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Tomac, Mehmet NazimInternal Fluid Dynamics and Frequency Characteristics of Feedback-Free Fluidic Oscillators
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

Keywords:

experimental fluid dynamics; refractive index matched particle image velocimetry; flow control; flow control actuators; fluidic oscillator; feedback-free fluidic oscillator; jet interactions; jet bifurcations

Hipp, Kyle DControl of a Post-Stall Airfoil Using Pulsed Jets
Master of Science, The Ohio State University, 2016, Aero/Astro Engineering
The performance of active flow control on a NACA 643-618 laminar airfoil at post-stall angles of attack is evaluated using discrete, wall-normal pulsed jets. Actuation is implemented near the leading edge of the airfoil. For actuation periods equal to one convective period, and two convective periods at a chord Reynolds number of 64,000, the time-average lift coefficient increases monotonically as the actuation duty cycle is reduced, for a given blowing ratio. Flow reattachment is achieved following the termination of a short duration pulse, enclosing a separation bubble. The reattachment point propagates towards the trailing edge at a rate three times slower than the convective period of the flow. Extended jet off-times can cause full separation to reoccur should the reattachment point reach the trailing edge, however optimal jet off-times can cause suction pressure to extend over much of the airfoil chord. Higher duty cycle actuation results in a phase shift of the dynamics that appears to be commensurate with the duration of the jet. A disturbance initiated by the termination of the jet causes a delay in the redevelopment of the shear layer and the reattachment of the flow, prohibiting high lift values from being attained. Data collected at multiple post-stall angles of attack show that the rate at which the reattachment point propagates downstream increases with higher angles of attack. At twice the Reynolds number, the same dynamics appear to persist however a weak suction pressure recovery over the extent of the separation bubble reduces the distinction in time-averaged lift between short and long jet pulse durations.

Committee:

Jeffrey Bons (Advisor); James Gregory (Committee Member)

Subjects:

Aerospace Engineering; Engineering; Experiments

Keywords:

aerodynamics; flow control; fluid dynamics

Clifford, Christopher J.An Investigation of Physics and Control of Flow Passing a NACA 0015 in Fully-Reversed Condition
Doctor of Philosophy, The Ohio State University, 2015, Mechanical Engineering

Flow control experiments were performed on a NACA 0015 airfoil in fully-reversed condition, which is anticipated to occur on the retreating blade side of advanced helicopters such as slowed-rotor compound rotorcraft. Control was achieved using nanosecond dielectric barrier discharge (NS-DBD) plasma actuators. The Reynolds number based on a chord length of 203 mm was fixed at 5.0 · 105, corresponding to a freestream velocity of ~38 m/s. Two angles of attack were considered: α = 0° and 15°, each of which is relevant to a particular implementation of slowed-rotor technology.

At α = 0°, the flow resembles that of a flow behind a cylinder. A von Karman vortex street formed in the wake where alternating vortex shedding occurred at a Strouhal number of 0.12. Excitation was performed using an NS-DBD on one side of the airfoil, with plasma formation just upstream of the separation line. However, there was no discernible influence upon the baseline behavior.

At α = 15°, fully separated flow on the suction side extended well beyond the airfoil with naturally shed vortices at a Strouhal number of 0.19. Plasma actuation was evaluated at both the aerodynamic leading-edge (ALE) and aerodynamic trailing-edge (ATE) of the airfoil. The flow responded to the plasma actuation at the ALE by generating organized coherent structures in the shear layer over the separated region. Moderate excitation around the natural shedding Strouhal number had the most significant effects: synchronizing the shedding from the ALE and ATE, creating moderately sized structures that convected far downstream, greatly reducing the separation area, increasing lift, and decreasing drag. Excitation at much higher Strouhal numbers resulted in the flow returning to its natural shedding state, but with less coherent structures that diffused in the wake. This reduced the separation area and significantly reduced drag. Plasma actuation at the ATE caused a reduction in the magnitude of the fundamental and harmonic peaks in pressure spectra over a broad range of excitation Strouhal numbers. Excitation at the ATE altered the structures over the separated region, suggesting an upstream communication. At excitation frequencies higher than the natural shedding frequency, the natural shedding process was disrupted, weakening the naturally shed structures in the wake. Synchronous excitation at the ALE and ATE was predominantly characterized by the associated ALE excitation. Two cases were found where ATE excitation in addition to ALE excitation had a significant effect, but in those two cases, the flow shared characteristics of individual excitation at the ALE and ATE. The resultant flow was somewhere between the two independent excitations. With asynchronous excitation, the addition of ATE excitation counteracted the lift benefits of ALE excitation. As the ATE excitation increased, the amount of lift decreased. The effect on drag was minimal, suggesting that ALE excitation has a much more significant influence on drag than ATE excitation, even at high ATE frequencies.

Committee:

Mo Samimy (Advisor); Igor Adamovich (Committee Member); Datta Gaitonde (Committee Member); James Gregory (Committee Member)

Subjects:

Aerospace Engineering; Fluid Dynamics; Mechanical Engineering

Keywords:

aerodynamics; fluid dynamics; rotorcraft; helicopter; naca 0015; reverse flow; flow control; plasma actuation; instability excitation

POONDRU, SHIRDISHLarge-Eddy Simulation and Active Flow Control of Low-Reynolds Number Flow through a Low-Pressure Turbine Cascade
PhD, University of Cincinnati, 2008, Engineering : Mechanical Engineering

The operating Reynolds numbers (Re) for a low-pressure turbine (LPT) in an aircraft engine can drop below 25,000 during high-altitude cruise conditions, resulting in massive separation and subsequent transition on the blade suction surface. This separation causes a significant loss in the engine efficiency. Hence, accurate prediction of the flow physics at these low-Re conditions is required to effectively implement flow control techniques which can help mitigate separation-induced losses. The present work investigated this low-Re transitional flow through a LPT cascade comprised of the generic Pratt & Whitney “PAKB” blades, using high-order accurate compact numerical schemes in conjunction with large-eddy simulation (LES), with and without subgrid-scale (SGS) models. The study examined the predictive capability of the explicit Smagorinsky and dynamic Smagorinsky SGS models, as compared to the Implicit LES (ILES) technique (LES without an explicit SGS model). The research also implemented active flow control on the LPT blades using momentum injection via surface blowing. All simulations utilized a dual-topology, multi-block, structured grid, and computations were performed on a massively parallel computing platform using MPI-based communications. The baseline cases (without control) were simulated at Re ~ 10,000, 25,000 and 50,000. The computed numerical results for all three cases showed good agreement with available experimental data. The Smagorinsky and dynamic Smagorinsky SGS model results provided no significant improvement over the ILES results because of the low level of energy in the subgrid-scales for the present low-Re flow conditions investigated, and hence, the ILES technique was used for all subsequent flow-control simulations.

