Testing done on the flat-back Ahmed Body and other bluff bodies has shown the existence of a bi-stable reflectional symmetry-breaking wake at Reynolds numbers ranging from 340 to 2.41 x 106. Several methods of flow control, both active and passive, have been used to improve the efficiency of the Ahmed body but their effect on the bi-stable nature of the wake has not been investigated. This work details the experimental investigation done to determine if piezoelectrically driven synthetic jet actuators are capable of suppressing the bi-stable wake effects observed behind the Ahmed Body. The synthetic jets were designed and manufactured to have a maximum total coefficient of momentum of 1.0E-3 with a frequency range up to 2000 Hz or F+ = 17.25. The piezoelectric actuators used were bimorph bending disks with no center shim and were driven by a square waveform. Pressure data was collected from 25 pressure ports on the rear of the model at 625 Hz for 600 seconds per run and filtered using a lowpass filter at 35 Hz to remove interference. Center of Pressure probability distributions and Principle Component Analysis were used to identify wake shapes and modes. Results with no jet actuation showed good agreement with previously published work on the Ahmed Body. It was found that the actuation frequency had an effect on the ability of the synthetic jets to affect the wake. Actuating at F+ = 1 (116 Hz) showed a bi-stable wake with an even distribution between wake modes. Higher actuation frequencies showed either a skewed distribution with a weakening of the bi-stable effects (4 < F+ < 8) or a complete removal of the bi-stable distribution (8 < F+ < 12). Frequencies higher than F+ = 12 did not show any effect on the bi-stable distribution. There was a negative correlation between actuation frequency and average wake pressure; it is theorized that the synthetic jets enhance mixing in the shear layer around the recirculation bubble in the wake to decrease average pressure.

An investigation was conducted to assess the usefulness of anti-optimization in the process of validating of aerodynamic codes. Anti-optimization is defined here as the intentional search for regions where the computational and experimental results disagree. Maximizing such disagreements can be a useful tool in uncovering errors and/or weaknesses in both analyses and experiments. The codes chosen for this investigation were an airfoil code and a lifting line code used together as an analysis to predict three-dimensional wing aerodynamic coefficients. The parameter of interest was the maximum lift coefficient of the three-dimensional wing, CL max. The test domain encompassed Mach numbers from 0.3 to 0.8, and Reynolds numbers from 25,000 to 250,000. A simple rectangular wing was designed for the experiment. A wind tunnel model of this wing was built and tested in the NASA Langley Transonic Dynamics Tunnel. Selection of the test conditions (i.e., Mach and Reynolds numbers) were made by applying the techniques of response surface methodology and considerations involving the predicted experimental uncertainty. The test was planned and executed in two phases. In the first phase runs were conducted at the pre-planned test conditions. Based on these results additional runs were conducted in areas where significant differences in CL max were observed between the computational results and the experiment — in essence applying the concept of anti-optimization. These additional runs were used to verify the differences in CL max and assess the extent of the region where these differences occurred. The results of the experiment showed that the analysis was capable of predicting CL max to within 0.05 over most of the test domain. The application of anti-optimization succeeded in identifying a region where the computational and experimental values of CL max differed by more than 0.05, demonstrating the usefulness of anti-optimization in process of validating aerodynamic codes. This region was centered at a Mach number of 0.55 and a Reynolds number of 34,000. Including considerations of the uncertainties in the computational and experimental results confirmed that the disagreement was real and not an artifact of the uncertainties. / Ph. D.

The purpose of the study was to evaluate the merits of the DiscRotor concept that combine the features of a retractable rotor system for vertical take-off and landing (VTOL) with an integral, circular wing for high-speed flight. Tests were conducted to generate basic aerodynamic characteristics of the DiscRotor in hover and in fixed-wing flight. To assess the DiscRotor during hover, small scale tests were conducted on a 3ft diameter rotor without the presence of a fuselage. A "hover rig" was constructed capable of rotating the model rotor at speeds up to 3,500 RPM to reach tip speeds of 500fps. Thrust and torque generated by the rotating model were measured via a two-component load cell, and time averaged values were obtained for various speeds and pitch angles. It has been shown that the DiscRotor will perform well in hover. Ground Effects in hover were examined by simulating the ground with a movable, solid wall. The thrust was found to increase by 50% compared to the ground-independent case. Pressure distributions were measured on the ground and disc surfaces. Velocity measurements examined the flow field downstream of the rotor by traversing a seven hole velocity probe. A wake behind the rotor was shown to contract due to a low pressure region that develops downstream of the disc. Wind tunnel experimentation was also performed to examine the fixed wing flight of the DiscRotor. These experiments were performed in the VA Tech 6â X6â Stability Tunnel. A model of the fuselage and a circular wing was fabricated based upon an initial sizing study completed by our partners at Boeing. Forces were directly measured via a six degree of freedom load cell, or balance, for free stream velocities up to 200fps. Reynolds numbers of 2 and 0.5 million have been investigated for multiple angles of attack. Low lift-to-drag ratios were found placing high power requirements for the DiscRotor during fixed-wing flight. By traversing a seven-hole velocity probe, velocities in a 2-D grid perpendicular to the flow were measured on the model. The strengths of shed vortices from the model were calculated. A method to improve fixed-wing performance was considered where two blades were extended from the disc. An increase of 0.17 in the CL was measured due to the interaction between the disc and blades. This research utilized a wide range of experiments, with the aim of generating basic aerodynamic characteristics of the DiscRotor. A substantial amount of quantitative data was collected that could not be included in this document. Results aided in the initial designs of this aircraft for the purpose of evaluating the merit of the DiscRotor concept. / Master of Science

