Aerodynamics - Normal Shock Waves Quiz(MCQ)

1 .

Which of these is incorrect for the normal shock wave analysis?

A)

Steady flow

B)

Adiabatic flow

C)

No viscous effects or body forces

D)

Differential form of conservation equations used

Correct Answer : Differential form of conservation equations used

Explanation : The normal shock wave analysis uses the control volume approach. The integral form of conservation equations are applied to the control volume. The flow is steady, adiabatic, without viscous effects and zero body forces.

2 .

The supersonic flow over a blunt body is given. Mark the area where the normal shock wave exists.

A)

i

B)

ii

C)

iii

D)

iv

Correct Answer : In the figure, a strong bow shock exists in front of the flow. The shock wave is curved around the blunt body, but the region of the shock closest to the nose is essentially perpendicular to the flow, i.e. region i is a normal shock.

3 .

Which is not true for continuity equation of normal shock?

A)

ρ₁u₁A₁=ρ₂u₂A₂

B)

ρ₁u₁=ρ₂u₂

C)

No viscosity

D)

Steady flow

Correct Answer : ρ₁u₁=ρ₂u₂

Explaination : The continuity equation of the normal shock equation is derived from the continuity equation of mass and is given as ρ₁u₁A₁=ρ₂u₂A₂. If the area of the control volume is same on both sides, it becomes ρ₁u₁=ρ₂u₂. The flow is steady and without viscosity.

4 .

The most important quantity that dominates the physical properties of compressible flow is _______

A)

Speed of light

B)

Speed of sound

C)

Distance of propagation

D)

Density of medium of propagation

Correct Answer : Speed of sound

Explanation : The speed of sound is the quantity that dominates the physical properties of compressible flow. Speed of light is not important in compressible flow problems. Density of medium and distance of propagation are secondary quantities.

5 .

Mach number is a measure of the directed motion of the gas compared with the random thermal motion of the molecules.

A)

True

B)

False

C)

Can Not Say

D)

None of the above

Correct Answer : True

Explanation : This comes from the physical meaning of the Mach number. When we take the ratios of per unit mass kinetic energy to the potential energy of the fluid particle moving along a streamline, it comes proportional to the square of the Mach number. Thus, this is true.

6 .

Select the incorrect statement out of the following.

A)

lim_{τ_s→0?} is the case of an incompressible flow

B)

Incompressible flow is theoretically zero-Mach number flows

C)

Speed of sound in incompressible medium is zero

D)

Mach number for finite velocity object, in incompressible flow is zero

Correct Answer : Speed of sound in incompressible medium is zero

Explaination : Incompressible flow is the limiting case of isentropic compressibility being zero. This gives an infinite speed of sound and a zero Mach number in that medium. Thus, theoretically, zero-Mach number flows are incompressible.

7 .

For a perfect gas, if we half the density, keeping the temperature same the speed of sound changes. This is due to the change in pressure.

A)

True

B)

False

C)

Can Not Say

D)

None of the above

Correct Answer : False

Explanation : The given statement is false. For a perfect gas, the speed of sound is a pressure of temperature only. And hence, by changing pressure or density but keeping the temperature same, speed of sound will not change.

8 .

Which is not the correct reason for sound of speed being higher in helium than in air at the same temperature?

A)

Helium is lighter than air

B)

Y for helium is higher than air

C)

Gas constant is same for both

D)

R for helium is much larger than for air

Correct Answer : Gas constant is same for both

Explanation : The sound of speed depends on gamma, T and the gas constant R for the respective gas. In case of helium and air, the gamma for helium is higher than air. Also, helium is lighter than air thus, R for helium being higher. This gives, at the same temperature, speed of sound more in helium than air.

9 .

The correct statement for a point where the speed of sound is a, is_______

A)

a* is the stagnation speed of sound

B)

a* is the maximum speed of sound

C)

a₀ is the stagnation speed of sound

D)

a₀ is the characteristic speed of sound

Correct Answer : a₀ is the stagnation speed of sound

Explaination : For a point in the flow, where the speed of sound is a, a₀ is the stagnation speed of sound associated with that point. While a* is the sonic characteristic value associated with that same point. None of these is the maximum speed of sound in the flow.

10 .

Select the incorrect statement for the properties concerned with a flow.