Separation control of the LPT flow was implemented using synthetic normal jets, synthetic vortex-generator jets, and pulsed vortex-generator jets (VGJs) at Re ~ 10,000, for four blowing ratios ranging from 0.5 to 4.7, where the blowing ratio is defined as the ratio of the jet-exit velocity magnitude to the local free-stream velocity. All the jets were implemented by specifying an analytical boundary condition at the jet exit surface. The effectiveness of the jets was assessed in terms of the integrated wake loss coefficient values, and the modified Zweifel coefficient values. The Zweifel coefficient represents the component of the integrated blade Cp distribution contributing to the direction of rotation. Among the three types of control jets implemented, the synthetic normal jets were found to be more effective than the synthetic or pulsed VGJs. For pulsed VGJs, the effective blowing ratio was found to be 2.0 in the present study, compared to the value of 0.4 documented in the literature for control at a Re = 25,000, indicating a strong dependence of the effective blowing ratio on Re. The study also examined the flow control mechanisms of the synthetic normal jets and vortex-generator jets. It was found that the mechanism for effectiveness of synthetic jets was a combination of instability-triggered transition and free-stream momentum entrainment. Finally, the synthetic jets and synthetic VGJs were found to be more effective when the jets were located just upstream of the natural separation point.

Committee:

Urmila Ghia (Advisor)

Subjects:

Engineering, Mechanical

Keywords:

Large-eddy simulation; Low-pressure turbines; Active flow control

Gompertz, Kyle AdlerSeparation Flow Control with Vortex Generator Jets Employed in an Aft-Loaded Low-Pressure Turbine Cascade with Simulated Upstream Wakes
Master of Science, The Ohio State University, 2009, Aeronautical and Astronautical Engineering
Detailed pressure and velocity measurements were acquired at Rec = 20,000 with 3% inlet free stream turbulence intensity to study the effects of position, phase and forcing frequency of vortex generator jets employed on an aft-loaded low-pressure turbine blade in the presence of impinging wakes. The L1A blade has a design Zweifel coefficient of 1.34 and a suction peak at 58% axial chord, making it an aft-loaded pressure distribution. At this Reynolds number, the blade exhibits a non-reattaching separation region beginning at 60% axial chord under steady flow conditions without upstream wakes. Wakes shed by an upstream vane row are simulated with a moving row of cylindrical bars at a flow coefficient of 0.91. Impinging wakes thin the separation zone and delay separation by triggering transition in the separated shear layer, although the flow does not reattach. Instead, at sufficiently high forcing frequencies, a new time-mean separated shear layer position is established which begins at approximately 72%Cx. Reductions in area-averaged wake total pressure loss of more than 75% were documented. One objective of this study was to compare pulsed flow control using two rows of discrete vortex generator jets (VGJs). The VGJs are located at 59%Cx, approximately the peak Cp location, and at 72%Cx. Effective separation control was achieved at both locations. In both cases, wake total pressure loss decreased 35% from the wake only level and the shape of the Cp distribution indicates that the cascade recovers its high Reynolds number (attached flow) performance. The most effective separation control was achieved when actuating at 59%Cx where the VGJ disturbance dominates the dynamics of the separated shear layer, with the wake disturbance assuming a secondary role only. On the other hand, when actuating at 72%Cx, the efficacy of VGJ actuation is derived from the relative mean shear layer position and jet penetration. When the pulsed jet actuation (25% duty cycle) was initiated at the 72%Cx location, synchronization with the wake passing frequency (8.7Hz) was critical to produce the most effective separation control. A 20% improvement in effectiveness over the wake-only level was obtained by aligning the jet actuation between wake events. A range of blowing ratios was investigated at both locations to maximize separation reduction with minimal mass flow. The optimal control parameter set for VGJ actuation at 72%Cx does not represent a reduction in required mass flow compared to the optimal parameter set for actuation at 59%Cx. Differences in the fundamental physics of the jet interaction with the separated shear layer are discussed and implications for the application of flow control in a full engine demonstrator are reviewed. Evidence suggests that flow control using VGJs will be effective in the highly unsteady LPT environment of an operating gas turbine, provided the VGJ location and amplitude are adapted for the specific blade profile.

Committee:

Jeffrey Bons, PhD (Advisor); James Gregory, PhD (Committee Member)

Subjects:

Aerospace Materials; Engineering; Fluid Dynamics; Mechanical Engineering

Keywords:

flow control; vortex generator jets; negative jet; low pressure turbine aerodynamics; simulating wakes

Rethmel, Christopher C.Airfoil Leading Edge Flow Separation Control Using Nanosecond Pulse DBD Plasma Actuators
Master of Science, The Ohio State University, 2011, Mechanical Engineering
This work continues an ongoing development and use of dielectric barrier discharge (DBD) plasma actuators driven by repetitive nanosecond pulses for high Reynolds number aerodynamic flow control. These actuators are believed to influence the flow via a thermal mechanism which is fundamentally different from the more commonly studied AC-DBD plasmas. Leading edge separation control on an 8-inch chord NACA 0015 airfoil is demonstrated at various post-stall angles of attack (α) for Reynolds numbers (Re) and Mach numbers (M) up to 1,150,000 and 0.26, respectively (free stream velocity = 93 m/s). The nanosecond pulse driven DBD can extend the stall angle at low Re by functioning as an active trip. At post-stall α, the device generates coherent spanwise vortices that transfer momentum from the freestream to the separated region, thus reattaching the flow. This is observed for all Re and M spanning the speed range of the subsonic tunnel used in this work. A comparison of leading edge separation control between NS-DBD and AC-DBD plasma actuation demonstrates the increased control authority of NS-DBD plasma at higher flow speeds. The actuator is also integrated into a feedback control system with a stagnation-line-sensing hot film on the airfoil pressure side. A simple on/off type controller that operates based on a threshold of the mean value of the power dissipated by the hot film is developed for this system. A preliminary extremum seeking controller is also investigated for dynamically varying Re. Several challenges typically associated with integration of DBD plasma actuators into a feedback control system have been overcome. The most important of these is the demonstration of control authority at realistic takeoff and landing Re and M.

