# Aerodynamics – II BE Fifth Sem

1. What is a De Laval nozzle and state the function of the same.
2. What are stagnation conditions?
3. What is a shock wave?
4. Define oblique shock wave.
5. State small Perturbation theory.
6. What is meant by Prandtl Glauert transformation?
7. What is lower and upper critical Mach number?
8. Name the characteristics of swept wings.
9. What are the main points to be considered while designing the tunnel
layouts?
10. What is a shock tube?
PART – B (5 x 12 = 60)
11. Derive the relation between pressure and temperature at inlet and
outlet for compressible fluids.
P2 / P1 = (T2 / T1)g / g-1
(or)
12. A De Laval nozzle has to be designed for an exit Mach number of 1.5
with exit diameter of 200 mm. Find the ration of throat area to exit
area necessary. The reservoir conditions are given as:
Po = 1 atm. (gauge) To = 20° c
Find also the maximum flow rate through the nozzle. What will be
the exit pressure and temperature.

13. A normal shock wave moves at a constant speed of 500 m/s into still
air at 0° c and 0.7 atm. Determine the static and stagnation conditions
present in the air after the passage of wave.
(or)
14. Air flows above a frictionless surface having a sharp corner. The
flow angles and mach number in downstream from the corner are –
60° and 4.0 respectively. Calculate the upstream Mach number for
flow angle of 15° clockwise and 15° counter clockwise.
15. A missile has a conical nose with a semi vertex angle of 4° and is
subjected to a Mach number of 12 under actual conditions. A model
of the missile has to be tested in a supersonic wind tunnel at a test
section Mach number of 2.5. calculate the semi vertex angle of the
conical nose of the model.
(or)
16. The upper and lower surfaces of a symmetrical 2-D aerofoil are given
by Z= ± c (1-/c)2 where c is the chord and  <<1. The aerofoil is
at zero distance in a steady supersonic stream of Mach number M in
the positive  direction.
(a) Find the velocity components according to the linear theory
in the upper region of disturbance.
(b) Show that the drag coefficient of the aerofoil is given by
2 1/ 2
2
2
15 ( 1)
8

=
¥
M
Cd
e
e
17. (a) What is transonic region (2)
(b) Derive the equation for transonic flow (10)
(or)
18. Explain the effects of thickness, camber and aspect ration of wings.
19. What are the various tunnel layouts and explain with design features?
(or)
20. What are the flow visualization methods and explain optical methods.