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Casaday, Brian PatrickInvestigation of Particle Deposition in Internal Cooling Cavities of a Nozzle Guide Vane
Doctor of Philosophy, The Ohio State University, 2013, Aero/Astro Engineering
Experimental and computational studies were conducted regarding particle deposition in the internal film cooling cavities of nozzle guide vanes. An experimental facility was fabricated to simulate particle deposition on an impingement liner and upstream surface of a nozzle guide vane wall. The facility supplied particle-laden flow at temperatures up to 1000°F (540°C) to a simplified impingement cooling test section. The heated flow passed through a perforated impingement plate and impacted on a heated flat wall. The particle-laden impingement jets resulted in the buildup of deposit cones associated with individual impingement jets. The deposit growth rate increased with increasing temperature and decreasing impinging velocities. For some low flow rates or high flow temperatures, the deposit cones heights spanned the entire gap between the impingement plate and wall, and grew through the impingement holes. For high flow rates, deposit structures were removed by shear forces from the flow. At low temperatures, deposit formed not only as individual cones, but as ridges located at the mid-planes between impinging jets. A computational model was developed to predict the deposit buildup seen in the experiments. The test section geometry and fluid flow from the experiment were replicated computationally and an Eulerian-Lagrangian particle tracking technique was employed. Several particle sticking models were employed and tested for adequacy. Sticking models that accurately predicted locations and rates in external deposition experiments failed to predict certain structures or rates seen in internal applications. A geometry adaptation technique was employed and the effect on deposition prediction was discussed. A new computational sticking model was developed that predicts deposition rates based on the local wall shear. The growth patterns were compared to experiments under different operating conditions. Of all the sticking models employed, the model based on wall shear, in conjunction with geometry adaptation, proved to be the most accurate in predicting the forms of deposit growth. It was the only model that predicted the changing deposition trends based on flow temperature or Reynolds number, and is recommended for further investigation and application in the modeling of deposition in internal cooling cavities.

Committee:

Jeffrey Bons (Advisor); Ali Ameri (Committee Member); Michael Dunn (Committee Member); Datta Gaitonde (Committee Member); Sandip Mazumder (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

particle deposition; turbomachinery; internal cooling, engine fouling

Heberling, BrianA Numerical Analysis on the Effects of Self-Excited Tip Flow Unsteadiness and Upstream Blade Row Interactions on the Performance Predictions of a Transonic Compressor
MS, University of Cincinnati, 2017, Engineering and Applied Science: Aerospace Engineering
Computational fluid dynamics (CFD) simulations can offer a detailed view of the complex flow fields within an axial compressor and greatly aid the design process. However, the desire for quick turnaround times raises the question of how exact the model must be. At design conditions, steady CFD simulating an isolated blade row can accurately predict the performance of a rotor. However, as a compressor is throttled and mass flow rate decreased, axial flow becomes weaker making the capturing of unsteadiness, wakes, or other flow features more important. The unsteadiness of the tip clearance flow and upstream blade wake can have a significant impact on a rotor. At off-design conditions, time-accurate simulations or modeling multiple blade rows can become necessary in order to receive accurate performance predictions. Unsteady and multi- bladerow simulations are computationally expensive, especially when used in conjunction. It is important to understand which features are important to model in order to accurately capture a compressor’s performance. CFD simulations of a transonic axial compressor throttling from the design point to stall are presented. The importance of capturing the unsteadiness of the rotor tip clearance flow versus capturing upstream blade-row interactions is examined through steady and unsteady, single- and multi-bladerow computations. It is shown that there are significant differences at near stall conditions between the different types of simulations.

Committee:

Paul Orkwis, Ph.D. (Committee Chair); Shaaban Abdallah, Ph.D. (Committee Member); Mark Turner, Sc.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

computational fluid dynamics;turbomachinery;transonic compressor;cfd;compressor stall

., Abhay SrinivasNovel Compressor Blade Design Study
MS, University of Cincinnati, 2015, Engineering and Applied Science: Aerospace Engineering
Jet engine eciency goals are driving compressors to higher pressure ratios and engines to higher bypass ratios, each one driving to smaller cores. This is leading to larger tip gaps relative to the blade height. These larger relative tip clearances would negate some of the cycle improvements, and ways to mitigate this eect must be found. A novel split tip blade geometry has been created which helps improve the eciency at large clearances while also improving operating range. Two identical blades are leaned in opposite directions starting at 85% span. They are cut at mid chord and the 2 halves then merged together so a split tip is created. The result is similar to the alula feathers on a soaring bird. The concept is that the split tip will energize the tip ow and increase range. For higher relative tip clearance, this will also improve eciency. The 6th rotor of a highly loaded 10 stage machine was chosen as the baseline for this study. Three dimensional CFD simulations were performed using CD Adapco\’s Star-CCM+ at 5 clearances for the baseline and split tip geometry. The choking ow and stall margin of the split tip blade was higher than that of the baseline blade for all tip clearances. The pressure ratio of the novel blade was higher than that of the baseline blade near choke, but closer to stall it decreased. The sensitivity of peak eciency to clearance was improved. At tight clearances of 0.62% of blade height, the maximum eciency of the new design was less than the baseline blade, but as the tip clearance was increased above 2.5%, the maximum eciency increased. Structural analysis was also performed to ascertain the feasibility of the design.

