Practice Test 2:
Exam will cover: from Area change through shock interactions
Will not cover: expansion waves
One multipart short answer
Three calculator type problems

Nozzles

Diffusers

Oblique shocks

Shock interactions

Shock interactions with walls
1.
(T2.1)
a.
Sketch a Mach reflection and describe the condition at which it will occur.
b.
Describe the ‘jump’ conditions across the slip line.
c.
A windtunnel is designed for a test section Mach number of 4.0. What is the smallest
ratio of the diffuser throat area that will enable the tunnel to start.
2.
(T2.2) Consider the airfoil shape shown below.
a.
Derive the equation for the lift coefficient.
b.
Calculate the lift coefficient using shock/expansion theory
3.
(T2.3) A supersonic windtunnel is designed for a test section Mach number of 3.0. Given that
the nozzle throat area is 0.1 m
2
, the reservoir stagnation pressure is 20 atm, and the diffuser exit
area is 0.59 m
2
and assuming that the diffuser is properly designed and the tunnel is started.
a.
Find the diffuser Mach number.
b.
Find the back pressure required to stabilize the terminating shock exactly at the
diffuser throat.
c.
At what back pressure would the terminating shock sit exactly at the diffuser exit?
4.
(T2.4) Consider the picture below.
a.
Find the angle so that the reflected shock just touches the leading edge of the flat
plate.
b.
Find the Mach number and the flow direction (relative to the horizontal) in Region 4.
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(H.6.1) Air flows isentropically at a rate of 0.5 kg/s through a convergingdiverging nozzle. At the
inlet, the pressure is 650 kPa, the temperature is 285 K, and the area is 0.00065 m
2
. If the exit
area is 0.0013 m
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 Spring '11
 Dr.Edwards
 Fluid Dynamics, Aerodynamics, 1 m, Shock wave, 0.5 kg

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