Doctor of Philosophy, The Ohio State University, 2025, Aerospace Engineering

Bladed disks play a vital role in the functionality of all turbomachinery systems. Their design is pivotal for optimizing performance in applications that include aerospace, power generation, and industrial processes. Bladed disks are subject to high vibrational amplitudes due to upstream flow obstructions, and thus are a large focus of active research. Two key factors that influence vibrational performance in bladed disks and are essential for accurate modeling are damping and mistuning. Damping is important to accurately model because it directly influences the amplitude of oscillations by dissipating vibrational energy. By reducing the amount of energy in the system over time, damping helps control excessive vibrations, which can lead to structural fatigue, instability, or failure. For this reason, turbomachinery designers are continually trying to increase damping with technologies such as under-platform dampers, ring dampers, and coatings. Mistuning is blade-to-blade variations that break the cyclic symmetry of the bladed disk and typically occurs due to damage, manufacturing variations, or wear. Mistuning leads to mode localization that results in higher vibrational amplitudes in some sectors. Some examples of mistuning that have been previously studied are stiffness, damping, geometric, and interface mistuning. Due to the random nature of the mistuning variations, statistical analyses are performed to understand the sensitivity to mistuning. For industrial models, this is difficult, unless the model is reduced in size. To this end, reduced order models (ROMs) have been developed that project mistuning to a reduced space to save on computational expenses. This dissertation focuses on several major objectives. The first objective is to improve on previously developed ROM methods that accurately capture damping and mistuning. This objective is captured twofold: (1) a parametric ROM method was extended to capture the structural dynamics of a bladed disk with (open full item for complete abstract)

Committee: Kiran D'Souza (Advisor); Ahmet Kahraman (Committee Chair); Jack McNamara (Committee Member); Randall Mathison (Committee Member) Subjects: Aerospace Engineering; Experiments; Mechanical Engineering; Mechanics

PhD, University of Cincinnati, 2024, Engineering and Applied Science: Aerospace Engineering

The move towards renewable energy has highlighted the need for large-scale, environmentally friendly energy storage solutions. Among these, the Supercritical Carbon Dioxide (sCO2) power cycle is emerging as a promising technology for advanced energy conversion. The effectiveness of such systems depends heavily on the compressor's performance. Using optimization-based methods, a multistage axial compressor has been designed, and its first stage has been built and tested experimentally. Through a series of detailed design iterations and optimization strategies, 3D CFD analyses, the compressor's geometric and operational parameters were fine-tuned to address the unique challenges posed by operation using CO2. Key findings highlight the successful implementation of design optimization that significantly reduces aerodynamic losses and improves the overall efficiency of the compressor stages. The optimized compressor demonstrates robust performance across a range of operating conditions, particularly focusing on improved stall margin, which emphasizes the potential of sCO2 technology in contributing to efficient and sustainable energy systems. Further detailed studies using CFD to analyze cavity effects in shrouded configurations, tip clearance effects, real gas effects, and Reynolds number effects were performed. Experimental validations, conducted at the University of Notre Dame Turbomachinery Laboratory, confirm the CFD predictions and showcase the practical viability of the compressor design and the approach used. This work not only advances the state-of-the-art in turbomachinery design for supercritical fluids but also lays a foundation for future research into the integration of sCO2 and real gas based compressors in renewable energy systems and industrial applications. The insights gained from this study underscore the critical importance of tailored design and optimization strategies in overcoming the thermophysical challenges associated (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Chair); Daniel Cuppoletti Ph.D. (Committee Member); Kelly Cohen Ph.D. (Committee Member); Jeong-Seek Kang Ph.D. (Committee Member); Shaaban Abdallah Ph.D. (Committee Member) Subjects: Aerospace Engineering

