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Kheniser, Issam E.Film Cooling Experiments in a Medium Duration Blowdown Facility
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

Keywords:

Blowdown; Film; cooling; film-cooling

Davis, Shanon MarieHeat-Flux Measurements for a Realistic Cooling Hole Pattern and Different Flow Conditions
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

Keywords:

film cooling; flat plate; heat transfer; heat-flux; engine design; CFD; medium-duration; blowdown;

Peterson, Blair AA Study of Blockage due to Ingested Airborne Particulate in a Simulated Double-Wall Turbine Internal Cooling Passage
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

Keywords:

Gas Turbine; Deposition; Film Cooling; Impingement Cooling; Particulate Ingestion; Flow Blockage; Internal Cooling; Double-Wall;

Kahveci, Harika SenemThe Influence of Film Cooling and Inlet Temperature Profile on Heat Transfer for the Vane Row of a 1-1/2 Stage Transonic High-Pressure Turbine
Doctor of Philosophy, The Ohio State University, 2010, Mechanical Engineering
The goal of this research was to establish an extensive database for typical engine hardware with a film-cooled first stage vane, which represents the foundation for future turbomachinery film cooling modeling and component heat transfer studies. Until this time, such a database was not available within the gas turbine industry. Accordingly, the study focuses on determination of the local heat flux for the airfoil and endwall surfaces of the vane row of a fully-cooled turbine stage. The measurements were performed at the Ohio State University Gas Turbine Laboratory using the Turbine Test Facility. The full-scale rotating 1 and 1/2 turbine stage is operated at the proper corrected engine design conditions: Flow Function (FF), corrected speed, stage Pressure Ratio (PR), and temperature ratios of gas to wall and gas to coolant. The primary measurements of temperature, pressure, and heat flux are repeated for different vane inlet temperature profiles and different vane cooling flows to establish an understanding of the influence of film cooling on local heat transfer. Double-sided Kapton heat-flux gauges are used for heat-flux measurements at different span locations along the airfoil surfaces and along the inner endwall. The cooling scheme consists of numerous cooling holes located on the endwalls, at the airfoil leading edge, on the airfoil pressure and suction surfaces, and at the trailing edge, resulting in a fully cooled first stage vane. The unique film-cooled endwall heat transfer data demonstrated in contour plots reveals insight to the complex flow behavior that is dominant in this region, which becomes even more complicated with the addition of coolant. Varying profile shapes resulted in significant heat transfer variations in a growing fashion towards the trailing edge region, which increased in magnitude when there is no coolant supply. The largest cooling effect is observed on 5% span pressure surface and at the inner endwall region. Heat transfer decreases from tip towards hub with addition of cooling. However, a similar decrease is not observed at the inner endwall region by doing so, which suggests excess coolant once beyond an optimum blowing ratio. Cooling flow rate and temperature profile shape affect the distributions on the airfoil surface very similarly, the latter observed more clearly at the endwall region. The vane outer cooling effect is comparable to the combined coolant effect at all surfaces, while no impact of purge flow is observed. Aligning the hot streaks with the vane leading edge lowered heat transfer compared to mid-passage alignment at the mid-span suction surface and through the endwall passage, and increased it at the endwall exit, while the pressure surface is found to be insensitive to this switch. Comparison with a previous research program with the un-cooled version of the vane gave good agreement on the pressure surface and at the endwall, but significantly lower heat transfer on the suction surface due to ingestion of the hot flow through the cooling holes when there is no cooling.

Committee:

Michael G. Dunn, PhD (Advisor); Charles W. Haldeman, PhD (Committee Member); J. William Rich, PhD (Committee Member); Mohammad Samimy, PhD (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

Film cooling; rotating rig; heat transfer; high-pressure vane

LITZLER, JEFFREY WCOMPUTATIONAL SIMULATION AND ANALYSIS OF FILM COOLING FOR THE LEADING-EDGE MODEL OF A TURBINE BLADE
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

Keywords:

film-cooling; cooling effectiveness; leading-edge; bluff-body; heat transfer

Aghasi, Paul P.Dependence of Film Cooling Effectiveness on 3D Printed Cooling Holes
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 delayed film hole blow off and higher film cooling effectiveness at increased blowing ratios compared to the geometries with lower measured surface roughness.

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

Keywords:

Film Cooling;Heat transfer;mass transfer;pressure sensitive paint;roughness;3D printing

Webb, Joshua J.The Effect of Particle Size and Film Cooling on Nozzle Guide Vane Deposition
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

Keywords:

Deposition; Film Cooling; Roughness; Stokes Number

Stenger, DouglasThree-Dimensional Numerical Simulation of Film Cooling on a Turbine Blade Leading-Edge Model
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 the cooling-pipe exit. The conductive-wall results show a much closer agreement with experimental data for film effectiveness as compared to the adiabatic-wall predictions. Simulations were also performed with higher values of turbulence intensity at the cooling-hole inlet, and these predicted the coolant-mainstream interaction and the film-cooling effectiveness more accurately.

