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

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

The goal of this work is to design and develop an injector with a novel porous injection technology (PIT) for dry low NOx combustor (DLN). One of the key factors that is essential for lowering NOx levels is the efficient mixing of fuel-air in both spatial and temporal domains. The porous injection technology has the potential to reduce the spatial and temporal gradients to a minimum. This novel injector design utilized different concepts such as lean premixing, micro-mixing and straight flow with bluff bodies' stabilization mechanism. The micro-mixer is a multi-injector block with nine injectors arranged in an equally spaced rectangular 3 by 3 array. Each injector in the multi-injector block has a porous tube through which fuel is injected. The porous tube is made of stainless steel with 30 µm porosity. Each porous tube is surrounded by eight smaller tubes through which compressed air is passed. The centerbody is mounted above the porous tube. The fuel and air mix in the annular space between the injector wall and the porous tube. The reacting and non-reacting flows of the micro-mixer under atmospheric conditions and a pressure drop of 4% were investigated as part of the injector development process. To evaluate the fuel-air mixing quality, two measurement techniques were used. The CO2 mixing technique - developed in-house, was used to quantify the spatial variations in the fuel mass fraction. Planar Laser Induced Fluorescence (PLIF) was used to obtain both spatial and temporal fuel mass fractions. The CO2 mixing measurements were used to validate the PLIF data for quantification. The RMS fluctuations in spatial and temporal domains were quantified from PLIF data. The length of the upper block was optimized and decided based on the mixing quality. Furthermore, Particle Image Velocimetry (PIV) measurements were conducted to study the injector's aerodynamics under the same operating conditions. The PIV measurements showed a Central Toroidal Recirculation Zone (CT (open full item for complete abstract)

Committee: San-Mou Jeng Ph.D. (Committee Chair); Jun Cai Ph.D. (Committee Member); Jongguen Lee Ph.D. (Committee Member); Bassam Mohammad Abdelnabi Ph.D. (Committee Member) Subjects: Aerospace Materials

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

As global warming becomes a cause of serious concern worldwide, stricter and stricter emission regulations are being imposed on gas turbine engines. Flameless combustion is a novel combustion technique that offers a significant reduction in NOx and CO emissions. The presence of a flameless flame is indicated by uniform temperature distribution in the combustor, which leads to simultaneous reductions in NOx and CO emissions. Nick Overman conducted tests on a swirl stabilized flameless burner in the GDPL lab at the University of Cincinnati. A well-distributed flameless flame was observed for an overall equivalence ratio of 0.36. But as the equivalence ratio was increased, the flame in the combustor switched to a diffusion flame, and non-uniform temperature distribution was observed, which led to an increase in NOx emissions. The work in this thesis aims to improve the operational range of flameless combustion by modifying the swirl stabilized setup used by Nick Overman to include a cavity upstream of the swirler. ANSYS Fluent is used to numerically investigate the performance of such a cavity-swirler setup. The k-epsilon realizable and Laminar Finite Rate model is used to model turbulence and combustion, respectively. Multiple cavity designs which lead to a final successful design are described in detail in this thesis. The final successful design consists of 8 fuel injectors surrounded by 24 air injectors introducing fresh reactants to the cavity. Modifications were then made to this design to include 8 injectors in the second stage of the swirler. The cavity injectors aligned at an angle to the cavity also possess a swirl angle to impart a tangential component of velocity to the reactants being introduced in the cavity. The performance of two designs, the Swirler Reduced air and Swirler Fuel cases, are investigated at different equivalence ratios. Parameters such as temperature, OH distribution, NOx, CO, CO2, H2O, and combustion efficiency are used to compare the tw (open full item for complete abstract)

Committee: Ephraim Gutmark Ph.D. (Committee Chair); Prashant Khare Ph.D. (Committee Member); Rodrigo Villalva Gomez Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2019, Engineering and Applied Science: Aerospace Engineering