Committee:

Dr. Mo Samimy, PhD (Advisor); Dr. James Gregory, PhD (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

flow control; airfoil separation; aerodynamics; plasma

Kim, YootaiControl of physics-based fluid animation using a velocity-matching method
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

Keywords:

computer graphics; computer animation; fluid animation; physics based animation; fluid simulation; smoke simulation; smoke animation; fluid animation control; path based flow control; linear feedback control; velocity matching; natural phenomena

Atkinson, Michael D.Control of Hypersonic High Angle-Of-Attack Re-Entry Flow Using a Semi-Empirical Plasma Actuator Model
Doctor of Philosophy (Ph.D.), University of Dayton, 2012, Aerospace Engineering

The aim of this dissertation was to explore the possibility of using flow control to stabilize re-entry flight at very high angle-of-attack. This was carried out in three steps: 1) study the structure of representative high angle-of-attack re-entry flows; 2) develop a semi-empirical plasma actuator model that can be applied to control high angle-of-attack re-entry flows; 3) application of the plasma actuator model to study the control of representative re-entry flows. The calculations include viscous and thermochemical non-equilibrium effects, and a high-fidelity physical model to resolve complex flow structure.

The contribution of this dissertation was to provide a detailed description of hypersonic viscous flow around blunt-nosed elliptical cone at very high angle-of-attack. High-fidelity, thermochemical non-equilibrium numerical solutions of high angle-of-attack re-entry flows were not published prior to this research, and thus this research can provide a foundation to calculate, analyze, and describe very high angle-of-attack hypersonic re-entry flows.

Paramount to this dissertation was the development of a new phenomenological MHD plasma actuator model. A semi-empirical actuator model was developed by adding source terms to the momentum equation, vibrational energy equation, and total energy equation, employing an exponential decay function based on the formulations of Kalra et al. and Poggie. This new plasma actuator model was extended from Poggie's model to include thermochemical non-equilibrium effects and expanded from Kalra's et al. two-dimensional model to include three-dimensional effects. Development, validation, and calibration of the plasma actuator model was based on a qualitative comparison to the experiment of Kalra et al. on manipulating turbulent shock-wave/bounday layer interaction using plasma actuators. The effect of the plasma actuators on turbulent shock-wave/boundary-layer interaction was simulated numerically and a detailed description of the complex flow structure with and without actuation was provided.

Finally, application of the plasma actuators to control the complex flow structure of high angle-of-attack re-entry flight vehicles was investigated. To the best of the author's knowledge, no prior research on high angle-of-attack re-entry vehicle control using plasma actuators has been published. Lastly, this dissertation serves as a foundation to compute, analyze, and control complex flow generated around re-entry vehicles at high angle-of-attack.

Committee:

José A. Camberos, PhD (Committee Chair); Jonathan Poggie, PhD (Committee Co-Chair); Aaron M. Altman, PhD (Committee Member); Youssef N. Raffoul, PhD (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Hypersonic, Reentry; CFD; Plasma actuators; Flow Control; High angle of attack

Singhal, Achal SudhirUnsteady Flow Separation Control over a NACA 0015 using NS-DBD Plasma Actuators
Master of Science, The Ohio State University, 2017, Mechanical Engineering
Flow field surrounding a moving body is often unsteady. This motion can be linear or rotary, but the latter will be the primary focus of this thesis. Unsteady flows are found in numerous applications, including sharp maneuvers of fixed wing aircraft, biomimetics, wind turbines, and most notably, rotorcraft. Unsteady flows cause unsteady loads on the immersed bodies. This can lead to aerodynamic flutter and mechanical failure in the body. Flow control is hypothesized to reduce the load hysteresis, and is achieved in the present work via nanosecond pulse driven dielectric barrier discharge (NS-DBD) plasma actuators. These actuators have been effective in the delay or mitigation of static stall. The flow parameters were varied by Reynolds number (Re=167,000-500,000), reduced frequency (k=0.025-0.075), and excitation Strouhal number (Ste=0-10). It was observed that the trends of Ste were similar for all combinations of Re and k, and three major conclusions were drawn. It was first observed that low Strouhal number excitation (Ste<0.5) results in oscillatory aerodynamic loading in the stalled stage of dynamic stall. At high Strouhal number excitation (Ste>2), this behavior is not observed, as in the static stall cases. Second, all excitation resulted in earlier flow reattachment. Lastly, it was shown that excitation resulted in reduced aerodynamic hysteresis and dynamic stall vortex strength. The decrease in the strength of the dynamic stall vortex is achieved by the formation of excited structures that bleed the leading edge vorticity prior to the ejection of the dynamic stall vortex. At sufficiently high excitation Strouhal numbers (Ste˜10), the dynamic stall vortex was suppressed.

Committee:

Mo Samimy (Advisor); Datta Gaitonde (Committee Member); James Gregory (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Flow Control; Aerospace; Dynamic Stall; Plasma Actuators

Crawley, Michael BUnderstanding the Aeroacoustic Radiation Sources and Mechanisms in High-Speed Jets
Doctor of Philosophy, The Ohio State University, 2015, Mechanical Engineering
It has been well-known within the aeroacoustic community that the dominant noise sources in high-speed turbulent jets are related to the large-scale structures which are generated in the initial shear layer by instabilities and rapidly grow, interact, and disintegrate as they convect downstream. However, the exact dynamics of these large-scale structures which are relevant to the noise generation process are less clear. This work aims to study the dynamics of, and noise generated by, the large-scale structures in high-fidelity in a Mach 0.9 turbulent jet using simultaneous pressure and velocity data acquisition systems alongside plasma-based excitation to produce either individual or periodic coherent ring vortices in the shear layer. In the first phase, the irrotational near-field pressure is decomposed into its constitutive acoustic and hydrodynamic components, and two-point cross-correlations are used between the acoustic near-field and far-field in order to identify the dominant noise source region. Building upon the work of previous researchers, the decomposition is performed using a spatio-temporal wavelet transform, which was developed during the current work and found to be more robust than previous techniques. Results indicated that for both individual as well as periodic large-scale structures, the dominant noise reaching the far-field at low angles to the jet axis is being generated in the upstream region of the jet, ending just before the end of the potential core (in a time-averaged sense) in the unexcited jet. This is not to say that no noise is generated outside of this region, just that the most energetic and coherent acoustic radiation is emitted here. The large-scale structure interactions were then investigated by stochastically-estimating the time-resolved velocity fields from time-resolved near-field pressure traces and non-time-resolved planar velocity snapshots. For computational efficiency, the ensemble velocity snapshots were first decomposed into orthogonal modes, and a mapping from the near-field pressure to the expansion coefficients was then produced using a feedforward neural network using backpropagation for learning. The coherent structures generated by the excitation were then identified and tracked using standard vortex identification routines. When exciting the jet at very low frequencies, an individual structure quickly rolled up into a coherent structure within two jet diameters and then advected until roughly four jet diameters downstream, at which point it underwent a rapid disintegration. For the periodically-excited jet, multiple smaller-scale structures are initially apparent just downstream of the nozzle exit. These structures quickly undergo multiple mergings to produce a single large-scale structure with a separation distance that matches the excitation wavelength. Similar to the impulsively-excited structures, these now large-scale structures advect downstream and undergo a rapid disintegration near the end of the potential core. Finally, from Ribner's dilatation-based acoustic analogy the aeroacoustic source terms were computed using the time-resolved velocity field produced by the stochastic estimation. Interpretation of the results is challenging however, due to the number of assumptions and simplifications necessary for the computations given the limitations of the current experimental capabilities. Analysis of the computed source fields found that the coherent structures produced a convected wavepacket-like event, centered on the jet lipline though reaching into the potential core. For the individual vortex rings, a clear modulation of the spatial extent and amplitude was observed as the vortex began to break down just upstream of the end of the potential core. This behavior is also present for the periodic train of vortices observed at higher excitation frequencies, however it is obscured by an amplification of the source in the upstream region where the multiple smaller-scale structures merge. As the excitation frequency was increased, and multiple vortex mergings occurred before the end of the potential core, the aeroacoustic source associated with the merging amplified such that it was distinct from the vortex disintegration source. The results from this work indicate that the disintegration of the coherent ring vortices are the dominant aeroacoustic source mechanism for the Mach 0.9, high Reynolds number jet studied here. However, the merging of vortices in the initial shear layer was also identified as a non-trivial noise source mechanism in high-speed, turbulent jets. Future work will focus on improving the source localization by utilizing acoustic beamforming techniques to identify the source region from the acoustic near-field, in place of the two-point correlations used in this work. Additionally, the structure dynamics and noise generation process will be explored in high-order azimuthal modes.