Committee:

Mark Turner, Sc.D. (Committee Chair); Shaaban Abdallah, Ph.D. (Committee Member); Paul Orkwis, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Compressors;Turbomachinery;StarCCM;CFD

Sharpe, Jacob Andrew3D CFD Investigation of Low Pressure Turbine Aerodynamics
Master of Science in Mechanical Engineering (MSME), Wright State University, 2017, Mechanical Engineering
A 3-D Reynolds-Averaged Navier Stokes (RANS) model of a highly-loaded blade profile has been developed using a commercial CFD code with an unstructured/structured grid and several different turbulence models. The ability of each model to predict total pressure loss performance is examined in terms of the spanwise loss distribution and the integrated total pressure loss coefficient. The flowfield predicted by each model is investigated through comparisons of isosurfaces of Q criterion to previous Implicit Large Eddy Simulation (ILES) results. The 3-equation k-kl-¿ model was shown to provide the most accurate performance predictions for a baseline 3-D LPT geometry, and was then used to analyze the effect a new 3D contoured geometry. The model accurately predicted the qualitative improvement made by the contour by weakening the various vortex structures.

Committee:

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

Subjects:

Aerospace Engineering; Fluid Dynamics

Keywords:

Computational fluid dynamics; aerodynamics; turbomachinery; low pressure turbines

Holden, Jacob R.Experimental Testing and Computational Fluid Dynamics Simulation of Maple Seeds and Performance Analysis as a Wind Turbine
MS, University of Cincinnati, 2016, Engineering and Applied Science: Aerospace Engineering
Descending maple seeds generate lift to slow their fall and remain aloft in a blowing wind; have the wings of these seeds evolved to descend as slowly as possible? A unique energy balance equation, experimental data, and computational fluid dynamics simulations have all been developed to explore this question from a turbomachinery perspective. The computational fluid dynamics in this work is the first to be performed in the relative reference frame. Maple seed performance has been analyzed for the first time based on principles of wind turbine analysis. Application of the Betz Limit and one-dimensional momentum theory allowed for empirical and computational power and thrust coefficients to be computed for maple seeds. It has been determined that the investigated species of maple seeds perform near the Betz limit for power conversion and thrust coefficient. The power coefficient for a maple seed is found to be in the range of 48 - 54% and the thrust coefficient in the range of 66 - 84%. From Betz theory, the stream tube area expansion of the maple seed is necessary for power extraction. Further investigation of computational solutions and mechanical analysis find three key reasons for high maple seed performance. First, the area expansion is driven by maple seed lift generation changing the fluid momentum and requiring area to increase. Second, radial flow along the seed surface is promoted by a sustained leading edge vortex that centrifuges low momentum fluid outward. Finally, the area expansion is also driven by the spanwise area variation of the maple seed imparting a radial force on the flow. These mechanisms result in a highly effective device for the purpose of seed dispersal. However, the maple seed also provides insight into fundamental questions about how turbines can most effectively change the momentum of moving fluids in order to extract useful power or dissipate kinetic energy.

Committee:

Mark Turner, Sc.D. (Committee Chair); Shaaban Abdallah, Ph.D. (Committee Member); Paul Orkwis, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Maple seed;Computational Fluid Dynamics;Turbomachinery;Wind Turbine;Betz Limit;Boimimicry

Stahl, Brian JamesThermal Stability and Performance of Foil Thrust Bearings
Master of Sciences (Engineering), Case Western Reserve University, 2012, EMC - Aerospace Engineering
The performance map of torque, power, and load capacity as a function of speed for three open source and geometrically identical air lubricated compliant foil thrust bearings is extended from a maximum speed of 40,000 rpm to 70,000 rpm. The bearings are tested against Inconel thrust runners coated with PS400. A second open source foil design is used to make two additional bearings – one of which is also mapped to 70,000 rpm. A non-dimensional torque based on hydrodynamic lubrication consolidates the data within ±15%. The four bearings tested on the high speed rig held their maximum load capacities at speeds between 50,000 and 60,000 rpm, which corresponds to approximately 3.8 to 4.6 million DN. Load capacity diminishes above these speeds. Separately heat treated top foils, in combination with previously conditioned bump foils and backing plate, offer reduced variability when compared to a new matched set.

Committee:

Dr. Joseph M. Prahl (Committee Chair); Dr. J. Iwan D. Alexander (Committee Member); Dr. Paul J. Barnhart (Committee Member); Dr. Christopher DellaCorte (Committee Member); Dr. Robert J. Bruckner (Committee Member)

Subjects:

Aerospace Engineering; Engineering; Mechanical Engineering

Keywords:

foil bearing; thrust bearing; thermal stability; oil-free turbomachinery; performance map

Dykas, Brian DavidFACTORS INFLUENCING THE PERFORMANCE OF FOIL GAS THRUST BEARINGS FOR OIL-FREE TURBOMACHINERY APPLICATIONS
Doctor of Philosophy, Case Western Reserve University, 2006, Mechanical Engineering
The operating characteristics of foil gas thrust bearings are explored experimentally and analytically to ascertain the physical mechanisms that limit bearing performance. Measurements of bearing power loss and load capacity made in a variety of configurations highlight several important factors which influence performance. Consistent with conventional hydrodynamic theory, surface condition of the foil and surface condition of the runner have a large influence on bearing performance. Furthermore, active thermal management via cooling air flow and passive thermal management via conduction through the runner have a large influence. Thermal effects are shown to be more pronounced at higher loads where gas film heat generation and resulting thermoelastic distortion are larger, but smooth lubricious surfaces are needed to achieve these loads. With non-optimal surface conditions such as high levels of roughness, it is shown that asperity contact dominates over thermal deformation. This dissertation quantifies the effects of these non-ideal surface conditions on the load capacity of foil thrust bearings. It is determined that both smooth, low friction surfaces combined with adequate thermal management are necessary to support large loads at high speeds. Furthermore, analysis and modeling suggest that enhanced thermal management is possible by optimizing the thermal characteristics of the runner, an approach not yet exploited by the foil bearing community.