Finally, a novel concept of pulsing the coolant flow was implemented so as to achieve film-cooling effectiveness equivalent to that with constant cooling, but with reduced overall coolant air, thereby enhancing turbine efficiency. Pulsed cooling with pulsing frequency PF = 5 and 10Hz, and duty cycle DC = 50%, shows the greatest cooling effects. The three-hole cooling results indicate that the 49 mm spanwise distance used for computing the spanwise-averaged values for film-cooling effectiveness accounts for all of the film-coolant spreading provided by the single hole. Also, the neighboring cooling holes contribute little film cooling to the 49 mm spanwise distance. The most significant new finding in this work is that the inclusion of wall conductance is the main factor responsible for reproducing the experimental data.

Committee:

Urmila Ghia, PhD (Committee Chair); Karman Ghia, PhD (Committee Member); Milind Jog, PhD (Committee Member)

Subjects:

Engineering

Keywords:

Turbine Engine; CFD; Fluent; Film Cooling; Simulation

Chen, LiangInfrared Thermography Technique for Measuring Heat Transfer to a Film Cooled Object
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 10% for the regions that are not affected by backside and cooling channel walls boundary conditions. When the boundary conditions and through-wall conduction do impact the results, the adiabatic wall assumption causes the infrared thermography to predict a lower heat flux than the heat-flux gauges. After the method had been clearly demonstrated, it was used to make comparisons between cooled and uncooled cases to clearly identify how the film cooling spreads and mixes from the rows of cooling holes.

Committee:

Randall Mathison, Dr (Advisor); Michael Dunn, Dr (Committee Member); Herman Shen, Dr (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Infrared thermography; Heat transfer; Film Cooling

Boehler, Michael DavidTransient Aerothermodynamics of Flow Initialization for a Flat Plate Film Cooling Experiment in a Medium Duration Blowdown Wind Tunnel Facility
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

Keywords:

film cooling; blowdown; flat plate; compression heating; ingestion

Bonilla, Carlos HumbertoThe Effect of Film Cooling on Nozzle Guide Vane Ash Deposition
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

Keywords:

Deposition;film cooling;ash deposition;nozzle guide vane;deposit;heat transfer;turbine deposition;engine deposition

Nickol, Jeremy BHeat Transfer Measurements and Comparisons for a Film Cooled Flat Plate with Realistic Hole Pattern in a Medium Duration Blowdown Facility
Master of Science, The Ohio State University, 2013, Mechanical Engineering
Recent advances have for the first time made it possible to perform experiments using fully-cooled rotating turbine stages operating at design-corrected conditions. These experiments produce realistic measurements of the time-averaged and time-accurate local heat transfer for the stage. Despite the value of these results, however, it is still not possible to provide full-coverage measurements due to the difficulties inherent in the view angle and instrumentation of a rotating turbine. Computational simulations show promise in filling in these gaps, but these calculations are not yet able to accurately model the complicated flow physics of these experiments. Experiments utilizing simplified geometries, the most common being a flat plate with a single row of cooling holes, have been run for decades and reported in the literature in literally thousands of papers. They have given valuable insight regarding the more fundamental aspects of film cooling, and the results of these works are used by the industry (with some proprietary modifications) in the design of current engine hardware. These experiments succeed in showing the influence of individual variables with a good deal of detail and clarity; however, due to excessive geometric simplification, the results of these studies fail to capture important flow characteristics observed in data from rotating experiments. In an attempt to help bridge the gap between the common single and double-row flat-plate experiments and the fully cooled rotating experiments, a flat-plate experiment is performed using a medium-duration blowdown facility with a real cooling hole pattern representative of the pressure side of a high-pressure turbine blade. A computational prediction for the experiment is also performed and presented. New heat transfer and film cooling performance parameters are defined for use in the unsteady blowdown facility. Measurements are made to investigate the influence of blowing ratio spanning the range of interest to industry. These results are compared to both simplified hole pattern experiments and rotating experiments to help quantify the effect of the simplified patterns used in the majority of film cooling experiments. Finally, lessons for future experimentation at The Ohio State University Gas Turbine Laboratory are outlined.

Committee:

Michael Dunn, Ph.D (Advisor); Mohammad Samimy, Ph.D (Committee Member)

Subjects:

Aerospace Engineering; Mechanical Engineering

Keywords:

film cooling; flat plate; heat transfer; realistic hole pattern; CFD; gas turbine

Kartuzova, Olga ValeryevnaA computational study for the utilization of jet pulsations in gas turbine film cooling and flow control
Doctor of Engineering, Cleveland State University, 2010, Fenn College of Engineering

Jets have been utilized in various turbomachinery applications in order to improve gas turbines performance. Jet pulsation is a promising technique because of the reduction in the amount of air removed from compressor, which helps to increase turbine efficiency. In this work two areas of pulsed jets applications were investigated, first one is film cooling of High Pressure Turbine (HPT) blades and second one is flow separation control over Low Pressure Turbine (LPT) airfoil using Vortex Generator Jets (VGJ)

The inlet temperature to the HPT significantly affects the performance of the gas turbine. Film cooling is one of the most efficient methods for cooling turbine blades. This technique is simply employing cool air discharged from rows of holes into the hot stream. Using pulsed jets for film cooling purposes can help to improve the effectiveness and thus allow higher turbine inlet temperature without affecting the blade's life. Engine cost will thus be reduced by providing the same capacity from smaller, lighter engines. Fuel consumption will be lowered, resulting in lower fuel cost. Effects of the film hole geometry, blowing ratio and density ratio of the jet, pulsation frequency and duty cycle of blowing on the film cooling effectiveness were investigated in the present work.