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

NASA Glenn Research Center has laid out goals to meet the needs of modern day commercial air travel. The goals laid out by NASA aim to reduce the fuel consumption while improving upon the performance parameters of future aircraft. One approach that has been researched to meet these standards is the implementation of a boundary layer ingesting (BLI) inlet. It has been shown that BLI inlet designs can achieve much of the targets set out by NASA. However, BLI inlets can harm the performance of the engine due to the large boundary layer growth which develops at the bottom surface of the inlet. Later studies have shown that the placement of a fan at the aerodynamic interface plane can attenuate the losses associated with the boundary layer growth. However, the severe changes in total pressure and high incidence angles have shown to hinder the performance of the fan and provided risk for the fan to fail due to stall conditions. The goal of this study is to optimize the design of the fan to survive the harsh conditions of the BLI inlet. The optimization was broken up into three primary phases. Phase 1 utilized a 2D optimization approach which sought to independently optimize cross-sections of a single passage under tangentially averaged inlet flow profile, and then restacked the optimized cross-sections to form a new rotor geometry. Phase 2 built upon the optimized designs of the first phase and optimized a single passage of the rotor modeled with a tangentially averaged inlet flow profile using a high end computational fluid dynamics tool called TURBO. A toolbox, labeled the Blade Altering Toolbox, was created for the second phase in order to manipulate the shape and design of the 3D rotor. Phases 1 and 2 were both used to produce a narrowed down design in order to save computational costs. Models of the Phase 2 optimized geometries were analyzed as both a single passage model using a tangentially averaged inlet profile which contained information of the distortion field, (open full item for complete abstract)

Committee: Datta Gaitonde (Committee Member); Herman Shen (Committee Member); Randall Mathison (Committee Member); Jen-Ping Chen (Advisor) Subjects: Aerospace Engineering; Design

Doctor of Philosophy (Ph.D.), University of Dayton, 2023, Aerospace Engineering

Turbomachinery components, such as the low pressure turbine, are highly complex rotating machines, therefore, conducting fundamental fluid mechanics studies in them is exceedingly difficult. For this reason, testing is generally completed in facilities such as linear cascades, like the one present in the Low Speed Wind Tunnel Facility at AFRL, which typically utilize a low freestream turbulence intensity, when in reality, the freestream turbulence intensity in a full, rotating low pressure turbine is likely much higher. Slightly elevating the freestream turbulence intensity (e.g., 3%) typically improves the Reynolds-lapse characteristics of a blade profile by affecting the transition process, reducing the detrimental effects of laminar boundary layer separation, and shifting the knee in the loss curve. Front loaded blades are more resistant to separation, however, they can experience high losses in the endwall region due to the complex vortical structures present. Therefore, a better understanding whether high levels of freestream turbulence intensity will increase the overall losses generated in the passage is important. An intial study with a jet based active grid was completed on the L2F blade. Based of the insight gained from that study, a new mechanical agitator based active grid was implemented into a linear cascade of L3FHW-LS blades in order to more effectively study how elevated FSTI impacts the endwall flow behavior and loss production. Coefficient of pressure measurements, three planes of SPIV, two additional planes of flow visualization, and three planes of total pressure loss measurements were collected. Impacts of incoming turbulence on the endwall losses as well as the endwall flow structures were assessed.

Committee: Markus Rumpfkeil (Advisor); Christopher Marks (Committee Member); Sidaard Gunasekaran (Committee Member); John Clark (Committee Member) Subjects: Aerospace Engineering

MS, University of Cincinnati, 2022, Engineering and Applied Science: Aerospace Engineering

In a world that is constantly changing and adapting, so is aviation. Within aviation, one of the most complex tasks has been the design of the propulsion device (the propulsors). The work of this thesis aimed to optimize and design an electric fan, an area of which there is little published information. For this thesis, a 0.3556 m (14 in) diameter propulsor was designed for standard day, sea level conditions. The design and effects of the inlet and outlet were not included in this study, and a constant hub was used to account for the electric motor diameter. First, the flow and design characteristics were used to develop the initial flow path with T-C DES, a mean-line compressor flow path creator. These flow paths were then used by T-AXI, an axisymmetric flow solver that uses loss models to account for losses. T-AXI created the flow condition files needed in the optimization. T-Blade3, an open-source parametric 3D blade builder, was used to create the initial blades. These blades were improved using both a 2D and 3D Optimization approach. Both processes solved multi-objective weighted functions, where the design was balanced between two operating conditions: one at the design point (DP) and the other at high incidence or low mass flow off-design point (offDP). This approach built in stall margin into the process. These processes were driven by Python scripts, optimized with OpenMDAO, an opensource tool developed by NASA; the processes modified control parameters in T-Blade3 input to create blade geometry that was used in the CFD simulations. Mises was used for the 2D optimization; Mises is a quasi-3D flow solver developed at MIT. Cadence's (Numeca) FINE/Turbo was used for the full 3D optimization. The 2D optimization minimized the loss coefficient across airfoils from five sections of each blade row. One inlet flow angle was set to the design condition, and the off design had 7 degrees of incidence. This resulted in a midspan loss coefficient for the two inlet f (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Member); Prashant Khare Ph.D. (Committee Member); Daniel Cuppoletti Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2021, Engineering and Applied Science: Aerospace Engineering