Committee:

Joseph Prahl (Advisor)

Keywords:

foil bearings; gas bearings; air bearings; thrust bearings; oil-free turbomachinery

Kulkarni, SameerDevelopment of a Methodology to Estimate Aero-Performance and Aero-Operability Limits of a Multistage Axial Flow Compressor for Use in Preliminary Design
Master of Sciences (Engineering), Case Western Reserve University, 0, EMC - Aerospace Engineering
The preliminary design of multistage axial compressors in gas turbine engines is typically accomplished with mean-line methods. These methods, which rely on empirical correlations, estimate compressor performance well near the design point, but may become less reliable off-design. For land-based applications of gas turbine engines, off-design performance estimates are becoming increasingly important, as turbine plant operators desire peaking or load-following capabilities and hot-day operability. The current work implements a one-dimensional stage stacking procedure, including a new blockage term, which is used to estimate off-design compressor performance and operability range of a 13-stage axial compressor used for power generation. The procedure utilizes stage characteristics which are constructed from computational fluid dynamics (CFD) simulations of groups of stages. The stage stacking estimates match well with CFD results. These CFD results are used to assess a metric which estimates the stall limiting stages.

Committee:

Jaikrishnan R. Kadambi, PhD (Committee Chair); J. Iwan D. Alexander, PhD (Committee Member); John J. Adamczyk, PhD (Advisor); Joseph M. Prahl, PhD (Committee Member); Mark L. Celestina, PhD (Advisor)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

axial compressors; computational fluid dynamics; turbomachinery

GUIDOTTI, EMANUELEAnalysis of the Unsteady Flow in an Aspirated Counter-Rotating Compressor Using the Nonlinear Harmonic Balance Method
MS, University of Cincinnati, 2008, Engineering : Aerospace Engineering
A multistage frequency domain (nonlinear harmonic) Navier-Stokes unsteady flow solver has been used to analyze the flow field in the MIT (rotor/rotor) counter-rotating compressor. The numerical accuracy and computational efficiency of the nonlinear harmonic method implemented in the CFD code Numeca Fine/Turbo has been demonstrated by comparing predictions with experimental data and nonlinear time-accurate solutions for the test case. The comparison is good, especially considering the big saving in time with respect to a time accurate simulation. Details of the flow field are presented and a physical explanation is provided. Also, suggestions and recommendations on the use of the harmonic balance method are provided. It is concluded that the development of efficient frequency domain approaches enables unsteady flow predictions to be used in the design of modern turbomachines.

Committee:

Mark Turner (Committee Chair); Paul Orkwis (Committee Member); Shaaban Abdallah (Committee Member)

Subjects:

Energy

Keywords:

cfd; turbomachinery; harmonic balance

Wroblewski, Adam ChristopherHealth Monitoring of Cracked Rotor Systems using External Excitation Techniques
Master of Science in Mechanical Engineering, Cleveland State University, 2008, Fenn College of Engineering
Cracked rotors present a significant safety and loss hazard in nearly every application of modern turbomachinery. This thesis focuses on the health monitoring, modeling, and analysis of machines with transverse breathing cracks, which open and close due to the self-weight of the rotor. After considering the modeling of cracked rotors, the thesis investigates an active structural health monitoring approach, focusing on the application of an active magnetic actuator to apply a specially designed external force excitation to the rotating shaft. Extensive experimental data has been collected and analyzed utilizing advanced diagnostic techniques. The presented results demonstrate that the use of a magnetic force actuator to apply external excitation has potential in the diagnostics of cracked rotors. The observed unique crack signatures demonstrate the ability of the method for early diagnosis of transverse rotor cracks.

Committee:

Jerzy Sawicki, PhD (Advisor); John Frater, PhD (Committee Member); Ana Stankovic, PhD (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

Rotordynamics; Turbomachinery; Active Magnetic Bearings; Health Monitoring; Shaft Crack; Cracked Rotor; Breathing Crack; Magnetic Actuator; External Excitation; Transverse Rotor Crack;

Garafolo, Nicholas GordonAN EXPERIMENTAL INVESTIGATION OF MULTIPLE MODE EXCITATION OF AN INTEGRALLY BLADED DISK
Master of Science, University of Akron, 2006, Mechanical Engineering
High cycle fatigue of jet aircraft engine turbomachinery components is a serious problem for aircraft engine designers. State-of-the-art compressor designs have reduced mechanical damping and increased unsteady aerodynamic interaction between blade rows. New manufacturing techniques have enabled the rotor to be comprised of an integrally bladed disk, called a blisk. Advanced compressor designs have blades with low aspect ratios, highly complex geometry and higher tip speeds. Inherent in the single disk stage designs is a reduction in mechanical damping; no longer do the connections of blade to disk offer a viable source of damping. Thus, it is important to accurately portray the structural response of the component in a laboratory environment. In an engine environment, a blade row is simultaneously excited from a number of sources. This multiple mode environment was simulated by the modification of the Traveling Wave test rig at the Turbine Engine Fatigue Facility of the Air Force Research Laboratory. The objective of the research presented herein is to measure response of the ADLARF fan rotor excited in a multiple mode environment. The ADLARF fan is significantly structurally mistuned through the replacement of two opposing metallic blades with composite blades done in a past experiment. In this substantially mistuned environment, the assumption of the linear superposition theory for these multiple mode excitations was validated through spatial discrete Fast Fourier transforms and a data point superposition. It has been shown that the response of a bladed disk depends on the engine order of the excitation. Showing the necessity for matching the spatial mode order of the response with that of the engine order of the excitation has provided valuable insight towards fatiguing a complete bladed disk in a laboratory setting.

