Download GTU BE/B.Tech 2019 Summer 6th Sem New 2160101 Aerodynamics Ii Question Paper

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GUJARAT TECHNOLOGICAL UNIVERSITY
BE - SEMESTER- VI(NEW) — EXAMINATION - SUMMER 2019

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Subject Code: 2160101 Date: 10/05/2019
Subject Name: Aerodynamics II
Time: 10:30 AM TO 01:00 PM Total Marks: 70

Instructions:

1. Attempt all questions.

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2. Make suitable assumptions wherever necessary.
3. Figures to the right indicate full marks.

MARKS

Q.1 (a) Plot Variation of C_L with Re, and aspect ratio. 03
(b) Define mean aerodynamic chord [MAC] and procedure to find MAC. 04

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(c) Derive expression for classical thin airfoil theory. 07

Q.2 (a) Define:- Turbulent flow, Transition and dihedral. 03
(b) State Kelvin Circulation Theorem. 04
(c) Explain The Vortex lattice method. 07
OR

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(c) Explain Kutta-Joukowsky Transformation and details procedure about transformation of circle into symmetric airfoil. 07

Q.3 (a) Write Kutta Condition and explain with sketch. 03
(b) Differentiate between finite and infinite wing. Derive expression for ‘a’ (C_L-a slope) and ‘a₀’. 04
(c) What is Transformation? Explain transformation of a circle into a straight line with a neat sketch. 07
OR

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Q.3 (a) What is bound vortex and Horse shoe vortex? 03
(b) Explain Numerical Nonlinear Lifting Line Method 04
(c) Explain Prandtl lifting line theory and derive expression for downwash velocity. 07

Q.4 (a) Explain Prandtl-Glauert Compressibility correction. 03
(b) Explain formation of primary and secondary vortex on delta wing. 04

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(c) Derive Linearised Supersonic Pressure Coefficient formula. 07
OR
Q.4 (a) Explain the area rule with neat sketch. 03
(b) 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 C_p at this point. 04
(c) Derive expression for induced velocity by finite and infinite vortex tube 07

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Q.5 (a) What is whitcomb airfoil 03
(b) A straight finite wing has aspect ratio 8 operating at angle of attack 4°.calculate C_Di (induced drag coefficient) and C_L assume that wing identical to flat plate 04
(c) Using KJT transform cylinder into symmetrical airfoil 07

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Q.5 (a) State Helmholtz’s theorem 03

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(b) Consider flow over the NACA 2412 airfoil where Re = 3240000. Calculate the boundary layer thickness at 1.5 m from leading edge 04
(c) Define Critical Mach number and Drag Divergence Mach Number? Explain both the terms in detail with a proper diagram. 07

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