A Jetcat P400 turbojet was used as a baseline benchmark to design a compressor stage which is capable of increasing mass flow by 50% without growing in size, in an effort to increase thrust to weight ratio. An axisymmetric through-flow design script was developed to create a two-stage mixed flow compressor design. Parametrization was utilized with a Single Objective Function Genetic Algorithm and an axisymmetric solver to optimize the flow-path and work input for efficiency. The design was materialized through GTSL developed tools T-Blade3 and geomturbo to create a baseline geometry. The baseline geometry was analyzed and improved upon with a Multi Objective Function Genetic Algorithm optimizing for mass flow and efficiency. Design parameters controlling blade geometry were defined and the genetic algorithm was paired to a 3D CFD meshing and analysis tool (FINE/TURBO). The first of two rotor-stator stages in the design is completed through aerodynamic evaluation. A 50% increase in mass-flow was achieved while maintaining the original size envelope and a 2:1 pressure ratio. The isentropic efficiency of the design is 80.667% at the design point and 81.79% peak efficiency at lower mass flow and pressure ratio.

Committee: Mark Turner Sc.D. (Committee Chair); Michael List Ph.D. (Committee Member); Paul Orkwis Ph.D. (Committee Member); Fred Schauer Ph.D. (Committee Member) Subjects: Aerospace Materials

Doctor of Philosophy, The Ohio State University, 2021, Mechanical Engineering

Turbomachinery, like jet engines and industrial gas turbines in power plants, are very advanced and complex machines. Due to the complexity and cost of modern turbomachinery, there is active research in accurately predicting the physical system dynamics using computational models. Two big mechanisms that affect the structural response are the prestress effects from high rotational speeds and mistuning effects from tolerance deviations, wear, or damage. Understanding the role these two mechanisms play in the computational modeling of these systems is an important step toward a complete digital twin of an entire jet engine. There previously existed modeling methods that enabled each to be analyzed independently, but not simultaneously in an efficient manner. This will be one of the focus points of this dissertation. The other focus being an experimental investigation into exciting system resonances of a rotating bladed disk using air jets. These experiments will be used to validate the computational modeling method developed. This dissertation has three primary objectives. The first objective is to present reduced order modeling methods that allow for the efficient modeling of coupled systems and rotating systems, both with small or large mistuning. By efficiently including these mechanisms, more realistic boundary conditions can be used to help validate the reduced order models (ROMs) with experimental data. Both modeling methods create models a fraction of the size of the full model while retaining key dynamic characteristics of the full model. The second objective of this work is to show the capability of air jets in exciting synchronous and non-synchronous vibrations in a rotating bladed disk. Much previous research in this field focused on experiments with stationary systems. These tests can help isolate specific mechanisms that may be present in bladed disks, but may limit the applicability of the results to actual rotating systems. This work presents a method (open full item for complete abstract)

Committee: Kiran D'Souza (Advisor); Randall Mathison (Committee Member); Manoj Srinivasan (Committee Member); Herman Shen (Committee Member) Subjects: Mechanical Engineering

MS, University of Cincinnati, 2020, Engineering and Applied Science: Aerospace Engineering