Committee:

Mo Samimy (Advisor); Datta Gaitonde (Committee Member); James Gregory (Committee Member); Mei Zhuang (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

aeroacoustics; high-speed jets; flow control

Bhattacharya, SamikInvestigation of Three Dimensional Forcing of Cylinder Wake with Segmented Plasma Actuators and the Determination of the Optimum Wavelength of Forcing
Doctor of Philosophy, The Ohio State University, 2013, Aero/Astro Engineering
The wake of a circular cylinder was forced in a three dimensional manner with the help of spanwise non-uniformly located dielectric barrier discharge plasma actuators. The segmented actuators created a square wave forcing profile due to alternate existence of plasma and no plasma region on the cylinder at ±80 degree from the forward stagnation point. The wavelength (lambda) of the actuators was varied from 1d to 6d (d = diameter of the cylinder) and the induced velocities from all the actuators were matched. Below a threshold power level, vortex shedding was not significantly attenuated for any wavelengths, although distinct patterns of streamwise vortices formed for lambda > 2d. Due to this, the near wake developed a spanwise wavy structure: in the no-plasma region the spanwise vortex came closer to the centerline, while it was shifted away from the centerline in the plasma region. In this low-power forcing regime, the drag in the wake was not significantly affected as compared to the case when the forcing power level was above the threshold. Forcing with the lambda = 4d actuator at this high power level resulted in substantial drag reduction and an increase in three-dimensionality of the near wake accompanied by a drastic attenuation of vortex shedding. The energy of the shedding was distributed over a wide range of frequencies, with the most prominent frequency in the velocity spectra being lower than the natural shedding frequency indicating reduced formation length. For lambda > 4d, the wake behind the no-plasma region was much wider compared to that of the plasma region along with a clear difference in formation length, which resulted in higher drag than the lambda = 4d case. This finding lead to the recognition of lambda = 4d as the optimum wavelength of forcing. In the high power-forcing regime, counter-rotating vortices were formed in the horizontal plane at the edge of each buried electrode, which created an alternate zone of backflow in the no-plasma region and downstream flow in the plasma region. Appearance of a saddle point marked the boundary of the backflow region indicating increased level of strain in the no-plasma region in the high power forcing case. The transition from a lower power level below the threshold to that above it was marked by a change in the sense of rotation of the streamwise vortices for lambda > 3d. This change was the result of the dominance of vortices created by high power forcing over a secondary vorticity whose sign matched that of the low power case. It is concluded that this change in the sense of rotation of streamwise vortices with power level of actuation is an inherent feature of the segmented forcing for any wavelength; however, it is the optimum wavelength for which this transition is achieved with minimum induced velocity. The streamwise vortices in the low power case could not disrupt the shedding process, whereas in the high power regime, existence of strong counter-rotating vortices created backflow in front of no plasma region, which diverted flow from the spanwise vortex and thus stymied its growth.

Committee:

James W. Gregory (Advisor)

Subjects:

Aerospace Engineering; Engineering; Experiments; Fluid Dynamics; Mechanical Engineering

Keywords:

cylinder wake, flow control, plasma actuator, stream wise vortex, three dimensional forcing, active forcing

Hirt, Stefanie MExperimental Study of Fillets to Reduce Corner Effects in an Oblique Shock-Wave/Boundary-Layer Interaction
Master of Sciences, Case Western Reserve University, 2015, EMC - Aerospace Engineering
A series of tests were conducted in the 15 cm x 15 cm supersonic wind tunnel at NASA Glenn Research Center that focused on corner effects of an oblique shock-wave/boundary-layer interaction. In an attempt to control the interaction in the corner region, eight corner fillet configurations were tested. Three parameters were considered for the fillet configurations; the radius, the fillet length, and the taper length. Fillets effectively reduced the boundary-layer thickness in the corner; however, there was an associated penalty in the form of increased boundary-layer thickness at the tunnel centerline. Larger fillet radii caused greater reductions in boundary-layer thickness along the corner bisector. To a lesser, but measureable, extent, shorter fillet lengths resulted in thinner corner boundary layers. Overall, of the configurations tested, the largest radius resulted in the best combination of control in the corner, evidenced by a reduction in boundary-layer thickness, coupled with minimal impacts at the tunnel centerline.

Committee:

Jaikrishnan Kadambi (Advisor); Paul Barnhart (Committee Member); Yasuhiro Kamotani (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Shock-wave boundary-layer interaction; flow control; corner fillets

Bloxham, Matthew JonA Global Approach to Turbomachinery Flow Control: Loss Reduction using Endwall Suction and Midspan Vortex Generator Jet Blowing
Doctor of Philosophy, The Ohio State University, 2010, Aeronautical and Astronautical Engineering

A flow control scheme using endwall suction and vortex generator jet (VGJ) blowing was employed to reduce the turbine passage losses associated with the endwall flow field and midspan separation. Unsteady midspan control at low Re had a significant impact on the wake total pressure losses, decreasing the area-average losses by 54%. The addition of leading edge endwall suction resulted in an area-average total pressure loss reduction of 57%. The minimal additional gains achieved with leading edge endwall suction showed that the horseshoe vortex was a secondary contributor to endwall loss production (primary contributor- passage vortex).

