Anna University, Chennai
SRINIVASAN ENGINEERING COLLEGE, PERAMBALUR
DEPARTMENT OF AERONAUTICAL ENGINEERING AERODYNAMICS-II QUESTION BANK
PART - A
UNIT – I ONE DIMENSIONAL COMPRESSIBLE FLOW
1) Differentiate between compressible and incompressible flow
Compressible – variable Density
In compressible – Constant density
2) Write the Bernoulli’s equation for incompressible flow.
| |
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. 0.1 ≥ > 0.8
Transonic mach no. 0.8 ≥ > 1.2
Supersonic mach no 1.2 ≥ > 5
Hypersonic mach no. ≥ 5
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?
1 2
∗ + 1
− 1
!"# $"#
10) Write the Bernoulli’s equation for compressible flow.
%
&
)
( +
=
UNIT – II NORMAL, OBLIQUE SHOCKS AND EXPANSION WAVES
1) What is meant by Normal Shock?
If the shock wave is perpendicular to the free stream velocity, then it is normal shock wave
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3) What is meant by shock tube?
* $*" =
+ − +"#
2
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.
4) 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.
5) 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. |
6) Give the relation between Shock angle (β), Mach number and Flow deflection angle (θ).
" 0 3 1 − 1
tan / = 20 1 2
7
" + cos 21# + 2
7) What is meant by Shock Polar?
Shock polar is the graphical representation of the oblique shock properties.
8) 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.
9) Define characteristic Mach number?
If the Mach number is equal to one, then it is characteristic Mach number, M*
UNIT – III DIFFERENTIAL EQUATIONS OF MOTION FOR STEADY COMPRESSIBLE FLOWS
1) Write the equation of linearised potential theory.
81 − ∞ 9∅;; + ∅<< + ∅== = 0
2) 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
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3) What is perturbation potential function?
It is the small increment in the velocity potential function.
4) 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.
5) Write the prandtl Glauret relation.
6) Define method of characteristics?
@ = @B,D
>1 − ∞
@ = @E,D
>1 − ∞
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@ = @?,D
>1 − ∞
It is the numerical methods for solving the full non linear equations of motion for in viscid, ir rotational, flow.
UNIT – IV AIRFOIL IN HIGH SPEED FLOWS
1) Define Critical Mach number.
It is free stream Mach number, when the sonic condition is first attained at any point of the body.
2) 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.
3) 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.
4) 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.
5) 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.
6) 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.
7) 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.
8) 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.
UNIT – V HIGH SPEED WIND TUNNELS
1) What is a wind tunnel and classify the wind tunnels?
Wind tunnel is the flow device used for the calibration of the flow properties over the body.
2) What is meant by Blow down tunnel?
It is a intermittent type in which the energy is stored in the form of pressure or vacuum or both and is allowed to drive the wind tunnel only for few seconds out of each pumping hour
3) What is meant by flow visualizations?
It is the process of visualizing the flow patterns behind the body.
4) What are the optical flow visualization methods?
a. Interferometer
b. Schlieren system c. Shadow graph
PART – B
UNIT – I ONE DIMENSIONAL COMPRESSIBLE FLOW
1) . Explain Quasi-one dimensional flow and area velocity relation.
2) Define De Laval Nozzle and derive the Area Mach number relation.
3) With neat sketch briefly explain the flow in a Convergent-divergent nozzle.
4) For an aircraft flying at a speed of 1000kmph, find the variation of speed of sound a, and
Mach number M, with sea level and 11km altitude.
5) During a flight, a fighter aircraft attains its cruise speed of 600 m/s at 10km altitude after taking off at 150 m/s from sea level. Assuming the speed to have increased linearly with altitude during the climb, compute the Mach number variation with altitude.
6) A fighter aircraft attains its maximum speed of 2160 kmph at an altitude of 12 km. The take-off speed at sea level is 270 kmph. If the flight speed increases linearly with altitude, compute the variation of stagnation temperature with altitude for a climb up to the maximum speed.
7) Air flows through a duct. The pressure and temperature at station 1 are P1 = 0.7 atmand T1 = 300C, respectively. At a second station, the pressure is 0.5 atm. Calculate the temperature and density at the second station. Assume the flow to be isentropic.
