Dissertation (M.S.)--University of Toledo, 2007 / Typescript. "Submitted as partial fulfillment of the requirements for the Doctor of Philosophy in Engineering." Bibliography: leaves 223-237.

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Thesis(M.S.)--Auburn University, 2005. / Abstract. Vita. Includes bibliographic references.

Thesis (M.S. in Aerospace Engineering) -- University of Dayton. / Title from PDF t.p. (viewed 10/06/09). Advisor: Aaron Altman. Includes bibliographical references (p. 140-142).

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An experimental investigation was conducted in a static ground test facility to determine the flow quality of a serpentine inlet duct incorporating active flow control for several simulated flight conditions. The total pressure distortion at the aerodynamic interface plane (AIP) was then used to predict the resulting stability for a compression system. This study was conducted using a model of a compact, low observable, engine inlet duct developed by Lockheed Martin. A flow control technique using air injection through microjets at 1% of the inlet mass flow rate was developed by Lockheed Martin to improve the quality of the flow exiting the inlet duct. Both the inlet duct and the flow control technique were examined at cruise condition and off-design simulated flight conditions (angle of attack and asymmetric distortion). All of the experimental tests were run at an inlet throat Mach number of 0.55 and a resulting Reynolds number of 1.76*105 based on the hydraulic diameter at the inlet throat. For each of the flight conditions tested, the flow control scheme was found to improve the flow uniformity and reduce the inlet distortion at the AIP. For simulated cruise condition, the total pressure recovery was improved by ~2% with the addition of flow control. For the off-design conditions of angle of attack and asymmetric distortion, the total pressure recovery was improved by 1.5% and 2% respectively. All flight conditions tested showed a reduction in circumferential distortion intensity with flow control. The cruise condition case showed reduced maximum circumferential distortion of 70% with the addition of flow control. A reduction in maximum circumferential distortion of 40% occurred for the angle of attack case with flow control, and 30% for the asymmetric distortion case with flow control. The inlet total pressure distortion was used to predict the changes in stability margin of a compression system due to design and off-design flight conditions and the improvement of the stability margin with the addition of flow control. A parallel compressor model (DYNTECC) was utilized to predict changes in the stability margin of a representative compression system (NASA Stage 35). Without flow control, all three cases show similar reduced stability margins on the order of 30% of the original stability margin for NASA Stage 35 at 70% corrected rotor speed. With the addition of flow control, the cruise condition tested improved the stability margin to 80% of the original value while the off-design conditions recover to 60% of the original margin. Overall, the flow control has been found to be extremely beneficial in improving the operating range of a compression system for the same inlet duct without flow control. / Ph. D.

Approved for public release; distribution is unlimited / The flowfield downstream of a vertically-launched surface-to-air missile model at an angle of attack of 50° and a Reynolds number of 1.1 x 10(5) was investigated in a wind tunnel of the Naval Postgraduate School. The goal of this thesis is to experimentally validate the pressure measurement system for flowfield variables with elevated levels of turbulence; to determine the location and intensity of the asymmetric vortices in the wake of the VLSAM model at a raised level of freestream turbulence; and to display the asymmetric vortices by velocity mapping and pressure contours. The purpose is to correlate the results with the force measurements of Rabang to provide a greater understanding of the vortex flowfield. The body-only configuration was tested. Two flowfield conditions were treated: the nominal ambient wind tunnel condition, and a condition with grid­ generated turbulence of 3.8% turbulence intensity and a dissipation length scale of 1.7 inches. The following conclusions were reached: 1) The relative strengths of the asymmetric vortices can be noted by the sharp spike shape in the ambient condition; this condition becomes diffused and becomes fatter in the turbulent condition; 2) The right side vortex has greater strength than the left side one as seen by the diffusion in the total pressure coefficient and static pressure coefficient contours with and without a turbulent condition; 3) an increase in turbulence intensity tends to reduce the strength of the asymmetric nose-generated vortices; also pushes the two asymmetric vortices closer together; 4) and crossflow velocities were examined and were found to indicate the behavior denoted by the pressure contours. / http://archive.org/details/flowfieldmeasure00lung / Lieutenant, Republic of China Navy

Approved for public release; distribution is unlimited / A low speed wind tunnel investigation was conducted to examine the aerodynamic characteristics of the flowfield around a three percent scale YF-17 lightweight fighter prototype model at high angles of attack using flow visualization and force and moment measurements. Smoke filaments, injected into the wind tunnel test section, were illuminated by a laser sheet to highlight flow phenomena about the model. Force and moment measurements were made using a precision six-component strain gage balance. The investigation marked the first attempt at qualitative flow analysis using the laser sheet flow visualization system recently installed in the Naval Postgraduate School low speed wind tunnel facility. The investigation was undertaken to specifically identify flow phenomena and/or regions of interest that may have bearing on the design and performance of supermaneuverable aircraft. The data indicate a good correlation between the observed flow phenomena and force and moment measurements at various angles of attack, thus establishing the credibility of such experimental investigations for high angle of attack aerodynamic research. / http://archive.org/details/experimentalinve00leed / Lieutenant Commander, United States Navy

Aerodynamic devices, such as wings, are used in higher levels of motorsport (Formula-1 etc.) to increase the contact force between the road and tyres (i.e. to generate downforce). This in turn increases the performance envelope of the race car. However the extra downforce increases aerodynamic drag which (apart from when braking) is generally detrimental to lap-times. The drag acts to slow the vehicle, and hinders the effect of available drive power and reduces fuel economy. Wings, in automotive use, are not constrained by the same parameters as aircraft, and thus higher angles of attack can be safely reached, although at a higher cost in drag. Variable geometry aerodynamic devices have been used in many forms of motorsport in the past offering the ability to change the relative values of downforce and drag. These have invariably been banned, generally due to safety reasons. The use of active aerodynamics is currently legal in both Formula SAE (engineering compet ition for university students to design, build and race an open-wheel race car) and production vehicles. A number of passenger car companies are beginning to incorporate active aerodynamic devices in their designs. In this research the effect of ground proximity on the lift, drag and moment coefficients of inverted, two-dimensional aerofoils was investigated. The purpose of the study was to examine the effect ground proximity on aerofoils post stall, in an effort to evaluate the use of active aerodynamics to increase the performance of a race car. The aerofoils were tested at angles of attack ranging from 0° - 135°. The tests were performed at a Reynolds number of 2.16 x 105 based on chord length. Forces were calculated via the use of pressure taps along the centreline of the aerofoils. The RMIT Industrial Wind Tunnel (IWT) was used for the testing. Normally 3m wide and 2m high, an extra contraction was installed and the section was reduced to form a width of 295mm. The wing was mounted between walls to simulate 2-D flow. The IWT was chosen as it would allow enough height to reduce blockage effect caused by the aerofoils when at high angles of incidence. The walls of the tunnel were pressure tapped to allow monitoring of the pressure gradient along the tunnel. The results show a delay in the stall of the aerofoils tested with reduced ground clearance. Two of the aerofoils tested showed a decrease in Cl with decreasing ground clearance; the third showed an increase. The Cd of the aerofoils post-stall decreased with reduced ground clearance. Decreasing ground clearance was found to reduce pitch moment variation of the aerofoils with varied angle of attack. The results were used in a simulation of a typical Formula SAE race car.

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