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Design and development of a rectangular supersonic wind tunnel facility for the study of shock/boundary layer interactions

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Title: Design and development of a rectangular supersonic wind tunnel facility for the study of shock/boundary layer interactions
Author(s): Chang, Wilbur
Advisor(s): Elliott, Gregory; Dutton, J. C.
Department / Program: Aerospace Engineering
Discipline: Aerospace Engineering
Degree Granting Institution: University of Illinois at Urbana-Champaign
Degree: M.S.
Genre: Thesis
Subject(s): high speed flow shockwave boundary layer interactions supersonic viscous flow Greg Elliott Craig Dutton Wilbur Chang Gulfstream National Aeronautics and Space Administration (NASA) inlet business jet wind tunnel experimental fluid mechanics micro vortex generators
Abstract: The research work of this thesis was part of the low-boom supersonic inlet project conducted by NASA, Gulfstream Aerospace Corporation, Rolls-Royce, and the University of Illinois. The low-boom supersonic inlet project itself was part of a new supersonic business jet design. The primary goal of the low-boom supersonic inlet group was to reduce the sonic boom of the new jet by improving the performance of the inlets. The objective of the experimental team at the University of Illinois was to provide testing support to evaluate simple, passive, and bleedless inlet shock wave/boundary layer interaction control options, called micro-vortex generators. A new supersonic wind tunnel was designed and built on the University of Illinois campus to enhance high-speed flow testing capabilities used for studying these flow-control devices. A newer and larger tunnel will also contribute to expand the College of Engineering’s tools for studying and understanding high-speed fluid mechanics and applications to aerodynamics and propulsion technologies. The new wind tunnel is a rectangular testing facility with a 5” by 5” cross-sectional area in the test section. It is a blowdown, intermittent, open-loop facility, capable of operating at Mach numbers of 1.4 and 2, with the theoretical capability of reaching Mach 3. The wind tunnel was assembled and installed in the west wing of Aeronautical Lab A, and shares the same air supply system with an existing axisymmetric supersonic wind tunnel and an open jet anechoic testing chamber. The run time for the tunnel at Mach 1.4 is approximately 120 seconds, with a turnaround time of approximately 10 minutes. A brief study and experimental testing of a set of micro-vortex generators, in the form of ramped vanes, were performed to demonstrate the capabilities of the new tunnel. A 5° expansion diffuser was situated downstream of the vortex generators to model the effects of a supersonic external compression inlet and diffuser. The incoming boundary layer profiles of the top and bottom walls were characterized through particle image velocimetry. Visual diagnostic techniques of schlieren imaging, surface oil flow visualization, and particle image velocimetry were conducted to study the effects of the ramped vanes on shock wave/boundary layer interactions. The vanes were observed to produce two distinct vortex pairs that entrained and mixed higher momentum air into the boundary layer. Higher velocity air was observed to lie closer to the diffuser surface from the turbulent mixing caused by the vortex generators.
Issue Date: 2012-02-06
URI: http://hdl.handle.net/2142/29604
Rights Information: Copyright 2011 Wilbur Chang
Date Available in IDEALS: 2012-02-06
Date Deposited: 2011-12
 

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