Doctor of Philosophy, The Ohio State University, 2020, Mechanical Engineering

Gas turbine is an integrated part of modern aviation and power generation industry. The thermal efficiency of a gas turbine strongly depends on the turbine inlet temperature (TIT), and the turbine designers are continuously pushing the TIT to a higher value. Due to the increased freedom in additive manufacturing, the complex internal and external geometries of the turbine blade can be leveraged to utilize innovative cooling designs to address some of the shortcomings of current cooling technologies. The sweeping jet film cooling has shown some promise to be an effective method of cooling where the coolant can be brought very close to the blade surface due to its sweeping nature. A series of experiments were performed using a row of fluidic oscillators on a flat plate. Adiabatic cooling effectiveness, convective heat transfer coefficient, thermal field, and discharge coefficient were measured over a range of blowing ratios and freestream turbulence. Results were compared with a conventional shaped hole (777-hole), and the sweeping jet hole shows improved cooling performance in the lateral direction. Numerical simulation also confirmed that the sweeping jet creates two alternating vortices that do not have mutual interaction in time. When the jet sweeps to one side of the hole exit, it acts as a vortex generator as it interacts with the mainstream ow. This prevents the formation of the counter-rotating vortex pair (CRVP) and allows the coolant to spread in the lateral direction. The results obtained from the low speed at plate tests were utilized to design the sweeping jet film cooling hole for more representative turbine vane geometry. Experiments were performed in a low-speed linear cascade facility. Results showed that the sweeping jet hole has higher cooling effectiveness in the near hole region compared to the shaped hole at high blowing ratios. Next, a detailed experimental investigation of sweeping jet film cooling on the suction surface of a near engine scale (open full item for complete abstract)

Committee: Jeffrey Bons (Advisor); James Gregory (Committee Member); Randall Mathison (Committee Member); Ali Ameri (Committee Member) Subjects: Aerospace Engineering; Aerospace Materials; Engineering; Mechanical Engineering

Doctor of Philosophy, The Ohio State University, 2017, Aero/Astro Engineering

Reverse-oriented film cooling, which consists of film cooling holes oriented to inject coolant in the opposite direction of the freestream, is experimentally and numerically investigated. Tests are conducted at various blowing ratios (M = 0.25, 0.5, and 1.0) under both low and high freestream turbulence (Tu = 0.4% and 13%), with a density ratio near unity. The interesting flow field that results from the reverse jet-in-crossflow interaction is characterized using flow visualization, particle image velocimetry, and thermal field measurements. Heat transfer performance is evaluated with adiabatic film effectiveness and heat transfer coefficient measurements obtained using infrared thermography. Adiabatic effectiveness results show that reverse film cooling produces very uniform and total coverage downstream of the holes, with some reduction due to increased freestream turbulence. The reverse film cooling holes are evaluated against cylindrical holes in the conventional configuration, and were found to perform better in terms of average effectiveness and comparably in terms of net heat flux reduction, despite augmented heat transfer coefficient. Compared to shaped hole data from the current study as well as previous literature, the reverse film cooling holes generally exhibited worse heat transfer performance. The aerodynamic losses associated with the film cooling are characterized using total pressure measurements downstream of the holes. Losses from the reverse configuration were found to be higher when compared to cylindrical holes in the conventional and compound angle configurations. To investigate the unsteady three-dimensional flow physics, large eddy simulations were conducted to replicate the experiment at all three blowing ratios, under low and high freestream turbulence. The models were first validated against the experimental measurements, before being used to provide insight into the complicated flowfield associated with the interaction between the reve (open full item for complete abstract)