8) Air is allowed to expand from an initial state A (where PA = 2.068 x 105 N/m2 and TA =
333K) to state B (where PB = 1.034 x 105 N/m2 and TB = 305 K). Calculate the change in the specific entropy of the air, and show that the change in entropy is the same for (a) an isobaric process from A to some intermediate state C followed by an isovolumetric change from C to B, and (b) an isothermal change from A to some intermediate state D followed by an isentropic change from D to B.
9) A ramjet flies at 11 km altitude with a flight mach number of 0.9. In the inlet diffuser, the air is brought to the stagnation condition so that it is stationary just before the combustion chamber. Combustion takes place at constant pressure and a temperature increase of
15000C results. The combustion products are then ejected through the nozzle. (a) Calculate the stagnation pressure and temperature. (b) What will be the nozzle exit velocity? ( At inlet Pα = 0.3 atm and Tα = 213 K, at exit Pexit = 0.3 atm).
10) A De Laval Nozzle has to be designed for an exit Mach number of 1.5 with exit diameter of 200mm. Find the ratio of throat area to exit area necessary. The reservoir conditions are given as P0 = 1 atm; T0 = 200C. Find also the maximum mass flow rate through the nozzle. What will be the exit pressure and temperature?
UNIT – II NORMAL, OBLIQUE SHOCKS AND EXPANSION WAVES
1) Derive the Prandtl Normal Shock relation for a Perfect gas.
2) Derive the Hugoniot equation and explain the Hugoniot Curve.
3) The flow Mach number, pressure, and temperature ahead of a normal shock are given as
2.0, 0.5 atm and 300 K respectively. Determine M2, P2 , T2, and V2 behind the wave.
4) A re-entry vehicle (RV) is at an altitude of 15,000 m and has a velocity of 1850 m/s. A bow shock wave envelops the RV. Neglecting dissociation, determine the stagnation pressure and temperature just behind the shock wave on the RV center line where the shock wave may be treated as normal shock.. Assume that the air behaves as perfect gas, with γ = 1.4 and R = 287 J/kg-K.
5) A normal shock moves in a constant area tube as shown in figure. In region 1, V1 = 100 m/s, T1 = 300C and P1 = 0.7 atm. Shock speed CS with respect to a fixed coordinate system is 600 m/s. Find fluid properties in region 2.
6) Write short notes on
(i) Supersonic flow over a wedge
(ii) Weak Oblique shocks
(iii) Supersonic Compression
(iv) Supersonic Expansion by Turning
UNIT – III DIFFERENTIAL EQUATIONS OF MOTION FOR STEADY COMPRESSIBLE FLOWS
1) Derive the linearised small perturbation potential theory.
2) Write short notes on. (i) Mach waves
(ii) Mach angles
(iii) Solutions for Supersonic flows.
3) Derive the Linearised two-dimensional supersonic flow theory.
4) Derive the Prandtl Glauret affine transformation relations for subsonic flows.
5) Explain Small perturbation equation for compressible flows.
UNIT – IV AIRFOIL IN HIGH SPEED FLOWS
1) Briefly explain the characteristics features of the lower critical mach number and upper critical mach number.
2) Briefly explain the characteristics features of swept wings.
3) Briefly explain the effects of thickness, camber and aspect ratio over the performance of wings in high speed flows.
4) Briefly explain the need and characteristic features of Transonic area rule.
5) Explain the supersonic airfoil in detail.
6) Explain the following;
(iii) Disadvantages of Swept wings
(iv) Delta wing
UNIT – V HIGH SPEED WIND TUNNELS
1) Explain the Hypersonic wind tunnel (Helium) with sketches.
2) Sketch a typical shock tunnel and explain its principle of operation. What are the advantages and limitations of shock tunnel?
3) Draw a neat sketch of a supersonic wind tunnel circuit and explain the function of each component.
4) Draw a neat sketch of a transonic wind tunnel circuit and explain the function of each component.
5) Draw a neat sketch of a hypersonic wind tunnel circuit and explain the function of each component.
6) Briefly explain the Blow down, indraft and induction tunnel layouts and their design features.
7) Briefly explain the Helium and gun tunnels and its applications.
8) Briefly explain the various optical methods of flow visualization.
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