Committee:

Scott Sawyer (Advisor)

Keywords:

Traveling Wave; Multiple Mode Excitation; Engine Order; Integrally Bladed Disk; Bladed Disk; Blisk; IBR; Modal Analysis; Turbomachinery; rotor; ADLARF; Force Response; Forced Response; Linear Superposition Theory; High Cycle Fatigue

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

Committee:

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

Subjects:

Aerospace Engineering; Fluid Dynamics

Keywords:

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

Szabo, IstvanA Numerical Study of Water Injection on Transonic Compressor Rotor Performance
PhD, University of Cincinnati, 2008, Engineering : Aerospace Engineering
In this study, numerical simulations of two-phase flow in a transonic compressor rotor (NASA Rotor 37) were performed. Both flow and droplets governing equations were formulated and solved in the reference frame of the rotating blades. An Eulerian-Lagrangian approach was used for the continuous and discrete phases with a two-way interaction model to simulate the mass-, momentum- and energy exchange between the different phases. Water particles were injected at the inlet with uniform particle mass flux, fully evaporating inside the rotating blade row. The phase change was most intense in areas adjacent to shock waves, where the slip velocity of the droplets was the highest. Results show decreased circumferentially averaged total temperature ratio of the air-vapor mixture across the span, which is the direct result of inter-phase energy coupling. An entropy based approach to calculate the isentropic efficiency of a wet compression process in a transonic compressor rotor was also presented. Under the proposed method, the viscous dissipation function was calculated everywhere in the domain in the post-processing phase of the numerical simulation and integrated to the wall, with special treatment in the near-wall regions where high rates of entropy generation occur. For a water to air mass flow ratio of 1% results show increased entropy production across the span, resulting in a 5% drop in compressor isentropic efficiency. Analytical integration of wall functions and numerical integration of the viscous dissipation function allows for reasonable results even with relatively coarse grids and can also be applied for single-phase flows. A parametric study of the effect of initial particle parameters on the wet compression process was also performed. Several speedlines have been computed with different amounts of water, especially near the tip. Results show that numerical stall can be delayed with injection of water near the tip, due to the increase of the axial momentum of the fluid in the endwall region, which is the direct result of phase change.

Committee:

Mark G. Turner, ScD (Committee Chair); Paul D. Orkwis, PhD (Committee Member); Awatef Hamed, PhD (Committee Member); Joseph P. Veres (Committee Member)

Subjects:

Engineering

Keywords:

multiphase; compressor; turbomachinery; fogging; injection

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

Green, Brian RichardTime-Averaged and Time-Accurate Aerodynamic Effects of Rotor Purge Flow for a Modern, Rotating, High-Pressure Turbine Stage and Low-Pressure Turbine Vane
Doctor of Philosophy, The Ohio State University, 2011, Mechanical Engineering

Rotor purge flow cavity seals are used in gas turbine engines to prevent ingestion of the mainstream gas flow into the purge cavity. Ingestion into this cavity leads to an increase in the cavity air temperature and subsequently to the rotor disk and stator metal temperatures leading to higher thermal stresses and reduced disk and stator fatigue life. An over designed cavity seal with an excess amount of purge flow has the downside of increasing engine fuel consumption through reduced turbine efficiency. The opposite approach of strengthening the hardware to withstand the higher stress and temperatures would increase the weight of the propulsion system. Understanding how the purge flow cavity and cavity seals interact with the mainstream gas is important to producing a balanced design between weight, fuel consumption, efficiency, and fatigue life of surrounding hardware.

The main objective of this research was to perform an experimental and computational study of a one and one half stage high-pressure turbine installed at The Ohio State University Gas Turbine Laboratory Turbine Test Facility with emphasis on the rotor purge cavity. The rig housing the turbine stage incorporated many features found in a typical commercial high-pressure turbine such as a cooled high-pressure vane row with hub and shroud cooling, a downstream blade row followed by a downstream vane row, the ability to created elevated radial inlet temperature profiles using a combustor emulator, and a cooling supply line to the purge cavity. Multiple runs were performed to study the effects of cooling flows from both an aerodynamic and heat transfer perspective and incorporated instrumentation throughout the rig in order to capture time-accurate temperature, pressure, and heat flux measurements. The run matrix included cold rig configurations with no cooling flow, high-temperature uniform inlet profiles at the vane inlet for cases with and without cooling flows, and high-temperature radial inlet profiles with and without cooling flows. The computational study was performed using the Numeca FINE/Turbo code utilizing a multiple blade row model with both a steady and harmonic unsteady technique in order to simulate the physics of the experiments.

Comparisons between the data and the computational results were performed for five different operational conditions: cold inlet and no cooling flow run, an elevated radial inlet temperature profile with purge and without purge cooling, and an elevated flat inlet temperature profile with purge and without purge cooling flow. The solutions were found to match very well to the exit rake measurements, the leading edge blade profile in the rotating frame of reference, the time-averaged and time-accurate static pressure on the vane hub, the stationary cavity wall, and the rotating cavity. Time-average comparisons were shown for the thermocouples located on the blade platform and in the stationary and rotating side of the cavity. For the two radial inlet cases and for the cold inlet case, these comparisons showed very good agreement while the two flat inlet profile cases showed that the computational models general under-predicted the static temperature levels both on the rotor platform and in the cavity.