Research presented in this thesis focuses on the development of a novel, open-source, automated, and fully parametric non-axisymmetric turbomachinery aero-design-optimization system, NAX. This research forms a part of a collaboration between NASA Glenn Research Center and University of Cincinnati for the development of a Boundary Layer Ingestion based Turbo-electric propulsion system under the NASA N+3 initiative. For the present design, a 1.5 stage aft-mounted Tail Cone Thruster (TCT) unit is considered to ingest the boundary layer flow from the air-frame to increase the overall propulsion efficiency of the aircraft. This results in circumferential distortion (r, θ)in total pressure, swirl, and meridional flow (PT, α,ϕ) at the TCT inlet. at the TCT inlet. These non-uniformities causes partial off-design operation of the TCT and a significant departure in its component efficiency and aero-mechanics integrity. A key feature of the NAX design system is its capability to allow for the design/optimize of spanwise (r) and/or circumferentially (θ) non-axisymmetric 3D blade shapes. It uses a harmonics-based design space parametrization, which offers an extended control in r, θ on blade parameters like blade angles, sweep, lean, chord, and thickness distribution. These non-uniformities (Σ ai sin(θ + ϕi)) can be controlled by modifying the magnitude and phase (i) at arbitrary span. NAX is demonstrated with the Inlet Guide Vanes (IGV) for TCT propulsor. This non-axisymmetric IGV is optimized to reduce the downstream rotor incidence and increase IGV performance under 2D distortion. A novel approach of transforming relative flow angle ( β) from IGV to rotor frame of reference is used to include the uncoupled rotor effects in the optimization of IGV. The typical multi-fidelity design framework is used to develop TCT using a PT radial distortion at throughflow level and then further optimized for 2D distortion at 3D (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Chair); Mark Celestina Ph.D. (Committee Member); Paul Orkwis Ph.D. (Committee Member) Subjects: Aerospace Materials

Master of Science, The Ohio State University, 2021, Mechanical Engineering

This document presents the development and implementation of a new generation of double-sided heat-flux gauges at The Ohio State University Gas Turbine Laboratory (GTL) along with heat transfer measurements for film-cooled airfoils in a single-stage high-pressure transonic turbine operating at design corrected conditions. Double-sided heat flux gauges are a critical part of turbine cooling studies, and the new generation improves upon the durability and stability of previous designs, while also introducing high-density layouts that provide better spatial resolution. These new customizable high-density double-sided heat flux gauges allow for multiple heat transfer measurements in a small geometric area such as immediately downstream of a row of cooling holes on an airfoil. Two high-density designs are utilized: Type A consists of 9 gauges laid out within a 5 mm by 2.6 mm (0.20 inch by 0.10 inch) area on the pressure surface of an airfoil, and Type B consists of 7 gauges located at points of predicted interest on the suction surface. At the same time, improvements to the manufacturing and installation processes for single gauges increased the survival rate of the gauges from 47% to 84%. Both individual and high-density heat flux gauges are installed on the blades of a transonic turbine experiment for the second build of the High-Pressure Turbine Innovative Cooling program (HPTIC2). Run in a short duration facility, the single-stage high-pressure turbine operated at design-corrected conditions (matching corrected speed, flow function, and pressure ratio) with forward and aft purge flow and film-cooled blades. Gauges are placed at repeated locations across different cooling schemes in a rainbow rotor configuration. Airfoil film-cooling schemes include round, fan, and advanced shaped cooling holes in addition to uncooled airfoils. Both the pressure and suction surfaces of the airfoils are instrumented at multiple wetted distance locations and percent spans from roughly (open full item for complete abstract)

Committee: Randall Mathison PhD (Advisor); Michael Dunn PhD (Committee Member) Subjects: Aerospace Engineering; Engineering; Experiments; Mechanical Engineering

MS, University of Cincinnati, 2020, Engineering and Applied Science: Aerospace Engineering

One-way fluid structure interaction (FSI) is explored in two independent scenarios. First, a rectangular single expansion ramp nozzle (SERN) is analyzed through CFD. The SERN was simulated at three conditions: over-expanded, perfectly-expanded, and under-expanded (NPR 3.0, 3.671, and 4.75, respectively). A Schlieren image comparison between CFD results and experiment revealed that there were discrepancies in the location of the shock emanating from the nozzle throat and reflecting through the divergent ramp section. Structural analysis showed that the top plate deforms under load, deflecting outward from the flow by 31 micrometers when over-expanded and nearly 0.3 millimeters when under-expanded. Modal analysis was performed for unloaded, over- and under-expanded conditions, with the first several modes in good agreement with experimental results at the over-expanded condition. For the first four modes found with FEA, all were within 0.7% relative error, except the second mode that was 4.5% different. The second scenario in which FSI is applied is analysis and design of a rotor in a three-row stage with an inlet guide vane (IGV) and outlet guide vane (OGV) of a boundary layer ingesting (BLI) propulsor. The blade shape of the rotor was altered by adjusting the spanwise parameters of thickness distribution, sweep, lean, and chord multiplier in order for the design to satisfy structural requirements. Pressure loads were extracted from CFD and applied to the rotor blade surface for a thorough structural and modal analysis to gather an understanding of how the blade behaved under load. At design speed, the rotor achieved a safety factor of 1.6911 and deformed 0.88 millimeters radially and 5.05 millimeters tangentially. Several modes present possibilities of resonance from perturbations from both IGV and OGV. Although these conditions exist, tuning blade modes to avoid resonance is left to future work.

