Master of Science, The Ohio State University, 2025, Aerospace Engineering

Delta wing planforms are dominated by the presence of a leading-edge vortex (LEV). Due to the typically larger sweep angle, the pressure surface shear layer more readily sheds onto the suction surface. The stable LEV retains a low-pressure core which contributes to increased lift. However, when facing an adverse pressure gradient, the vortex is likely to separate and burst. This leads to a loss of lift and further increase in drag. Swept wing planforms suffer from the buildup of spanwise flow at moderate angles of attack. Induced by the offset suction peak, a spanwise pressure gradient bends flow outboard creating a flow separation near the wingtip. As angle of attack is increased the separation region propagates inboard towards the root. At large enough angles of attack, the flow (already prone to separation) forms a LEV similar to a nominal delta wing flow. These flow physics can be altered through the use of flow control. Rather than employing passive strategies which permanently increase parasitic drag and radar cross-section, the current study investigates active flow control (AFC). A multi fence-type AFC method is applied to a 30° swept NACA 643-618 wing and a leading-edge row of outboard-oriented jets is applied to a 45° swept NACA 0012 cropped delta wing. Each wing is detailed to represent its respective planform grouping. Wind tunnel testing of the two geometries is conducted at M = 0.05, correlating to Reynolds numbers on the order of 10^5. Load cell testing and surface oil flow visualizations contributed to both a quantitative and qualitative understanding of AFC's effect. The AFC was found to increase lift performance on both wing planforms, while in certain cases also reducing the drag. Delta wing flow control increased performance through multiple mechanisms. Lift was increased by an elongation of the stable LEV, general outboard flow reattachment, and likely, a local reduction in pressure. Certain jet configurations shifted the LEV separat (open full item for complete abstract)

Committee: Nathan Webb (Committee Member); Jeffrey Bons (Advisor) Subjects: Aerospace Engineering

Master of Science, The Ohio State University, 2023, Mechanical Engineering

The application of supersonic rectangular twin jets (RTJ) is of interest to current and future generations of tactical aircraft. Previous studies and application of supersonic twin jets have shown that increased far-field (FF) noise severely affect the personnel near these aircraft, while their near-field (NF) pressure fluctuations are of concern due to potential structural fatigue of aircraft aft components. The primary objective of this work in the study of RTJ, with an aspect ratio of 2 and a design Mach number of 1.5, was twofold: to characterize unheated baseline flows and to apply active control via localized arc filament plasma actuators (LAFPAs) for understanding of the key physics, and secondly to apply the knowledge gained from unheated jets toward heated flow experiments with a target total temperature ratio (TTR) of 2. The use of NF microphone linear and azimuthal arrays allowed for measurement of NF pressure fluctuations using the calculated overall sound pressure level (OASPL) as well as the detection of the coupling mode by processing signals using a wavelet-based phase and coherence technique. The FF microphone array was utilized for assessment of FF acoustics at various polar and azimuthal angles. High-speed schlieren imagery was used to visualize the flow field and calculate spectral proper orthogonal decomposition (SPOD) energy modes to observe the dynamics of large scale structures (LSS) within the jets' shear layers. Baseline unheated and heated flows showed that the jets couple along the minor axis out-of-phase (OOP) and in-phase (IP) at overexpanded and underexpanded regimes respectively. Under both heated and unheated flow conditions, the LAFPAs showed significant control authority to alter coupling mode, greatly affecting NF internozzle pressure fluctuations and confirming previous literature. Higher excitation frequencies in the unheated jet campaign showed a reduction of both near- and far-field levels dependent of jet Mach number (Mj), an (open full item for complete abstract)

Committee: Mo Samimy (Advisor); Nathan Webb (Committee Member); Lian Duan (Committee Member) Subjects: Acoustics; Aerospace Engineering; Mechanical Engineering