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

Master of Science, The Ohio State University, 2010, Mechanical Engineering

As gas turbine engines are driven to be more efficient, quiet, and to produce less pollutant the turbine inlet temperature has a tendency to be driven upwards. The life of a turbine engine component decreases dramatically as the metal temperature increases. Because film cooling of high-pressure turbine airfoils has become common practice, improving the ability to predict film-cooling effectiveness is a critical problem of interest. Finding better, more efficient ways to use the cooling air is far preferable to using more of it. However, even if a given cooling-hole configuration proves to be effective in a flat-plate environment (which is the test article of interest in this thesis), it may not be effective on a turbine blade that is exposed to dynamic conditions that cannot be easily replicated. The goal of the experiment reported here is to measure the film effectiveness for a blowing ratio, temperature ratio and free stream Mach number, all similar to those experienced by the pressure surface of a rotating blade turbine blade with the same cooling-hole configuration, but for the flat-plate test article noted above. The cooling gas flow will be initiated earlier than the main flow to allow for proper setup of the cooling flow. This data will be used as a comparison to simulation results obtained using the CFD code Fine TURBO. It is shown in this work that the cooling-gas supply system interaction with the external gas supply associated with the blowdown facility process is not simple, and the current model used to design the experiment is not as good as it could have been. The effect of cooling was observed and the data closely resembled the simulations done using the CFD code Fine TURBO. Unfortunately, due to problems with the double-sided Kapton heat-flux gauges, heat flux data was not obtained in the immediate vicinity of the cooling holes. Solutions to the problems encountered in this experiment are relatively straightforward and are presented.

Committee: Michael Dunn Prof. (Advisor); Charles Haldeman Prof. (Committee Member) Subjects: Mechanical Engineering

Master of Science (M.S.), University of Dayton, 2022, Aerospace Engineering

Rotating Detonation Engines (RDEs) and pressure gain combustion (PGC) present a pathway to increased performance and fuel savings due to improved thermal efficiency and power density. RDEs utilize detonations to combust reactants, which provides higher thermal efficiencies than deflagration combustion. This increase in efficiency comes from increases in total pressure achieved across the detonation front, whereas deflagrations produce losses in total pressure. However, high thermal loads have limited uncooled and conventionally manufactured RDE test duration. Currently there is a need to develop novel cooling schemes that minimize the associated performance penalty, provide adequate cooling to extend test duration, and characterize changes in RDE performance and operability. This investigation was aimed at quantifying film cooling when applied to the unsteady and adverse pressure gradient of a RDE. Two film cooled outer-body combustion liners were manufactured and tested using a H2-air operated 6-inch RDE with an aerospike plug nozzle, heat sink center-body, and a 0.64 inch detonation channel width. Additionally, a control liner without holes was manufactured and tested. The two film cooled liners varied film pressure drop to characterize changes in RDE operability, temperature response, and cooling manifold pressure unsteadiness. All liners used approximately equivalent total flow area, as well as diameter weighted axial and circumferential spacing to allow comparison. The combustion air injection area ratio was set to 0.33, and the nozzle area ratio set to 1.0 and 0.66 relative to the channel area. Combustion air manifold pressures, cooling air manifold pressures, cooling air temperature, combustion liner temperature, operating mode, detonation stability, and detonation wave speed were analyzed for an array of combustion air mass flow rates, equivalence ratios, cooling air mass flow rates, and liner geometries. A high-speed camera was utilized to confirm operating (open full item for complete abstract)

Committee: Matthew Fotia (Committee Chair); Frederick Schauer (Committee Member); Adam Holley (Committee Member) Subjects: Aerospace Engineering; Mechanical Engineering

Doctor of Philosophy, The Ohio State University, 2022, Aerospace Engineering

Gas turbine engines are an essential technology in aviation and power generation. One of the challenges associated with increasing the efficiency of gas turbines is the thermal loading experienced by the engine components downstream of the combustors especially the high-pressure turbine blades. High temperatures and rotational velocities can cause blade failures in numerous ways such as creep or stress rupture. Technologies like film cooling are implemented in these components to lower the thermal loading and reduce the risk of failure. However, these introduce complexities into the flow which in turn increases the difficulty of predicting the performance of film cooled turbines. Accurately predicting the capabilities of these components is essential to prevent failure in gas turbine engines. Engineers use a combination of experiments and computational simulations to understand how these technologies perform and predict the operating conditions and lifespan of these components. A combined experimental and numerical program is performed on a single stage high-pressure turbine to increase understanding of film cooling in gas turbines and improve computational methods used to predict their performance. The turbine studied is a contemporary production model from Honeywell Aerospace with both cooled and uncooled turbine blades. The experimental work is performed at The Ohio State University Gas Turbine Laboratory Turbine Test Facility, a short duration facility operating at engine corrected conditions. The experiments capture heat flux, temperature, and pressure data across the entire blade, but this work will focus on the turbine blade tip data. Tip temperature data are captured using a high-speed infrared camera providing a unique data set unseen in the current literature. In addition to the experiments, transient conjugate heat transfer simulations of a single turbine passage are performed to recreate the experiments and give insight into the flow field in the tip (open full item for complete abstract)