A)

a* and a₀ are constant at a point

B)

a* and a₀ are related but not equal

C)

a* and a₀ are constant along the streamline for an adiabatic, inviscid, steady flow

D)

a* and a₀ are constant along the entire flow for an adiabatic, inviscid, steady flow

Correct Answer : a* and a₀ are constant along the entire flow for an adiabatic, inviscid, steady flow

Explaination : a* and a₀ are the defined properties of the flow, constant at a point. They are related to each other but not same. For an adiabatic, inviscid, steady flow they are constant along the streamline. And if all the streamlines are coming from same uniform free-stream, they are constant along the entire flow.

11 .

For a calorically perfect gas, the ratio of stagnation to static temperature depends upon_______

A)

γ

B)

γ , M

C)

M

D)

R, M

Correct Answer : γ , M

Explaination : The stagnation and static temperatures for a calorically perfect gas are related with the equation

Thus, it is clearly seen that this ratio depends on the gamma and the Mach number in the medium. It’s a very important relationship.

12 .

The definitions of P₀ and ρ₀ involve______

A)

Isentropic compression

B)

Isothermal compression

C)

Adiabatic compression

D)

Any process gives the same value

Correct Answer : Isentropic compression

Explaination : The defined properties of the flow P₀ and ρ₀ involve isentropic compression, which is adiabatic and reversible both. These properties are the values of the flow parameters when the flow is isentropically compressed to zero velocity, at a point with properties P and ρ.

13 .

The characteristic Mach number is constant for a given value of temperature and gamma.

A)

True

B)

False

C)

Can Not Say

D)

None of the above

Correct Answer : False

Explanation : The defined M* or the characteristic Mach number depends upon the Mach number, for any given gamma. It changes if the Mach number is changed, which changes by changing the flow speed at any temperature. Hence, the given statement is false.

14 .

At Mach number M = 6.3, the gas can be termed as a perfect gas.

A)

True

B)

False

C)

Can Not Say

D)

None of the above

Correct Answer : False

Explanation : At M = 6.3, the flow is termed as hypersonic. Due to this the additional flow properties related to hypersonic flow such as high temperature and low density effects comes into account. And because of these parameters the flow cannot be treated as perfect gas anymore.

15 .

If the downstream Mach number behind the shock is 0.8, calculate the strength of the Shock.

A)

0.523

B)

0.692

C)

0.724

D)

0.956

Correct Answer : 0.724

16 .

What will be the density ratio for air traveling at Mach number M₁ = infinite?

A)

3

B)

4

C)

5

D)

6

Correct Answer : 6

17 .

If the change in entropy across the shock is 425 J/K, calculate the total pressure ratio across the shock for an air.

A)

1.23

B)

2.71

C)

2.98

D)

3.16

Correct Answer : 2.71

18 .

Which kind of flow is observed cross the normal shock wave?

A)

Adiabatic

B)

Isentropic

C)

Reversible

D)

Isothermal

Correct Answer : Adiabatic

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.

19 .

How does the flow velocity changes across the normal shock wave?

A)

increases

B)

Decreases

C)

Becomes zero

D)

Remains the same

Correct Answer : Decreases

20 .

If the shock wave is the infinitely weak normal shock, then what will be the value of downstream Mach number after the shock?

A)

M = 0.8

B)

M = 1.0

C)

M = 1.7

D)

M =1.9

Correct Answer : M = 1.0

21 .

The strength of the shock is defined by _____________

A)

P₁/ΔP

B)

M₂/M₁

C)

ΔP/P₁

D)

M₁/M₂

Correct Answer : ΔP/P₁

Explaination : As the flow passes through shock, its pressure increases. Hence the strength of the shock is governed by the ratio of the difference in pressure across the shock to upstream pressure and is defined as,
(P₂-P₁)/P₁ = ΔP/P₁
Therefore the larger the pressure difference across the shock, the stronger the shock is and vice versa.

22 .

Select the false statement for the normal shock wave.

A)

The total pressure remains constant

B)

The total temperature remains constant

C)

Normal shock wave is a compression wave

D)

Entropy increases across normal shock wave

Correct Answer : The total pressure remains constant

Explanation : Across the normal shock wave, the entropy increases, by second law of thermodynamics. It is a compression wave and the total pressure also changes. It is related to the entropy change inversely. Also, being an adiabatic flow, total temperature remains constant.

23 .

The total temperature across the normal shock wave is constant. Which statement gives the wrong reason for this?

A)

Calorifically perfect gas

B)

Thermal dissipation there

C)

Adiabatic flow across normal shock wave

D)

Normal shock wave has an isentropic flow

Correct Answer : Thermal dissipation there

Explanation : The flow across the normal shock wave is isentropic i.e. it is adiabatic also. Since, we are concerned with calorically perfect gases in an adiabatic, inviscid, steady flow the total temperature remains constant. Thermal dissipation is not a reason for total temperature being constant.

