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Friday, 27 November 2015

ME 6604 Gas Dynamics and Jet propulsion University question bank

ANNA UNIVERSITY-CHENNAI

REGULATION 2013

DEPARTMENT OF MECHANICAL ENGINEERING

Click to download: ME6602 Automobile Engg Unit wise important university questions

Click ME6602 Automobile Engg Anna University Question April May 2014

ME6604-GAS DYNAMICS AND JET PROPULSION

Unit IV- Jet propulsion

Part A (2 Marks)

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 (16 Marks)

1.Differentiate turbojet and turboprop propulsion engines with suitable diagrams

(AU: May2012)

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

(AU: May 2012)

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.

(AU: May 2012)

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

(AU: May 2011)

5.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)

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

7.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)

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

9.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)

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

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The question bank for unit 1,2,3,4,5 were uploaded from the following link or copy paste the following link

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http://annaunivstudymaterials.blogspot.in/2015/11/me6604-gas-dynamics-and-jet-propulsion_91.html

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http://annaunivstudymaterials.blogspot.in/2015/11/university-questions-gas-dynamics-and.html

ME6604 Gas dynamics and Jet propulsion lecture notes

ANNA UNIVERSITY-CHENNAI

REGULATION 2013

DEPARTMENT OF MECHANICAL ENGINEERING

ME6604-GAS DYNAMICS AND JET PROPULSION

Unit-III Normal and oblique shocks

Part A (2 Marks)
Click:DME- Previous 5 years Anna University questions

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)

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

(AU: May 2012)

2.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)

3.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)

4.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)

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

6.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)

7.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)

8.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)

9.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)

10.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)

11.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 36 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)

12.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 60'. Determine the deflection angle on the wall, pressure and temperature ratios and final Mach number.

(AU: Dec 2007)

The question bank for unit 1,2,3,4,5 were uploaded from the following link or copy paste the following link

Click:DME- Previous 5 years Anna University questions

ME6604 Gas Dynamics and Jet propulsion question bank PDF

ANNA UNIVERSITY-CHENNAI

REGULATION 2013

DEPARTMENT OF MECHANICAL ENGINEERING

Click to download: ME6602 Automobile Engg Unit wise important university questions

Click ME6602 Automobile Engg Anna University Question April May 2014

Click:DME- Previous 5 years Anna University questions

ME6604-GAS DYNAMICS AND JET PROPULSION

Unit I – Basic concepts and Isentropic flows

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.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)

2.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)

3.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)

5.4.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)

6.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)

7.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, 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)

8.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)

9.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)

10.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)

11.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)

12.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)

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

(AU: Dec 2008)

Click:DME- Previous 5 years Anna University questions

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Friday, 27 November 2015

ME 6604 Gas Dynamics and Jet propulsion University question bank

ME6604 Gas dynamics and Jet propulsion lecture notes

ME6604 Gas Dynamics and Jet propulsion question bank PDF