Committee: Randall Mathison (Advisor); Sandip Mazumder (Committee Member); Michael Dunn (Committee Member); Jeffrey Bons (Committee Member) Subjects: Aerospace Engineering; Mechanical Engineering

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

Gas turbine performance is highly dependent on turbine inlet temperature, which often exceeds the working limitations of the materials involved. Film cooling is a widely used technology enabling highly efficient gas turbine cycles, where relatively cold air is injected as a film on the airfoil surfaces protecting the airfoils from the hot combustion gasses. Film cooled turbines exist in highly unsteady environments due to interactions between stationary and rotating components, and film cooling further complicates the flow. There is limited understanding of the unsteady nature of film cooling flows, resulting in limited ability to predict heat transfer and metal temperature on the components of a gas turbine. The goal of this work is to increase understanding of turbine cooling technology by examining time-accurate and time-averaged behaviors of the cooling flows. This dissertation incorporates experimental and computational analysis of pressure and heat transfer on an industry scale high-pressure turbine stage. Experimental measurements of pressure and heat transfer were performed on a turbine stage installed in the Turbine Test Facility at the Gas Turbine Laboratory. This facility is uniquely equipped to examine unsteady pressure and heat transfer on turbine stages operating at design corrected conditions. Heat transfer measurements are compared for multiple different cooling configurations on the rotating airfoils. Data are analyzed on time-averaged and time-resolved bases, and the results highlight cooling benefit differences among the various cooling hole shapes and coolant flow rates. Computational models of the turbine stage are also employed with steady and unsteady RANS modeling techniques. Experimental data are used for boundary conditions in the computational models as well as to evaluate the accuracy of the models. Comparisons of experimental and steady computations of film cooled turbines often result in poor agreement due to the complexity of film co (open full item for complete abstract)

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

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

Gas turbine engines are one of the most widely used propulsion systems and are commonly encountered in commercial and defense aircraft.Increasing the turbine inlet temperatures is one of the convenient approaches to increase the efficiency of the engine, but this process is restricted by the material threshold temperatures and the structural integrity of the turbine blade. In order to keep the blade temperatures within limits for seamless working of the engine, various cooling techniques are employed. One such technique is film cooling using holes wherein the air from the compressor is bled and fed into the film cooling holes. A part of this computational study on film cooling involves the numerical validation of the LES computational technique with Dynamic Smagorinsky model to predict the adiabatic mixing phenomenon of the cold air exiting from the 7-7-7 baseline shaped hole with the mainstream hot air for the BR and DR values of 1.5 against the experimental data provided by PSU. A single hole from the experimental array of 5 holes is considered with periodic boundary conditions and the parameters of comparison are the centerline and lateral averaged adiabatic effectiveness. The non-dimensional velocity parameter shows good comparison with the PIV measurements in the boundary layer at the leading and the trailing edges of the hole and the flat plate adiabatic effectiveness after compound averaging show decent comparison with the Infrared image data. The CRVP structure is weak in this case as expected for a laid-back fan shaped hole and the centerline and lateral averaged adiabatic effectiveness as compared to the experimental data lies within 25% deviation. The overall accuracy of the solution based on the lateral average adiabatic effectiveness is calculated to be 74.46% for this perfectly adiabatic case. The second part of this computational study involved employing the RANS technique with the K-Omega turbulence model to simulate the multiple perforation cy (open full item for complete abstract)

Committee: Mark Turner Sc.D. (Committee Chair); Milind Jog Ph.D. (Committee Member); Jay Kim Ph.D. (Committee Member); Paul Orkwis Ph.D. (Committee Member) Subjects: Aerospace Materials