Committee: Mark Turner Sc.D. (Committee Chair); Ephraim Gutmark Ph.D. (Committee Member); Gui-Rong Liu Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2020, Engineering and Applied Science: Aerospace Engineering

The first stage of an axial compressor was designed and optimized using 1D results from 3D CFD analysis. The compressor was based off of a small commercial jet engine compressor. Structural and modal analysis were performed on the compressor to ensure safety during testing. Modal analysis results were compared to Scanning Laser Vibrometry results. A custom test rig was designed and built utilizing a COTS turbocharger. Additive manufacturing was used to produce custom parts for the test rig, including the compressor test articles. Measurements of the flow were taken to find the performance characteristics of the turbine and custom compressor. The data was acquired using Labview and various COTS measurement devices. Issues with minimum thickness and high centrifugal loads led to multiple design changes which limited the expected performance of the compressor. Complications with additive manufacturing may have also reduced the performance of the compressor. Experimentation was performed at low speeds first. The test rig failed at around 60% design speed, likely due to issues with the custom compressor mounting plate. Performance characteristics were calculated with the limited data and speed lines were generated. The data was compared to the results from CFD analysis. The optimization system was shown to be useful in increasing the performance of a compressor stage. It was concluded that a low cost custom test rig using a turbocharger is a viable solution for small compressor testing. Additive manufacturing was shown to be useful in producing one-off custom designs, though is not a perfect solution. The overall process is a useful process for rapid development and testing of custom compressor designs for small engine applications.

Committee: Mark Turner Sc.D. (Committee Chair); Paul Orkwis Ph.D. (Committee Member); Fred Schauer Ph.D. (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2020, Engineering and Applied Science: Aerospace Engineering

The need for quick and low-cost numerical aerodynamic shape optimization capabilities has become of primary importance in recent years for aircraft and jet engine manufacturers to improve their product performance and reduce design costs. This performance enhancement process is very complex and might involve a few objective functions but, it could readily consider hundreds of design parameters. The success of the fast gradient-based approaches for optimization is due to the introduction of the Adjoint Method, which allowed the computation of the objective function gradient in a very computationally efficient fashion compared to alternative existing methods. The Adjoint Method made it possible to quickly calculate the derivatives of the objective function with respect to the volume grid nodes thanks to a particular differentiation of the Computational Fluid Dynamic solver, whereas several techniques could be used to obtain the design parameter derivatives of volume grid nodes. Despite its popularity, the Adjoint Method for design optimization is still not a standard tool among designers. Possible explanations include the lack of the functional gradient computation capability in some numerical codes for Computational Fluid Dynamics, the need of a well-converged flow solution (sometimes difficult or impossible to achieve), the quite common instability of the adjoint solver, the inability to easily link the derivatives to the actual design variables (especially for turbomachinery cases) and the inability to properly and robustly morph the geometry and the computational domain during the optimization phase. This dissertation details the implementation of the discrete Adjoint Method and its application in the design optimization of turbomachinery blades. The intent is to tackle some of the aforementioned drawbacks of the Adjoint Method and advance the applicability of this approach for automatic design optimization. Some of the contributions of the present work to this a (open full item for complete abstract)

Committee: Paul Orkwis Ph.D. (Committee Chair); Markus Rumpfkeil Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member); Benjamin Vaughan Ph.D. (Committee Member) Subjects: Aerospace Materials

