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Rice, Thomas StuartA comprehensive investigation into the supersonic viscous flow about a slender cone at high angle of attack : experimental and theoretical results /
Doctor of Philosophy, The Ohio State University, 1980, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Viscous flow;Aerodynamics;Cone--Aerodynamics

Disotell, Kevin JamesLow-Frequency Flow Oscillations on Stalled Wings Exhibiting Cellular Separation Topology
Doctor of Philosophy, The Ohio State University, 2015, Aero/Astro Engineering
One of the most pervasive threats to aircraft controllability is wing stall, a condition associated with loss of lift due to separation of air flow from the wing surface at high angles of attack. A recognized need for improved upset recovery training in extended-envelope flight simulators is a physical understanding of the post-stall aerodynamic environment, particularly key flow phenomena which influence the vehicle trajectory. Large-scale flow structures known as stall cells, which scale with the wing chord and are spatially-periodic along the span, have been previously observed on post-stall airfoils with trailing-edge separation present. Despite extensive documentation of stall cells in the literature, the physical mechanisms behind their formation and evolution have proven to be elusive. The undertaken study has sought to characterize the inherently turbulent separated flow existing above the wing surface with cell formation present. In particular, the question of how the unsteady separated flow may interact with the wing to produce time-averaged cellular surface patterns is considered. Time-resolved, two-component particle image velocimetry measurements were acquired at the plane of symmetry of a single stall cell formed on an extruded NACA 0015 airfoil model at chord Reynolds number of 560,000 to obtain insight into the time-dependent flow structure. The evolution of flow unsteadiness was analyzed over a static angle-of-attack range covering the narrow post-stall regime in which stall cells have been observed. Spectral analysis of velocity fields acquired near the stall angle confirmed a low-frequency flow oscillation previously detected in pointwise surface measurements by Yon and Katz (1998), corresponding to a Strouhal number of 0.042 based on frontal projected chord height. Probability density functions of the streamwise velocity component were used to estimate the convective speed of this mode at approximately half the free-stream velocity, in agreement with Yon and Katz. Large-amplitude streamwise Reynolds stresses in the separated shear layer were found to be manifested by the low-frequency oscillation through inspection of the spectral energy distribution. Using the method of Proper Orthogonal Decomposition to construct reduced-order models of the acquired time sequences, the low-frequency unsteadiness appeared to be linked to an interaction between the separated and trailing-edge shear layers, in contrast to a bubble-bursting mechanism which has been observed for different stall behaviors. As the static angle of attack was increased further, the separated flow structure was seen to transition to a faster eddy motion expected for bluff-body wakes. A novel scaling study was conducted to evaluate the potential role of low-frequency unsteadiness in producing the spanwise wavelengths associated with cell formation, which was found to be in favorable agreement with scaling trends in the literature. Finally, instantaneous pressure-sensitive paint measurements were demonstrated on a DU 97-W-300 wind turbine airfoil at chord Reynolds number of 225,000 with leading-edge trip applied, in which the development of spiral node structures associated with cell formation were captured in the trailing-edge separation. The contributed work suggests that further study into the influence of large-scale unsteadiness on the three-dimensional organization of stall cells is merited.

Committee:

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

Subjects:

Aerospace Engineering; Fluid Dynamics

Keywords:

wing aerodynamics; stall cells; high-angle-of-attack airfoil aerodynamics; turbulent separated flows; three-dimensional separation; large-scale unsteadiness; vortical wakes; low-frequency flow oscillations; time-resolved particle image velocimetry

Flegel, Ashlie BrynnAerodynamic Measurements of a Variable-Speed Power-Turbine Blade Section in a Transonic Turbine Cascade
Master of Science in Mechanical Engineering, Cleveland State University, 2013, Fenn College of Engineering
The purpose of this thesis is to document the impact of incidence angle and Reynolds number variations on the 3-D flow field and midspan loss and turning of a 2-D section of a variable-speed power-turbine (VSPT) rotor blade. Aerodynamic measurements were obtained in a transonic linear cascade at NASA Glenn Research Center in Cleveland, OH. Steady-state data were obtained for ten incidence angles ranging from +15.8&00B0; to -51.0&00B0;. At each angle, data were acquired at five flow conditions with the exit Reynolds number (based on axial chord) varying over an order-of-magnitude from 2.12 &00D7; 10^5 to 2.12 &00D7; 10^6. Data were obtained at the design exit Mach number of 0.72 and at a reduced exit Mach number of 0.35 as required to achieve the lowest Reynolds number. Midspan total-pressure and exit flow angle data were acquired using a five-hole pitch/yaw probe surveyed on a plane located 7.0 percent axial-chord downstream of the blade trailing edge plane. The survey spanned three blade passages. Additionally, three-dimensional half-span flow fields were examined with additional probe survey data acquired at 26 span locations for two key incidence angles of +5.8&00B0; and -36.7&00B0;. Survey data near the endwall were acquired with a three-hole boundary-layer probe. The data were integrated to determine average exit total-pressure and flow angle as functions of incidence and flow conditions. The data set also includes blade static pressures measured on four spanwise planes and endwall static pressures.