Master of Science, The Ohio State University, 2018, Aero/Astro Engineering

The effect of particle size on particle accumulation within an effusion cooling geometry common to gas turbines was investigated experimentally and computationally. A flat plate with an effusion cooling hole array based on a gas turbine combustor liner was subjected to particulate laden flow in an accelerated deposition facility. The tests were conducted at an engine relevant plate temperature of 1118 K, coolant temperature of 950 K, and held at a constant pressure ratio of 1.03 (cavity to ambient). To elucidate the effect of particle size, six unique size distributions of Arizona Road Dust (ARD) smaller than 20 micron were introduced to the flow independently. Experiments were also conducted in which two different dust size distributions were sequentially delivered to the test article. These experiments clearly indicate that the smallest particles within the range tested accumulate within the hole creating deposit structures and blocking the effusion holes. They also indicate that larger particles within this range can have an erosive effect on the deposit structures as they build, changing the structure's morphology and blockage behavior. Computational fluid domains were developed to replicate the test article geometry before and after a test to investigate the effect that deposit structures have on deposition. Particles were introduced to these domains after reaching a steady solution and were tracked through the solution with a Lagrangian trajectory solver. The impact locations of the particles were recorded and a particle sticking model was employed to determine if the particles stick or rebound. The domain with the clean hole showed that the smallest particles impact and were prone to sticking in the area where deposits form experimentally. As the particles increase in size, the number of impacts and likelihood of sticking decreased. The domain with scoop deposit structures showed that these structures can change the particle impact trajectories influenci (open full item for complete abstract)

Committee: Jeffrey Bons PhD (Advisor); Jen-Ping Chen PhD (Committee Member) Subjects: Aerospace Engineering

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

Film cooling effectiveness can be measured by thermal and mass transfer analysis. We have designed and developed a flat plate adiabatic wind tunnel facility to study film cooling effectiveness of selected film hole geometries by mass transfer analysis using a heavy gas (CO2) as a substitute for the coolant, measuring the mixing by both gas sampling and full field measurements by an optical measurement technique using pressure/oxygen sensitive paint. The flat plate adiabatic wind tunnel consists of a 24” long by 1.5” x 4” rectangular duct with a 0.5” radiused inlet mounted on a 15.5” ID x 24.5” long mainstream air plenum. Air was supplied to one end of the mainstream air plenum and forced through flow straightening components such as a perforated metal plate and honeycomb layers exiting through the wind tunnel.Cooling flow (CO2) was injected into the mainstream air through film cooling holes in test coupons that are attached to the top surface of 1.5” x 4” x 3” high cooling plenum, mounted on an opening in the bottom surface of the wind tunnel, 5” from the duct inlet. The test coupons are 3” x 5.5” x 0.375” thick with a centrally reduced pocket of 2” x 4” x 0.17” thickness. The film cooling holes are of 0.1” diameter. The tests were run for three cases of round (cylindrical), inclined film cooling holes (25°, 30° and 35°) at a coolant to mainstream density ratio of 1.5 and a mainstream Mach number, M# of 0.14 with a Blowing ratio, M ranging from 0.5 to 2.0 in increments of 0.25. The mass concentrations of the coolant gas were measured at discrete locations downstream of coolant injection by bleeding gas samples through instrumentation taps and passing them through a gas (CO2) analyzer. Film cooling effectiveness was then calculated by mass transfer analysis using the coolant gas concentrations. The test facility was validated by comparing the results with published data. The effects of blowing ratio and other flow parameters on film cooling effectiveness were studied (open full item for complete abstract)

Committee: Ephraim Gutmark Ph.D. (Committee Chair); Shaaban Abdallah Ph.D. (Committee Member); Jeffrey Kastner Ph.D. (Committee Member) Subjects: Aerospace Materials