As for the low-pressure turbine (LPT) stages, the boundary layer separation on the suction side of airfoils can occur due to strong adverse pressure gradients. The problem is exacerbated as airfoil loading is increased. If the boundary layer separates, the lift from the airfoil decreases and the aerodynamic loss increases, resulting in a drop in an overall engine efficiency. A significant increase in efficiency could be achieved if separation could be prevented, or minimized. Active flow control could provide a means for minimizing separation under conditions where it is most severe (low Re), without causing additional losses under other conditions (high Re). Minimizing separation will allow improved designs with fewer stages and fewer airfoils per stage to generate the same power. The effects of the jet geometry, blowing ratio, density ratio, pulsation frequency and duty cycle on the size of the separated region were examined in this work. The results from Reynolds Averaged Navier-Stokes and Large Eddy Simulation computational approaches were compared with the experimental data.

Committee:

Dr. Mounir Ibrahim (Committee Chair); Dr. Asuquo Ebiana (Committee Member); Dr. Hanz Richter (Committee Member); Dr. Miron Kaufman (Committee Member); Dr. Petru Fodor (Committee Member); Dr. Ralph Volino (Committee Member)

Subjects:

Mechanical Engineering

Keywords:

CFD; flow control; film cooling; LES; airfoil; separation; Vortex Generator Jets

Nickol, Jeremy BAirfoil, Platform, and Cooling Passage Measurements on a Rotating Transonic High-Pressure Turbine
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 compressibility effects within the cooling holes which are not present in low speed experiments. Another major focus of this work is on the forward purge cavity and rotor and stator inner endwalls. Pressure and heat transfer measurements are taken at several locations, and compared as both forward and aft purge flow rates are varied. It is seen that increases in forward purge rates result in a flow blockage and greater pressure on the endwalls both up and downstream of the cavity. Thus, even in locations where the coolant does not directly cover the metal surface, it can have a significant impact on the local pressure loading and heat transfer rate. The heat transfer on the platform further downstream, however, is unchanged by variations in purge flow rates.

Committee:

Randall Mathison (Advisor); Michael Dunn (Committee Member); Sandip Mazumder (Committee Member); Jeffrey Bons (Committee Member)

Subjects:

Aerospace Engineering; Engineering; Mechanical Engineering

Keywords:

gas turbine; heat transfer; aerodynamics; film cooling; shaped holes; purge; cavity; cooling; blowing; serpentine; rotating; transonic; high pressure turbine; platform; endwall

MISHRA, SUMANSingle-Hole Film Cooling on a Turbine-Blade Leading-Edge Model
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

Keywords:

Film cooling blowing ratio turbine blade leading edge cylindrical model

O'Neil, Alanna R.Chemiluminescence and High Speed Imaging of Reacting Film Cooling Layers
Master of Science (M.S.), University of Dayton, 2011, Aerospace Engineering

The demand for more efficient and compact gas turbine engines has resulted in an increase in the operating temperatures and pressures and a decrease in combustor weight and size. These advances may result in incomplete combustion products entering the turbine section. The products can react with the air intended to cool the turbine vanes, and the resulting flame can cause damage to the engine. This study reports chemiluminescence measurements of flames and correlates these to heat release rate and the measured heat flux to a surface. To accomplish this, fuel rich combustion products are generated in a well-stirred reactor. The flow of products is directed over a flat plate with cooling air jets normal to the flow. Chemiluminescence data of the flames is obtained, along with high speed images, and temperature measurements of the flow inside the test section. Three film cooling geometries are studied: normal holes, fan shaped holes, and slot. Measurements are acquired at three equivalence ratios (1.3, 1.4, and 1.5) at three different blowing ratios (M = 1, 4, and 7).

It is found that the rate of heat release from the flame does not always trend the same as the heat transfer to the surface. It is also seen that a large reaction region does not always equate to high heat flux to the surface. If enough cooling air is present the surface is protected from the heat released from the flame.

Committee:

Dilip Ballal, PhD (Committee Chair); Scott Stouffer, PhD (Committee Member); David Blunck, PhD (Committee Member); Sukh Sidhu, PhD (Committee Member)

Subjects:

Aerospace Engineering

Keywords:

Film Cooling; Chemiluminescence; Reacting Boundary Layers; Flames; Image Analysis;

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