Committee:

Mounir Ibrahim, PhD (Committee Chair); Miron Kaufman, PhD (Committee Member); Ralph Volino, PhD (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

Turbomachinery; aerodynamics; rotorcraft; cascade; power-turbine; turbines; experimental aerodynamics

Ross, Ian JonathanWind Tunnel Blockage Corrections: An Application to Vertical-Axis Wind Turbines
Master of Science (M.S.), University of Dayton, 2010, Aerospace Engineering

An investigation into wake and solid blockage effects of Vertical-Axis Wind Turbines (VAWTs) in closed test-section wind tunnel testing is described. Static wall pressures have been used to derive velocity increments along a wind tunnel test-section which in-turn are applied to provide evidence of wake interference characteristics of rotating bodies interacting within this spatially restricted domain. Vertical-axis wind turbines present a unique aerodynamic obstruction in wind tunnel testing whose blockage effects have not been extensively investigated.

The flow-field surrounding these wind turbines is asymmetric, periodic, unsteady, separated and highly turbulent. Static pressure measurements are taken along a test-section sidewall to provide a pressure signature of the test models under varying rotor tip-speed ratios (freestream conditions and model RPM’s). To provide some guidance on the scaling of the combined effects of wake and solid blockage, wake characteristics and VAWT performance produced by the same vertical-axis wind turbine concept have been tested at different physical scales in two different wind tunnels. This investigation provides evidence of the effects of large wall interactions and wake propagation caused by these models at well below generally accepted standard blockage figures.

Committee:

Aaron Altman, PhD (Committee Chair); Jewel Barlow, PhD (Committee Member); Eric Lang, PhD (Committee Member)

Subjects:

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

Keywords:

Low Speed Wind Tunnel; Wind Tunnel Blockage Corrections; Vertical-Axis Wind Turbine; Aerodynamics; Bluff-Body Aerodynamics; Savonius

Krumanaker, Matthew LeeAerodynamics and Heat Transfer for a Modern Stage and One-Half Turbine
Master of Science, The Ohio State University, 2003, Aeronautical and Astronautical Engineering
This paper describes measurements obtained for a state of the art one and one-half stage turbine with emphasis on the experimental results. As part of the experimental effort, the position of the HPT vane was clocked relative to the downstream LPT vane to determine the influence of vane clocking on the unsteady pressure loadings on the LPT vane and the HPT blade. In addition, the axial location of the HPT vane relative to the HPT blade was changed to investigate the combined influence of vane/blade spacing and clocking on the unsteady pressure loading. In recent years, other investigators have also examined clocking and vane/blade spacing (but not within the same data set) using different turbine stages and those works will be put into perspective with this work. Time-averaged and time-accurate surface pressure will be presented for several spanwise locations on the vanes and blade. Results were obtained at four different clocking positions for the HPT vane and for two different vane/blade axial spacings at three (of the four) clock positions. This thesis also describes the heat transfer results for a measurement program utilizing the same state of the art one and one-half stage transonic turbine. Both aerodynamic data (surface pressure data) and heat transfer data were obtained at the 50% span location on the HPT vane, HPT blade, and LPT vane. The heat flux data are normalized and presented as time-averaged Stanton numbers and compared to flat plate correlations. Time accurate Stanton numbers are also presented for selected locations on the HPT blade and on the blade outer air seal.

Committee:

Michael Dunn (Advisor)

Subjects:

Engineering, Aerospace

Keywords:

Gas Turbine Heat Transfer and Aerodynamics

Hammer, Patrick RichardA Discrete Vortex Method Application to Low Reynolds Number Aerodynamic Flows
Master of Science (M.S.), University of Dayton, 2011, Aerospace Engineering
Although experiments and CFD are very powerful tools in analyzing a niche of fluid dynamics problems relevant to developing Micro Aerial Vehicles (MAVs), reduced order methods have shown to be very capable in helping researchers achieve a basic understanding of flow physics with application to highly iterative design processes due to the less computationally expensive nature of the low order models. The current study used one low order method, the Discrete Vortex Method, to model the aerodynamic flow fields and forces around a thin airfoil undergoing a variety of flows, as well as parametric studies to determine the important factors that had to be adjusted to make the results more representative of the physical phenomenon being modeled. Initial investigations validated the code’s use in steady flow and low amplitude unsteady flow cases by comparing it with circulation distributions of various airfoil shapes, the Wagner function, and Theodorsen’s function. The results showed a strong dependency on bound vortex number and time step size. The code was then used to capture the flow behavior around the airfoil for various AIAA Fluid Dynamics Technical Committee Low Reynolds iv Number Working Group (FDTC-LRWG) canonical cases. Implementing the Uhlman method in the Discrete Vortex Method allowed for the calculation of the pressure at the airfoil surface and in the flow field during high angle attack maneuvers. This method proved very capable in calculating the pressures, forces, and force coefficients around the airfoil post-flow separation in the canonical cases where other methods (such as the Unsteady Bernoulli Method) fall short. The code was also tuned with respect to the results with respect to vortex size, leading edge separation strength factor, desingularization function, wake radius size factor, and in the Uhlman method itself to yield an optimal comparison with experimental and CFD results. The study found a bound vortex number of 30, a leading edge separation strength factor of 1.0, the planetary desingularization function, a wake radius size factor of 1.0, and using just the volume integral term on the RHS of the Uhlman method gave the best results for the geometry analyzed. An investigation then determined the dependency of reduced frequency on the lift and drag coefficients for the canonical cases. Finally, the code was used to model a “true perch” by implementing a curve fit function which caused the horizontal free stream velocity to decrease to zero. In this context, the forces were of more interest than the force coefficients since the coefficients experienced anomalous behavior as the free stream velocity approached zero. It was also interesting to find that the code modeled behavior very similar to shear layer instabilities in the LE and TE shear layers, caused by a rippling effect as the bound circulation changed in strength and sign as the LEV and TEV interacting with it. Recommendations were then made to apply the code to airfoils with either fixed or variable camber since camber acts as a high lift device and could prove very beneficial in the design and development of MAVs