Master of Science, The Ohio State University, 2016, Aero/Astro Engineering

An experimental investigation has been performed to verify heat flux measurements on a metallic film-cooled flat plate by using infrared thermography in a transient facility. Infrared thermography provides high-resolution temperature distributions and makes it easy to locate hot spots on the object of interest. Previous work has shown that infrared thermography can produce accurate measurement for an uncooled flat plate. The goal of this thesis is to show that infrared thermography can also measure heat flux for film-cooled components and lay the foundation for its use in full-scale rotating experiments. A stainless steel plate with rows of cooling holes was built for testing in a blowdown facility. During the experiment, the plate was exposed to hot main flow and supplied with low temperature air through the cooling holes. The heat flux on the plate surface was determined by performing a 3D ANSYS transient thermal analysis using the infrared temperature distribution as time dependent top surface boundary conditions. The challenge of this technique for the film-cooled plate is that the boundary conditions are not known for the cooling channel walls and backside walls that exposed to cooling air since it is not practical to obtain measurements in those regions. Although strong jet impingement and forced convection are taking place on these walls, they are treated as adiabatic for the numerical analysis and the analysis time window is kept short so that through-wall conduction will not affect the top surface during the run. It is shown that the through-wall conduction from the backside and cooling channel walls only impact the regions right upstream the cooling holes and only after a relatively long run time. The fidelity of this technique is verified by comparing the results calculated based on the infrared images to the results obtained from traditional heat-flux gauges. The infrared thermography and the heat-flux gauges measurements agree within (open full item for complete abstract)

Committee: Randall Mathison Dr (Advisor); Michael Dunn Dr (Committee Member); Herman Shen Dr (Committee Member) Subjects: Aerospace Engineering

Doctor of Philosophy, The Ohio State University, 2016, Mechanical Engineering

An experiment was performed at The Ohio State University Gas Turbine Laboratory for a film-cooled high-pressure turbine stage operating at design-corrected conditions, with variable rotor and aft purge cooling flow rates. Several distinct experimental programs are combined into one experiment and their results are presented. Pressure and temperature measurements in the internal cooling passages that feed the airfoil film cooling are used as boundary conditions in a model that calculates cooling flow rates and blowing ratio out of each individual film cooling hole. The cooling holes on the suction side choke at even the lowest levels of film cooling, ejecting more than twice the coolant as the holes on the pressure side. However, the blowing ratios are very close due to the freestream massflux on the suction side also being almost twice as great. The highest local blowing ratios actually occur close to the airfoil stagnation point as a result of the low freestream massflux conditions. The choking of suction side cooling holes also results in the majority of any additional coolant added to the blade flowing out through the leading edge and pressure side rows. A second focus of this dissertation is the heat transfer on the rotor airfoil, which features uncooled blades and blades with three different shapes of film cooling hole: cylindrical, diffusing fan shape, and a new advanced shape. Shaped cooling holes have previously shown immense promise on simpler geometries, but experimental results for a rotating turbine have not previously been published in the open literature. Significant improvement from the uncooled case is observed for all shapes of cooling holes, but the improvement from the round to more advanced shapes is seen to be relatively minor. The reduction in relative effectiveness is likely due to the engine-representative secondary flow field interfering with the cooling flow mechanics in the freestream, and may also be caused by shocks and other compr (open full item for complete abstract)

Committee: Randall Mathison (Advisor); Michael Dunn (Committee Member); Sandip Mazumder (Committee Member); Jeffrey Bons (Committee Member) Subjects: Aerospace Engineering; Engineering; Mechanical Engineering

MS, University of Cincinnati, 2016, Engineering and Applied Science: Aerospace Engineering

To investigate the viability of using additive manufacturing technology for flat plate film cooling experiments a new experiential facility was constructed using gas analysis and oxygen sensitive paint as a method of measuring and characterizing film cooling effectiveness for various additive manufacturing technologies as well as aluminum. The ultimate objective of this work is to assess whether these technologies can be a replacement for traditional aluminum CNC machining. Film Cooling Effectiveness is closely dependent on the geometry of the hole emitting the cooling film. These holes are sometimes quite expensive to machine by traditional methods so 3D printed test pieces have the potential to greatly reduce the cost of film cooling tests. What is unknown is the degree to which parameters like layer resolution and the choice of 3D printing technologies influence the results of a film cooling test. A new flat-plate film cooling facility employing the mass transfer analogy (introduction of foreign gas as coolant, not to be confused with the sublimation method) and measurements both by gas sample analysis and oxygen-sensitive paint is first validated using gas analysis and oxygen sensitive paint cross correlation. The same facility is then used to characterize the film cooling effectiveness of a diffuser shaped film cooling hole geometry. These diffuser holes (film hole diameter, D of 0.1 inches) are then produced by a variety of different manufacturing technologies, including traditional machined aluminum, Fused Deposition Modeling (FDM), Stereo Lithography Apparatus (SLA) and PolyJet with layer thicknesses from 0.001D (25 µm) to 0.12D (300 µm). Tests are carried out at mainstream flow Mach number of 0.30 and blowing ratios from 1.0 to 3.5. The coolant gas used is CO2 yielding a density ratio of 1.5. Surface quality is characterized by an Optical Microscope that calculates surface roughness. Test coupons with rougher surface topology generally showed dela (open full item for complete abstract)