A similar flow control strategy was employed with an emphasis on passage vortex (PV) control. During the design, a theoretical model was used that predicted the trajectory of the passage vortex. The model required inviscid results obtained from two-dimensional CFD. It was used in the design of two flow control approaches, the removal and redirection approaches. The emphasis of the removal approach was the direct application of flow control on the endwall below the passage vortex trajectory. The redirection approach attempted to alter the trajectory of the PV by removing boundary layer fluid through judiciously placed suction holes. Suction hole positions were chosen using a potential flow model that emphasized the alignment of the endwall flow field with inviscid streamlines. Model results were validated using flow visualization and particle image velocimetry (PIV) in a linear turbine cascade comprised of the highly-loaded L1A blade profile.

Detailed wake total pressure losses were measured, while matching the suction and VGJ massflow rates, for the removal and redirection approaches at ReCx=25000 and blowing ratio, B, of 2. When compared with the no control results, the addition of steady VGJs and endwall suction reduced the wake losses by 69% (removal approach) and 68% (redirection approach). The majority of the total pressure loss reduction resulted from the spanwise VGJs, while the suction schemes provided modest additional reductions (<2%). At ReCx=50000, the endwall control effectiveness was assessed for a range of suction rates without midspan VGJs. Area-average total pressure loss reductions of up to 28% were measured in the wake at ReCx=50000, B=0, with applied endwall suction employed using the removal scheme (compared to no suction at ReCx=50000). At which point, the total pressure loss core was almost completely eliminated. PIV showed that the endwall suction changed the location of the PV eliminating its influence on the suction surface of the turbine blade. Suction with the removal approach removed the corner vortex (CV) increasing the available span by more than 10%. The redirection approach was less effective at higher suction rates due to the continual presence of the CV.

A system analysis was performed that compared the power needed to operate the flow control system to the power gained by the system. The power gains were assessed by comparing the change in lift and wake total pressure losses with and without flow control. The resultant power ratio showed that only 23% of the total power gained was needed to operate the flow control system for an L1A rotor at ReCx=50000, B=2.

Committee:

Jeffrey Bons, PhD (Advisor); James Gregory, PhD (Committee Member); Jen-Ping Chen, PhD (Committee Member); Mohammad Samimy, PhD (Committee Member)

Subjects:

Engineering

Keywords:

turbomachinery; vortex generator jet; flow control; turbine; gas turbine; horseshoe vortex; passage vortex; corner vortex; separation

Marks, Christopher R.Surface Stress Sensors for Closed Loop Low Reynolds Number Separation Control
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 close the control loop.

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

Keywords:

low Reynolds number; fluid dynamics; surface stress sensitive film; flow control; separation control; S3F; plasma actuator; dielectric barrier discharge;

QUET, Pierre-Francois DA ROBUST CONTROL THEORETIC APPROACH TO FLOW CONTROLLER DESIGNS FOR CONGESTION CONTROL IN COMMUNICATION NETWORKS
Doctor of Philosophy, The Ohio State University, 2002, Electrical Engineering
In this dissertation a control theoretic approach is taken to design and analyze various flow controllers for congestion control in communication networks. First, a robust controller is designed for explicit-rate congestion control in single-bottleneck network. The controller guarantees stability robustness with respect to uncertain time-varying multiple time-delays in different channels, brings the queue length of the bottleneck node to a desired value asymptotically and satisfies a weighted fairness condition. The use of the outgoing link capacity is further investigated to improve performance. Also, some variations on a linear model of Active Queue Management supporting Transmission Control Protocol flows are used to design a robust AQM controller, and to analyse the performance and stability of Multi-Level ECN and Traffic-load based AQM schemes.

Committee:

Hitay Ozbay (Advisor)

Keywords:

Communication networks; Flow control; Uncertain time-varying multiple time-delays; Active queue management; Fluid flow model

Moore, Kenneth JayLarge Scale Visualization of Pulsed Vortex Generator Jets
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 was performed on the jet in crossflow, it was clear that both the beginning and ending events of the jet cycle were keys to eliminating or delaying flow separation.The effect of the beginning and ending events can be used to keep the flow attached for longer periods of time by increasing the frequency of the jet pulse. Due to limitations of the setup, higher frequency cases could not be studied. However, the experiment was successful in controlling a separated crossflow for blowing ratios greater than unity. The larger blowing ratios resulted in larger attachment size, and were able to sustain attachment for longer time periods.

Committee:

Mitch Wolff (Advisor)

Subjects:

Engineering, Mechanical

Keywords:

PIV; Flow Control; Separation Control; VGJ; Pulsed Jets

Lakhamraju, Raghava RajuCharacterization of the jet emanating from a self-exciting flexible membrane nozzle
PhD, University of Cincinnati, 2012, Engineering and Applied Science: Aerospace Engineering
The present research investigates the development and characterization of a novel self-exciting flexible membrane nozzle. Upon excitation (oscillations that are produced by exerting tension at the nozzle exit and passing air through it), the flexible nozzle is capable of producing time-dependent flow that is fairly consistent at a flow condition (a particular tension and volume flow rate of air). The fluidic device is a passive means of enhancing mixing as there is no external excitation mechanism. The resultant flow is self-excited over a range of conditions and produces pulsatile flow that is excited by the motion of the flexible membrane. The baseline configuration of the flexible membrane nozzle involves symmetrical placement of the edges at the nozzle exit. The exit of the nozzle offers variable area geometry, with the shape approximately resembling a variable aspect ratio ellipse. Particle Image Velocimetry (PIV) is employed to illustrate and characterize the large-scale flow structures of the jet motion and the eduction of coherent structures was performed using Proper Orthogonal Decomposition (POD). For a particular nozzle diameter, the flow conditions are controlled by the tension applied to the flexible nozzle and volume flowrate of air through it. PIV measurements have been conducted mainly along the mid-minor axis plane since the crucial flow structure interactions occur in this plane due to the nozzle operation. Based on a set of experiments conducted within the physical limitations of the nozzle, the near field of the nozzle exit was found to be governed by the interactions of two sets of large-scale vortical structures - starting vortices and entrainment vortices (features of pulsatile flows) and the exact nature of their evolution is dependent on the operating conditions. As in elliptic jets, the near field of the nozzle is found to be extremely sensitive to the initial conditions (nozzle configuration). A cross-spectral analysis is also performed in the near field of the jet using two hot-wire anemometers to characterize the evolution of large-scale flow structures for the various flow conditions. For a baseline nozzle in fully closed configuration, for a given tension at the nozzle exit, increase in volume flow tends to produce higher jet spread (more prominent at lower tensions) and increased range of turbulence production. For a particular flow rate, increase in tension results in a more symmetric jet along the centerline and high turbulence production in the near field. Under certain flow conditions, the dynamic flapping of the jet generated half-width spreading rates that exceeded that of slot nozzles. The flow characteristics are compared to that of existing nozzles that generate high mixing rates at the exit. The application of POD on the PIV information shows that the reconstructed images from few modes provide decent illustrations of the flow structures with filtered effect on the turbulent flow field. In essence, this analysis separates the oscillation of the jet from the velocity fluctuations due to the turbulent flow behavior. In the current study, with certain operating conditions, increased half-width spreading rates and enhanced centerline velocity decay can be generated. A predominant flapping jet or highly turbulent jet at the nozzle exit can be achieved by modifying the flow conditions.

