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Can you please give me the Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics previous years question papers as it is very urgent for me?
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Old May 1st, 2014, 02:49 PM
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Default Re: Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics previo

As you want to get the Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics previous years question papers so here is the information of the same for you:

Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics previous years question papers

PART – A (10 x 2 = 20)
Answer All the Questions
1. What is a continuum?
2. What are Reynolds and mach number?
3. Write down the four basic equations which satisfy the state points
before and after a normal shock wave.
4. What is a shock polar?
5. What is meant by drag force?
6. What is mach angle?
7. Differentiate between normal shock and oblique shock.
8. Define critical mach number.
9. What is supersonic tunnel?
10. Define specific Thrust.

PART – B (5 x 12 = 60)
Answer All the Questions
11. For the adiabatic flow or a perfect gas show that

12. 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.
(= 1.3, R = 0.469 kJ/kg K.
13. 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 = 290k is 2.2. There is a standing normal shock
wave just outside the diffuser entry. The flow in the diffuser is
sentropic. Determine at the exit of the diffuser (a) mach number
(b) Temperature (c) pressure what is the stagnation pressure loss
between the initial and final states of the flow.

14. Derive the rankine – Hugoniot relation for an oblique shock.
Compare graphically the variation of density ratio with the
initial mach number in isentropic flow and flow with oblique

15. Describe with the aid of sketches the development of a finite
amplitude rarefaction wave show the directions of flow and the
wave propagation.

16. The density ratio across a steep pressure wave moving into
stagnant air in a constant area duet is 2.0 calculate (a) the
pressure and temperature ratios across the wave and (b) the wave
mach number and the mach number of the induced flow.

17. Discuss: (a) Shock induced separation.
(b) Lift and drag divergence.
In high speed flows of Airfoil.

18. Discuss about the characteristics of swept wings in Airfoils

19. Discuss about ht eindraft and induction tunnel layouts with their
design features in high sped wind tunnels.

20. Explain (a) Transonic tunnels
(b) Supersonic tunnels
(c) Hypersonic tunnels
With their peculiarities.

PART – A (10 x 2 = 20)
Answer All the Questions
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

10. What is a shock tube?

PART – B (5 x 12 = 60)
Answer All the Questions
11. Derive the relation between pressure and temperature at inlet and
outlet for compressible fluids.
P2 / P1 = (T2 / T1)/ -1

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.

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.

16. The upper and lower surfaces of a symmetrical 2-D aerofoil are given
by Z= _(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

17. (a) What is transonic region (2)
(b) Derive the equation for transonic flow (10)

18. Explain the effects of thickness, camber and aspect ration of wings.

19. What are the various tunnel layouts and explain with design features?

20. What are the flow visualization methods and explain optical methods.

Contact Details:
Sathayabama University
Jeppiaar Nagar,
Rajiv Gandhi Road,
Tamil Nadu 600 119 ‎
044 2450 3150

Map Location:
Answered By StudyChaCha Member

Last edited by AdityaV; February 16th, 2015 at 04:50 PM.
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Old February 9th, 2015, 11:02 AM
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Join Date: Apr 2013
Default Re: Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics previo

Here I am providing the list of few questions of Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics exam question paper which you are looking for .
Sathyabama University-B.E in Aeronautical Engineering-5th Sem Aerodynamics exam question paper
PART - A (10 X 2 = 20)
Answer ALL the Questions
1. What are the equations required to define completely the
compressible flow?
2. Why should passage area increase with velocity in supersonic
3. What is shock polar? How is it useful?
4. What is Rayleigh flow?
5. What are the rules for reflection of shock and expansion waves?
6. Write the wave equation for supersonic small perturbation theory.
What are its solutions?
7. What is the need for the linear theory of supersonic flows?
8. Explain shock stall.
9. What is the basic optical principle of Schlieren technique?
10. What is the specific advantage of an induction tunnel?
PART – B (5 x 12 = 60)
Answer All the Questions
11. Discuss the performance of a convergent-divergent nozzle under
constant stagnation pressure and variable back pressure.
12. Starting from the energy equation for adiabatic flow derive a
relation between the flow Mach number and the characteristic
Mach number.
13. Derive the Rankine-Hugoniot pressure density relationship for
the shock and explain its significance.
14. Derive Prandtl relation for a normal shock and explain its
15. The exit pressure and Temperature are 125kPa and 20°C
respectively as air flows through a 25cm diameter pipe at
1200m3/min. The pipe is 60m long. Assuming the friction factor
f=0.005 estimate the inlet pressure and temperature.
16. The pressure and temperature of the air at inlet to a constant area
duct are 120kPa and 150°C respectively, with an inlet Mach
number of 3.0. Heat is transferred to the air as it flows through
the duct leading to an exit Mach number of 1.5. Find (i) the
pressure and temperature at exit. (ii) the maximum amount of
heat that can be transferred to the air if no shocks occur in the
flow. (iii) the exit pressure and temperature with the maximum
heat transfer.
17. Derive an expression for the CL and CD of a symmetric diamond
profile in supersonic flow at small angle of attack.
18. (a) Write short note on critical Mach number and suggest
methods to improve the critical Mach number.
(b) Explain Transonic and supersonic area rules.
19. What are the special problems of operating a wind tunnel in
Transonic, Supersonic and hypersonic Mach numbers?
20. Write a note on the optical methods of flow visualization in
supersonic wind tunnel and explain any one in detail.
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