Committee: Ephraim Gutmark Ph.D. D.Sc. (Committee Chair); Pepe Palafox Ph.D. (Committee Member); David Munday Ph.D. (Committee Member); Mark Turner Sc.D. (Committee Member) Subjects: Aerospace Materials

Master of Science, The Ohio State University, 2015, Aero/Astro Engineering

The development of flow blockage by particulate accumulation in the internal flow passages of a gas turbine double wall cooling scheme was studied experimentally. This parametric investigation focused on the effects of particle concentration, flow temperature, and particle size on the deposition characteristics in a cylindrical impingement/film cooling geometry. The impingement and film cooling hole layout is based on the leading edge cooling scheme of a modern nozzle guide vane (NGV). Tests were run at a constant pressure ratio of 1.02 (cavity pressure to exhaust) and the mass flow rate was permitted to decrease throughout the test as the cooling passages became obstructed. Particulate concentration was varied by holding the mass injected constant while adjusting the test injection time and rate. Particles consisted of Arizona Road Dust with distributions of 0-5, 0-10, and 0-20 µm. Flow blockage increased by 4% over a range of two orders of magnitude in particulate concentration for the smallest particle size distribution. At 452 °C the blockage levels increased to nearly four times that of the ambient conditions. Similar amounts of particulate deposited on the film cooling wall at ambient and high temperature, but the high temperature particulate caused greater blockage to the film holes. The effect of particle size was difficult to discern due to clumping of the smallest particles into large agglomerations. This clumping effect was coupled with the trend of increasing temperature. Implications for continued internal deposition research are discussed.

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

MS, University of Cincinnati, 2009, Engineering : Mechanical Engineering

The present study is a three-dimensional numerical investigation of the effectiveness of film cooling for a turbine blade leading-edge model with both a single and a three-hole cooling configuration. The model used has the same dimensions as those in the experimental investigation of Ou and Rivir (2006). It consists of a half cylinder with a flat after-body, and well represents the leading edge of a turbine blade. The single coolant hole is situated approximately at the spanwise center of the cylindrical model, and makes an angle of 21.5 degrees to the leading edge and 20 degrees to the spanwise direction. For the three-hole configuration, the center hole is positioned the same as the single hole in the single-hole configuration, with the adjacent holes located at a spanwise distance of 37.4 mm on either side of the center hole. Multi-block grids were generated using GridGen, and the flows were simulated using the flow solver Fluent. A highly clustered structured C-grid was developed around the leading edge of the model. The outer unstructured-grid domain represents the wind tunnel as used in the experimental study of Ou and Rivir (2006), and the leading-edge model is located at the center of the domain. Simulations were carried out for blowing ratios, M, ranging from 0.75 to 2.0. Turbulence was represented using the k-? shear-stress transport (SST) model, and the flow was assumed to have a free-stream turbulence intensity of 0.75%. Two types of boundary conditions were used to represent the blade wall: an adiabatic surface, and a conductive surface. The adiabatic-wall results over-predicted the film-cooling effectiveness in the far downstream region for low blowing ratios. Also, in the vicinity of the cooling hole, an increase in blowing ratio resulted in higher film cooling effectiveness than observed in the experiments. It should be noted that the steady RANS-based turbulence model used under-predicts the interaction between the coolant and mainstream flow near (open full item for complete abstract)

Committee: Urmila Ghia PhD (Committee Chair); Karman Ghia PhD (Committee Member); Milind Jog PhD (Committee Member) Subjects: Engineering