24 .

Specifying the temperature ratio across the normal shock wave will yield which of the following across the normal shock wave?

A)

Upstream flow velocity

B)

Upstream Mach number

C)

Downstream temperature

D)

Downstream flow velocity

Correct Answer : Upstream Mach number

Explanation : Specifying a dimensionless quantity across the normal shock wave will specify all the other ratios and both the upstream and downstream Mach numbers. But for finding the velocity, temperature also needs to be specified and vice versa.

25 .

Select the correct statement for a Mach wave.

A)

M > 1

B)

\(\frac {P_2}{P_1}\)=0.528

C)

\(\frac {T_2}{T_1}\)=1

D)

\(\frac {\rho_2}{\rho_1}\)=∞

Correct Answer : \(\frac {T_2}{T_1}\)=1

Explaination : A Mach wave is a normal shock wave of diminishing strength. It occurs for M=1 upstream. Then downstream M=1 also. And all the ratios are equal to 1, i.e \(\frac {P_2}{P_1}\)=1, \(\frac {T_2}{T_1}\)=1, \(\frac {\rho_2}{\rho_1}\)=1. This can be found by calculation when we put m=1. The properties across Mach wave do not change.

26 .

For a particular gas, Mach number behind the shock wave is a function of which all parameters ahead of the shock wave. Choose the correct option.

A)

Mach number only

B)

Mach number, pressure

C)

Mach number, temperature

D)

Mach number, temperature, pressure

Correct Answer : Mach number only

Explaination : The remarkable result for the normal shock wave is that for a given gas (given gamma), the Mach number ahead of the normal shock wave is a function of the Mach number ahead of the normal shock wave only, irrespective of pressure, density, temperature etc. The values are tabulated and found in gas tables for reference.

27 .

The definition of stagnation pressure can help us calculate the velocity of sound for an incompressible flow. What else do we need for the calculation?

A)

Static velocity

B)

Total pressure

C)

Ambient temperature

D)

Static pressure

Correct Answer : Static pressure

Explaination : The velocity of sound can be calculated using the Pitot tube which measures the stagnation, also called total pressure and measuring the static pressure as well. There is no need to find temperature. And static velocity is nothing but zero velocity.

28 .

The Rayleigh Pitot tube formula does not relate which of these quantities to the others?

A)

Pitot pressure

B)

Upstream flow velocity

C)

Free-stream Mach number

D)

Free-stream static pressure

Correct Answer : Free-stream static pressure

Explanation : The Rayleigh Pitot tube formula helps to calculate the free- stream Mach number using the measured values of free-stream static pressure and Pitot pressure. It is based on the fact that a bow shock is formed at the mouth of the Pitot tube for a compressible supersonic flow. The Pitot pressure is the stagnation pressure behind the bow shock.

29 .

For supersonic compression flows, there is a bow shock formed at the mouth of the Pitot tube.

A)

True

B)

False

C)

Can Not Say

D)

None of the above

Correct Answer : True

Explanation : The supersonic flow has a very high velocity. When it finds an obstruction in the form of a Pitot tube, a bow shock is formed for the supersonic flow over a blunt body. As a result, the stagnation pressure at the mouth of the Pitot tube is the pressure behind the bow shock.

30 .

In the Hugoniot relation, the changes across the shock are related in terms of ________

A)

Electrical properties

B)

Chemical properties

C)

Thermodynamic properties

D)

Mechanical properties

Correct Answer : Thermodynamic properties

Explanation : Since the shock can be visualized as a thermodynamic device that compresses the gas, the Hugoniot relation only relates the thermodynamic quantities across the shock, without any reference to velocity and Mach number.

31 .

Which of the following relation represents the Hugoniot relation?

A)

Δs = -P_av/Δv

B)

Δe = -P_av/Δv

C)

Δe = -P_avΔv

D)

Δs = -P_avΔv

Correct Answer : Δe = -P_avΔv

Explaination : The Hugoniot relation implies that “the change in internal energy is equal to the mean pressure across the shock times the change in specific volume.”
i.e. e₂-e₁ = \(\frac {p_1+p_2}{2}\)(v₁-v₂)
e₂-e₁ = –\(\frac {p_1+p_2}{2}\)(v₂-v₁)
Δe = -P_avΔv

32 .

