Hall Ticket No Question Paper Code: AAE008
INSTITUTE OF AERONAUTICAL ENGINEERING
(Autonomous)
Four Year B.Tech V Semester End Examinations (Supplementary) January, 2019
Regulation: IARE R16
HIGH SPEED AERODYNAMICS
Time: 3 Hours Max Marks: 70
Answer ONE Question from each Unit
All Questions Carry Equal Marks
All parts of the question must be answered in one place only
UNIT I
1. Write short notes on
Thermo dynamics systems ii) Enthalpy
iii) Calorifically perfect gas iv) Perfect gas
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
variations in stagnation temperature with altitude for a climbing up to the maximum speed in
steps of 3 Km.
2. Define the principle of momentum equation and derive the equations for the conservations of
momentum in integral form.
Consider a Boeing 747 airliner cruising at a velocity of 885.14 Km/hr at a standard altitude of
11582.5 where the free stream pressure and temperature are 20.68 kPa and 487:50C, respectively.
A one-fiftieth scale model of the 747 is tested in a wind tunnel where the temperature is
537:50C. Calculate the required velocity and pressure of the test airstream in the wind tunnel
such that the lift and drag coefficients measured for the wind-tunnel model are the same as for
free flight. Assume that both and a are proportional to T 1/2.
UNIT II
3. What is shock expansion theory? How it is applicable for super sonic airfoils.
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. Explain the theta-Beta-M relation for wide range of supersonic flow.
Consider a supersonic flow with M p 1 atm, and T 288 K. This flow is deflected at
a compression corner through 200. Calculate T p0, and T0 behind the resulting oblique
shock wave.
UNIT III
5. Write a short notes on
Fanno flow
ii) Rayleigh flow.
At a given point on the surface of an airfoil, the pressure coefficient is -0.3 at very low speeds. If
the free stream Mach number is 0.6, calculate Cp at this point.
Page 1 of 2
6. Consider a flow through constant area pipe entering fanno flow and derive expression for ideal
gas equation to calculate the density ratio from pressure and temperature ratio.
Consider the isentropic flow through a convergent-divergent nozzle with an exit-to-throat area
ratio of 2. The reservoir pressure and temperature are 1 atm and 288 respectively. Calculate
the Mach number, pressure, and temperature at both the throat and the exit for the cases where
the flow is supersonic at the exit and the flow is subsonic throughout the entire nozzle
except at the throat, where M 1.
UNIT IV
7. Explain with neat sketches transonic area rule and explain in detail.
A supersonic nozzle is designed to operate at Mach 2.0. Under a certain operating condition,
however, an oblique shock making a 450 angle with the flow direction is observed at the nozzle
exit plane, as in Figure 1. What percent of increase in stagnation pressure would be necessary
to eliminate this shock and maintain supersonic flow at the nozzle exit?
Figure 1
8. How important is Prandtl-Glauert compressibility correction and derive the equations for Cl and
Cm
At a given point on the surface of an airfoil, the pressure coefficient is −0.3 at very low speeds
in an incompressible flow cl If the free stream Mach number is 0.6, calculate Cp and lift
coefficient for at this point.
UNIT V
9. Draw a neat sketch of a low speed wind tunnel circuit and explain the function of each component.

What is the reservoir pressure for the tunnel if The nozzle of a supersonic wind tunnel has an
exit to throat area ratio of 6.79 when the tunnel is running, a pitot tube mounted in the test
section, measures 1.448 atm.

10. Draw a neat sketch of a free-piston shock wind tunnel and explain the function of each component.

Draw a neat sketch of blow down type supersonic wind tunnel and explain the function of each
of the component.