MS, University of Cincinnati, 2008, Engineering : Mechanical Engineering

The present study numerically investigates the effectiveness of film cooling on a turbine blade leading-edge model through a single-hole coolant exit. The model used in this study has the same dimensions as that of an earlier experimental investigation by Ekkad et al., 2006. The cylindrical model, a half cylinder with a flat after-body attached to it, provides a good representation of the leading edge of a turbine blade. The coolant hole is situated approximately at the center of the cylindrical model along the spanwise direction and makes an angle of 21.5 degrees to the true leading edge and 20 degrees to the spanwise direction. A multi-block grid is generated using GridGen, and the flow is simulated using the flow solver FLUENT. Blowing ratio, M, is defined as the mass flux ratio of the coolant and the mainstream. Simulations are carried out for different blowing ratios M, ranging from 0.25 to 1.0, with the k-ε realizable turbulence model and the k-ω SST model. The flow is assumed to have a free-stream turbulence intensity of 0.75%. Additionally, the enhanced wall function approach is used as the near-wall treatment in the computational model for the simulation with k-ε realizable turbulence model. Results obtained indicate an increased film-effectiveness for low blowing ratio in the far downstream region. Also, in the vicinity of the coolant jet exit, it is observed that increase in blowing ratio increases the film cooling effectiveness. It is also concluded that steady RANS-based turbulence models under-predict the interaction between the jet and the mainstream at the jet exit, and the spreading of coolant downstream.

Committee: Dr. Urmila Ghia (Advisor) Subjects: Engineering, Mechanical

MS, University of Cincinnati, 2007, Engineering : Mechanical Engineering

The application of interest is the cooling of turbine blades in large gas combustion engines where hot gases from the combustor cause thermal deterioration of the metal turbine blades. A thin-film of coolant flow buffers the hottest parts of the blade surface. Heat transfer on a bluff body and, subsequently, a single-hole cooling problem is solved numerically in two-dimensions. The flow is assumed to be incompressible, and the laminar, steady Navier-Stokes equations are used to obtain the flow solution. Results for the bluff-body heat transfer agree very well with experimental data up to the separation point, and are within 20% of the data thereafter. The film-cooling simulation yielded higher cooling effectiveness due in large part to the use of the two-dimensional model, which treats the hole as a slot with higher coolant mass. Results from the simulations indicate that the Cobalt flow solver is capable of solving complex heat transfer problems.

Committee: Dr. Urmila Ghia (Advisor) Subjects: Engineering, Mechanical

Master of Science, The Ohio State University, 2012, Aero/Astro Engineering

An accelerated deposition test facility was used to study the relationship between film cooling, surface temperature, and particle temperature at impact on deposit formation. Tests were run at gas turbine representative inlet Mach numbers (0.1) and temperatures (1090°C). Deposits were created from lignite coal fly ash with median diameters of 1.3 and 8.8µm. Two CFM56-5B nozzle guide vane doublets, comprising three full passages and two half passages of flow, were utilized as the test articles. Tests were run with different levels of film cooling back flow margin and coolant temperature. Particle temperature upon impact with the vane surface was shown to be the leading factor in deposition. Since the particle must traverse the boundary layer of the cooled vane before impact, deposition is directly affected by the film and metal surface temperature as well. Film coolant jet strength showed only minor effect on deposit patterns on the leading edge. However, larger Stokes number (resulting in higher particle impact temperature) corresponded with increased deposit coverage area on the shower head region. Additionally, infrared measurements showed a strong correlation between regions of greater deposits and elevated surface temperature on the pressure surface. Thickness distribution measurements also highlighted the effect of film cooling by showing reduced deposition immediately downstream of cooling holes. A set of secondary tests were also conducted to briefly study the effect of Stokes number on leading edge deposition with no cooling, in order to support conclusions from the primary tests. It was found that larger Stokes number led to an increase in rate of deposition due to a greater number of particles being able to follow their inertial trajectories and impact the vane. Implications for engine operation in particulate-laden environments are discussed.