Which curve is the locus of all possible pressure-volume conditions behind the normal shocks of different strengths for given upstream p₁ and v₁?

A)

Hugoniot curve

B)

Prandlt curve

C)

Isentropic curve

D)

Kelvin-Plank curve

Correct Answer : Hugoniot curve

Explaination : As in thermodynamic equilibrium state the specific energy can be expressed as a function of pressure and specific volume i.e.
e = e(p, v)
Then from the Hugoniot relation, p₂ = f(p₁, v₁, v₂)
This means that for given p₁ and v₁, it represents p₂ as a function of v₂. And the plot of this relation in pv diagram is called Hugoniot curve, and such curve is the locus of all possible pressure-volume conditions behind the normal shocks of different strengths for given upstream p₁ and v₁.

33 .

If the air flow having an upstream pressure 1.89 bar passes through a shock wave, compressing to pressure 2.6 bar across the shock, determine the ratio os specific volume.

A)

0.528

B)

1.254

C)

3.687

D)

4.932

Correct Answer : 1.254

Explaination : According to Hugoniot curve the pressure ratio and density ratio are related by;
\(\frac {p_2}{p_1} = \frac {(\frac {γ+1}{γ-1})(\frac {v_1}{v_2})-1}{(\frac {γ+1}{γ-1})-(\frac {v_1}{v_2} )} \)
Therefore for given flow properties \(\frac {2.6}{1.89}=\frac {6(\frac {v_1}{v_2})-1}{6-(\frac {v_1}{v_2}) }\)
i.e. \(\frac {v_2}{v_1}\) = 1.254

34 .

Across the detonation wave,__________

A)

Pressure decreases and density increases

B)

Pressure increases and density decreases

C)

Both pressure and density decreases

D)

Both pressure and density increases

Correct Answer : Both pressure and density increases

Explanation : A supersonic wave is termed as a detonation wave. Therefore in detonation heat and radial diffusion do not control the velocity, but the shock wave structure developed by the supersonic waves causes the flow to slow down, increasing its pressure, density, and temperature across the shock.

35 .

Along the Hugoniot curve, the pressure is bounded by which of the following boundaries?

A)

∞ ≤ p ≤ 0

B)

0 ≤ p ≤ ∞

C)

1 ≤ p ≤ 0

D)

0 ≤ p ≤ 1

Correct Answer : 0 ≤ p ≤ ∞

Explaination : The Hugoniot curve represented by pv diagram asymptotically approaches to the line
p = -(γ-1)/(γ+1)
As a result of this, the entire range of pressure ratio occurs between zero and infinite, i.e. the pressure on the Hugoniot curve is bounded by 0 ≤ p ≤ ∞.

36 .

The change in flow properties across the sound wave can be assumed as _______

A)

Irreversible

B)

Isentropic

C)

Nonisentropic

D)

Non-adiabatic

Correct Answer : Isentropic

Explanation : A sound wave is an infinitesimally small pressure wave. Hence the changes across the wave are very small and the speed of the process corresponding to these changes is very fast. Hence if there is no heat transfer in the system, the change in flow properties across the sound wave can be assumed as isentropic.

37 .

The temperature ratio across the moving shock is a function of which of the following?

A)

Entropy ratio

B)

Enthalpy ratio

C)

Density ratio

D)

Pressure ratio

Correct Answer : Enthalpy ratio

38 .

A shock of strength 0.528 is generated by piston traveling in air at 288 K inside the shock tube. Determine the shock speed. Assume perfect gas.

A)

343 m/s

B)

379 m/s

C)

410 m/s

D)

543 m/s

Correct Answer : 410 m/s

39 .

For a very small pressure jump across the shock, the shock is termed as _______

A)

Weak shock

B)

Jump shock

C)

Strong shock

D)

Curved shock

Correct Answer : Weak shock

Explanation : For a very small pressure jump across the shock, P₂ is just slightly higher than P₁, and according to the relation of the strength of the shock
(P₂-P₁)/P₁ << 1
Hence since there is a minor change in flow properties across the shock, such shock is called the weak shock.

40 .

What will be the density ratio across the strong shock moving in the fluid medium for a monoatomic gas?

A)

0

B)

4

C)

6

D)

8

Correct Answer : 4

41 .

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)

100 m/s

B)

165.39 m/s

C)

214.59 m/s

D)

242.85 m/s

Correct Answer : 242.85 m/s

42 .