Committee:

Michael Dunn (Advisor); Sandip Mazumder (Committee Member); Mohammad Samimy (Committee Member); Jen-Ping Chen (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

Turbomachinery;High-pressure Turbine Rotor Purge Flow;Unsteady CFD;

DRENSKY, GEORGE KERILOVAMBIENT AND HIGH TEMPERATURE EROSION INVESTIGATION OF MATERIALS AND COATINGS USED IN TURBOMACHINERY
MS, University of Cincinnati, 2002, Engineering : Aerospace Engineering
The design and development of high performance turbomachinery operating in an ambient environment with solid particles require a thorough knowledge of the fundamental phenomenon associated with particulate flows. Because of the serious consequences of turbomachinery erosion on their performance and life, it is important to have reliable methods for predicting their erosion when solid particles are ingested with the incoming flow. The ingestion of these solid particles over a period of time will reduce the efficiency of the propulsion system, causing increased fuel consumption and decrease in thrust. Several studies, which are essential to predicting blade surface erosion intensity and pattern, have been conducted at the University of Cincinnati's Propulsion Laboratory over the past (30ty) years. This particular work describes only some investigation done on erosion behavior of materials and coatings exposed to different types of solid particles, velocities, temperatures, and impingement angles. For the present work the following materials and coatings were evaluated: (AA-22) Ceramic Material, Tungsten Carbide Cobalt (WC-Co) coating, Chromium Carbide Nickel (Cr 3C 2Ni) coating, and Titanium Nitride (TiN) coating. The erosive wear of the samples was studied experimentally by exposing them to particle-laden flow at velocities from (100 to 1000 ft/sec). The studied temperatures were between ambient (70°) and (1400°F) and impingement angles from (15° to 90°) degrees. The erosive particles used for the ambient and high temperature testing were: Runway Sand (100 and 1000 microns), Aluminum Oxide (Al 2 0 3 -27.5 microns), Crushed Quartz (60 microns), Arizona Test Dust (20 and 100 microns), and Silica Carbide (30 microns).

Committee:

Dr. Widen Tabakoff (Advisor)

Keywords:

erosion; coatings; turbomachinery; erosion tunnel facility; Titanium Nitride protective coats

LIST, MICHAEL GREGORYQUARTER ANNULUS SIMULATIONS OF BLADE ROW INTERACTION AT SEVERAL GAPS AND DISCUSSION OF LOW PHYSICS
MS, University of Cincinnati, 2007, Engineering : Aerospace Engineering
In order to accurately model the physics associated with losses in a transonic compressor, a time-accurate simulation of a transonic compressor rig was developed. Parameters for grid resolution and time-stepping found in previous phase-lag analyses were used to develop quarter annulus simulations to investigate the physics involved in the rotor bow shock interaction with a highly loaded upstream blade row and its effect on the compressor. Three different axial spacings between the rotor and the upstream blade row have been simulated corresponding to the experimental rig. Details of the unsteady and time-averaged flow field are presented and described. The numerical simulations indicate the same trends as the experiment for efficiency at various spacings. The interaction mechanisms which govern entropy generation and losses are explored. These include wake-shock, wake-boundary layer, and shock-boundary layer interactions.

Committee:

Dr. Mark Turner (Advisor)

Subjects:

Engineering, Aerospace

Keywords:

computational fluid dynamics; turbomachinery; blade row interaction; flow physics

Swar, RohanParticle Erosion of Gas Turbine Thermal Barrier Coating
MS, University of Cincinnati, 2009, Engineering : Aerospace Engineering
The purpose of this research is to examine and understand the complex phenomenon associated with the particle impacts on turbine blades and the associated erosion of Thermal Barrier Coated (TBC) turbine vane and blade surfaces by ingested solid particle impacts. Both experimental and computational techniques were used to find out the parameters relevant to rebound characteristics of particles and erosion rate of TBC coatings.

In the experimental study, tests were conducted in the erosion wind tunnel facility at University of Cincinnati for TBC coated and uncoated blade materials to determine the erosion rates and particle restitution characteristics under different impact conditions. Particle Image Displacement Velocimetry (PIDV) technique was used to determine particle rebound characteristics for different impact conditions. From the experimental results, empirical erosion rate models and restitution coefficient models for Alumina particles impacting on TBC coated blade surface are developed using non-linear regression analysis technique to predict the erosion rate and restitution coefficients for various impact conditions.

In the computational analysis, numerical simulations were conducted for the three-dimensional flow field and particle trajectories through a high pressure single stage axial gas turbine. The solution to the Reynolds Averaged Navier Stokes (RANS) equations for turbulent compressible flow were obtained numerically using ANSYS CFX solver for unsteady N-S equations in their conservation form. In gas turbine applications generally the particle loadings that are encountered are sufficiently low hence a one-way gas-particle interaction model was used to simulate the particle dynamics involved. This does not take into consideration the effects of dispersed particles’ momentum exchange with the gas flow field.

The experimentally based particle surface restitution models were incorporated in the simulations to determine particle rebound conditions after each surface impact. The computed particle surface impact statistics were combined with experimentally based erosion models to predict the stator vanes and rotor blades coated surface erosion pattern and intensity.

The experimental results reveal that the erosion rate increases with increase in impingement angle, impact velocity, and temperature. The trajectories are determined for 26 µm and 500 µm alumina particles. The simulation results show that the particle velocity in the stator is reduced by the surface impacts, which causes the particles to enter the rotor with negative incidence compared to the flow. The rotor impacts reduce the particle velocities in the rotating frame, but could increase their absolute velocity. Inertia dominates the 500µm particle trajectories that reenter the stator row after rebounding from the rotor leading edge impacts. The simulation results predicted the intensity and pattern of TBC erosion over the stator and rotor blade surfaces and the variation in the overall blade surfaces erosion with ingestion velocity.

