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)

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