The velocity of the fluid behind the reflected shock wave inside the tube is _______

A)

Zero

B)

Infinite

C)

> velocity ahead of the shock

D)

= velocity ahead of the shock

Correct Answer : Zero

Explaination : As the shock gets reflected from the tub walls and travels back, the strength of the shock is such that the originally induced fluid motion with velocity behind the shock is completely stopped. Hence the fluid velocity behind the reflected shock becomes zero.

43 .

The strength of the reflected shock is _______ compared to the incident shock.

A)

Weaker

B)

Equal

C)

Nullified

D)

Stronger

Correct Answer : Weaker

Explanation : As the flow passes through the incident wave its Mach number decreases as it moves downstream, i.e. M2 < M1. Now this M2 becomes the upstream Mach number M1 for the reflected wave, and since the deflection angles are the same for both incident and reflected waves, while reflected shock has lower upstream Mach number, the strength of the reflected wave is weaker than the incident wave.

44 .

As per the below figure, what will be the velocity of gas ahead of the shock?

A)

C_R

B)

V_p

C)

C_R + V_p

D)

C_R – V_p

Correct Answer : C_R + V_p

Explaination : Since the shock represented in the figure is the reflected shock wave, as the shock moves upstream the velocity of gas relative to the shock wave is C_R + V_p.

45 .

As per the below figure, what will be the velocity of gas ahead of the shock?

A)

C_R

B)

V_p

C)

C_R + V_p

D)

C_R – V_p

Correct Answer : C_R + V_p

Explaination : Since the shock represented in the figure is the reflected shock wave, as the shock moves upstream the velocity of gas relative to the shock wave is C_R + V_p.

46 .

If the shock angle of the incident wave and reflected wave is β1 and β2 respectively, then __________

A)

β2 = 0

B)

β1 > β2

C)

β1 < β2

D)

β1 = β2

Correct Answer : β1 > β2

Explaination : As the wave gets reflected from the wall according to general shock characteristics, its Mach number decreases. Hence the reflected wave now has a lower upstream Mach number than the incident wave, due to which the reflected wave has a lower shock angle than an incident wave, i.e. β1 > β2.

47 .

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

Correct Answer : Pressure 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.

48 .

Choose the correct statement from the following.

A)

Mach number doesn’t depend upon the internal energy

B)

To increase the Mach number, we need to increase both the internal energy and kinetic energy

C)

To increase Mach number, we need to either increase the internal energy or decrease the kinetic energy.

D)

To increase the Mach number, we need to either decrease the internal energy or increase the kinetic energy

Correct Answer : To increase the Mach number, we need to either decrease the internal energy or increase the kinetic energy

Explaination : According to relation, \(\frac {KE}{Internal \, Energy} = \frac {γ(γ-1) M^2}{2}\); it clearly shows that to increase Mach number we need to either decrease internal energy or increase kinetic energy.

49 .

Which of the following is true about the regions on the either sides of the contact surface inside a shock tube?

A)

Both regions have same temperature and pressure but different velocity

B)

Both regions have same velocity and pressure but may have different density and temperature

C)

Both regions have same density and pressure but different temperature

D)

Both regions have same temperature and density but may have different velocity and pressure

Correct Answer : Both regions have same velocity and pressure but may have different density and temperature

Explanation : On the either side of the contact surface, the temperature and density may be different but it is necessary that the pressure and fluid velocity be the same, since contact surface is an imaginary boundary which separates high pressure gases from mixing with low pressure gases.

50 .

When is flow considered to be compressible?

A)

If the flow Mach number is greater than one

B)

If the percentage change in the density of the flow is more than 15%

C)

If the flow is more than Mach number 0.3 or the percentage change in density is more than 5%

D)

If the flow is less than Mach number 0.5 or the percentage change in density is more than 10%

Correct Answer : If the flow is more than Mach number 0.3 or the percentage change in density is more than 5%

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%.

51 .

When are gases said to be ideal?

A)

When gases are thermally perfect

B)

When gases are calorically perfect

C)

When gases are thermally and physically perfect

D)

When gases are thermally and calorically perfect

Correct Answer : When gases are thermally and calorically perfect

Explanation : Gases are said to be ideal when they are thermally and calorically perfect. When the specific heat capacity of the gas does not change and is constant, it is considered as calorically perfect. If internal energy and enthalpy is a function of temperature only then gas is said to be thermally perfect.

52 .

If \(\frac {p_2}{p_1}\) = 29.65 and \(\frac {p_4}{p_1}\) = 35.92, then what is the value of expansion strength inside a shock tube?