Master of Science, The Ohio State University, 2020, Aero/Astro Engineering

This thesis covers the intensive research effort to elucidate the role of elevated temperature in deposition. Several experimental campaigns were conducted in this pursuit. The testing explored high temperature deposition with 0-10 micron Arizona Road Dust (ARD) with the intent of creating a yield strength model that included temperature effects and could be incorporated into the existing OSU deposition model. Experimental work was first conducted in the impulse kiln facility where small amounts of the test dust were placed on ceramic targets and rapidly exposed to temperatures between 1200K and 1500K. Trends in the packing factor confirmed the existence of two threshold values (1350K and 1425K) that could be linked to strength characteristics of the dust when exposed to high temperatures. Using the information obtained from the kiln experiments, HTDF testing was conducted between 1325K and 1525K. Exit temperatures were set at 25K intervals in this region with a constant jet velocity of 150 m/s. The capture efficiency data showed this trend with temperature and indicated a softening temperature and melting temperature of 1362K and 1512K respectively. With these critical values in hand, the Ohio State University Molten Model was created to modify yield strength with particle velocity and temperature. The model was tested using CFD and showed a good capability for capturing particle temperature effects in deposition from an impinging particle-laden jet. A subsequent test campaign was conducted to explore the effect of varying surface temperature on deposition. Hastelloy coupons with Thermal Barrier Coatings (TBCs) were subjected to a constant jet at 1600K jet and 200 m/s while being cooled via a backside impingement jet. Surface temperatures between 1455K and 1125K were impacted with 0-10 micron ARD while an IR camera monitored the surface. Coupons with higher coolant flowrates (lower surface temperature) saw significantly lower deposition rates than the higher surf (open full item for complete abstract)

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

Master of Science in Mechanical Engineering (MSME), Wright State University, 2019, Mechanical Engineering

Blending is a method of fan and compressor blade repair. The goal of the blending process is to remove stress concentration points such as cracks and nicks along the leading, trailing, or tip edges of the blade. The stressed areas are typically removed by grinding or cropping away the surrounding material. For integrally bladed rotor (IBR) disks, repairing a damaged blade is much more economical than replacing the entire disk. However, the change in shape of the blade will change the local aerodynamics and result in mistuning, both structurally and aerodynamically. In a worst case scenario, the change in the aerodynamic forces acting on the blades could lead to either flutter or resonant fatigue failure. This study focused on the effects of the unsteady aerodynamic loading of blended blades. To examine the phenomena in detail a commercial computational fluid dynamics (CFD) code was used to predict the loading of a first-stage, transonic turbofan subjected to varying degrees of blending. The study revealed the effects of the blend are not limited to the blended blade, as changes in steady and unsteady pressure loading was predicted on the other nonblended blades on the rotor. A maximum steady sectional loading change of 7% at 99% span on the blended blade for the maximum sized blend case was predicted. The unsteady analysis found an 18% maximum change in unsteady 1st mode sectional loading on the blade adjacent to the blended blade for the 18 engine order excitation at 90% span. Therefore, the potential for a serious aeromechanics issue exists which requires an unsteady aerodynamic analysis to be performed. It is difficult to determine the true aeromechanical effects without a coupled aerodynamics/structural analysis, but this aerodynamic loading investigation suggests larger blends could be safely used and should be analyzed.

Committee: J. Mitch Wolff Ph.D. (Advisor); David A. Johnston Ph.D. (Committee Member); Rolf Sondergaard Ph.D., P.E. (Committee Member) Subjects: Aerospace Engineering; Fluid Dynamics; Mechanical Engineering

Master of Science in Mechanical Engineering (MSME), Wright State University, 2019, Mechanical Engineering

A primary source of periodic unsteadiness in low-pressure turbines is the wakes shed from upstream blade rows due to the relative motion between adjacent stators and rotors. These periodic perturbations can affect boundary layer transition, secondary flow, and loss generation. In particular, for high-lift front-loaded blades, the secondary flowfield is characterized by strong three-dimensional vortical structures. It is important to understand how these flow features respond to periodic disturbances. A novel approach was taken to generate periodic unsteadiness which captures some of the physics of turbomachinery wakes. Using stationary pneumatic devices, pulsed jets were used to generate disturbances characterized by velocity deficit, elevated turbulence, and spanwise vorticity. Prior to application in a turbine flow environment, the concept was explored in a small developmental wind tunnel using a single device. The disturbance flowfield for different input settings was measured using hot-film anemometry and Particle Image Velocimetry. Insight was also garnered on how to improve later design iterations. With an array of devices installed upstream of a linear cascade of high-lift front-loaded turbine blades, settings were found which produced similar disturbances at varying frequencies that periodically impinged upon the leading-edge region. These settings were used to conduct an in-passage secondary flow study using high-speed Stereoscopic Particle Image Velocimetry. Results demonstrated the application of the periodic unsteadiness generator but found minor changes to the passage vortex. The vortex rotational strength decreased, and migration increased with increased perturbation frequency. Fourier analyses found the PV to be responsive at the actuation frequency with phase-locked ensemble-averaged data revealing that the disturbance periodically caused the PV to lose rotational strength. However, at the tested discrete frequencies, the vortex did not become locked (open full item for complete abstract)

