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JEPPIAAR ENGINEERING COLLEGE

DEPARTMENT OF AERONAUTICAL ENGINEERING

AE2303/Aerodynamics II/Important 15 questions

UNIT-I

1. i) Derive the area-mach number relation and explain why convergent-divergentnozzle is needed for supersonic flow. (8Marks)

ii) Explain what is choking in a C-D nozzle and show that the expression for

choked mass flow rate for an isentropic flow of duct through a duct is

(8Marks)

2. i) Derive the one dimensional adiabatic steady state energy equation and deduce theisentropic relations for a perfect gas. (8Marks)

ii) Obtain an expression for the speed of sound and show that the speed of sound is

proportional to the square root of the absolute temperature of air. (8Marks)

3. Problems based on C-D nozzle to find out the mass flow rate, throat area cross sectionalarea, pressure, velocity, and Mach number at the nozzle exit

UNIT-II

1. i) Show that the strength of a normal shock in a perfect gas depends only on Machnumber ahead of the shock or For an oblique shock wave bring out proper relationships

between the flows parameters in front of the shock and behind the shock. (8Marks)ii) Derive Rayleigh supersonic Pitot formula. Why is Rayleighs correction for total

pressure required in supersonic flows? (8Marks)

2. i) Derive Prandtl- Meyers Expansion waves for a flow over a convex corner.ii) Derive Prandtl relation for a normal shock in a perfect gas.

3. Problems based on shocks to find out the downstream(total pressure &temperature,density) conditions of the shocks, deflection angle, strength of the shock,

UNIT-III

1. i)Explain the procedure to obtain supersonic nozzle contour for a given Mach numberusing Method of characteristics. Also draw neat sketches for continuous and centred

expansion supersonic nozzles. (8Marks)

ii) Show that the local Mach number is unity at the point of maximum entropy on the

Rayleigh line. Or For the Rayleigh flow, show that the mach number

at which

temperature is maximum. Further find the value of

for =1.4 (8Marks)

2. i)Derive the Rankine-Hugoniot relation for a shock. (10Marks)ii) What is the difference between Rayleigh and fanno flow.(6Marks)

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3. Problems based on the double wedge to calculate the flow Mach numbers and theslipstream using shock expansion theory or Problems based on two dimensional wedge to

find out CLand CDusing shock expansion theory

UNIT-IV

1. Problems based on two-dimensional wing profile or for a flat plate to calculate CLand CDusing linearized theory.

2. State the assumptions and limitations made in the small-perturbation potential theory andshow that the linearized pressure coefficient is a function of the perturbation velocity in

the Main flow direction only.

CP= - .

3. i) Describe the Prandtl-Glauert affine transformation for subsonic flow over airfoils andhighlight its significance.

ii) Explain about thin aerofoil theory for subsonic and supersonic flows.

UNIT-V

1. i) Explain about Shock induced separationii)Explain in detail about the effect of thickness, camber and aspect ratio on the

characteristics of wings

2. i)Briefly discuss transonic area rule and supercritical airfoilii)Write short notes on reflection of shock wave and expansion waves from free- surface

boundary.

3. i) Write brief notes on critical Mach number and drag divergence Mach number. ii) Explain how large drag increase takes place at transonic flow. What are the control

measures adopted at the design stage?

iii) Explain the advantages and disadvantages of the effect of sweepback

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