Committee:

Awatef Hamed, PhD (Committee Chair); Widen Tabakoff, PhD (Committee Member); Robert Miller, PhD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Thermal Barrier Coating;Gas Turbine Erosion;CFD;Particle Trajectory;Turbomachinery

Giuliani, James EdwardJet Engine Fan Response to Inlet Distortions Generated by Ingesting Boundary Layer Flow
Doctor of Philosophy, The Ohio State University, 2016, Aero/Astro Engineering
Future civil transport designs may incorporate engines integrated into the body of the aircraft to take advantage of efficiency increases due to weight and drag reduction. Additional increases in engine efficiency are predicted if the inlets ingest the lower momentum boundary layer flow that develops along the surface of the aircraft. Previous studies have shown, however, that the efficiency benefits of Boundary Layer Ingesting (BLI) inlets are very sensitive to the magnitude of fan and duct losses, and blade structural response to the non-uniform flow field that results from a BLI inlet has not been studied in-depth. This project represents an effort to extend the modeling capabilities of TURBO, an existing rotating turbomachinery unsteady analysis code, to include the ability to solve the external and internal flow fields of a BLI inlet. The TURBO code has been a successful tool in evaluating fan response to flow distortions for traditional engine/inlet integrations. Extending TURBO to simulate the external and inlet flow field upstream of the fan will allow accurate pressure distortions that result from BLI inlet configurations to be computed and used to analyze fan aerodynamics and structural response. To validate the modifications for the BLI inlet flow field, an experimental NASA project to study flush-mounted S-duct inlets with large amounts of boundary layer ingestion was modeled. Results for the flow upstream and in the inlet are presented and compared to experimental data for several high Reynolds number flows to validate the modifications to the solver. Once the inlet modifications were validated, a hypothetical compressor fan was connected to the inlet, matching the inlet operating conditions so that the effect on the distortion could be evaluated. Although the total pressure distortion upstream of the fan was symmetrical for this geometry, the pressure rise generated by the fan blades was not, because of the velocity non-uniformity of the distortion. Total pressure profiles at various axial locations are computed to identify the overall distortion pattern, how the distortion evolves through the blade passages and mixes out downstream of the blades, and where any critical performance concerns might be. Stall cells are identified that are stationary in the absolute frame and are fixed to the inlet distortion. Flow paths around the blades are examined to study the stall mechanism. Rather than a static airfoil stall, it is observed that the non-uniform pressure loading promotes a three-dimensional dynamic stall. The stall occurs at a point of rapid incidence angle oscillation, observed when a blade passes through the distortion, and re-attaches when the blade leaves the distortion.

Committee:

Jen-Ping Chen, Ph.D. (Advisor); Jeffrey Bons, Ph.D. (Committee Member); Dunn Mike, Ph.D. (Committee Member); Gaitonde Datta, Ph.D. (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

turbomachinery;CFD;inlets

Nemnem, Ahmed M. F.A General Multidisciplinary Turbomachinery Design Optimization system Applied to a Transonic Fan
PhD, University of Cincinnati, 2014, Engineering and Applied Science: Aerospace Engineering
The blade geometry design process is integral to the development and advancement of compressors and turbines in gas generators or aeroengines. A new airfoil section design capability has been added to an open source parametric 3D blade design tool. Curvature of the meanline is controlled using B-splines to create the airfoils. The curvature is analytically integrated to derive the angles and the meanline is obtained by integrating the angles. A smooth thickness distribution is then added to the airfoil to guarantee a smooth shape while maintaining a prescribed thickness distribution. A leading edge B-spline definition has also been implemented to achieve customized airfoil leading edges which guarantees smoothness with parametric eccentricity and droop. An automated turbomachinery design and optimization system has been created. An existing splittered transonic fan is used as a test and reference case. This design was more general than a conventional design to have access to the other design methodology. The whole mechanical and aerodynamic design loops are automated for the optimization process. The flow path and the geometrical properties of the rotor are initially created using the axi-symmetric design and analysis code (T-AXI). The main and splitter blades are parametrically designed with the created geometry builder (3DBGB) using the new added features (curvature technique).The solid model creation of the rotor sector with a periodic boundaries combining the main blade and splitter is done using MATLAB code directly connected to SolidWorks including the hub, fillets and tip clearance. A mechanical optimization is performed with DAKOTA (developed by DOE) to reduce the mass of the blades while keeping maximum stress as a constraint with a safety factor. A Genetic algorithm followed by Numerical Gradient optimization strategies are used in the mechanical optimization. The splittered transonic fan blades mass is reduced by 2.6% while constraining the maximum stress below 50% material yield strength using 2D sections thickness and chord multipliers. Once the initial design was mechanically optimized, a CFD optimization was performed to maximize efficiency and/or stall margin. The CFD grid generator (AUTOGRID) reads 3DBGB output and accounts for hub fillets and tip gaps. Single and Multi-objective Genetic Algorithm (SOGA, MOGA) optimization have been used with the CFD analysis system. In SOGA optimization, efficiency was increased by 3.525% from 78.364% to 81.889% while only changing 4 design parameters. For MOGA optimization with higher weighting efficiency than stall margin, the efficiency was increased by 2.651% from 78.364% to 81.015% while the static pressure recovery factor was increased from 0.37407 to 0.4812286 that consequently increases the stall margin. The design process starts with a hot shape design, and then a hot to cold transformation process is explained once the optimization process ends which smoothly subtracts the mechanical deflections from the hot shape. This transformation ensures an accurate tip clearance. The optimization modules can be customized by the user as one full optimization or multiple small ones. This allows the designer not to be eliminated from the design loop which helps in taking the right choice of parameters for the optimization and the final feasible design.