Committee:

Aaron Altman (Committee Chair); Frank Eastep (Committee Member); Greg Reich (Committee Member)

Subjects:

Aerospace Engineering; Fluid Dynamics; Mechanical Engineering

Keywords:

Unsteady Aerodynamics; High Angle of Attack; Discrete Vortex Method; Vortex Particle Method; Perching

Rizzetta, Donald PasqualeAsymptotic solution for two-dimensional viscous supersonic and hypersonic flows past compression and expansion corners /
Doctor of Philosophy, The Ohio State University, 1976, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Supersonic compressors;Asymptotes;Aerodynamics

Betts, Robert WilliamThe effects of hypersonic viscous interaction on static stability of slender bodies in simulated non-equilibrium flows /
Doctor of Philosophy, The Ohio State University, 1973, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Aerodynamics;Pressure

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

Mustafa, MansoorInvestigation into Offset Streams for Jet Noise Reduction
Master of Science, The Ohio State University, 2015, Aero/Astro Engineering
This effort investigates the near field behavior of two ideally-expanded subsonic dual-stream jets. One case implements a traditional symmetric, concentric dual-stream nozzle configuration while the other imposes an asymmetric, eccentric layout to model the behavior of an offset stream. The essence of an offset stream is to force an uneven azimuthal distribution of the secondary coflow and create an outside stream that varies in thickness. Past studies have shown a benefit in acoustic propagation in the direction of the thickest coflow and the present work further analyzes this phenomenon. A LES (Large Eddy Simulation) approach is implemented to run the simulations for both cases and a number of qualitative and quantitative analyses tools are used for post-processing. A reduction in the noise levels for the lower, thicker side of the eccentric nozzle is observed in comparison to the baseline concentric case. Examination of the mean flow behavior shows a shorter, thinner primary potential core for the offset case and a faster axial velocity decay rate. The asymmetric distribution of the coflow causes varying velocity profiles in the radial direction for the top and bottom regions and consequently produces unique flow features on either side. Lower levels of shear stress and slower decay rates lead to less turbulence production on the lower side of the eccentric nozzle. An investigation into the flow structures reveals lower vorticity and weaker convective structures on the bottom which influences propagation in that direction. Two-point correlation analysis reveals the presence of smaller turbulence scales in the lower, thicker portion of the eccentric case. This is further confirmed by an Empirical Mode Decomposition (EMD) study that shows lower frequency ranges dominate the concentric near field in comparison to the eccentric. The combination of these unique features demonstrate the principles behind the acoustic benefit of implementing offset stream flows in dual-stream nozzle configurations.

Committee:

Datta Gaitonde (Advisor); Mei Zhuang (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

offset; noise reduction; jets; aerospace; aerodynamics; LES; computational simulation; jet noise; eccentric; concentric; dual-stream; nozzle

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

Allenstein, Jacob T.An Investigation in Gold-Plating Scaled Turbofan Engine Simulators through Means of Aerodynamic and Load Cell Thrust Measurements with Comparisons to Full-Scale Engine Results
Master of Science, The Ohio State University, 2013, Aero/Astro Engineering
Model testing offers investigators a useful tool that can provide good insight into the aerodynamics behind full-scale engines and test facilities. Understanding the aerodynamics of a full-scale engine can help investigators update old test facilities and help design new facilities and engines. Gold-plating an engine is a process that the industry uses to compare the performance of an engine from the same family or class of engines in various facilities. The gold-plated engine can be used to determine a correlation factor of a testing facility or to determine the similarities in performance between a new or old engine to the gold-plated engine. The use of a correlation or correction factor can be used to correct a deviation in a measurement made in a test facility to bring the engine’s performance back to readings performed in a “free-air” environment. Model testing consists of the use of a simulator or scaled version of an engine that generates performance close to full-scale operation conditions. The simulator used in this study was constructed to not have any moving parts but rather be driven by a high pressure air system, providing a safe alternative to testing the full-scale engine. The objective of the study was to understand how to perform the gold-plate process to an engine simulator through the means of the aerodynamic and load cell thrust readings. The study examined the thrust that can be calculated from the exhaust mass flow of the engine simulator, or aerodynamic thrust, and compared the results to the load cell thrust from the load cell. This investigation will provide the framework to build and expand the knowledge of how an engine simulator works and how it can be used to understand the aerodynamics of an engine. The analysis was conducted at The Ohio State University’s Aeronautical and Astronautical Research Laboratories’ JETS facility. Two simulators went through the gold-plating process: the GE HF-120 and the GE90-B4 turbofan simulators. Each simulator used a high pressure, ejector driven system which does not require any moving parts. Both simulators were set up in a free-air environment and were tested at a variety of different pressures. The first simulator investigated was the GE HF-120 turbofan engine. The HF-120 used two types of inlet cowls: a flight cowl and a bellmouth testing cowl. The second simulator investigated was the GE90-B4 turbofan engine, which has two ejector stages: the fan and core. These tests have contributed to an improved understanding of the thrust measurements obtained from these engine simulators. These results will create the foundation of operating the simulator to verify data produced in a model scale test of a full-scale facility. The setup provided a reliable test environment for the engine simulator as seen through multiple, repeatable test runs. This knowledge and these techniques could help predict full-scale test results with use of scaled engine simulators with a higher degree of accuracy, providing investigations an inexpensive and safe environment for determining the full-scale facility’s performance.

