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The measurement of wind tunnel flow quality for the transonic flow regime has been plagued by the inability to interpret complex unsteady flow field information obtained in the free stream. Traditionally hot wire anemometry and fluctuating pressure techniques have been used to quantify the unsteady characteristics of a wind tunnel. This research focuses on the application of these devices to the transonic flow regime. Utilizing hot wire anemometry, one can decompose the unsteady flow field with a three sensor technique, to obtain fluctuations associated with the velocity, density, and total temperature. Implementing thermodynamic and kinematic equations, new methods for expanding the measured velocity, density, and total temperature fluctuations to obtain additional fluctuations are investigated. The derived static pressure fluctuations are compared to the static pressure fluctuations obtained with a conventional fluctuating static pressure probe. The results of this comparison are good, which implies that the individual velocity, density, and total temperature components are time accurate. In the process of obtaining a high quality fluctuating flow field information, it was necessary to evaluate the calibration of the hot wire sensors. A direct calibration approach was compared to a conventional non-dimensional technique. These two calibration techniques should have resulted in the same hot wire sensitivities. There were significant differences in the hot wire sensitivities as obtained from the two approaches. The direct approach was determined to have less errors due to the added heat transfer information required of the indirect approach. Both calibration techniques demonstrated that the velocity and density sensitivities were in general not equal. This suggests that the velocity and density information cannot be combined to form a mass flow. A comparison of several hot wire techniques was included to highlight the errors obtained when assuming that these sensitivities are the same. An evaluation of the free stream flow quality associated with a Laminar Flow Control experiment was carried out in the Langley Research Center 8-Foot Transonic Pressure Tunnel (8' TPT). The facility was modified with turbulence manipulators and a liner that provided a flow field around a yawed super-critical airfoil that is conducive to transition research. These devices are evaluated to determine the sources of disturbances associated with the LFC experiment. / Ph. D.

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Approved for public release, distribution unlimited / Flow around second-generation controlled-diffusion blades in cascade at stall was investigated using fivehole probe surveys. Wake pressure surveys were conducted at various locations between the cascade centerline and endwall region. Corresponding inlet pressure surveys were also performed to allow the total pressure loss distribution to be calculated across the blades. A fully automated traverse mechanism was implemented for the probe surveys ahead of and aft of the cascade blades. The number of wake surveys conducted allowed the pressure distribution profiles to show pressure deficiencies at the endwall region in greater detail than previous studies.

Approved for public release, distribution unlimited / The Naval Postgraduate School (NPS) Micro-Air Vehicle (MAV) wind tunnel was refurbished in this study. This wind tunnel has a 61 by 38 centimeter test section with a contraction ratio of 6.75, and testing speeds up to 9 meters per second (m/s). The objectives of this work were to create a high quality, customized facility for further MAV study, and to fully characterize the wake of the MAV at a plane downstream of the model. Extensive repairs were made to the wind tunnel intake, test section, and fan. The flow field produced in the tunnel was analyzed with hot wire anemometry. The turbulence intensity ranged from .7% to .5% for freestream velocities of 2 m/s to 5 m/s. The velocity variation across the test section ranged from .251% to .125%. A model of the NPS MAV was placed in the test section and the wake was studied with and without the main wing, and in vertical and horizontal orientations. The wake was characterized with hot wire anemometry and flow visualization techniques at various flapping frequencies and freestream velocities. / US Navy (USN) author.

A research report submitted to the Faculty of Engineering and the Built Environment, University of the Witwatersrand, Johannesburg, in partial fulfilment of the requirements for the degree of Master of Science in Engineering 2016 / The nozzle contour profiles of the CSIR’s supersonic wind tunnel (high speed wind tunnel) were designed to produce smooth, uniform and shock-free flow in the operating section of the facility. The existing profiles produce weak waves in the test section region which induces flow gradients and flow angularities in the air flow, effectively degrading the air flow quality, which in turn perturbs the wind tunnel data. The wind tunnel geometry and tunnel constraints were employed in accordance with the method of characteristics technique to design the supersonic nozzle profiles. The Sivells’ nozzle design method was deemed the most feasible which calculates the profile downstream of the inflection point. The throat block profile was amalgamated with this profile to yield a profile from the throat to the test section. A boundary layer correction was applied to the profiles to account for viscous effects which cause a Mach number reduction from the desired test section Mach number. An automatic computation was used for the profile design and a computational method analysed the Mach distribution, flow angularity and density gradient (to determine the occurrence of shocks and expansions) of the profiles implemented in the tunnel, for the full Mach number range of the HSWT. The methods used, achieved uniform and shock-free flow such that the Mach number and flow angularity were within the acceptable quality limits of the HSWT.

Thesis (M.Phil.)--Hong Kong University of Science and Technology, 2007. / Includes bibliographical references (leaves 184-189). Also available in electronic version.

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Wind tunnel tests were conducted to investi gate t he effect of placing a propulsion fan behind the tip of a li.tting ring for the purpose of recovering the energy in the trailing vortex. Results indicated tha~ significant increase of thrust could be obtained as a result of vortex energy recovery. In addition, the pre-swirling effect provided b.r the vortex allowed the tan to absorb more power and develop more thrust at a lower rotational speed. / Thesis (M.S.)--Wichita State University, Dept. of Aeronautical Engineering.

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