Committee:

Ephraim Gutmark, PhD DSc (Committee Chair); Shaaban Abdallah, PhD (Committee Member); Jeffrey Kastner, PhD (Committee Member); Paul Orkwis, PhD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

flow control;flexible nozzle;self-excitation;pulsatile jet;noncircular jet;starting and entrainment vortices;

Frankhouser, Matthew WilliamNanosecond Dielectric Barrier Discharge Plasma Actuator Flow Control of Compressible Dynamic Stall
Master of Science, The Ohio State University, 2015, Aero/Astro Engineering
Dynamic stall is a performance-limiting phenomenon experienced by rotorcraft in directional and maneuvering flight. Dynamic stall occurs on the retreating blade due to the high angles of attack that are experienced by the blades. Increasing the angle of attack is required to overcome the asymmetry of lift across the rotor disk that is a result from the velocity disparities between the advancing and retreating blade. This works sets out to study and improve the performance of a dynamically pitching NACA 0015 airfoil. The airfoil is subjected to both an incompressible and compressible flow field to simulate the dynamics of a rotor blade with cyclic pitching. In this experimental investigation of dynamic stall flow control, the effectiveness of nanosecond dielectric barrier discharge (NS-DBD) plasma actuation will be evaluated as a means to exert control authority. The NS-DBD plasma actuation is generated by a high-voltage magnetic compression pulsed power supply that was designed and built at The Ohio State University. To measure the influence of plasma actuation on the flow, surface pressures on the airfoil were measured through discrete pressure taps located on both the suction and pressure surfaces. The surface pressures are used to calculate the lift and moment during the dynamic pitching cycle. To visualize the compressibility effects in the outer flow, shadowgraph imagery was used to capture features in the flow around the leading edge of the test article. Tests were conducted at static and oscillating angles of attack at both Mach 0.2 and 0.4, and Reynolds numbers of 1.2 million and 2.2 million respectively. Pitch oscillations were conducted at reduced frequencies of k = 0.05. Actuation frequencies varied from non-dimensional frequencies (F + ) of 0.78 to 6.09. Surface pressures acquired at Mach 0.2 without actuation applied agreed with historical data at static angles of attack, validating that the application of the actuator had limited intrusiveness to the flow. When subjected to pitch oscillations, plasma actuation reduced the severity of lift and moment stall by altering the development of the dynamic stall vortex at Mach 0.2. At Mach 0.4, marginal improvements were gained through actuation. Excitation resulted in a strong dynamic stall vortex that convected more slowly in comparison to the baseline case. Shadowgraph imagery revealed lambda shock waves forming over the first 15 percent of the airfoil chord in the same proximity of th

Committee:

James Gregory, PhD (Advisor); Jeffrey Bons, PhD (Committee Member); Mo Samimy, PhD (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

dynamic stall; nanosecond dielectric barrier discharge plasma actuation; flow control

Benton, Stuart IraCapitalizing on Convective Instabilities in a Streamwise Vortex-Wall Interaction
Doctor of Philosophy, The Ohio State University, 2015, Aero/Astro Engineering
Secondary flows in turbomachinery and similar engineering applications are often dominated by a single streamwise vortex structure. Investigations into the control of these flows using periodic forcing have shown a discrete range of forcing frequency where the vortex is particularly receptive. Forcing in this frequency range results in increased movement of the vortex and decreased total pressure losses. Based on the hypothesis that this occurs due to a linear instability associated with the Crow instability, a fundamental study of instabilities in streamwise vortex-wall interactions is performed. In the first part of this study a three-dimensional vortex-wall interaction is computed and analyzed for the presence of convective instabilities. It is shown that the Crow instability and a range of elliptic instabilities exist in a similar form as to what has been studied in counter-rotating vortex pairs. The Crow instability is particularly affected by the presence of a solid no-slip wall. Differences in the amplification rate, oscillation angle, Reynolds number sensitivity, and transient growth are each discussed. The spatial development of the vortex-wall interaction is shown to have a further stabilizing effect on the Crow instability due to a “lift-off” behavior. Despite these discoveries, it is still shown that amplitude growth on the order of 20% is possible and transient growth mechanisms might result in an order-of-magnitude of further growth if properly initiated. With these results in mind, an experiment is developed to isolate the streamwise vortex-wall interaction. Through the use of a vortex generating wing section and a suspended splitter plate, a stable interaction is created that agrees favorably in structure to the three-dimensional computations. A small synthetic jet actuator is mounted on the splitter plate below the vortex. Phase-locked stereo-PIV velocity data and surface pressure taps both show spatial amplification of the disturbance in a frequency range which agrees well with the prediction for the Crow instability. An analysis of the vortex response shows a primarily horizontal oscillation of the vortex column which strongly interacts with the secondary vortex structure that develops in the boundary layer.

Committee:

Jeffrey Bons, Ph.D. (Advisor); Mohammad Samimy, Ph.D. (Committee Member); James Gregory, Ph.D. (Committee Member); Jen-Ping Chen, Ph.D. (Committee Member)

Subjects:

Aerospace Engineering; Fluid Dynamics

Keywords:

active flow control; vortex; linear stability; synthetic jet actuator; wind tunnel; particle image velocimetry; turbomachinery; low pressure turbine;

Yugulis, Kevin LeeHigh Subsonic Cavity Flow Control Using Plasma Actuators
Master of Science, The Ohio State University, 2012, Aero/Astro Engineering

Localized arc filament plasma actuators have been used to control pressure fluctuations in a cavity with a length to depth ratio of 4.86. The rear wall of the cavity is inclined 30° above the horizontal plane and the cavity length is 61.7 mm, measured from the leading edge of the cavity to the mid-plane of the ramp. Five actuators have been uniformly distributed along the span of the wind tunnel at 1 mm upstream to the cavity leading edge. Experiments were conducted at Mach 0.6 and a Reynolds number of approximately 2x105 based on cavity depth. Forcing was conducted quasi-two-dimensionally and three-dimensionally. With this Mach number and geometry, the cavity was strongly resonating at the 2nd Rossiter mode corresponding to a frequency of 2.5 kHz.