Committee:

Richard Freuler (Advisor); Clifford Whitfield (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

model testing; engine simulator; aerodynamics; simulator thrust; gold-plating

Shyam, Vikram3-D Unsteady Simulation of a Modern High Pressure Turbine Stage: Analysis of Heat Transfer and Flow
Doctor of Philosophy, The Ohio State University, 2009, Aeronautical and Astronautical Engineering
This is the first 3-D unsteady RANS simulation of a highly loaded transonicturbine stage and results are compared to steady calculations and experiments. A low Reynolds number k-ε turbulence model is employed to provide closure for the RANS system. Phase-lag is used in the tangential direction to account for stator-rotor interaction. Due to the highly loaded characteristics of the stage, inviscid effects dominate the flowfield downstream of the rotor leading edge minimizing the effect of segregation to the leading edge region of the rotor blade. Unsteadiness was observed at the tip surface that results in intermittent 'hot spots'. It is demonstrated that unsteadiness in the tip gap is governed by both inviscid and viscous effects due to shock-boundary layer interaction and is not heavily dependent on pressure ratio across the tip gap. This is contrary to published observations that have primarily dealt with subsonic tip flows. The high relative Mach numbers in the tip gap lead to a choking of the leakage flow that translates to a relative attenuation of losses at higher loading. The efficacy of a new tip geometry is discussed to minimize heat flux at the tip while maintaining choked conditions. Simulated heat flux and pressure on the blade and hub agree favorably with experiment and literature. The time-averaged simulation provides a more conservative estimate of heat flux than the steady simulation. The shock structure formed due to stator-rotor interaction is analyzed. A preprocessor has also been developed as a conduit between the unstructured multi-block grid generation software GridPro and the CFD code TURBO.

Committee:

Jen-Ping Chen, PhD (Advisor); Ali Ameri, PhD (Committee Member); Meyer Benzakein, PhD (Committee Member); Jeffrey Bons, PhD (Committee Member); Henry Busby, PhD (Committee Member)

Subjects:

Engineering; Physics

Keywords:

rotor-stator interaction; Transonic turbine stage; unsteady heat transfer; Tip leakage; unsteady aerodynamics; computational fluid dynamics

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

Snyder, Troy A.A Coupled Wake-Integral/Vorticity Confinement Technique for the Prediction of Drag Force
Master of Science in Engineering, University of Akron, 2012, Mechanical Engineering

This work couples the two enabling technologies of wake-integral drag prediction and Vorticity Confinement (VC) for the improved prediction of drag from an Euler CFD simulation. Induced drag computations of a thin wing are shown to be more accurate than the more common method of surface pressure integration when compared to Prandtl lifting-line theory. Furthermore, the Vorticity Confinement method is shown to improve trailing vortex preservation and counteract the shift from induced to entropy drag as the distance with which the vortex convects downstream of the wing increases.

The desired application of this work is drag prediction, most notably induced drag, for aerodynamic design. High-fidelity Euler CFD simulations are desirable as they provide the most complete inviscid flow field solution. However, accurate induced drag prediction via the surface integration of pressure is generally intractable barring a sufficiently refined surface grid and resultant increase in computational load. Furthermore, the alternative wake-integral technique for drag prediction suffers from numerical dissipation. VC is shown to control the numerical dissipation with very modest computational overhead.

VC is implemented in both a two-dimensional finite-volume Euler code written by the author as well as the commercial cfd code ANSYS FLUENT. The two-dimensional research code is used to test specific formulations of the VC body force terms and illustrate the computational efficiency of the method compared to a 'brute force' reduction in spatial step size. For a three-dimensional wing simulation, ANSYS FLUENT is employed with the VC body force terms added to the solver with user-defined functions (UDFs).