Doctor of Philosophy, The Ohio State University, 2007, Electrical Engineering

Flow control refers to the ability to manipulate fluid flow so as to achieve a desired change in its behavior, which offers many potential technological benefits, such as reducing fuel costs for vehicles and improving effectiveness of industrial processes. An interesting case of flow control is cavity flow control, which has been the motivation of this study: When air flow passes over a shallow cavity a strong resonance is produced by a natural feedback mechanism, scattering acoustic waves that propagate upstream and reach the shear layer, and developing flow structures. These cause many practical problems including damage and fatigue in landing gears and weapons bays in aircrafts. Presently there is a lack of sufficient mathematical analysis and control design tools for flow control problems. This includes mathematical models that are amenable to control design. Recently reduced-order modeling techniques, such as those based on proper orthogonal decomposition (POD) and Galerkin projection (GP), have come to interest. However, a main issue with these models is that the effect of boundary conditions, which is where the control input is, gets embedded into system coefficients. This results in a form quite different from what one deals with in standard control systems framework, which is a set of ordinary differential equations (ODE) where the input appears as an explicit term. Another issue with the standard POD/GP models is that they do not yield to systems that have any apparent structure in their coefficients. This leaves one with little choice other than to neglect the nonlinearities of the models and employ standard linear control theory based designs. The research presented in this thesis makes an effort at closing the gaps mentioned above by 1) presenting a reduced-order modeling method utilizing a novel technique for input separation on POD/GP models, 2) introducing a technique based on averaging theory and center manifold theory so as to reveal certain struct (open full item for complete abstract)

Committee: Andrea Serrani (Advisor) Subjects:

Doctor of Philosophy, The Ohio State University, 2013, Aero/Astro Engineering

In this work, the internal fluid dynamics and frequency characteristics of feedback-free fluidic oscillators are investigated experimentally and numerically. The internal flow field of various scale oscillators was extracted using a refractive index-matched Particle Image Velocimetry (PIV) technique with the help of a PIV phase averaging method and a new sensor setup for simultaneous frequency measurements in refractive index matching fluid. Three different operating regimes (low flow rate, transition and high flow rate regions) and the fluid dynamics of the oscillating behavior in these regimes are revealed with PIV measurements. Flow topologies extracted with PIV measurements differ in these three flow regimes and were found to exhibit various flow features. Frequency measurements were conducted with the use of various experimental techniques including a method that allows non-intrusive measurement. The frequency characteristics were varied depending on properties such as the working fluid, scale and cavity geometry of the fluidic oscillator. Non-dimensional parameters were defined by taking the effects of these variables into account to allow effective comparison of fluidic oscillator designs. Furthermore, 33 modified designs were also tested to provide support for future fluidic oscillator modifications.

Committee: James W. Gregory PhD (Advisor); Mohammad Samimy PhD (Committee Member); Mei Zhuang PhD (Committee Member) Subjects: Aerospace Engineering

Doctor of Philosophy, The Ohio State University, 2016, Electrical and Computer Engineering

Electric power grid is experiencing a major paradigm shift toward a more reliable, efficient, and environmentally friendly grid. The concept of microgrid is introduced to integrate distributed renewable generation in proximity to demands for both environmental and power-efficient promises. A microgrid can be disconnected, or "islanded", from the main grid and operates on its own, providing energy to remote areas or during faults of the main grid for better reliability. Islanded microgrids inherit several different properties from traditional power grids, including uncertain and limited generation, mixed R/X ratio lines, and lack of power inertia from synchronous generators. Those properties pose new challenges for the stable operation of islanded microgrids. The dissertation is dedicated to addressing the control challenges of islanded microgrids. The contribution is twofolds. First, we propose a polynomial time optimal power flow (OPF) solver which finds an optimal operating point for the inverters of the distributed energy resources. The proposed algorithm can account for the cost functions on the reactive generation that are common in microgrids. It also brings new understanding on the conjectures of exact semidefinite programming (SDP) convex relaxation on the OPF problem. Furthermore, we show that without the load over-satisfaction assumption usually seen in the literature, a near global optimum can be found for the OPF problem with arbitrary convex quadratic cost functions. The results are important to both microgrids and the classical OPF problem. Our second major contribution is developing a novel distributed controller that addresses the control challenges originated from limited generation, mixed R/X ratio lines, and lack of power inertia properties of islanded microgrids. The proposed controller can ensure proportional active and reactive power sharing and frequency synchronization while respecting the voltage constraints. Variances of the distributed (open full item for complete abstract)

