65m and diameter is 80mm. The inlet stagnation conditions are 3.5 bar and 37⁰C. The friction co-efficient is 0.008. What is exit stagnation pressure γ =1.4? (8) [MayJune-2006]

7. The flow of air in a long pipe is under isothermal condition. At the inlet the ratio of static to total pressure is 0.962. The inlet temperature is 300⁰K and the inlet static pressure is 1 bar. Find the total static pressure, static temperature and the Mach number at the exit. The total temperature ratio across the duct is 1.2. Take γ =1.4.  (16) [MayJune-2006]

8. Air flowing in an insulated duct with friction coefficient f= 0.002. At inlet the velocity is 130m/sec, temperature 400K and pressure is 250KPa. The diameter of the duct is 16cm. (i) find the length of the pipe required that gives 20% drop in stagnation pressure. (ii) find the properties of air at a section 3.5m from inlet and (iii) find the maximum length of the pipe. (16) [Nov/Dec-2006]

9. A combustion chamber delivers the gases at a Mach number of 0.9 at 250KPa and 1273K. If the ratio of the stagnation temperatures at the exit and entry is 3.75, determine the Mach number, pressure and temperature of the gas at entry. What is the amount of heat added and the maximum heat that can be added? (16) [Nov/Dec-2006]

10. Show that flow after normal shock is always subsonic. (8) [Nov/Dec-2006]

11. The ratio of the 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, T=290K 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, Mach number, temperature and pressure. What is the stagnation pressure loss between the initial and final states of the flow? (8) [Nov/Dec-2006], (16) [May/June-2013]

12. Plot and explain the wave pattern at the exit of a overexpanded nozzle. (6) [Nov/Dec-2006]

13. Determine the temperature and pressure field around a symmetric double wedge of 20⁰ included angle kept at 15⁰ angle of attack to a supersonic stream of Mach number 2.5, by the shock-expansion theory. (10) [Nov/Dec-2006]

14. A jet of air at 275K and 0.69bar has an initial Mach number of 2.0. If it passes through a normal shock wave, determine

                                i.            Mach number

                              ii.            Pressure and temperature,

                            iii.            Speed of sound and

                            iv.            Jet velocity downstream of the shock. (16) [April/May-2008]

15. An air stream at a Mach number of 2.0 is isentropically deflected by 10⁰ in the clockwise direction. If the initial pressure and temperature are 98KN/m² and 97⁰C, determine the final state of air after expansion. (16) [April/May-2008]

16. The Mach number at 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 the gas at exit are 2.5bar and 1273K 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. (16) [Nov/Dec-2008]

17. Air flows with negligible friction in a constant area duct. At section one, the flow properties are t₁=60.4⁰C, p₁=135KPa absolute and velocity 732m/s. Heat is added to the flow between section one and section two, where the Mach number is 1.2. Determine the flow properties at section two, the heat transfer per unit mass and entropy change. (16) [May/June-2009]

18. Discuss under what condition a compression wave changes into a shock wave. (4) [Nov/Dec-2009]

19. The state of a gas (r=1.3; R=0.469KJ/Kg-k) up stream of a normal shock wave is given by the following data.

            M_x=2.5; P_x=2bar; T_x=275⁰K.

Calculate the Mach number, pressure, temperature of the gas downstream of the shock. (12) [Nov/Dec-2009]

20. What is meant by normal shock in a flow through convergent and divergent nozzle? Explain. (8) [Nov/Dec-2009]

21. If a normal shock in a flow of Nitrogen at a velocity of 700meter/sec, find the Mach number and the properties at the downstream of shock. The temperature and pressure before shock are 30⁰C and 2bar respectively. (8) [Nov/Dec-2009]

22. Define the Rankine-Hergonoit relations. (6) [April/May-2010]

24. Air flows through a constant area duct whose walls are kept at a low temperature. The air enters the pipe at a Mach number of 0.52, a pressure of 200KPa, and a temperature of 623K. The rate of heat transfer from the air to the walls of pipe is estimated to be 400KJ/Kg of air. Find the Mach number, temperature and pressure at the exit of the pipe. Assume that the flow is steady, that the effects of wall friction are negligible, and that the air behaves as a perfect gas. (16) [Nov/Dec-2010]

25. Derive the equation for Mach number in the downstream of the normal shock wave. (8) [May/June-2012]

26. The velocity of a normal shock wave moving into stagnant air (p=1.0bar, t=17⁰C) is 500m/s. If the area of cross section of the duct is constant determine pressure, temperature, velocity of air, stagnation temperature, and Mach number imparted upstream of the wave front. (8) [May/June-2012]

27. Air approaches a symmetrical wedge (angle of deflection δ=15⁰) at a Mach number of 2. Consider strong waves conditions. Determine the wave angle, pressure ratio, density ratio, temperature ratio and downstream Mach number. (8) [May/June-2012]

28. Derive the equation for static pressure ratio across the oblique shock waves. (8) [May/June-2012]

29. Air is flowing in an insulated duct with a Mach number of M₁=0.25. At a section downstream entropy is greater by an amount of 0.124KJ/Kg-K as a result of friction. What is the Mach number at this section? The static properties at inlet are 700KPa and 60⁰C. Find velocity, temperature and pressure at exit. Find properties at the critical section. (16) [May/June-2012]

30. Adiabatic flow of air takes place in a constant area duct. Because of friction, the Mach number increases from 0.3 to 0.7. The initial temperature of air is 400K and the pressure is 20bar. Determine (i) the final pressure (ii) the final temperature (iii) the ratio of density and (iv) mass flow per unit cross section. (16) [Nov/Dec-2012]

31. What are the effects of heat addition and removal from a gas during Rayleigh flow? (6) [Nov/Dec-2012]

32. Prove that in a Rayleigh line at maximum entropy point Mach number is unity. (10) [Nov/Dec-2012]

33. Air is flowing in an insulated duct with friction coefficient f=0.002. At inlet, velocity of air is 130m/s, pressure is 250KPa, and temperature is 400K. Determine the following:

                                i.            Length of the pipe required so as to give 20% drop in stagnation pressure

                              ii.            Properties of air at a section 3.5m from inlet, and

                            iii.            Maximum pipe length. (16) [May/June-2013]

34. Air enters a constant area pipe with velocity 150m/s, temperature 60⁰C and pressure 0.5MN/m². If 180KJ/Kg of heat is added to the pipe, find

                    i.            The final pressure

                  ii.            The final Mach number and

                iii.            The change in stagnation pressure and entropy. (16) [May/June-2013]

35. A normal shock occurs in the diverging section of a convergent-divergent air nozzle. The throat area is 1/3 times exit area and the static pressure at exit is 0.4 times the stagnation pressure at the entry. The flow is throughout isentropic except through the shock. Determine:

                                i.            Mach number M_x and M_y

                              ii.            The static pressure and

                            iii.            The area of cross section of the nozzle at the section of nozzle where the normal shock occurs. (16) [May/June-2014]

36. A gas (γ=1.3) at p₁=345 mbar, T₁=350K and M₁=1.5 is to be isentropically expanded to 138 mbar. Determine:

                                i.            The deflection angle

                              ii.            Final Mach number

                            iii.            The temperature of the gas. (16) [May/June-2014]

37. Derive the Prandtl’ equation for flow through an oblique shock

            α^X2 - (r-1/r+1) C_t²=

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