Committee:

Alex Povitsky, Dr. (Advisor); Abilash Chandy, Dr. (Committee Member); Scott Sawyer, Dr. (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

drag; vorticity confinement; aerodynamics; wake integration

Elsharnoby, Mohamed A.On the stability of the supersonic boundary layer /
Doctor of Philosophy, The Ohio State University, 1987, Graduate School

Committee:

Not Provided (Other)

Subjects:

Engineering

Keywords:

Aerodynamics;Boundary layer

Kao, Yi-HuanExperimental Investigation of Aerodynamics and Combustion Properties of a Multiple-Swirler Array
PhD, University of Cincinnati, 2014, Engineering and Applied Science: Aerospace Engineering
An annular combustor is one of the popular configurations of a modern gas turbine combustor. Since the swirlers are arranged as side-by-side in an annular combustor, the swirling flow interaction should be considered for the design of an annular gas turbine combustor. The focus of this dissertation is to investigate the aerodynamics and the combustion of a multiple-swirler array which features the swirling flow interaction. A coaxial counter-rotating radial-radial swirler was used in this work. The effects of confinement and dome recession on the flow field of a single swirler were conducted for understanding the aerodynamic characteristic of this swirler. The flow pattern generated by single swirler, 3-swirler array, and 5-swirler array were evaluated. As a result, the 5-swirler array was utilized in the remaining of this work. The effects of inter-swirler spacing, alignment of swirler, end wall distance, and the presence of confinement on the flow field generated by a 5-swirler array were investigated. A benchmark of aerodynamics performance was established. A phenomenological description was proposed to explain the periodically non-uniform flow pattern of a 5-swirler array. The non-reacting spray distribution measurements were following for understanding the effect of swirling flow interaction on the spray distribution issued out by a 5-swirler array. The spray distribution from a single swirler/ fuel nozzle was measured and treated as a reference. The spray distribution from a 5-swriler array was periodically non-uniform and somehow similar to what observed in the aerodynamic result. The inter-swirler spacing altered not only the topology of aerodynamics but also the flame shape of a 5-swirler array. As a result, the distribution of flame shape strongly depends on the inter-swirler spacing.

Committee:

San-Mou Jeng, Ph.D. (Committee Chair); Shanwu Wang, Ph.D. (Committee Member); Awatef Hamed, Ph.D. (Committee Member); Milind Jog, Ph.D. (Committee Member); Jongguen Lee, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

swirling flow interaction;annular gas turbine combustor;swirler;aerodynamics;combustion

Jensen, Christopher DouglasGlobal Pressure and Temperature Surface Measurements on a NACA 0012 Airfoil in Oscillatory Compressible Flow at Low Reduced Frequencies
Master of Science, The Ohio State University, 2012, Aero/Astro Engineering
A co-axial contra-rotating helicopter in forward flight has stall mechanisms that are intrinsically different from those on a traditional helicopter. Traditional helicopters need a cyclic pitch mechanism to balance the rotor lift about each rotation, which leads to dynamic stall from the rapid oscillations in pitch. Co-axial contra-rotating helicopters, which have a fixed pitch about each cycle, encounter a sinusoidal oscillation in Mach number with the mean velocity seen as the rotational velocity while the half amplitude is equal to the forward flight speed. The fluid dynamic mechanism limiting the forward flight speed is entirely different from that of traditional dynamic stall studies. This work sets out to design, create, and study the application of an oscillatory compressible flow field on a NACA 0012 airfoil in order to experimentally model this flow situation. A few different pressure-sensitive paints and imaging techniques were developed for investigation of this oscillatory effect. Ultimately a fast-acting bi-luminophore pressure- and temperature-sensitive paint was chosen for the investigation which uses a polymer-ceramic basecoat and a mixture of luminescent elements. These measurements were made using a two-camera, single-shot, intensity-based pressure-sensitive paint technique. Temperature-corrected pressure measurements were made and accounted for the intrinsic temperature sensitivity of pressure-sensitive paint. This dual-luminophore technique allows for accurate unsteady pressure measurements in a non-uniform and varying temperature environment; however, due to inadequate unsteady surface pressure tap measurements the pressure results were limited to the steady runs. This work involved the design and creation of a modification to Ohio State’s 6” x 22” Transonic Wind Tunnel to enable oscillations of the freestream Mach number. The current configuration produces Mach number oscillations between 0.44 and 0.64 for a Reynolds number range of 17 – 43 million per meter at frequencies up to 21 Hz. Unsteady shock location measurements were made at angles of attack of 9, 10 and 11 degrees and frequencies of 2.1, 9.5, 15.25 and 21 Hz on the NACA 0012 airfoil. Detailed measurements of the shock movement were made with these advanced measurement techniques in order to investigate unsteady effects of the oscillatory freestream flow. Unsteady effects were pronounced for a reduced frequency of 0.037 which is below the typical quasi steady to unsteady threshold of 0.05. Coefficient of pressure measurements for the steady runs were validated with historical data. It was found that the coefficient of lift measurements were in very good agreement, while the coefficient moment had significant errors. Furthermore, the PSP measurements were compared with particle image velocimetry data of other researchers in order to form a comparison of the on- and off-body fluid dynamics at a frequency of 9.5 Hz and angles of attack of 9 and 10 degrees. The single-shot pressure-sensitive paint technique was able to measure buffeting, which was found to be highly three dimensional over the span of the airfoil. Similar shock location unsteadiness due to buffeting was also measured in the forced oscillation cases at lower Mach numbers than steady runs, causing aperiodic behavior at certain azimuth locations with higher Mach numbers.