Committee: Wei Zhang (Advisor); Kevin Passino (Committee Member); Andrea Serrani (Committee Member); Krishnaswamy Srinivasan (Other) Subjects: Electrical Engineering; Mechanical Engineering

Master of Science, The Ohio State University, 2015, Mechanical Engineering

Aerodynamic drag is an increasingly important factor in ground vehicle design due to its large impact on overall fuel economy. The average vehicle drag coefficient has improved significantly since the advent of the automobile, however the marginal gains possible with traditional shape optimization are beginning to decrease. There is increased need to improve the drag coefficient as a means of reducing global fossil fuel consumption, which prompts the automotive industry to investigate additional methods of drag mitigation. One method may be the use of active flow control (AFC) aimed at large scale changes in the flowfield through the introduction of energy perturbations at strategic locations on the vehicle surface. In this study, separation control with fluidic oscillators was examined on a modified square-back Ahmed vehicle model to advance the possibility of AFC application to production vehicles. A fluidic oscillator is a simple pneumatic device that converts a steady flow input into a spatially oscillating jet. This AFC actuator was selected due to its proven separation control efficiency and robustness. Studies involving the application of fluidic oscillator separation control to simplified vehicle models have been conducted by other researchers, however the large parameter space related to oscillator effectiveness yields many unanswered questions. The goal of this work was to answer more of the relevant questions needed to bridge the gap between lab and application. The majority of this experimental study was done in a scale wind tunnel facility owned and operated by a North American automaker at a Reynolds number based on model length of 1.4x10^6 or higher. A modified aft section containing boat-tail flaps and fluidic oscillators was added to the square-back Ahmed model and various parameter sensitivity trends were examined. Parameters of interest included flap angle, oscillator jet location, jet velocity, jet spacing, jet size, moving ground (open full item for complete abstract)

Committee: James Gregory Dr. (Advisor); Jeffrey Bons Dr. (Committee Member) Subjects: Aerospace Engineering; Automotive Engineering; Mechanical Engineering

Doctor of Philosophy (PhD), Wright State University, 2011, Engineering PhD

Low Reynolds number boundary layer separation causes reduced aerodynamic performance in a variety of applications such as MAVs, UAVs, and turbomachinery. The inclusion of a boundary layer separation control system offers a way to improve efficiency in conditions that would otherwise result in poor performance. Many effective passive and active boundary layer control methods exist. Active methods offer the ability to turn on, off, or adjust parameters of the flow control system with either an open loop or closed loop control strategy using sensors. This research investigates the use of a unique sensor called Surface Stress Sensitive Film (S3F) in a closed loop, low Reynolds number separation control system. S3F is an elastic film that responds to flow pressure gradients and shear stress along its wetted surface, allowing optical measurement of wall pressure and skin friction. A new method for installing the S3F sensor to assure a smooth interface between the wall and wetted S3F surface was investigated using Particle Image Velocimetry techniques (PIV). A Dielectric Barrier Discharge (DBD) plasma actuator is used to control laminar boundary layer separation on an Eppler 387 airfoil over a range of low Reynolds numbers. Several different DBD plasma actuator electrode configurations were fabricated and characterized in an open loop configuration to verify separation control of the Eppler 387 boundary layer. The open loop study led to the choice of a spanwise array of steady linear vertical jets generated by DBD plasma as the control system flow effecter. Operation of the plasma actuator resulted in a 33% reduction in section drag coefficient and reattachment of an otherwise separated boundary layer. The dissertation culminates with an experimental demonstration of S3F technology integrated with a control system and flow effecter for closed loop, low Reynolds number separation control. A simple On/Off controller and Proportional Integral (PI) controller were used to clos (open full item for complete abstract)

Committee: Mitch Wolff PhD (Advisor); Rolf Sondergaard PhD (Committee Member); James Menart PhD (Committee Member); Mark Reeder PhD (Committee Member); Joseph Shang PhD (Committee Member) Subjects: Aerospace Engineering; Engineering; Fluid Dynamics; Mechanical Engineering

Master of Science in Engineering (MSEgr), Wright State University, 2005, Mechanical Engineering