A)

0.2357

B)

0.8254

C)

1.2114

D)

2.4229

Correct Answer : 0.8254

Explaination : We know that p₂ = p₃ in a shock tube, so expansion strength is given by, \(\frac {p_3}{p_4} = \frac {p_2}{p_1} \times \frac {p_1} {p_4} = \frac {29.65}{35.92}\) = 0.8254.

53 .

Which relation is of prime importance to find for shock tube operation?

A)

The relation between shock strength and diaphragm pressure ratio

B)

Normal shock relations and Isentropic relations

C)

The relation between diaphragm pressure ratio and reservoir conditions

D)

The relation between normal shock strength and expansion wave pressure ratio

Correct Answer : The relation between shock strength and diaphragm pressure ratio

Explanation : For shock tube operation, it is of prime importance to develop a relation between shock strength \(\frac {p_2}{p_1}\) and diaphragm pressure ratio \(\frac {p_4}{p_1}\). Once the shock strength is known, all other glow quantities can be found from the normal shock relations.

54 .

Which non-equilibrium conditions are created due to severe gradients of flow properties across a shock?

A)

Vibrational stress and heat transfer

B)

Viscous stress and heat transfer

C)

Viscous stress and vibrational stress

D)

Only viscous stress as the flow is adiabatic

Correct Answer : Viscous stress and heat transfer

Explanation : Viscous stress and heat transfer are two non-equilibrium conditions observed across the flow due to high-temperature gradient and pressure gradient.

55 .

What is the Mach wave angle for flow at Mach 1.43?

A)

0.774°

B)

30°

C)

44.37°

D)

45.6°

Correct Answer : 44.37°

56 .

What is the region outside the Mach cone called?

A)

Zone of action

B)

Control volume

C)

Zone of silence

D)

Zone of reaction

Correct Answer : Zone of silence

Explanation : The region with all the disturbances inside the Mach cone of the moving body is called the zone of action. The region outside the Mach cone is known as the zone of silence where the objects are unaware of the disturbances. A control volume is any volume in space where analysis is conducted.

57 .

What happens to the Mach number across a normal shock?

A)

The Mach number will become transonic

B)

The Mach number will become subsonic

C)

The Mach number will become supersonic

D)

The Mach number will become hypersonic

Correct Answer : The Mach number will become subsonic

58 .

What makes an expansion wave?

A)

Collection of Mach waves

B)

Collection of shock waves

C)

Collection of normal shock waves

D)

Collection of oblique shock waves

Correct Answer : Collection of Mach waves

Explanation : Expansion waves are collection of infinite Mach waves which are diverging from a convex corner. They together form an expansion fan and the flow is isentropic across it.

59 .

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 increases

C)

Shock angle is not affected by Mach number

D)

Shock angle decreases

Correct Answer : Shock angle decreases

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.

60 .

Among the following, choose the irreversible process.

A)

Across a Mach wave

B)

Across a subsonic flow

C)

Across a shock wave

D)

Across an expansion wave

Correct Answer : Across a shock wave

Explanation : Entropy increase across a shock wave makes it an irreversible process. The entropy increase can be found out as it is a function of static temperature and pressure ratios across the normal shock.

61 .

Which is not a property of Mach Wave?

A)

Adiabatic

B)

Reversible

C)

Isentropic

D)

Non-isentropic

Correct Answer : Non-isentropic

Explanation : Flow across a Mach wave follows an isentropic process. Isentropic process is characterized by its adiabatic and reversible process. Stagnation properties remains constant across the wave.

62 .

What is true about a supersonic flow around a wedge?

A)

The flow changes its direction abruptly and pressure decreases with acceleration

B)

There is a sudden change in flow direction at the body and pressure increases downstream of the shock

C)

The flow changes its direction smoothly and pressure decreases with acceleration

D)

The flow changes its direction smoothly and pressure increases downstream with acceleration

Correct Answer : 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.

63 .

What will happen when the wedge angle is finite?

A)

Streamlines over the wedge are parallel to the flow ahead of the wedge

B)

Finite disturbances will be introduced and the wave will be called a shock wave

C)

The deviation of streamlines will be finite and pressure will decrease across the shock wave

D)

Small disturbances will be introduced and we can consider these to be identical to sound pulses

Correct Answer : Finite disturbances will be introduced and the wave will be called a shock wave

Explanation : When the wedge angle is finite, direction of flow changes rapidly and this deviation causes shock wave. There is pressure increase across the shock wave but total pressure reduces since velocity reduces.