Committee:

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

Subjects:

Aerospace Engineering

Keywords:

PSP; TSP; pressure sensitive paint; aerodynamics; Mach oscillation; freestream oscillation; NACA 0012

Smith, Christopher StephenExperimental Validation of a Hot Gas Turbine Particle Deposition Facility
Master of Science, The Ohio State University, 2010, Aeronautical and Astronautical Engineering
A new turbine research facility at The Ohio State University Aeronautical and Astronautical Research Lab has been constructed. The purpose of this facility is to re-create deposits on the surface of actual aero-engine Nozzle Guide Vane (NGV) hardware in an environment similar to what the hardware was designed for. This new facility is called the Turbine Reacting Flow Rig (TuRFR). The TuRFR provides air at temperatures up to 1200 °C and at inlet Mach numbers comparable to those found in an actual turbine (~0.1). Several validation studies have been undertaken which prove the capabilities of the TuRFR. These studies show that the temperature entering the NGV cascade is uniform, and they demonstrate the capability to provide film cooling air to the NGV cascade at flow rates and density ratios comparable to the NGV design. Deposition patterns have also been created on the surface of actual NGV hardware. Deposition was created at different flow temperatures, and it was found that deposition levels decrease with decreasing gas temperature. Also, film cooling levels were varied from 0% film cooling to 4% film cooling. It was found that with increased rates of film cooling deposition decreased. With the TuRFR capabilities demonstrated, research on the effects of deposition on the aerodynamic performance of the NGV hardware was conducted. Integrated non-dimensional total pressure loss values were calculated in an exit Rec range of 0.2x106 to 1.7x106 for a deposit roughened NGV cascade and a smooth cascade. The data suggests that deposition causes increased losses across the NGV cascade and possibly earlier transition. The data also suggests a possible region of separated flow in the NGV cascade which disappears at higher exit Reynolds numbers. These results are similar to those found in the literature.

Committee:

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

Subjects:

Mechanical Engineering

Keywords:

turbine; deposition; gas turbine; turbine deposition; turbine aerodynamics

FU, YONGQIANGAerodynamics and Combustion of Axial Swirlers
PhD, University of Cincinnati, 2008, Engineering : Aerospace Engineering
A multipoint lean direct injection (LDI) concept was introduced recently in non-premixed combustion to obtain both low NOx emissions and good combustion stability. In this concept, a key feature is the injection of finely atomized fuel into the high-swirling airflow at the combustor dome that provides a homogenous, lean fuel-air mixture. In order to achieve the fine atomization and mixing of the fuel and air quickly and uniformly, a good swirler design should be studied. The focus of this dissertation is to investigate the aerodynamics and combustion of the swirling flow field in a multipoint lean direct injector combustor. A helical axial-vaned swirler with a short internal convergent-divergent venturi was used. Swirlers with various vane angles and fuel nozzle insertion lengths have been designed. Three non-dimensional parameter effects on non-reacting, swirling flow field were studied: swirler number, confinement ratio and Reynolds number. Spray and combustion characteristics on the single swirler were studied to understand the mechanism of fuel-air mixing in this special configuration. Multi-swirler interactions were studied by measuring the confined flow field of a multipoint swirler array with different configurations. Two different swirler arrangements were investigated experimentally, which include a co-swirling array and a counter-swirling array. In order to increase the range of stability of multipoint LDI combustors, an improved design were also conducted. The results show that the degree of swirl and the level of confinement have a clear impact on the mean and turbulent flow fields. The swirling flow fields may also change significantly with the addition of a variety of simulated fuel nozzle insertion lengths. The swirler with short insertion has the stronger swirling flow as compared with the long insertion swirler. Reynolds numbers, with range of current study, will not alter mean and turbulent properties of generated flows. The reaction of the spray dramatically changes the gas phase velocity distribution, while the convergent-divergent nozzle strongly affects the spray velocity profiles. The multipoint flow field has a very complicated structure, especially for the flow structure near the swirler exit, where very strong interactions exit among the adjacent swirlers. Multipoint swirler arrays with the recessed center swirler will alter flow structure significantly. There is a short strong central recirculation zone in both co-swirler and counter-swirler recessed arrays, which may increase the operability range of the multipoint swirl-venturi LDI combustor.

Committee:

Dr. San-Mou Jeng (Advisor)

Subjects:

Engineering, Aerospace

Keywords:

Swirling Flow; Axial swirler; Multipoint; LDI; Aerodynamics; NOx

Vytla, Veera Venkata Sunil KumarMultidisciplinary Optimization Framework for High Speed Train using Robust Hybrid GA-PSO Algorithm
Doctor of Philosophy (PhD), Wright State University, 2011, Engineering PhD

