# Gas Dynamics MCQ (Multiple Choice Questions)

1. What is the branch of fluid dynamics that deals with the behavior of gases, particularly under high-speed and compressible conditions?
a) Aerodynamics
b) Gas Dynamics
c) Fluid Mechanics
d) Thermodynamics

Explanation: Gas Dynamics specifically focuses on the study of the behavior of gases, particularly under conditions where compressibility effects are significant, such as high speeds and changes in pressure.

2. How the condensation affects the flow Mach number and pressure at supersonic speeds?
a) Both the static pressure and Mach number Increases
b) The static pressure increases and the Mach number decreases
c) Both the static pressure and Mach number decreases
d) The static pressure decreases and Mach number Increases

Explanation: The changes in flow properties due to condensation depends upon the amount of heat released through condensation. At supersonic speeds, the flow temperature at divergent portion is low compared to the inlet temperature. This cause the air moisture to condense inside the nozzle and hence at supersonic speed with increase in condensation Mach number increase and in turn pressure decreases.

3. Which of the following C-D nozzle has minimum divergence losses?
a) Bell nozzle
b) Annular nozzle
c) Aerospike nozzle
d) Conical nozzle

Explanation: The Bell nozzle out of other nozzles have more divergence angle at the throat that moderately decreases up to exit nearly to about 2° to 5°, while the conical nozzle has a constant divergence angle. As a result of this, the divergence losses at throat decreases compared to other nozzles.

4. In gas dynamics, what term describes the ratio of the speed of an object to the speed of sound in the surrounding medium?
a) Froude Number
b) Knudsen Number
c) Mach Number
d) Reynolds Number

Explanation: The Mach number represents the ratio of the speed of an object to the speed of sound in the medium. It is a dimensionless quantity commonly used to characterize flow regimes in gas dynamics.

5. What type of gas flow occurs when the flow velocity is less than the speed of sound and disturbances propagate upstream?
a) Supersonic Flow
b) Subsonic Flow
c) Hypersonic Flow
d) Transonic Flow

Explanation: Subsonic flow refers to gas flow where the velocity is less than the speed of sound, and disturbances propagate upstream, commonly encountered in many engineering applications.

6. The gas stored in the reservoir having pressure 7 bar and temperature 300 K exhausts from the nozzle which has an exit area of 0.025 m2 to surroundings at velocity 570 m/s. The mass flow rate through the nozzle is 23 Kg/s and surrounding having pressure of 1 bar. Calculate the thrust produced by the nozzle.
a) 33.69 KN
b) 26.45 KN
c) 18.83 KN
d) 28.11 KN

Explanation: The nozzle thrust equation is defined as;
T = $$\dot{m}$$Ve + (Pe – Pb) Ae
Now, Me = Ve/a = 1.67
Where a = $$\sqrt {γRT} = \sqrt {1.4*287*288}$$ = 340.17 m/s
Thus, Ve = Me*a = 1.67*340.17 = 568 m/s
And T = 23*568+(7*105-1*105)*0.025
T = 28.11 KN

7. Which is the desired angle for a diffuser or diffuser walls?
a) 7°
b) 8°
c) 5°
d) 6°

Explanation: A supersonic wind tunnel should be designed to have a small angle diffuser, with an angle of about 5° between opposite walls. As a result of this diffuser length increases which allows incorporating the devices that help to prevent flow separation.

8. What will be the value of the characteristics Mach number for air if the local Mach number tends to reach infinity?
a) Infinite
b) 2.45
c) 1.37
d) 1

Explanation: The relation between characteristics Mach number and local Mach number is;
M2=$$\frac {2}{((γ+1/M^{*2} )-(γ-1) )}$$
Now at M → ∞, M* → $$\sqrt {\frac {γ+1}{γ-1}}$$ → 2.45 (As for air, γ = 1.4)

9. Which kind of flow is observed cross the normal shock wave?
a) Reversible
b) Isothermal
c) Isentropic

Explanation: The flow across the normal shock wave is adiabatic as there is no heat transfer takes place during the process and the temperature increases due to the conversion of kinetic energy into internal energy across the shock wave.