Committee:

Mark Turner, Sc.D. (Committee Chair); Anthony J. Gannon, Ph.D. (Committee Member); Shaaban Abdallah, Ph.D. (Committee Member); Paul Orkwis, Ph.D. (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Multidisciplinary;Optimization;Turbomachinery;Aero-mechanical;Design

Longo, Joel JosephUnsteady Turbomachinery Flow Simulation With Unstructured Grids Using ANSYS Fluent
Master of Science, The Ohio State University, 2013, Aero/Astro Engineering
In the realm of CFD, simulations can be very time consuming and the ability to obtain a solution quickly to a given problem is extremely valuable. Specialized solvers that use structured grids often have a lot of time spent in preprocessing to generate a grid if the geometry is relatively complex. Alternative software exists that makes use of unstructured grids, but often are general solvers and may not be able to obtain as accurate of a solution. In an attempt to see how well an unstructured alternative solver would compare to a structured specialized solver, a verified solution taken from the structured MSU-TURBO turbomachinery code was duplicated in ANSYS Fluent with an unstructured grid. Despite some limitations with ANSYS Fluent's turbomachinery setup, a close, but not exact, solution was obtained and compared to the MSU-TURBO solution.

Committee:

Jen-Ping Chen (Advisor); Mei Zhuang (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

fluent; turbomachinery; turbine; unstructured grid; TURBO

Mathison, Randall MelsonExperimental and Computational Investigation of Inlet Temperature Profile and Cooling Effects on a One and One-Half Stage High-Pressure Turbine Operating at Design-Corrected Conditions
Doctor of Philosophy, The Ohio State University, 2009, Mechanical Engineering
As the demand for greater efficiency and reduced specific fuel consumption from gas turbine engines continues to increase, design tools must be improved to better handle complicated flow features such as vane inlet temperature distortions, film cooling, and disk purge flow. In order to understand the physics behind these features, a new generation of turbine experiments is needed to investigate these features of interest for a realistic environment.This dissertation presents for the first time measurements and analysis of the flow features of a high-pressure one and one-half stage turbine operating at design corrected conditions with vane and purge cooling as well as vane inlet temperature profile variation. It utilizes variation of cooling flow rates from independent circuits through the same geometry to identify the regions of cooling influence on the downstream blade row. The vane outer cooling circuit, which supplies the film cooling on the outer endwall of the vane and the trailing edge injection from the vane, has the largest influence on temperature and heat-flux levels for the uncooled blade. Purge cooling has a more localized effect, but it does reduce the Stanton Number deduced for the blade platform and on the pressure and suction surfaces of the blade airfoil. Flow from the vane inner cooling circuit is distributed through film cooling holes across the vane airfoil surface and inner endwall, and its injection is entirely designed with vane cooling in mind. As such, it only has a small influence on the temperature and heat-flux observed for the downstream blade row. In effect, the combined influence of these three cooling circuits can be observed for every instrumented surface of the blade. The influence of cooling on the pressure surface of the uncooled blade is much smaller than on the suction surface, but a local area of influence can be observed near the platform. This is also the first experimental program to investigate the influence of vane inlet temperature profile on a cooled turbine operating at design corrected conditions. The vane inlet temperature profile has a substantial effect on the temperature measured at the blade leading edge and the Stanton Numbers deduced for the uncooled blade airfoil. While the temperature profile is slightly reshaped passing through the vane, a radial or hot streak profile introduced at the vane inlet can still be clearly measured at the blade. Hot streak magnitude and alignment also influence the blade temperature and heat-flux measurements. A concurrent effort to predict the blade leading edge and platform temperatures for the uncooled portions of this experiment using the commercial code FINE/Turbo is also presented. This investigation is not intended as a detailed computational study but as a check of current code implementation practices and a sanity check on the data. The best predictions are generated using isothermal wall boundary conditions with the nonlinear harmonic method. This is a novel prediction type that could only be performed using a development version of FINE/Turbo.

Committee:

Dr. Michael Dunn, PhD (Advisor); Dr. Sandip Mazumder, PhD (Committee Member); Dr. William Rich, PhD (Committee Member); Dr. Mohammad Samimy, PhD (Committee Member)

Subjects:

Fluid Dynamics; Mechanical Engineering

Keywords:

turbomachinery; gas turbine; film cooling; temperature distortion; FINE/Turbo; short-duration; vane inlet temperature profile; hot streak; purge cooling; isothermal; blade; vane; jet engine; design-corrected conditions; computational fluid dynamics

Gutzwiller, DavidAutomated Design, Analysis, and Optimization of Turbomachinery Disks
MS, University of Cincinnati, 2009, Engineering : Aerospace Engineering
Turbomachinery disks are used in virtually all axially configured gas turbines. These components are both highly stressed and heavy. A single engine may contain over a dozen disks that represent a significant percentage of the total engine weight. Proper design and optimization of turbomachinery disks is an important topic that could yield a significant reduction in engine weight and improved engine system design. This paper focuses on rapid low fidelity design, analysis, and optimization of isotropic and transversely isotropic disks. Background discussion includes the development of a one dimensional plane stress model, common disk geometry parameterization methods, and the implementation of a genetic algorithm for thickness distribution optimization. Two new methods of parameterizing disk geometry are presented. One is a Continuous Slope (CS) method, and the other is an arbitrary control point method. Additional discussion focuses on the effectiveness and robustness of the geometry definition methods in relation to each other and in comparison to a higher fidelity finite element method. Hardware from the GE E3 high pressure compressor, GE E3 high pressure turbine, and the GE90 fan are shown as examples. It is shown that the two new geometry definitions methods produce optimum disks with less mass than the traditional methods. Lastly, the potential pitfalls of using a plane stress model are discussed along with recommendations on how to avoid erroneous results.