Time-resolved pressure measurements were used to assess the effectiveness of the actuators. Forcing quasi-two-dimensionally was found to be very effective, achieving a reduction in peak tone magnitude of over 20 dB and a reduction in broadband SPL of up to 5 dB. In general, the results for forcing in this manner were extremely sensitive to forcing frequency. The most effective forcing frequency was found at approximately 3300 Hz. Forcing was also conducted in several three-dimensional configurations. Overall certain three-dimensional configurations were found to be more effective than the quasi-two-dimensional forcing, and significantly less sensitive to frequency.

Particle image velocimetry was used to understand how the forcing affected the shear layer. Interesting vortex dynamics such as possible vortex merging was observed, the details of which help to understand why certain frequencies are more effective than others. It was determined that the vortices in the shear layer are significantly weaker under three-dimensional forcing compared to quasi-two-dimensional forcing. This could help to explain the overall increase in effectiveness seen with three-dimensional forcing.

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Committee:

Mo Samimy, Dr. (Committee Co-Chair); James Gregory, Dr. (Committee Co-Chair)

Subjects:

Aerospace Engineering

Keywords:

aerodynamics; plasma actuators; cavity flow; active flow control

Hollis, Rebecca M.The Effects of Localized Blade Endwall Suction on Surface Heat Transfer
Master of Science, The Ohio State University, 2009, Aeronautical and Astronautical Engineering
Two methods of flow control were designed to mitigate the effects of the horseshoe vortex structure (HV) at an airfoil/endwall junction. An experimental study was conducted in a low-speed wind tunnel to quantify the effects of localized boundary layer removal on surface heat transfer. A transient infrared technique was used to measure theconvective heat transfer values along the surrounding surface. Particle image velocimetry was used to collect the time-mean velocity vectors of the flow field adjacent to the endwall along three planes of interest. Boundary layer suction was applied through a thin slot, in the leading edge of the airfoil at two heights. The first height, referred to as Method 1, was immediately along the endwall, the second height, Method 2, was located at ~1/3 of the approaching boundary layer height. Five suction rates were tested, 0%, 6.5%, 11%, 15% and 20% of the boundary layer mass flow was removed at a constant rate. Both methods reduced the effects of the HV with increasing suction on the symmetry, 0.5-D and 1-D planes. Method 2 performed better at reducing the surface heat transfer but Method 1 outperformed Method 2 aerodynamically by completely removing the HV structure when the 11% suction rate was applied. This method however produced other adverse effects such as high surface shear stress and localized areas of high heat transfer near the slot edges at high suction rates.

Committee:

Jeffrey P. Bons, PhD (Committee Chair); James W. Gregory, PhD (Committee Member)

Subjects:

Fluid Dynamics

Keywords:

horseshoe vortex; flow control; boundary layer removal; heat transfer

STANEK, MICHAEL JOSEPHA NUMERICAL STUDY OF THE EFFECT OF FREQUENCY OF PULSED FLOW CONTROL APPLIED TO A RECTANGULAR CAVITY IN SUPERSONIC CROSSFLOW
PhD, University of Cincinnati, 2005, Engineering : Aerospace Engineering
Stabilization of turbulent free shear flows is a poorly understood, and recently discovered flow phenomenon,not described in modern textbooks on fluid dynamics. This dissertation describes the design and large-scale experimental test of one type of flow control actuator, a rod in crossflow, which is shown to pulse at high frequency (relative to the dominant instabilities of a turbulent free shear layer), and in the process, locally stabilizes that shear layer. The shear layer in question spans a cavity (representative of a 1/10th scale modern aircraft weapons bay) in supersonic (Mach 1.2)crossflow. Without the high frequency flow control, the cavity experiences acoustic resonance (and the creation of large coherent vortical structures), which creates sound pressure levels high enough to fatigue aircraft components. With the high frequency control (and the local shear layer stabilization), the sound pressure levels are rendered benign. Evidence of suppression due to other types of high frequency pulsing devices (primarily resonance tube type designs) is also presented. A numerical study is undertaken to investigate the nature of the stabilization and acoustic suppression. An implicit, 2nd-order in space and time flow solver, coupled with a recently-developed hybrid RANS - LES turbulence model by Nichols,is utilized in a Chimera-based parallel format, to numerically simulate both the unsuppressed cavity in resonance, as well as the effect of pulsing flow control. Due to the limited ability to vary frequency using a rod in crossflow type device, a pulsed jet device is simulated instead. Frequency (and in a limited number of cases, amplitude) of pulse is varied, from 0 Hz (steady) up to 5000 Hz. The change in the character of the flow control effect as pulsing frequency is changed is described, and linked to changes in acoustic levels. The observed local stabilization of the cavity turbulent shear layer is shown in simulation to be the result of a violent instability and breakdown of the injected vortical structure caused by the high frequency pulsing. This behavior is only observed in simulation above a certain critical frequency. Below this critical frequency, pulsing is shown in simulation to provide little benefit with respect to suppression of high cavity acoustic levels.

Committee:

Dr. Stanley Rubin (Advisor)

Subjects:

Engineering, Aerospace

Keywords:

Fluid Dynamics; Flow Control; High Frequency; Cavity; Weapons Bay

Gan, SubhadeepActive Separation Control of High-Re Turbulent Separated Flow over a Wall-Mounted Hump using RANS, DES, and LES Turbulence Modeling Approaches
PhD, University of Cincinnati, 2010, Engineering : Mechanical Engineering
Most practical flows in engineering applications are turbulent, and exhibit separation which is generally undesirable because of its adverse effects on performance and efficiency. Therefore, control of turbulent separated flows has been a topic of significant interest as it can reduce separation losses. Often, flow control work employs passive techniques to manipulate the flow. Passive-flow control does not require any additional energy source to achieve the control, but is accompanied by additional viscous losses. It is more desirable to employ active techniques as these can be turned on and off, depending on the flow control requirement. The primary goal of the present work is to numerically investigate a high Reynolds number turbulent separated flow. It is Case 3 of the 2004 CFD Validation on Synthetic Jets and Turbulent Separation Control Workshop, http://cfdval2004.larc.nasa.gov/case3.html, conducted by NASA for the flow over a wall-mounted hump. Followed by the baseline flow simulation, i.e, without flow control, active flow control will be investigated using both steady suction jet as well as a "synthetic" jet. The present work also implements the use of two jets (steady suction and synthetic jets) as have not been previously implemented for this flow model. For the synthetic-jets case, the work also studies the effect of two jets in opposite phase. The secondary goal of this work is to bring together a variety of turbulence models and simulation approaches for one flow problem. The flow is simulated using steady and unsteady-state three-dimensional RANS equations-based turbulence models and three-dimensional time-dependent DES and LES methods. Multiple turbulence modeling approaches help to ascertain what models are most appropriate for capturing the physics of this complex separated flow. The results will help us better decide what models to choose for flows with adverse pressure gradients, flow separation and control of separated flows. For the flow over the wall-mounted hump, the simulation results agree well with experiment. Significant computational-resources savings was realized by using an analytical exit velocity profile for the active flow control jets, instead of simulating the entire flow-control manifold without sacrificing the quality of the work. Results compared with experimental values were surface pressure coefficient, skin friction coefficient, mean velocity profiles, Reynolds stresses and flow reattachment locations. Simulation results show some degree of variation with experimental results in the separated flow region. The steady-suction active control was able to reduce the reattachment length the most. The region of negative streamwise velocity was the smallest in the active flow control with steady suction. The multiple jets cases, with steady suction and synthetic jets, were able to reduce the length of separation bubble in comparison to the corresponding single jet cases. The synthetic jets case, using two jets in opposite phase, was able to achieve the most uniform velocity field in the separation bubble region. The work shows great promise in implementing active flow control, using single and multiple jets, for separated flow at high Reynolds number.

Committee:

Urmila Ghia, PhD (Committee Chair); Milind Jog, PhD (Committee Member); Kirti Ghia, PhD (Committee Member); Donald French, PhD (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

CFD;Separated flow;Active flow control;Steady Suction;Synthetic Jet

Sinha, AniruddhaDevelopment of reduced-order models and strategies for feedback control of high-speed axisymmetric jets
Doctor of Philosophy, The Ohio State University, 2011, Mechanical Engineering

Localized arc filament plasma actuators have demonstrated significant potential in controlling high-speed and high Reynolds number jets in open-loop. The two primary goals of jet control are either noise reduction or bulk mixing enhancement. This research develops the tools for implementing feedback for this flow control system. The particular jet considered is a Mach 0.9 axisymmetric configuration with Reynolds number 670,000.

The jet near-field pressure is well-suited for real-time non-intrusive observation of the flow state. Its response to forcing is similar to that of the far acoustic field. Forcing near the jet column mode results in amplification; forcing close to the shear layer mode yields attenuation. As a preliminary effort, two model-free feedback control algorithms are developed and implemented for online optimization of the forcing frequency to extremize the near-field pressure fluctuations. The steady-state behavior of the jet under closed-loop control matches the optimal open-loop results. However, the responsiveness of the controllers is poor since the dynamics of the jet are neglected.

The first step in model-based feedback control is the development of a reduced-order model for the unforced jet. A cylindrical domain spanning the end of the jet potential core is chosen for the significance of its dynamics to the applications at hand. A combination of proper orthogonal decomposition and Galerkin projection is used to reduce the Navier-Stokes equations into a small set of ordinary differential equations employing empirical data. Extensive validation is performed on two existing numerical simulation databases of jets spanning low and high Reynolds numbers, and subsonic and supersonic speeds. Subsequently, a 35-dimensional model is derived from experimental data and shown to capture the most important dynamical aspects. The short-term prediction accuracy is found to be acceptable for the purpose of feedback control. The statistics from intermediate-term simulations also display good agreement with experimental observations. However, the simulated trajectories from the model grow unbounded beyond about 50 flow time steps.

The model of the unforced jet is augmented to incorporate the effects of plasma actuation. The periodic forcing is modeled as a deterministic pressure wave specified on the boundary of the modeling domain. Forcing of the 35-dimensional model around the jet column mode produces acceptable simulation of the nonlinear response that is observed in experiments. However, the sensitivity of the response is sharper than expected.

Various strategies are adapted and implemented for real-time estimation of the flow state from near-field pressure information. These are assessed using the numerical database of the low Reynolds number Mach 0.9 jet. The most useful strategy is the linear time invariant filter employing a linearized version of the dynamic model developed for the unforced jet. This is shown to be more accurate than the more common stochastic estimation techniques, while requiring minimal processing resources.

The actual design of the feedback laws in this very challenging problem remains an open question. The possible directions to take in addressing this in the future are discussed.

Committee:

Mo Samimy, PhD (Advisor); Andrea Serrani, PhD (Committee Co-Chair); Datta Gaitonde, PhD (Committee Member); Jeffrey Bons, PhD (Committee Member)

Subjects:

Acoustics; Aerospace Engineering; Applied Mathematics; Fluid Dynamics; Mechanical Engineering

Keywords:

reduced-order model; flow control; proper orthogonal decomposition; Galerkin projection; jet; turbulence

Webb, Nathan JosephControl of Supersonic Mixed-Compression Inlets Using Localized Arc Filament Plasma Actuators
Master of Science, The Ohio State University, 2010, Mechanical Engineering
Shock wave/boundary layer interactions (SWBLIs) occur in many supersonic internal flow applications, specifically in mixed compression inlets, as well as in external flows. In this study a nominally Mach 2 mixed compression inlet is modeled by two experimental setups: 1) A compression ramp-generated impinging SWBLI, and 2) a variable angle wedge (VAW) generated impinging SWBLI. The compression ramp and the wedge both serve to generate an oblique shock wave that impinges on the boundary layer on the opposite surface of the wind tunnel. This produces an impinging SWBLI within the test section that replicates the flow found in a mixed compression inlet. A SWBLI can cause flow separation and it is desirable to efficiently prevent this to avoid the many adverse consequences that may result otherwise. The goal of this study is to investigate the ability of localized arc-filament plasma actuators (LAFPAs) to effectively control the interaction. The LAFPAs show significant ability to beneficially affect the SWBLI depending on various operating parameters such as geometry and forcing Strouhal number. This ability apparently stems from a manipulation of instabilities naturally present in the flow. For the compression ramp facility the LAFPAs were most effective when located upstream of the shock foot, forcing with a Strouhal number of 0.03, and operated in-phase. The VAW facility is currently being debugged and will be used for future detailed experiments investigating the control authority of the LAFPAs.

Committee:

Mohammad Samimy, PhD (Advisor); Michael Dunn, PhD (Committee Member); Jeffrey Bons, PhD (Committee Member)

Subjects:

Engineering; Mechanical Engineering

Keywords:

flow control; LAFPA; plasma actuator; SWBLI; supersonic inlet

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