10. If the piston moving inside the shock tube is having a velocity of 100 m/s with a shock wave being created having a density ratio 1.7 bar across it. Calculate the shock speed.
a) 242.85 m/s
b) 214.59 m/s
c) 165.39 m/s
d) 100 m/s

Explanation: The piston speed in terms of density ratio is;
Vp = cs(1-$$\frac {\rho_1}{\rho_2}$$)
100 = cs(1-$$\frac {1}{1.7}$$)
Cs = 242.85 m/s

11. For the shock wave reflection from the open end, which of the following condition is satisfied?
a) Velocity boundary condition
b) Density boundary condition
c) Pressure boundary condition
d) Temperature boundary condition

Explanation: When the Shock wave is incident on the open end, it either gets reflected into compression or expansion wave, to adjust the flow pressure with the boundary pressure. Hence for a shock wave reflection from the open end, the pressure boundary condition is satisfied.

12. When is flow considered to be compressible?
a) If the flow Mach number is greater than one
b) If the flow is more than Mach number 0.3 or the percentage change in density is more than 5%
c) If the percentage change in the density of the flow is more than 15%
d) If the flow is less than Mach number 0.5 or the percentage change in density is more than 10%

Explanation: Compressible effects are observed in all flow regimes but its effects are only significant from Mach number 0.3 or if the percentage change in density of the flow is more than 5%.

13. What is true about a supersonic flow around a wedge?
a) The flow changes its direction abruptly and pressure decreases with acceleration
b) The flow changes its direction smoothly and pressure increases downstream with acceleration
c) The flow changes its direction smoothly and pressure decreases with acceleration
d) There is a sudden change in flow direction at the body and pressure increases downstream of the shock

Explanation: When the flow is supersonic around a wedge, it is observed that there is a sudden change in flow direction at the body and pressure increases downstream of the shock.

14. If the Mach number is increased keeping the same wall deflection angle, then what is going to happen to shock angle?
a) Shock gets detached
b) Shock angle is not affected by Mach number
c) Shock angle decreases
d) Shock angle increases

Explanation: In θ-β-M relation, where θ represents wall deflection angle, β represents shock angle and M represents Mach number, it shows that with increasing Mach number and constant wall deflection angle, the shock angle decreases.

15. What is true for smooth supersonic compression through weak shocks?
a) Change in flow velocity is achieved in a shorter distance when closer to the wall than further away
b) Velocity downstream of the smooth supersonic compression is subsonic
c) Increase in entropy by a number of weak shocks is greater than the increase by one single shock for a given deflection
d) Mach lines formed due to small deflection are divergent

Explanation: Increase in entropy by a number of weak shocks is smaller than a single shock for a given deflection angle. Reduction in entropy is of factor $$\frac {1}{n^2}$$, for ‘n’ weak shock waves. Also, the Mach lines are convergent in nature and hence, the change in flow velocity is achieved in a shorter distance when further away from the wall than closer to it. The deflection angle is taken to be small for smooth supersonic compression to obtain supersonic flow downstream.

16. What does the Prandtl-Meyer function physically represent in a turning flow?
a) Mach wave angle of the centered expansion fan wave
b) The wall deflection angle measured from the vertical axis perpendicular to the upstream flow
c) The flow inclination of the wave with respect to the horizontal axis parallel to the upstream flow
d) The inclination angle of the wave with respect to the Mach line corresponding to Mach 1

Explanation: The Prandtl-Meyer function denoted by ν(M) is a function of the Mach number of the flow. Physically it represents the flow inclination angle of the wave corresponding to the line with zero turning angle (∅ = 0). When Mach number is equal to 1 the function becomes zero and hence the angle is measured with respect to this line.

