JNTU
JNTU B.Tech II AERODYNAMICS I Previous Question Paper (April/May - 2023 - R13)
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 the advantage in approximating the flow to be incompressible?
Answer
Approximating flow as incompressible simplifies the governing equations by assuming constant density. This eliminates the need to solve the energy equation simultaneously with the continuity and momentum equations, significantly reducing computational complexity and allowing for the use of potential flow theory.
Write the Navier-Stokes equation for an incompressible fluid.
Answer
For an incompressible Newtonian fluid with constant viscosity, the Navier-Stokes equation is:
where ρ is density, V is the velocity vector, p is pressure, μ is dynamic viscosity, and g is the body force per unit mass.
State Kelvin theorem.
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What are the elementary flows to be superimposed to simulate the lifting flow over circular cylinder?
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Define boundary layer thickness.
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What is the difference between skin friction drag and pressure drag?
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State Helmholtz theorem.
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What is vortex panel method? Write its governing equation.
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What is the purpose of slats?
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What is the effect of geometric pitch on airscrew performance?
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Part B
Using Buckingham’s Pi theorem, show that aerodynamic force is a function of Reynolds number and Mach number.
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Derive the 3-D continuity equation in differential form with an appropriate fluid flow model.
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Derive the stream function for a doublet flow.
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Derive the expression for Aerodynamic Center and Center of Pressure. Write the applications of each.
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Explain the effect of transition and surface roughness on airfoils.
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The geometric and aerodynamic data for a wing of a large white butterfly is as follows: Flight speed, U∞ = 1.55 m/s; average chord, c = 28 mm; average span, s = 40 mm; air density = 1.12 kg/m3; air viscosity, μ = 1.8×10-6 Pa/s; drag at zero lift = 128 μN. Estimate the boundary-layer thickness at the trailing edge.
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Derive expressions for induced angle of attack and induced drag coefficient.
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The measured lift slope for the NACA 23012 airfoil is 0.105 degree-1, and αL=0 = -1.5 degrees. Consider a finite wing using this airfoil, with AR = 10 and taper ratio = 0.6. Assume that δ = τ. Calculate the lift and induced drag coefficients for this wing at a geometric angle of attack = 9 degrees.
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Derive an expression for Froude efficiency of the propulsive system.
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Using simple momentum theory, develop an expression for the thrust of a propeller in terms of disc area, air density, and axial velocities of the air a long way ahead, and in the plane, of the propeller disc.
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