JAWAHARLAL NEHRU TECHNOLOGICAL UNIVERSITY HYDERABAD
JAWAHARLAL NEHRU TECHNOLOGICAL UNIVERSITY HYDERABAD B.Tech III HIGH SPEED AERODYNAMICS Previous Question Paper (January/February - 2023 - R18)
Note: i) Question paper consists of Part A, Part B. ii) Part A is compulsory, which carries 25 marks. In Part A, Answer all questions. iii) In Part B, Answer any one question from each unit. Each question carries 10 marks and may have a, b as sub questions.
Part A
What is a bow shock and when does it occur?
Answer
A bow shock is a curved, detached shock wave that forms upstream of a blunt-nosed body in supersonic flow. It occurs when the Mach number is supersonic but the body geometry prevents the formation of an attached oblique shock, forcing the flow to decelerate to subsonic speeds through a normal shock portion at the stagnation point.
What is the basic law used in deriving momentum equation and its mathematical form.
Answer
The basic law used is Newton's Second Law of Motion, which states that the net force acting on a fluid element is equal to the rate of change of momentum. Its mathematical form for a control volume is:
Define shock detachment distance with respect to a flow over blunt body.
Answer
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Define oblique shock wave and the term wave angle.
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Define under expanded nozzle along with a diagram.
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What is the equation for ratio of second throat area to first throat area of a supersonic wind tunnel? Also specify the condition for choking of the nozzle.
Answer
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Define explicit finite difference solution along with its limitation.
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Define the terms ‘Domain of Dependence’ and ‘Region of Influence’ in a supersonic flow with the help of a diagram.
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Define detonation-driven shock tunnel.
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Illustrate a typical shock tunnel with the help of a neat diagram.
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Part B
Develop the equations for conservation of energy in integral form.
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An airplane is flying at an altitude of 20,000m. The pressure at a point on the wing is 19152 N/m2, assuming isentropic flow over the wing, calculate the temperature at this point.
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Explain the concept of shock wave and boundary layer interaction.
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Develop the expression for density relationship for normal shock for a perfect gas in terms of Mach number.
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Explain Quasi-one-dimensional flow and area velocity relation.
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Develop expressions for the pressure, area and temperature at any section of the nozzle in terms of throat critical conditions.
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Explain the significance of Prandtl-Glauert compressibility correction and derive the equations for Cl and Cm.
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For an airfoil, at a given point on its surface, the pressure coefficient is -0.3 at very low speed in an incompressible flow cl = 2πα. If the free stream Mach number is 0.6, calculate Cp and lift coefficient.
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Explain in detail, the working of a free-piston shock wind tunnel with the help of a neat diagram.
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Explain the characteristics of an expansion tube with the help of a neat diagram.
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