17. What is the shock wave angle of the reflected shock with respect to the horizontal, when a shock is reflected from the surface with shock wave angle given as β=40 and flow Mach number is 2?
a) 50.2°
b) 45.6°
c) 39.7°
d) 37.9°

Explanation: From θ-β-M graph, for M1=2 and β1=40, θ=10.5°. Now from oblique shock relation tables, we can find M2=1.63. Since the flow direction is unchanged after crossing the reflected wave, from the θ-β-M graph, for M2=2 and θ=10.5°, β2=50.2°. So the reflected shock wave angle w.r.t the ground is βr2-θ=50.2-10.5=39.7°.

18. A uniform supersonic flow at Mach number 2 expands through two convex corners of 10° each. What is the angle of the second expansion fan?
a) 46.38°
b) 14.16°
c) 36.38°
d) 26.38°

Explanation: From Prandtl-Meyer functions table for M1 = 2, ν1 = 26.38°. We have ν2 = ν1 + θ = 36.38°. Again referring to the table, we get M2 = 2.38. The Prandtl-Meyer function after the second fan is ν3 = ν2 + θ = 46.38° and corresponding Mach number M3 = 2.83. The Mach angles are μ1 = 30, μ2 = 24.84, μ3 = 20.69. The angle of the second fan is μ23 = θ + μ2 – μ3 = 14.16°.

19. What is the result of the interaction between a normal shock and right running oblique shock?
a) Formation of lower strength normal shock
b) Formation of higher strength normal shock
c) It gives rise to a reflected right running oblique shock
d) It gives rise to a reflected left running oblique shock

Explanation: Intersection of a normal shock and right running oblique shock gives rise to a left running reflected shock in order to bring the flow to its original direction. The reflected oblique shock is of lesser strength than the right running oblique shock.

20. What is the lift and drag value per unit span when a thin diamond wedge airfoil of thickness t and pressure above the top surface is p2 and p3 at zero angle of attack?
a) Lift = 0 and Drag = (p2 – p3)t
b) Lift = 0 and Drag = (p3 – p2)t
c) Drag = 0 and Lift = (p2 – p3)t
d) Drag = 0 and Lift = (p2 + p3)t

Explanation: Since the angle of attack is zero, the pressure on the top and bottom surface is equal which leads to zero lift. For drag, it is given by, Drag = 2(p2t – p3t) × $$\frac {1}{2}$$ = (p2 – p3)t.

## Chapterwise Multiple Choice Questions on Gas Dynamics

Our MCQs focus on all topics of the Gas Dynamics subject, covering all topics. This will help you to prepare for exams, contests, online tests, quizzes, viva-voce, interviews, and certifications. You can practice these MCQs chapter by chapter starting from the 1st chapter or you can jump to any chapter of your choice.

## 1. MCQ on Basic Equations of Compressible Flow

The section contains Gas Dynamics multiple choice questions and answers on the first law of thermodynamics, the second law of thermodynamics, and thermal properties of perfect gases.

## 2. Gas Dynamics Questions on Steady One-Dimensional Flow

The section covers questions and answers on steady one-dimensional flow, De Laval nozzle, supersonic flow generation, and diffusers.

## 3. Gas Dynamics MCQ on Normal Shock Waves

The section contains MCQs on normal shock waves, Hugoniot equation, propagating shock waves, reflected shock waves, shock tubes, and Mach angle.

## 4. Oblique Shock and Expansion Waves

The section contains multiple choice questions and answers on oblique shock and expansion waves, supersonic flow over a wedge, the Prandtl-Meyer expansion, reflection and intersection of shocks and expansion waves, detached shocks, and Mach reflection.

If you would like to learn "Gas Dynamics" thoroughly, you should attempt to work on the complete set of 1000+ MCQs - multiple choice questions and answers mentioned above. It will immensely help anyone trying to crack an exam or an interview.

Wish you the best in your endeavor to learn and master Gas Dynamics!

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