Moore, Kenneth Jay. M.S., Department of Mechanical and Materials Engineering, Wright State University, 2005. Large Scale Visualization of Pulsed Vortex Generator Jets. The use of small jets of air has proven to be an effective means of flow control on low Reynolds number turbine blades. Pulsing of these jets has also shown benefits in reducing the amount of air needed to achieve the same level of flow control. An experiment using Hot Wire Anemometry and Particle Image Velocimetry (PIV) has been used to investigate how these pulsed jets interact with the boundary layer to help keep the flow attached. A 25x scaled jet in a flat plate has been utilized. The 25.4 mm diameter jet has a pitch angle of 30° and a skew angle of 90°. Pitch angle is defined as the angle the jet makes with the surface of the plate, and the skew angle is the angle that the projection of the jet on the surface makes with the crossflow. The jet was pulsed at both 0.5 Hz and 4 Hz with varying pulse durations (duty cycles), as well as various blowing ratios (ratio of the jet velocity to the freestream velocity). Duty cycles of 10, 25, 50, and100 percent were implemented at a blowing ratio of unity. Blowing ratios of 0.5, 1, 2, and 4 were implemented at a 50% duty cycle and at 0.5 Hz. Velocity and vorticity planes were obtained at various spanwise locations and used in the characterization of the jetflow. Both the free jet as well as the jet in crossflow were studied. A calibration experiment was also performed using PIV on a rotating disk. The calibration experiment was successful and the PIV results averaged a 1.56% error. The hot wire experiment with the free jet showed that the starting vortex is a key event at the beginning of each cycle, and the end of each cycle included a “kick-back” and a suction effect that could also have an influence on the boundary layer. The PIV experiment was performed first on the free jet, and results were comparable to the hot wire results. When the PIV experiment w (open full item for complete abstract)

Committee: Mitch Wolff (Advisor) Subjects: Engineering, Mechanical

MS, University of Cincinnati, 2005, Engineering : Aerospace Engineering

The overall objective of this work was to develop an active fluidic control system that can effectively manipulate the vortex breakdown location over a highly swept delta wing. By moving the vortex breakdown fore or aft, a pitching moment can be induced on the delta wing without the use of any conventional control surfaces. The active control system can be incorporated into a feedback loop to input a desired pitching moment based on the real-time measured surface pressure. The type of active fluidic control system shown to be the most effective at delaying vortex breakdown was the along-core injection technique. In the previous studies using this technique, the control was an open loop system, that is, the flow field was observed and measured and then injection conditions were changed manually. The process was repeated until optimal conditions were observed. In order to move this technique closer to practical application, two important steps must be taken. The fist step is to close the control loop so that input changes can be instantly made according to the changing flow field. The second step is to simplify and optimize the control system. One advantage this control system has over conventional control surfaces is its simplicity. Conventional control surfaces employ complex and costly combinations of hydraulic and electronic components whereas the along-core injection system requires only solenoid actuators and a source of compressed air. Compressed air is readily available in the form of compressor bleed on most gas turbine engines. Other advantages of this system include elimination of drag due to deployed control surfaces and reduced radar vulnerability. A 60° delta wing model with a maximum span of 15.5 inches and a root chord of 13.5 inches was mounted in a subsonic wind tunnel. The wing is equipped with six control jets with variable azimuthal and pitch angles on the top surface approximately beneath the vortex core. A thorough optimization process was compl (open full item for complete abstract)

Committee: Ephraim Gutmark (Advisor) Subjects: Engineering, Aerospace

Doctor of Philosophy in Engineering, University of Toledo, 2007, Mechanical Engineering

