Aerodynamics II 2 Marks

DEPARTMENT OF AERONAUTICAL ENGINEERING AERODYNAMICS-II PART – A QUESTION WITH ANSWERS 1) Differentiate between compressible and incompressible flow Compressible – variable Density In compressible – Constant density 2) Write the Bernoulli’s equation for incompressible flow. = const 3) Write the adiabatic relation between pressure and density.

4) What is meant by Mach angle? It is the angle between mach line and the direction of motion of body. 5) Define (i) Zone of action (ii) Zone of silence (iii) Mach Waves (or) Mach lines. Region inside the mach cone – zone of action Region outside the cone – zone of silence The lines at which the pressure disturbance is concentrated and which generate the cone are called as mach waves or mach lines. 6) Classify the flow regimes in terms of Mach number. Subsonic mach no. Transonic mach no. Supersonic mach no Hypersonic mach no. 7) How velocity of the flow varies in convergent and divergent ducts for subsonic and supersonic condition. velocity Subsonic Supersonic Increases

Convergent duct

Divergent duct

Decreases

Divergent duct

Convergent duct

8) What is meant by ‘De Laval Nozzle’? It is the convergent Divergent nozzle. It is the only means to produce supersonic flow. 9) Write the Area Mach number relation? (

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(

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)]

10) Write the Bernoulli’s equation for compressible flow. ∫ 11) What is meant by Normal Shock?

If the shock wave is perpendicular to the free stream velocity, then it is normal shock wave 12) Write the Hugoniot equation and explain each terms involved in it. (

)

13) What is meant by shock tube?

It is a device to produce high speed flow with high temperatures, by traversing normal shock waves which are generated by the rupture of a diaphragm separating a high pressure gas from the low pressure gas. 14) Define Oblique shock?

If the shock developed due to the supersonic flow and if it is inclined at an angle, β to the free stream direction, then it is oblique shock. 15) Differentiate between shock wave and expansion wave.

Shock wave

Expansion wave.

Supersonic flow over the compression corner produces shock waves

Supersonic flow over the expansion corner produces expansion waves.

16) Give the relation between Shock angle (), Mach number and Flow deflection angle (). (

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17) What is meant by Shock Polar?

Shock polar is the graphical representation of the oblique shock properties. 18) Define sonic circle. The circle with radius M* =1 is called as the sonic circle. Inside the circle all the velocities are subsonic and outside the velocities are supersonic. 19) Define characteristic Mach number?

If the Mach number is equal to one, then it is characteristic Mach number, M* 20) Write the equation of linearised potential theory. (

)

21) Write the Prandtl Glauret Rule. a. Stream lines of the compressible flow are far apart from each other by √

than in incompressible flow.

b. The ratio between aerodynamic characteristics in compressible and in-

compressible flow is also

√

22) What is perturbation potential function? It is the small increment in the velocity potential function. 23) Give the general features of method of characteristics?

They exists only in supersonic flow field Characteristics are co incident with mach lines While the derivatives of the flow properties are discontinuous, the flow properties themselves are continuous on the characteristics. 24) Write the prandtl Glauret relation. √

√

√ 25) Define method of characteristics?

It is the numerical methods for solving the full non linear equations of motion for in viscid, ir rotational, flow.

26) Define Critical Mach number.

It is free stream Mach number, when the sonic condition is first attained at any point of the body. 27) Distinguish between Lower Critical Mach number and Upper Critical Mach number.

The free stream Mach number for which the entire flow around the body is subsonic is called the lower critical Mach number. The free stream Mach number for which the entire flow around the body is supersonic is called the upper critical Mach number. 28) What is the effect of thickness over the performance of wings?

a. The critical Mach number decreases with the increasing thickness of the body. b. The co-efficient of pressure for the thick airfoil is greater than the thin airfoil. 29) What is the effect of camber over the performance of wings?

The character of the thickness and the camber is proportional to each other. a. The critical Mach number decreases with the increasing camber of the body. b. The co-efficient of pressure for the high camber airfoil is greater than the less camber airfoil. 30) What is meant by transonic area rule?

Transonic area rule states that, the cross sectional area of the body should have smooth variation with the longitudinal distance along the body. 31) What are the characteristics of swept back wing? By sweeping the wing, we can reduce the thickness to chord ratio ie., it makes the airfoil section thinner. Thus increasing the critical Mach number and thereby increasing the drag divergence Mach number. 32) What is drag divergence Mach number?

The value of Mach number when there is a sudden increase in the coefficient of te drag starts is the drag divergence mach number. 33) Why drag increases drastically over sonic speed? The drag increases drastically over the sonic region because the extensive region of the supersonic flow over the airfoil will be terminated by the strong shock

wave. These shock waves cause the severe flow separation downstream the shock which results in large increase in drag.