The main objective of this research is to characterize the swirling flow in a gas fuel injector used in gas turbine engines. The experimental investigation conducted by implementing particle image velocimetry (PIV) measurements. The effect of confinement sizes are studied, as well as different Reynolds number effect. Unconfined swirling flow showed different flow characteristics than confined typical swirling flow. Unconfined flow features stronger axial jet spreads with short and thin recirculation zones, which is indicative of thin inner shear layer (ISL). The width of the vortex breakdown bubble was about the same size of the nozzle diameter in unconfined cases. Confining the swirling flow caused the width of vortex breakdown bubble to increase to become three times of the nozzle diameter with large confinement, and twice of the nozzle diameter in both medium and small confinements. In addition, the size and shape of inner recirculation zones significantly changed with confinements. Shear layer becomes thicker due the increased width of inner recirculation zones in confined cases. The axial velocity magnitude experienced reduction with confinements, indicating weaker axial jet spread. Furthermore, confinement forced the inlet jet to penetrate radially, which can be noticed by the increase radial velocity magnitude near the exit. In addition, increasing Reynolds number in confined flow induced greater radial jet dispersion. The large confinement had the lowest axial velocity magnitude and demonstrated a unique flow filed structure. Both medium and small confinements have similar axial and radial velocity values as well as similar flow characteristics. The axial centerline plots of axial velocity showed that the length of the reverse flow region or vortex breakdown are increased with confinements. The radial velocity profile, regardless of their considerably low magnitudes, illustrated non-monotonic behavior with increasing Reynolds number in all cases particula (open full item for complete abstract)

Committee: Ephraim Gutmark Ph.D. (Committee Chair); Mark Turner Sc.D. (Committee Member); Rodrigo Villalva Gomez Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2019, Engineering and Applied Science: Aerospace Engineering

This research provides an experimental investigation into the behavior of naturally occurring, fuel-rich, unstable combustion of liquid fuel in a gas turbine combustor. Testing was done using an acoustically isolated single nozzle combustion test rig using liquid Jet-A fuel, capable of operating under unstable combustion conditions. A test matrix was established to identify operating conditions that incited high combustion instability in fuel-rich conditions. The pressure and flame intensity emissions were analyzed across the test matrix in both time and frequency domains. The Rayleigh Criterion was applied to identify the presence of thermo-acoustic instabilities using combustor pressure and flame emission fluctuations. The flame response was established to further characterize the flame dynamics using using the combustor velocity and flame emission fluctuations. Periodic flame structure and behavior was determined using a high speed camera and compared to the convection time delays calculated via flame emission measurements. The investigation concluded that for cases with sufficiently high pressure fluctuations, the Rayleigh Criterion identified thermo-acoustic instability in fuel-rich combustion. The high-speed video and flame emission measurements revealed the spatial and temporal characteristics of emissions in the OH*-band, indicating that it serves as the best approximation for heat release in a fuel-rich flame. Additionally, the high speed camera video and flame emission measurements indicated that CH*-band emissions convectively lag OH*-band emissions.

Committee: Jongguen Lee Ph.D. (Committee Chair); San-Mou Jeng Ph.D. (Committee Member); Kwanwoo Kim Ph.D. (Committee Member) Subjects: Aerospace Materials

MS, University of Cincinnati, 2017, Engineering and Applied Science: Aerospace Engineering