This dissertation presents experimental investigations of several novel active flow control methodologies that have been implemented for aerodynamic control and maneuvering of slender bodies at low and high angles of attack through flow manipulation. For low angles of attack, a U.S. Army Smart Cargo projectile was examined. For high angles of attack a U.S. Air Force countermeasure concept projectile termed DEX (Destructive Expendable) was examined. Low angle of attack control was attempted using two novel separation control techniques: reconfigurable porosity and miniature deployable spoilers. Results show that significant aerodynamic forces are generated by implementing reconfigurable porosity and can be effectively used to steer and maneuver air vehicles. Porous patterns with a “saw-tooth” configuration seem to be the most effective in generating consistent control forces over a wide range of angles of attack. Miniature deployable spoilers successfully demonstrated their ability in producing both positive and negative pitch and yaw controls by modulating the spoiler height and length when used on the boattail and Aero Control Fins (ACFs) of a projectile. The effect of aftbody strake parameters such as shape, locations (axial and azimuthal), deployment height, and number of strakes implemented was examined on a short blunt-nose projectile. Large yaw control authority was attained for a > 40 deg. The largest yaw control authority was produced by a rectangular-shaped strake. A robust closed-loop feedback controller was successfully tested using dynamic wind tunnel experiments to control the coning motion of a projectile. The controller showed good control authority and was capable of attaining and maintaining the commanded roll angle with a tolerance of ± 10 deg. A study was conducted to gain some insights into the fluid mechanics of short blunt-nose bodies of revolution at high angles of attack. Off- and on-surface flow visualization records are collected to study t (open full item for complete abstract)

Committee: T. Ng (Advisor) Subjects:

Doctor of Philosophy, The Ohio State University, 2012, Aero/Astro Engineering

Detailed investigation of the NACA 643-618 is obtained at a Reynolds number of 6.4x104 and angle of attack sweep of -5° < α < 25°. The baseline flow is characterized by four distinct regimes depending on angle of attack, each exhibiting unique flow behavior. Active flow control is exploited from a row of discrete holes located at five percent chord on the upper surface of the airfoil. Steady normal blowing is employed at four representative angles; blowing ratio is optimized by maximizing the lift coefficient with minimal power requirement. The range of effectiveness of pulsed actuation with varying frequency, duty cycle and blowing ratio is explored. Pulsed blowing successfully reduces separation over a wide range of reduced frequency (0.1-1), blowing ratio (0.5–2), and duty cycle (0.6–50%). A phase-locked investigation, by way of particle image velocimetry, at ten degrees angle of attack illuminates physical mechanisms responsible for separation control of pulsed actuation at a low frequency and duty cycle. Temporal resolution of large structure formation and wake shedding is obtained, revealing a key mechanism for separation control. The Kelvin-Helmholtz instability is identified as responsible for the formation of smaller structures in the separation region which produce favorable momentum transfer, assisting in further thinning the separation region and then fully attaching the boundary layer. Closed-loop separation control of an oscillating NACA 643-618 airfoil at Re = 6.4x104 is investigated in an effort to autonomously minimize control effort while maximizing aerodynamic performance. High response sensing of unsteady flow with on-surface hot-film sensors placed at zero, twenty, and forty percent chord monitors the airfoil performance and determines the necessity of active flow control. Open-loop characterization identified the use of the forty percent sensor as the actuation trigger. Further, the sensor at twenty percent chord is used to distinguish between (open full item for complete abstract)

Committee: Jeffrey Bons Dr. (Advisor); Mohammad Samimy Dr. (Committee Member); Jen-Ping Chen Dr. (Committee Member); Andrea Seranni Dr. (Committee Member) Subjects: Aerospace Engineering; Fluid Dynamics

Doctor of Philosophy, The Ohio State University, 2006, Electrical Engineering

In this dissertation, we present and discuss development, implementation, and experimental results of reduced-order model based feedback control of subsonic cavity flows. Model based feedback control of subsonic flows have been studied and implemented by the flow control group at the Collaborative Center of Control Science (CCCS) at the Ohio State University (OSU). The team, composed of researchers from the departments of Electrical Engineering and Mechanical Engineering at the Ohio State, the Air Force Research Laboratory, and NASA Glenn Research Center, possesses synergistic capabilities in all of the required multidisciplinary areas of experimental data acquisition, computational flow simulation, low dimensional modeling, controller design, and experimental validation. The goal of the CCCS effort is to develop tools and methodologies for the use of closed-loop aerodynamic flow control to manipulate the flow over maneuvering air vehicles. The problem chosen for the initial study by the CCCS flow team is control of the resonant noise generated by subsonic flow past an open cavity. This phenomenon is characterized by a strong coupling between the flow dynamics and the flow-induced acoustic field that can lead to self-sustained resonance. Two approaches towards model development have been studied in this dissertation. One aims at representing the physical properties of the system by dynamical models in transfer function forms, referred to as the physics-based linear model in this dissertation. The other approach we have followed is based on proper orthogonal decomposition (POD) and Galerkin projection methods involving the flow governing equations, which is referred to as the nonlinear model or Galerkin model in the dissertation. Each model mentioned above can be further divided into two types: model derived from numerical simulation data and model derived from real time experimental data. Different types of feedback controllers have been designed for corresponding f (open full item for complete abstract)