Committee: Mitch Wolff Ph.D. (Advisor); Rolf Sondergaard Ph.D. (Committee Member); Christopher Marks Ph.D. (Committee Member) Subjects: Aerospace Engineering; Engineering; Experiments; Fluid Dynamics; Mechanical Engineering

Master of Science in Mechanical Engineering (MSME), Wright State University, 2019, Mechanical Engineering

In turbomachinery design, complex internal flows give rise to significant losses and blockage whose effects are difficult to properly analyze without detailed computational fluid dynamics (CFD) methods or experiments. In a typical design method, CFD is used in conjunction with simpler throughflow or cascade codes to hasten the process. However, the lesser physical accuracy of the design codes demands the inclusion of models to improve the accuracy of the throughflow codes. This thesis aims to use CFD data to generate improved loss and blockage models for a 2D compressor throughflow code by matching throughflow data to CFD data using optimizations. This analysis gives insight into the connection between 3D flow features and design blockages and losses. This in turn forms the foundation for the behaviors required in the models for shocks, separations, and tip clearance flows, the implementation of which will improve the physical accuracy of the design code.

Committee: Mitch Wolff Ph.D. (Advisor); Rolf Sondergaard Ph.D. (Committee Member); Rory Roberts Ph.D. (Committee Member); Michael List Ph.D. (Other) Subjects: Aerospace Engineering; Mechanical Engineering

Master of Science in Mechanical Engineering (MSME), Wright State University, 2019, Mechanical Engineering

The low-pressure turbine is an important component of a gas turbine engine, powering the low-pressure spool which provides the bulk of the thrust in medium- and high-bypass engines. It is also a significant fraction of the engine weight and complexity as it can comprise up to a third of the total engine weight. One way to drastically reduce the weight of the low-pressure turbine is to utilize high lift blades. To advance high-lift technology, the Air Force Research Laboratory (AFRL) designed the L2F blade profile, which was implemented in the linear cascade at AFRL/RQT's low speed wind tunnel facility. The L2F blade has very high lift and an excellent midspan performance, however, it was previously demonstrated to generate significant losses in the endwall region. These losses are primarily driven by the complex time-dependent three-dimensional vortical structures present in the region of the junction of the blade and the endwall, dominated by the Passage Vortex (PV). Aerodynamic flow control is one way to mitigate these losses. Previously, a pulsed endwall blowing system was implemented in the endwall region of the L2F blade which produced a loss reduction. This loss reduction was dependent on the pulsing frequency. In this research, the vortical structures for the baseline flow were characterized with respect to time. The time dependent behavior of the passage vortex motion, location, and strength were found for each pulsing frequency to determine a relationship with total pressure loss reduction. The flow through the passage of the tunnel was characterized with respect to time using high-speed stereoscopic particle image velocimetry. The flow for each test condition was characterized using Q-criterion to determine the strength of the passage vortex and its time dependent behavior. It was found that the passage vortex loses and gains strength in an unsteady manner at time scales between 1.9 < ΔT+ < 6.7. The largest total pressure loss reduction was found to corres (open full item for complete abstract)

Committee: Mitch Wolff Ph.D. (Advisor); Christopher R. Marks Ph.D. (Committee Member); Rolf Sondergaard Ph.D., P.E. (Committee Member) Subjects: Mechanical Engineering

PhD, University of Cincinnati, 2018, Engineering and Applied Science: Aerospace Engineering

