ANNA UNIVERSITY-CHENNAI
REGULATION 2013
DEPARTMENT OF MECHANICAL ENGINEERING
Unit-III Normal and oblique shocks
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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