Committee: Hitay Ozbay Andrea Serrani (Advisor) Subjects: Engineering, Aerospace

Doctor of Philosophy, The Ohio State University, 2006, Computer and Information Science

Fluid animation remains one of the most challenging problems in computer graphics. Research on methods using physics-based simulation for animation has recently increased since this method has the capability of producing realistic fluid behavior. However, the primary drawback to using a simulation method is control of the resulting flow field because it is computationally expensive and highly nonlinear. The main goal of this research is to help users produce physically realistic fluid effects along a NURBS curve that can be specified directly or derived from an image or video. A linear-feedback velocity matching method is used to control the fluid flow. A physically realistic smoke flow along a user-specified path is generated by first procedurally producing a target velocity field, and then matching the velocity field obtained from a three-dimensional flow simulation with the target velocity field. The target velocity field can include various effects such as the small scale swirling motion characteristic of turbulent flows. The swirling motion is achieved by incorporating a vortex particle method into the linear feedback loop. The method is flexible in that any procedurally-generated target velocity field may be integrated with the fluid simulation. The efficacy of this approach is demonstrated by generating several three-dimensional flow animations for complex fluid paths, two-dimensional artistic fluid effects, and realistic tornado animations.

Committee: Raghu Machiraju (Advisor) Subjects: Computer Science

Master of Science, The Ohio State University, 1968, Graduate School

Committee: Not Provided (Other) Subjects:

Master of Science, The Ohio State University, 2024, Aerospace Engineering

Various trailing-edge (TE) Coanda active flow control (AFC) actuator configurations were experimentally investigated in a propeller slipstream flow to simulate hover flight. This thesis will provide a comprehensive evaluation of the configuration factors—in terms of spanwise position and geometric parameters—which impact the aerodynamic and control authority performance of TE Coanda AFC actuators in propeller-driven flow with no freestream velocity. Two main types of TE Coanda AFC configuration experiments were conducted. The first type of testing involved varying the spanwise placement and size of TE AFC while maintaining constant internal geometry and circular Coanda profile shape with continuous slot blowing. The spanwise location testing's objectives were to determine the optimal Coanda AFC actuator location relative to the propeller slipstream and compare Coanda flow control effectiveness to that of a traditional deflection control surface. Two trailing-edge Coanda actuator sections of fixed spanwise length were designed, fabricated, and evaluated in terms of lift force and pitching moment generation at varying spanwise locations. Velocity profile measurements for this study for the case of no freestream flow indicated that the propeller slipstream is asymmetric over the NACA 0012 wing and contracts toward the side of the wing on which the propeller blade descends during rotation, where propeller downwash is experienced. This asymmetry indicated that there may be an optimum location for Coanda AFC actuators at the wing trailing edge which couple with momentum from the propeller in regions of peak slipstream velocity. The second type of testing involved varying TE AFC nozzle and surface profiles while maintaining constant spanwise location. These nozzles included both continuous and discrete slot blowing as well as sweeping jets. Surfaces investigated included circular, elliptical, and biconvex profiles. Each configuration was mounted to the trailing edge of (open full item for complete abstract)

Committee: Matthew McCrink (Committee Member); Jeffrey Bons (Advisor) Subjects: Aerospace Engineering