Today's high performance aircraft are operating in expanded flight envelopes, often maneuvering at high angular rates at high angles-of-attack, even above maximum lift. Current aerodynamic models are inadequate in predicting flight characteristics in the expanded envelope, such as rapid aircraft departures and other unusual motions. Unsteady flows of aircraft are of real concern. The ability to accurately measure aerodynamic loads directly impacts the ability to accurately model and predict flight. Current wind tunnel testing techniques do not adequately address the data fidelity of a test point under the influence of fluctuating loads and moments. Additionally, forced oscillation test techniques, one of the primary tools used to develop dynamic models, do not currently provide estimates of the uncertainty of the results during an oscillation cycle. Further, in testing models across a range of flight conditions, there are frequently parts of the envelope which are well behaved and require few data points to arrive at a sound answer, and other parts of the envelope where the responses are much more active and require a large sample of data to arrive at an answer with statistical significance. Currently, test methods do not factor changes of flow physics into data acquisition schemes, so in many cases data are obtained over more iterations than required, or insufficient data may be obtained to determine a valid estimate. Methods of providing a measure of data integrity for static and forced oscillation test techniques are presented with examples. A method for optimizing required forced oscillation cycles based on decay of uncertainty gradients and balance tolerances is also presented. / Master of Science

京都大学 / 新制・課程博士 / 博士(工学) / 甲第24293号 / 工博第5066号 / 新制||工||1791(附属図書館) / 京都大学大学院工学研究科社会基盤工学専攻 / (主査)教授 八木 知己, 教授 KIM Chul-Woo, 教授 高橋 良和 / 学位規則第4条第1項該当 / Doctor of Philosophy (Engineering) / Kyoto University / DGAM

Obtaining aerodynamic forces and moments about all three orthogonal axes for arbitrary shapes at arbitrary orientations in a fast manner via a measurement technique specific to Cal Poly’s low-speed wind tunnel to continually obtain the forces and moments under quasi-steady conditions is explored. A Continuous Rotation Technique (CR) uses a 6-DOF load cell and stepper motor to rotate an object about an axis for a complete rotation. The forces and moments acting upon the object pass through the stepper motor and interface plates and recorded by the load cell as the object is rotated continuously a finite number of rotations. An optical encoder installation tracks the progress of a rotation serving indicator between the subsequent rotations to determine starting angles and body slippage. Average loading of measurements at each timestep along the average time for rotation is found. This dataset along the average time is placed in 1-degree bins and averaged to a final dataset of load measurements at 0.5-degree intervals for a complete 360-degree load map. Testing of this technique is performed on aspect ratio 1 circular cylinder and for a sphere. These test cases are representative of the 3D flow features that are prevalent for bodies of finite aspect ratios. The airloads on these canonical shapes were used to calibrate the measurement technique deriving similarity with prior work performed at Georgia Tech's wind tunnel. This thesis showcases the repeatability of the results in a different and smaller wind tunnel with improved certainty on angle measurements. The scope of the thesis tackles the showcase of an initial proof of concept for the incorporation of such a novel measurement technique using Cal Poly resources.

As a part of the revitalization of the supersonic wind tunnel maintained by the Aerospace Engineering Department of Mississippi State University, a new data acquisition and control system became incumbent. Previous data acquisition and control systems used in the operation of the supersonic wind tunnel made use of now outdated hardware and functioned with two central processing units; one processor was used for recording system response, while the other monitored and controlled the tunnel. A new system is required to provide adequate rapid real-time control, along with rapid acquisition of raw tunnel feedback or response data and tunnel pressure data all implemented on one computer processor. The data acquisition and control program that was developed provides synchronized data storage and control output commands, while providing the operator with all essential information and checklists required for running the tunnel. The program is a general user interface developed from previously implemented data acquisition and control systems to perform all predetermined tasks while minimizing operator front panel inputs. When the supersonic wind tunnel repair is completed, the compiled data acquisition and control program will be implemented in the operation of the tunnel. This paper details the development of the supersonic wind tunnel data acquisition and control system employed in the revitalization project. Without the information on the details of the data acquisition and control program, any future editing of the control system source codes will be a laborious task.

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