Committee:

Mark Turner, ScD (Committee Chair); Stephen Shirooni, PhD (Committee Member); Kumar Vemaganti, PhD (Committee Member); Shaaban Abdallah, PhD (Committee Member)

Subjects:

Aerospace Materials

Keywords:

Turbomachinery;Disk;Optimization;Genetic Algorithm

Hutton, Timothy M.Innovative Forced Response Analysis Method Applied to a Transonic Compressor
Master of Science in Engineering (MSEgr), Wright State University, 2003, Mechanical Engineering
Hutton, Timothy. M.S. Egr., Department of Mechanical Engineering, Wright State University, 2003. Innovative Forced Response Analysis Method Applied To A Transonic Compressor. A set of inlet guide vane (IGV) unsteady surface pressure measurements of a transonic compressor is presented. Using a flexible pressure sensor array, unsteady IGV suction-surface and pressure-surface pressures are acquired for six spanwise by five chordwise locations for various speed lines and throttle settings. Measurements from this sensor array are used to investigate unsteady vane/blade interaction aeromechanical forcing functions in a modern, highly loaded compressor stage. A significant effect is shown on the unsteady forced response of the IGV with changes in compressor operating point and IGV/rotor axial spacing for various span and chord locations. In particular, variations in the compressor operating point (i.e., mass flow rate and pressure ratio) cause change in both the magnitude and phase of the forced response, with the near-stall operating point producing the highest response. Changes in the axial spacing between the IGV and rotor rows from 12% to 26% of the IGV chord resulted in a 50% reduction in the magnitude of the forced response. A significant variation in the forced response with span is noted, especially at the 5% span location where the rotor relative flow is subsonic. In this region, changes in the operating point and axial spacing had a negligible effect on the forced response of the IGV. An innovative data reduction/analysis method is presented to quantify and statistically analyze the degree of blade-to-blade variations in the measured aerodynamic forcing functions obtained by turbomachinery experimentation. This method is used to analyze experimental data of IGV surface unsteady pressure response due to the aerodynamic forcing function produced by the downstream transonic compressor rotor with (1) factory-whole blades and (2) trimmed (blended) blades resulting from the repair of crack damage on two of the rotor blades. Results from the variation metric and l2-norm analysis indicate that the scaled metric possesses large magnitude change versus blade index for the trimmed rotor compared to that of the untrimmed rotor, with the largest values occurring near the trimmed blades. Each method is nearly always able to detect the trimmed blades. Using the distance between cluster centroids from the K-means cluster analysis as a metric of variation within each rotor, cluster distances increased by as much as a factor of 4 for the trimmed rotor compared to the untrimmed rotor. Therefore, correctly identifying the trimmed rotor data as having a significantly higher amount of blade-to-blade variability. Finally using the cluster distance as a goodness parameter for variability, the non-trimmed data was investigated for trends with changes in compressor operating conditions. This analysis showed an increase in blade-to-blade variability with increases in the compressor flow rate. Therefore, this data reduction/analysis method has the potential to be utilized as an indicator of the compressor operating point for control methods.

Committee:

J. Wolff (Advisor)

Subjects:

Engineering, Mechanical

Keywords:

Surface Pressure Inlet Guide Vane Transonic Compressor Variation Cluster Distance Statistic Least Squares Fit Turbomachinery Forced Response Unsteady

Grannan, Nicholas D.Design and Structural Analysis of a Dual Compression Rotor
Master of Science (M.S.), University of Dayton, 2013, Aerospace Engineering
The Dual Compression Rotor (DCR) is a turbine engine component technology which enables the novel turbine engine concept titled the Revolutionary Innovative Turbine Engine (RITE). The DCR and RITE concept is an attempt to provide significant improvements over the traditional turbine engine design. The RITE concept, along with the DCR, represents a paradigm shift over the traditional turbine engine design. The design of the DCR features two compressor stages, one forward flow and one reversed flow, along with an outer turbine stage on a single rotor. The RITE concept offers the potential to decrease specific fuel consumption over the current state of the art, while maintaining thrust and decreasing turbine inlet temperature. The RITE concept will eliminate the need for cooling and improve performance during operation away from the design point. The DCR decreases engine axial length requirements, reducing weight, and features available turbine cooling flow inboard on the rotor. This thesis focuses on the development of a small scale demonstration of the DCR concept. An iterative design process was performed on the DCR until an aerodynamic design of the compressor and turbine stages aligned with the structural performance of available materials. Finite element analysis was performed on the DCR geometry for each iteration. Following the establishment of a preliminary design, additional design work was performed on static structures, dynamic face seals, bearings, and test fixtures. Lead time for the fabrication of the DCR and static structures prohibited the inclusion of experimental results; however, suggested testing procedures and conclusions based on the design being fabricated are included.

Committee:

Steven E. Olson (Advisor); Robert A. Brockman (Committee Member); Frederick R. Schauer (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

turbomachinery; structural analysis; gas turbine; propulsion; novel concept

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