High speed trains are the most efficient means of public transportation. However the speed of the train needs to be increased (> 350 km/hr) to cover large distances in a short time to make it accessible to large population. With the increase in speed, number of issues related to efficiency, safety and comfort like the aerodynamic drag, structural strength, as well as the noise levels inside and outside of the train etc. need to be considered in the design of the high speed trains. Hence making it a multi disciplinary design problem. There are a large number of parameters from different disciplines that need to be tuned to identify the best design. The parameters need to be optimized to identify the best design configuration that meets the design requirements. This requires the use of robust and efficient optimization algorithms. Evolutionary algorithms have been used extensively in the engineering design optimization problems, but they suffer from a drawback of lack of robustness. One of the objectives of this research is to address the robustness issue of currently available optimization algorithms. A hybrid GA-PSO algorithm combining the benefits of both the original algorithms GA and PSO is proposed in this research. The hybrid GA-PSO algorithm was observed to be robust and accurate based upon the tests. The computer simulations required to complete the optimization of this problem are expensive both in terms of computational resources as well as time. To minimize the computational effort an adaptive surrogate model based on kriging was used during optimization. The accuracy of the surrogate model was checked during the optimization process using the parameter called expected improvement value (EIV) and is updated whenever found to be inadequate. The optimization algorithm combined with the adaptive surrogate modeling technique is tested on Branin function and is found to be robust and efficient.

The optimization of a high speed train is an MDO problem. The MDO problem can be simplified significantly if the problem can be decoupled thereby reducing the complexity of the problem. The objectives considered while finding the optimum design of the high speed train are aerodynamic drag for efficiency, structural strength for safety, and generated noise for human comfort. The objective for comfort, noise levels both inside and outside the train can be used as a decoupling objective between the aerodynamic and structural optimization. The optimization is performed sequentially. First step involves performing the shape optimization which identifies the optimum aerodynamic shape and structural optimization is performed on the optimum shape to identify the structure strong enough to withstand the aerodynamic loads with the least mass. A multi objective shape optimization is performed to identify the aerodynamic shape which induces least drag and generates least aerodynamic noise. Aerodynamic shape optimization requires the construction of new CAD models and some preprocessing to generate the computational mesh before the shape is analyzed. This step becomes complicated and is a hurdle when trying to automate the optimization process. Shape optimization is performed by using the shape control parameters on computational mesh and deforming the mesh along with the surface to obtain the optimum shape using commercial mesh deformation software, Sculptor. This approach was tested on a 2-D model before using it on a 3-D train model. Shape optimization is performed using a commercial CFD solver SC/Tetra. Since shape optimization is performed using mesh deformation software, there is an additional step of preparing the structure after the shape optimization is completed. Time averaged pressure loads acting on the structure are simulated using the optimum shape of the train and are mapped onto the structure. Structural optimization is performed to identify the structure that supports the optimum shape, with least mass and least noise levels inside the train. This optimization is performed using structural solver Abaqus. The suggested sequential MDO approach for high speed train reduces the optimization time required to find the optimum shape and structure of the train.

Committee:

George Huang, PhD (Committee Chair); Ravi Penmetsa, PhD (Committee Co-Chair); Haibo Dong, PhD (Committee Member); Jonathan Black, PhD (Committee Member); Norihiko Watanabe, PhD (Committee Member)

Subjects:

Engineering; Mechanical Engineering

Keywords:

aerodynamics; acoustics; optimization; GA; PSO; Kriging

Lego, Zachary MichaelAnalysis of High Angle of Attack Maneuvers to Enhance Understanding of the Aerodynamics of Perching
Master of Science (M.S.), University of Dayton, 2012, Aerospace Engineering
Due to their vastly unsteady aerodynamics, mimicking the flight maneuvers of natural flyers requires a deep understanding of the unsteady flow physics while also demanding the ability to expand the vehicle's flight envelope beyond normal conditions. Experimental and computational investigations are performed to study and improve the flight performance of a perching maneuver. Flowfield images and force history data are acquired to investigate the performance of highly unsteady motions experienced during a perching maneuver. The perching motions studied are divided into two first order approximations, a rapid pitching motion and an impulsively started hold in order to follow the North Atlantic Treaty Organization Research & Technology Organization Applied Vehicle Technology 202 (NATO RTO AVT 202) guidelines. Both types of motions are investigated through experimental testing and using a computational 2D Discrete Vortex Method (DVM). Using the 2D DVM allows the ability to determine whether a low order, two dimensional code can accurately predict an actual 3D flowfield generated by highly unsteady motions. The pitching motions are studied experimentally using 3D plates in the University of Dayton Low Speed Wind Tunnel (UD-LSWT). Results show that despite some 3D instabilities in the experiments, which are not produced by the DVM, overall the DVM matches the experiments well. Results are completed for varying pitch rates and the DVM matches the experiment more closely for the faster, more inertially dominated pitch rates. Upper surface pressure data is also recorded in both the experiment and DVM. The results suggest that differing pressure profiles could have a substantial influence on aerodynamic forces even at high angles of attack. The impulsively started hold motions completed experimentally with 3D plates in the AFRL Horizontal Free Surface Water Tunnel (AFRL HFWT) again produce some three-dimensional instabilities not modeled in the 2D DVM. When comparing the impulsively started results, the DVM did not match the startup of the motion in the force histories due to the use of different startup transients and the constants in the smoothing function. The DVM does well when comparing the flowfield. In an attempt to expand the available performance envelope during these types of motions variable camber is studied for both types of maneuvers using the 2D DVM and in both cases the initial results show that variable camber has the potential to positively influence the boundaries of the flight envelope during a perching maneuver.