Numerical methods for computational physics have been applied for many years in the fields of turbomachinery and acoustics. The computational approach to addressing problems in these fields has strongly influenced improvements in performance and advancements in understanding for the systems and physical processes that govern such applications. Particularly within the turbomachinery community, the spectrum of numerical methods being applied is dominated by second-order accurate finite-volume and finite-difference discretizations of the governing equations. These have been quite successful and their robustness has been important in their adoption within industrial engineering as tools for design and analysis. At the same time, such approaches are very dissipative and require dense computational grids to resolve sharp features and capture wave propagation. As problems become more complex and inter-related, the numerical methodologies for analysis tools that are used to address such problems and inform solutions must improve in their fidelity, accuracy, and mathematical rigor. At the same time, high-order finite-element methods have experienced significant attention in the computational fluid dynamics community for their mathematical formalism, localized approach for obtaining high-order accuracy, and amenability to adaptation of the numerical grid and accuracy of the numerical approximation. The challenges of such approaches are to achieve efficiency and stability for the numerical method. The result of many research efforts in this area has pushed the applicability of high-order finite-element methods into many fields and they are approaching the point where they might soon find routine application to industrial problems for engineering design and analysis. This dissertation details the development and application of an implicit discontinuous Galerkin method to applications in turbomachinery and acoustics. This is carried out for the purpose of advancing the ap (open full item for complete abstract)

Committee: Paul Orkwis Ph.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); John Benek Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member); Eric Wolf Ph.D. (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2018, Engineering and Applied Science: Aerospace Engineering

Abstract Energy conversion is a tightly coupled thermodynamic and fluid dynamic process which involves transport of several properties like mass, momentum, vorticity and energy. In reality, there is always some irreversibility associated in the form of entropy rise and must be accounted and minimized for improved performance of energy converters. Unducted and ducted rotors convert kinetic energy of the fluid into power and/or thrust. For the first time, this dissertation analyzes kinetic energy based loss as entropy rise in all horizontal axis unducted rotors using viscous dissipation through multi-fidelity framework developed in-house. Complex flow physics and the effect of B-spline based smooth manipulation of blade shapes are some of the highlights which reveal the inner workings of exergy destruction.A general low fidelity design analysis tool, py_BEM for all types of solo and contra-rotating horizontal axis unducted rotors is developed using blade element momentum theory with several enhancements to airfoil properties including contra-rotating configurations making it a unique tool. A unified approach makes it easier to design, analyze and optimize propellers, helicopter rotors in hover/vertical ascent, wind, tidal and hydrokinetic turbines at lower fidelity level. Linear and angular momentum, kinetic energy and power, exergy and entropy transport at this level provides a platform to understand energy conversion from a simplified flow physics with a thermodynamic perspective. Steady 3D RANS calculations are performed to understand the flow physics in high fidelity using control volume approach to investigate assumptions and document differences with rigorous domain and mesh dependency study. A general multi-fidelity multi-disciplinary design analysis and optimization framework is developed in-house for both ducted and unducted rotors.Underlying flow physics of unducted rotors is explained with regards to streamtube expansion (turbines) and contraction (prope (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); Paul Orkwis Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2018, Engineering and Applied Science: Aerospace Engineering

The predictable effect of surface curvature on the contiguous streamtube allows for the use of geometric curvature as a direct and aerodynamically meaningful parametric model to generate airfoil geometry. A novel and parsimonious parameterization technique driven by specifications of normalized meanline second derivatives, which is related to curvature, and a superimposed thickness distribution which explicitly eliminates or minimizes unintentional oscillations in curvature, is resented. The focus on smooth curvature control is inherently relevant to both internal and external flow geometries, enabling a unified method for generating both isolated airfoils and cascade sections. This technique is implemented and included in T-Blade3 which is an existing in-house open-source code. The underlying methodology for construction of camber-line is entirely analytical ensuring speed of execution. Two different thickness distributions are presented, one based on specifications of thickness B-spline control points, and another based on specifications of exact thickness. The B-spline thickness method requires a simple implementation and executes faster, but cannot impose an exact value of thickness at a specified location. The exact thickness method is capable of imposing specified thickness values both chord-wise and span-wise, while quantifying and optimizing the quality of the thickness curve based on curvature. Consequently, unintentional bumps on the airfoil and oscillations in curvature are eliminated or minimized. The parameterization ensures curvature and slope of curvature continuity on the airfoil surface which are critical for smooth surface pressure distributions. Consequently, losses due to unintentional pressure spikes are minimized and likelihood of separation reduced, resulting in a class of high-performance airfoils. The direct relationship between the parameterization and surface aerodynamics is demonstrated for both isolated and cascade airfoils. A framework (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); Donald French Ph.D. (Committee Member) Subjects: Aerospace Materials