The altitude relight of a gas turbine combustor is an FAA and EASA regulation which dictates the successful re-ignition of an engine and its proper spool-up after an engine malfunction during flight. At cruising altitudes, this becomes problematic, as the ignition energy required is proportional to the inverse of the squared ambient pressure, and the ignition sources are limited operationally. Additionally, current trends in gas turbine combustor for the reduction of harmful emissions such as NOx, CO, and UHC are resulting in designs that are leaner, quicker, and smaller. A comparison of combustor designs across generations yields that the stability margin during a high altitude relight has narrowed, and the combustion efficiency upon ignition decreases as the volume becomes smaller, all else being constant. The goal of this research is then to study the relight process of the currently operational generation of RQL combustors, which have shown proven reliability. Pressure drop, ambient pressure, ambient temperature, and equivalence ratio were all studied on a 3-cup single-annular combustor sector to create an ignition map. The flame development process was studied through the implementation of high-speed video. The three swirlers were each a radial-jet design with pressure atomizing fuel nozzles of the same flow number. Nozzles were inserted such that the face was flush with the base of the swirler, and the fuel pre-filmed on the swirler venturi during normal operation. Sets of dilution holes on the upper and lower radius of the combustor walls lined the combustor liner. Testing was conducted by placing the three-cup sector horizontally upstream of an air jet ejector in a high altitude relight testing facility. Air was maintained at room temperature for varying pressure, and then liquid nitrogen was introduced to chill the air down to a limit of -50 deg F, corresponding with an altitude of 30,000 feet. Fuel was injected at consistent equivalence ratios across mult (open full item for complete abstract)

Committee: San-Mou Jeng Ph.D. (Committee Chair); Awatef Hamed Ph.D. (Committee Member); Samir Tambe Ph.D. (Committee Member) Subjects: Aerospace Engineering

PhD, University of Cincinnati, 2017, Engineering and Applied Science: Aerospace Engineering

This research validates a physics based model for the thermo-acoustic behavior of a liquid-fueled gas turbine combustor as a tool for diagnosing the cause of combustion oscillations. A single nozzle, acoustically tunable gas turbine combustion rig fueled with Jet-A was built capable of operating in the unsteady combustion regime. A parametric study was performed with the experimental rig to determine the operating conditions resulting in thermoacoustic instabilities. The flame transfer function has been determined for varying fuel injection and flame stabilization arrangements to better understand the feedback loop concerning the heat release and acoustics. The acoustic impedance of the boundaries of the combustion system was experimentally determined. The results were implemented in a COMSOL Multiphysics model as complex impedance boundary conditions at the inlet and exit and a source term to model the flame and heat release. The validity of that model was determined based on an eigenvalue study comparing both the frequency and growth rate of the eigenvalues with the experimentally measured frequencies and pressures of the stable and unstable operating conditions. The model demonstrated that it can accurately predict the instability of the examined operating conditions. The model also closely predicted the frequency of instability and demonstrated the usefulness of including the experimentally determined acoustic boundary conditions over idealized sound hard boundaries.

Committee: Jongguen Lee Ph.D. (Committee Chair); Jay Kim Ph.D. (Committee Member); Kwanwoo Kim Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member) Subjects: Aerospace Materials

PhD, University of Cincinnati, 2014, Engineering and Applied Science: Aerospace Engineering

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

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

MS, University of Cincinnati, 2006, Engineering : Mechanical Engineering

Combustion in a lean pre-mixed (LPM) combustor may become unstable due to small changes in geometry and the manner in which reactants are introduced. This may lead to excessive thermal loads and possible off-design operation. A comprehensive understanding of combustion instability is therefore needed. The present study aims to analyze the flow and flame dynamics in a model LPM gas turbine combustor in LPM combustion. Fluent is used as the flow solver for the present study. The 3-D Navier-Stokes equations are solved along with finite-rate chemical reaction equations and variable thermo-physical properties. Large-eddy-simulation (LES) technique is used to model turbulence. The dynamic version of the Smagorinsky-Lilly model is employed to describe subgrid-scale turbulent motions and their effect on large-scale structures. At first a non-reactive LES was performed in model round and LM6000 combustor. The results for time averaged mean velocity are compared with the previous LES work by Grinstein et al. and Kim et al. Using non-reacting case for LM6000, reactive simulation was initiated, with lean methane-air mixture with equivalence ratio 0.56. Species transport equation is solved for global methane-air two-step reaction with six volumetric species to predict the local mass fraction of each species. The reaction rates that appear as source terms in the species transport equation are computed using finite-rate/eddy-dissipation model, which computes both, the Arrhenius rate and the mixing rate and uses the smaller of the two. It is observed that as the flow enters the chamber, it bifurcates in two shear layers forming a prong like structure. The layers further tend to reattach to the wall at a distance approximately equal to 3D. Counter-clockwise recirculation zones are formed in the corners, whereas clock-wise toroidal vortex structure is formed in the center. The flame is located in between these vortex structures and thus experiences shear-layer instabilities. It is al (open full item for complete abstract)