Master of Science, The Ohio State University, 2024, Aerospace Engineering

Uncrewed aerial vehicles (UAVs) have seen an explosion in development over the last decade. Advances in computing and communication technologies have allowed for rapid innovation in remote and autonomous flight controls, leading to development of new aerospace platforms. Without humans on board, vehicles have restrictions removed that can lead to unique configurations; for example, UAVs can be much smaller and more maneuverable than other platforms. An uncrewed platform is not limited by the maximum G-force its pilot can sustain. Relatively low costs further provide opportunities for innovation in UAV design. Thanks to reduced costs per vehicle and autonomous control systems, UAVs could operate with smaller error margins near the regimes of failure. Swept wings are selected in aircraft design for their aerodynamic performance at high Mach number. By reducing the component of freestream velocity incident on the wing in the direction normal to the leading edge, wing sweep increases the critical Mach number where a marked increase in drag is observed. While wing sweep is desirable for flight at high Mach numbers, a swept wing geometry can pose aerodynamic challenges in low speed, high angle of attack flight. The swept geometry develops a spanwise pressure gradient that is not observed on straight wings. The resulting spanwise flow develops a natural boundary layer separation control that is most effective the wing root. As a swept wing's angle of attack is increased, flow will initially separate near the wingtip and trailing edge. As the angle of attack continues to increase, separation will propagate toward the wing root and leading edge. Taking advantage of these fluid mechanics, a flow control system that interrupts spanwise flow over a swept wing's suction surface can improve aerodynamic performance at high angles of attack. Because the interruption is not needed for high-speed flight, an active flow control method is desirable. The unsteady aerodynamic eff (open full item for complete abstract)

Committee: Jeffrey Bons (Advisor); Randall Mathison (Committee Member) Subjects: Aerospace Engineering

Doctor of Philosophy, The Ohio State University, 2023, Aero/Astro Engineering

The feasibility of an active flow control enabled variable area turbine was explored. Pressurized air was ejected from the nozzle guide vanes to reduce the effective choke area, and mass flow rate through, the turbine inlet. A set of experimental and computational studies were conducted with varying actuator types and parameters to determine their effectiveness and develop models of the flow physics. Preliminary results from a simple quasi-1D converging-diverging nozzle, with an injection flow slot upstream of the throat, showed a 2.2:1 ratio between throttled mass flow rate and injected mass flow rate at a constant nozzle pressure ratio. The penetration of the injection flow and corresponding reduction in the primary flow streamtube were successfully visualized using a shadowgraph technique. Building on this success, a representative single passage nozzle guide vane transonic flowpath was constructed to demonstrate feasibility beyond the quasi-1D converging-diverging nozzle. Both secondary slot blowing from the vane pressure surface and vane suction surface just upstream of the passage throat again successfully reduced primary flow. In addition, fluidic vortex generators were used on the adjacent suction surface to reduce total pressure loss along the midspan and further throttle the primary flow. Computational fluid dynamics simulations were used to explore the effects of a variety of parameters on the flow blockage and actuator effectiveness. Simplified models were developed to describe the relationships of various factors impacting flow blockage, turning angle, and total pressure loss. Finally, the active flow control systems were simulated at engine relevant pressures and temperatures and found to have only a minimal drop in total pressure recovery and effectiveness, which could be predicted by the simplified blockage model.

Committee: Jeffrey Bons (Advisor); Datta Gaitonde (Committee Member); Randall Mathison (Committee Member) Subjects: Aerospace Engineering

Doctor of Philosophy, The Ohio State University, 2023, Aerospace Engineering

The detrimental effects of shock-wave boundary-layer interactions (SBLI) plague many high-speed aerodynamic applications. A primary driver for SBLI studies is the development of cost-effective control strategies. One promising technique is to use passive or active suction to remove low-momentum fluid near vehicle surfaces. This technique is known as boundary-layer bleed, which is commonly applied in high-speed inlets. While the effectiveness of bleed-based control has been proven, bleed introduces its own inefficiencies into the system since it reduces the effective mass capture of the inlet. An optimal bleed system seeks to minimize bleed flow while still achieving sufficient SBLI control. Simulations play a crucial role in successfully optimizing this control methodology; however, resolving the complex flow physics is prohibitively expensive for parametric design studies. Therefore, simplified two-dimensional bleed models are employed nearly exclusively in the inlet design process. However, these simplified models sometimes display discrepancies with complementary experiments, even when localized comparisons are favorable. This fact reveals the first major knowledge gap and objective that this research seeks to address: a quantification of the uncertainties associated with simplified bleed models. The error introduced by bleed models, especially under off-design conditions, is difficult to quantify which results in compounding uncertainties in realistic configurations, where multiple bleed patches are typically deployed. In the present work, the uncertainties associated with bleed-hole modeling are identified by comparing two Reynolds-averaged Navier-Stokes simulations: a three-dimensional simulation with resolved bleed holes and a two-dimensional simulation with a bleed boundary condition. Notable differences are observed in the boundary-layer shape factor, turbulent kinetic energy, and eddy viscosity. The sensitivity of an impinging shock to these differe (open full item for complete abstract)

Committee: Datta Gaitonde (Advisor); John Slater (Committee Member); Lian Duan (Committee Member); Jack McNamara (Committee Member) Subjects: Aerospace Engineering

Master of Science (M.S.), University of Dayton, 2023, Mechanical Engineering

Adaptive materials with programmable surface topography control can be utilized for selective boundary-layer tripping. Liquid crystal elastomers (LCE) have lately gained significant attention to be leveraged to enable these changes via repeatable and controlled out-of-plane deformations. The LCE can be preferentially aligned with circumferential patterns through the thickness of the film, which yields a predictable conical out-of-plane deformation when thermally activated. These reversible and predictable deployments can be utilized to develop a multifunctional surface designed for bodies in flow. This thesis concentrates on the experimental research of LCE behavior for purposes of active flow control via controlled surface topography. First, the deformations of the 12.7-mm diameter patterned LCE samples were characterized using digital image correlation in a controlled pressure chamber under positive and negative gauge pressures. The LCE's performance was highly dependent upon boundary conditions, specimen dimensions, and imprinted defect location relative to the boundary conditions, thus leading to the refinement of the LCE formulation to allow for a higher modulus. Then, to exhibit the potential for flow control, varying arrangements of representative topographical features were 3D-printed and characterized in a preliminary wind tunnel experiment using particle image velocimetry (PIV). Results demonstrated that a two-row arrangement of 1.5-mm feature height produced an asymmetric wake about a 73-mm cylinder, reducing drag while generating lift. Subsequently, a proof of concept model with active LCE elements was fabricated and tested using a force-balance instead of PIV in a wind tunnel. The results of the conceptual model demonstrated that LCEs exhibit the necessary performance to be used in flow control applications.

Committee: Richard Beblo Ph.D (Advisor); Siddard Gunasekaran Ph. D (Committee Member); Gregory Reich Ph. D (Committee Member) Subjects: Aerospace Engineering; Aerospace Materials; Engineering; Materials Science; Mechanical Engineering

Master of Science, The Ohio State University, 2023, Aerospace Engineering

Swept wings are a crucial tool in the design of high-speed aircraft. The sweep on a wing offers many advantages, reducing drag and raising the critical Mach number. These advantages however come at a steep trade off in terms of stall issues at low speed. Many different studies have investigated ways to retain the benefits of swept wings while curtailing some of the negative attributes. Physical flow control devices such as vortex generators and passive boundary layer fences have gained popularity as light weight means to control separation. Once these devices are in place however, there is no turning them off, and as such the drag penalty incurred is permanent. Active flow control seeks to build on the successes of these passive devices by controlling the flow fluidically, keeping the benefits while only paying for them when they are needed. Extensive research has focused on replicating the benefits of a passive boundary layer fence starting with slot blowing, and discretizing to vortex generating jets. The majority of this research has focused on low Reynolds number flows, targeting applications on small unmanned aerial vehicles and control surfaces at landing speeds. As active flow control matures into a production ready technology, it must be able to buy its way on to aircraft. As such, the effects of active flow control need to be understood as Reynolds numbers increase. The work presented in this paper represents an experimental study of a streamwise row of vortex generating jets located at 70% span on a 45° swept wing using a NACA 643-618 airfoil. The jets were actuated with a steady mass flow rate of 100 and 300 SLPM across a Reynolds numbers range from around 50,000 to around 700,000 resulting in mass flow coefficients in a range from 0.003 to 0.02. The wing performance was analyzed at the incidence of stall and deep stall qualitatively using global force measurements collected via a load cell while flow structures were identified using surface oil fl (open full item for complete abstract)

Committee: Jeffrey Bons (Advisor); Randall Mathison (Committee Member) Subjects: Aerospace Engineering; Fluid Dynamics; Technology