Committee:

Aaron Altman, PHD (Committee Chair); Markus Rumpfkeil, PHD (Committee Member); Gregory Reich, PHD (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Aerodynamics; Perching; Variable Camber; High Reduced Frequency Pitching Motion

Nandamudi, SrihimajaAerodynamics of Vocal Vibrato
Doctor of Philosophy (Ph.D.), Bowling Green State University, 2017, Communication Disorders

Airflow vibrato is the fluctuation in average airflow while singing with vibrato. Understanding airflow vibrato relates to a deeper understanding of its importance to physiological, pedagogical, and clinical aspects. Two studies were performed to examine airflow vibrato. The subjects for Study 1 were four professional Western classical singers, and for Study 2 four highly trained amateur singers. Aerodynamic and acoustic measures were compared among vibrato, bleating (a primarily adductory gesture), and external epigastric pumping (EEP, a primarily subglottal pressure manipulation). Utterances included speaking (phonation and whisper) and singing (constant /a/ vowel, different pitches and loudness levels).

Study 1 demonstrated how airflow vibrato compares with fundamental frequency (F0) and intensity vibrato. The correlation between rates of airflow and F0 vibrato was moderately strong. Mean airflow vibrato extents were larger for the female singers than for the male singers, and increased with pitch increase for all four singers. For the males, average airflow extent was 30 and 75 cm3/s for their lower and higher pitch, respectively, and for the females, 47 cm3/s and 94 cm3/s for their lower and higher pitch, respectively.

Study 2 was undertaken to better understand sources of airflow vibrato. Airflow modulations were produced during singing with vibrato and also while bleating and with external epigastric pumping. Bleating had the fastest alteration rate (9.5-12 Hz), whereas the other types had similar rates (vibrato: 4.8-6.0 Hz; EEP: 6.0–7.5 Hz). During phonation (combining all conditions), bleating had the largest airflow modulation extents (on average 144 cm3/s, compared to 30 cm3/s for vibrato and 46 cm3/s for EEP).

Overall results suggest that airflow vibrato typically leads F0 vibrato, and often has a more complex waveshape than F0 vibrato. Hypotheses generated from the study include: (1) A primarily subglottal pressure driven vibrato may provide relatively consistent but wide extents for both F0 and airflow vibrato. (2) A primarily glottal adduction driven vibrato may provide relatively low and inconsistent F0 vibrato extent, and high and inconsistent airflow extent. (3) A primarily CT driven vibrato may result in moderate to large F0 vibrato extent, and low airflow vibrato extent, with variable consistency.

Committee:

Ronald Scherer, Ph.D. (Advisor); Alexander Goberman, Ph.D. (Committee Member); Jason Whitfield, Ph.D. (Committee Member); Mingsheng Li, Ph.D. (Other)

Subjects:

Health Sciences; Music; Speech Therapy

Keywords:

Airflow vibrato; Aerodynamics; Singing voice; Vibrato; Bleat; External epigastric pumping; Electroglottography; Intensity vibrato; Fundamental frequency vibrato; Western classical singing

Bear, Philip StevenOn the Experimental Evaluation of Loss Production and Reduction in a Highly Loaded Low Pressure Turbine Cascade
Master of Science in Mechanical Engineering (MSME), Wright State University, 2016, Mechanical Engineering
Improvements in turbine design methods have resulted in the development of blade profiles with both high lift and good Reynolds lapse characteristics. An increase in aerodynamic loading of blades in the low pressure turbine section of aircraft gas turbine engines has the potential to reduce engine weight or increase power extraction. Increased blade loading means larger pressure gradients and increased secondary losses near the endwall. Prior work has emphasized the importance of reducing these losses if highly loaded blades are to be utilized. The present study analyzes the secondary flow field of the front-loaded low-pressure turbine blade designated L2F with and without blade profile contouring at the junction of the blade and endwall. The current work explores the loss production mechanisms inside the low pressure turbine cascade. Stereoscopic particle image velocimetry data, total pressure loss data and oil flow visualization are used to describe the secondary flow field. The flow is analyzed in terms of total pressure loss, vorticity, Q-Criterion, Reynolds’ stresses, turbulence intensity and turbulence production. The flow description is then expanded upon using an Implicit Large Eddy Simulation of the flow field. The RANS momentum equations contain terms with static pressure derivatives. With some manipulation these equations can be rearranged to form an equation for the change in total pressure along a streamline as a function of velocity only. After simplifying for the flow field in question the equation can be interpreted as the total pressure transport along a streamline. A comparison of the total pressure transport calculated from the velocity components and the total pressure loss is presented and discussed. Peak values of total pressure transport overlap peak values of total pressure loss through and downstream of the passage suggesting that total pressure transport is a useful tool for localizing and predicting loss origins and loss development using velocity data which can be obtained non-intrusively.

Committee:

Mitch Wolff, Ph.D. (Advisor); Rolf Sondergaard, Ph.D. (Committee Member); Rory Roberts, Ph.D. (Committee Member)

Subjects:

Aerospace Engineering; Engineering

Keywords:

turbines; PIV; SPIV; particle image velocimetry; low pressure turbines; high lift turbines; total pressure loss; experimental measurements; aerodynamics; total pressure transport; turbulent flow; reynolds stress; turbulence production; deformation work

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