Committee: Jeffrey Bons PhD (Advisor); Micheal Dunn PhD (Committee Member) Subjects: Aerospace Engineering; Aerospace Materials; Engineering

Master of Science, The Ohio State University, 2011, Mechanical Engineering

For many years turbine designers have utilized advancements in film-cooling technology to allow for increased high-pressure turbine inlet temperature. Prediction tools, used to predict the cooling effectiveness of the representative cooling-hole and cooling-hole pattern designs have been successful in keeping the engines on wing for a large number of operational hours, but there is room for and a desire for improvement in the technology. Therefore, a study was undertaken at the OSU GTL to find a way to obtain basic data needed to help improve CFD prediction capability. The particular facility utilized for this work is a medium-duration blowdown facility to which significant improvements in the operational procedure of the cooling system have been made for the purposes of this work and both the facility and the improvements will be described in detail in this thesis. In order to keep the CFD validation simple, a flat plate configuration with a realistic cooling hole pattern, representative of a high-pressure turbine blade for which measurements obtained as part of a full-stage experiment were obtained, was utilized along with flow properties of current interest to the industry. The measurements reported in this thesis yielded high response heat-flux measurements along the axial direction of the plate, including locations between the individual rows of cooling holes. The influence of Reynolds numbers on heat transfer to the plate was also explored. Lastly, the temperature of the main flow and the test section walls were varied to determine the effect of cooling on the local adiabatic wall temperature.

Committee: Michael Dunn PhD (Advisor); Charles Haldeman PhD (Committee Member) Subjects: Aerospace Engineering; Mechanical Engineering

Master of Science, The Ohio State University, 2011, Aero/Astro Engineering

A thesis is presented that investigates the effect of particle Stokes number and film cooling on the character of coal fly ash deposition on a turbine nozzle guide vane. The Turbine Reacting Flow Rig or TuRFR at The Ohio State University was used to produce coal fly ash deposits on real turbine hardware at operating conditions. CFM56-5B nozzle guide vane doublets were subjected to inlet temperatures of 1080 °C and a Mach number of 0.08 while seeding the flow with a sub-bituminous coal fly ash. The ash was processed to provide two different size distributions, that with a median Stokes number of 0.3 and Stokes number 4.0 and each ash was exposed to a vane set with and without film cooling. The transient character of deposit growth was investigated by a camera positioned to view the vanes during test time. Post-test measurements included using sophisticated metrology techniques to provide plots of deposit thickness and structure. The results were then compared to computation. Deposits thickness was observed to be a large function of particle loading but in general small Stokes number ash deposits were observed to be half the thickness of the large Stokes number deposits for a given test condition. For those tests which involved film cooling, deposits only formed on the leading 50% of the vane pressure surface while those tests without film cooling had deposits on the entire pressure surface. Deposit location is thus observed to be a strong function of vane surface temperature. Values of average surface roughness and peak to peak roughness were calculated for all tests. Film cooling was found to have a negligible effect on surface roughness while increased Stokes number was found to double the calculated roughness. The computational results were found to accurately depict initial deposit location, for both the un-cooled and cooled cases, but lacked the ability to accurately represent deposit evolution over time.

Committee: Jeffrey Bons PhD (Advisor); James Gregory PhD (Committee Member) Subjects: Aerospace Engineering; Alternative Energy; Engineering; Experiments

Master of Science, The Ohio State University, 2010, Aero/Astro Engineering

The magnitude of the temperature increase as a result of compression heating during the starting process of a high Reynolds number film cooling experiment was more than initially anticipated, creating a mismatch in the design conditions. A review of the time-accurate data showed that two fluid mechanisms, ingestion into the cooling holes during the starting process and compression heating were the causes of the problem, but in different amounts based on the experimental apparatus and time scales involved. The goal of this thesis is to use one of the facilities, the Small Calibration Facility, in conjunction with an analytical model developed for this task to determine how to properly change the start procedure to rectify the problem. The solution to the problem involved adding extra cooling mass at specific times. Several methods attempted, and this particular approach yielded excellent results. In addition, the rework of the start-up procedure allowed for better tuning of the blowdown facility, which resulted in a more constant blowing ratio throughout the experiment.

Committee: Mike Dunn (Advisor); Charles Haldeman (Advisor) Subjects: Mechanical Engineering