Committee: Dr. Urmila Ghia (Advisor) Subjects:

PhD, University of Cincinnati, 2004, Engineering : Aerospace Engineering

To achieve single digit NOx emission from gas turbine combustors and prevent the combustion dynamics encountered in Lean Premixed Combustion, it is essential to understand the correlations among emission characteristics, combustion dynamics, and dynamics and characteristics of swirling flow field. The focus of this dissertation is to investigate the emission characteristics and combustion dynamics of multiple swirl dump combustors either in premixing or non-premixed combustion (e.g. Lean Direct Injection), and correlate these combustion characteristics (emissions, combustion instability and lean flammability) to the fluids dynamics (flow structures and its evolution). This study covers measurement of velocity flow field, temperature field, and combustion under effects of various parameters, including inlet flow Reynolds number, inlet air temperature, swirl configurations, downstream exhaust nozzle contraction ratios, length of mixing tube. These parameters are tested in both liquid and gaseous fuel combustions. Knowledge obtained through this comprehensive study is applied to passive and active controls for improving gas turbine combustion performance in the aid of novel sensor and actuator technologies. Emissions and combustion characteristics are shown closely related to the shape and size of central recirculation zone (CRZ), the mean and turbulence velocity and strain rate, and dynamics of large vortical structures. The passive controls, mostly geometry factors, affect the combustion characteristics and emissions through their influences on flow fields, and consequently temperature and radical fields. Air assist, which is used to adjust the momentum of fuel spray, is effective in reducing NOx and depress combustion oscillation without hurting LBO. Fuel distribution/split is also one important factor for achieving low NOx emission and control of combustion dynamics. The dynamics of combustion, including flame oscillations close to LBO and acoustic combustion insta (open full item for complete abstract)

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

MS, University of Cincinnati, 2002, Engineering : Aerospace Engineering

This thesis presents an experimental study on the effect of geometry on the flow structure of experimental swirl-stabilized gas turbine burners. The purpose of this project was to test a number of geometric changes to a burner and determine how these changes affected the flow field in order to evaluate how they might affect a combustion process. Tests were initially conducted in a water channel on several full-scale experimental burner models, which included the use of elliptic outlet geometries and a vortex breakdown stabilizer. Velocity and phase-averaged PLIF measurements are presented for the different burner configurations. More extensive measurements were then conducted in a cold flow air channel on a series of one-quarter scale experimental swirlers. The scale models were designed with different geometries in order to test the effect of different blade angles and spacings. The models were also tested with a vortex breakdown stabilizer and a mixing section located between the swirler outlet and the sudden expansion. A scale model burner with an elliptic outlet was also tested. Stereo PIV measurements are presented for the different configurations. Water channel measurements identified axial oscillation of the vortex breakdown bubble as the primary driving mechanism behind combustion instabilities. Tests showed that the elliptic geometry damped oscillation of the breakdown bubble and the vortex breakdown stabilizer physically anchored the breakdown bubble inside the burner cone. Tests in the air channel showed that wider blade spacing resulted in a reduced swirl number and a weaker vortex breakdown. A wider blade angle was observed to increase the swirl and the strength of the vortex breakdown. Certain configurations of the breakdown stabilizer were seen to increase the strength of the vortex breakdown while other configurations forced it to form in an unstable configuration. The use of a mixing section at the swirler outlet caused the vortex breakdown bubb (open full item for complete abstract)

Committee: Ephraim Gutmark (Advisor) Subjects: