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1
Hossain, Mohammad Arif. "Sweeping Jet Film Cooling." The Ohio State University, 2020. http://rave.ohiolink.edu/etdc/view?acc_num=osu1586462423029754.
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Strater, Kurt F. "Countercurrent cooling of blown film." Thesis, McGill University, 1985. http://digitool.Library.McGill.CA:80/R/?func=dbin-jump-full&object_id=66003.
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Fawcett, Richard James. "Coherent unsteadiness in film cooling." Thesis, University of Oxford, 2011. http://ora.ox.ac.uk/objects/uuid:57ec3da6-4946-4f66-8421-b01d53d7e0fc.
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Film cooling is vital for the cooling of the blades and vanes in the high temperature environment of a jet engine high pressure turbine stage. Previous research into film cooling has typically concentrated on its time-mean performance. However, results from other studies upon more simplified geometries, suggest that coherent unsteadiness is likely to also be present in film cooling flows. The research presented in this thesis, therefore, aims to characterise what coherent unsteadiness, if any, is present within film cooling flows. Cylindrical and shaped cooling holes, located upon the pressure surface of a turbine blade within a large scale linear cascade, have been investigated. A blowing ratio range of 0.5 to 2.0 has been investigated, with either a plenum or perpendicular crossflow at the cooling hole inlet. Particle Image Velocimetry, high speed photography and Hot Wire Anemometry have been used to investigate the jet downstream of both cooling holes. The impact of crossflow at the hole inlet upon the flowfield inside both cooling holes has been investigated using Hot Wire Anemometry and a further numerical model solved by applying TBLOCK. The results presented in the current thesis, show the existence of two coherent unsteady structures in the jet downstream of both the cylindrical and the shaped holes. These structures are called shear layer vortices and hairpin vortices, and their formation is dependent on the velocity difference across the jet shear layer. Inside the cooling hole coherent hairpin vortices also appear to occur, with their formation dependent on the direction and magnitude of the crossflow at the hole inlet. The coherent unsteadiness presented here is shown for the first time for film cooling flows, and recommendations to build on the current study, in what is potentially an interesting research area, are made at the end of this thesis.
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Aghasi, Paul P. "Dependence of Film Cooling Effectiveness on 3D Printed Cooling Holes." University of Cincinnati / OhioLINK, 2016. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1458893416.
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Nowlin, Scott Raymond. "The use of intersecting film cooling passages for nozzle guide vane cooling." Thesis, University of Oxford, 2009. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.670018.
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Licu, Dragos N. "Heat transfer characteristics in film cooling applications." Thesis, National Library of Canada = Bibliothèque nationale du Canada, 1998. http://www.collectionscanada.ca/obj/s4/f2/dsk2/tape17/PQDD_0005/NQ34581.pdf.
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Friedrichs, Stefan. "Endwall film-cooling in axial flow turbines." Thesis, University of Cambridge, 1997. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.627225.
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Telisinghe, Janendra C. "Film cooling of turbine blade trailing edges." Thesis, University of Oxford, 2013. http://ora.ox.ac.uk/objects/uuid:86c06246-16e9-4378-9a61-e09317d31a92.
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In modern gas turbine engines, film cooling is extensively used to cool the components exposed to the hot mainstream gas path. In implementing film cooling on modern gas turbine engines, the trailing edge film poses a particularly challenging design problem. From an aerodynamic point of view, the trailing edge of a blade is designed to be as thin as possible. However, this conflicts with the implementation of the cooling design. The most common method of film cooling the trailing edge is via late pressure surface discrete film cooling holes. Another method of cooling the trailing edge is by using discrete pressure surface slots. This thesis documents a comparative aerodynamic and heat transfer study of three trailing edge cooling configurations. The study was carried out using a large scale, low speed wind tunnel situated at the Southwell Laboratory. The three trailing edge cooling configurations considered were as follows. First is the common late pressure film cooling of the trailing edge via discrete film cooling holes. This configuration is designated as datum configuration. Second is the pressure surface slot coolant ejection. This configuration was designated as cast cutback configuration. The third is the pressure surface ejection via discrete film cooling holes within a step cutback. This configuration was designated the machined cutback configuration. The above configurations were incorporated into three flat plates manufactured using stereolithography. In the aerodynamic study, the static pressure distribution and discharge coefficient for the three configurations were compared. Furthermore, two dimensional total pressure measurements were carried out using a traverse mechanism downstream of the test plates. The total pressure measurements were used to compute the mixed out losses for the three configurations. It was found that the datum and machined cutback configurations have similar discharge coefficients and mixed out losses whilst the cast cutback configuration produces greater mixed out loss. The film effectiveness and heat transfer coefficient on the pressure surface downstream of the coolant ejection was obtained using a steady state heat transfer technique. The effectiveness measurements were compared with those from the literature and correlated against the two dimensional slot model. The heat transfer measurements show that the cast cutback configuration has the potential to give higher effectiveness at the trailing edge.
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Renze, Peter Clemens-August. "Large eddy simulation of film cooling flows /." Aachen : Shaker, 2008. http://d-nb.info/989959031/04.
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Kheniser, Issam E. "Film Cooling Experiments in a Medium Duration Blowdown Facility." The Ohio State University, 2010. http://rave.ohiolink.edu/etdc/view?acc_num=osu1276540410.
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Leblanc, Christopher N. "Design, Analysis, and Development of a Tripod Film Cooling Hole Design for Reduced Coolant Usage." Diss., Virginia Tech, 2012. http://hdl.handle.net/10919/19206.
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This research has a small portion focused on interior serpentine channels, with the primary focus on improving the effectiveness of the film cooling technique through the use of a new approach to film cooling. This new approach uses a set of three holes sharing the same inlet and diverging from the central hole to form a three-legged, or tripod, design. The tripod design is examined in depth, in terms of geometric variations, through the use of flat plate and cascade rigs, with both transient and steady-state experiments. The flat plate tests provide a simplified setting in which to test the design in comparison to other geometries, and establish a baseline performance in a simple flow field that does not have the complications of surface curvature or mainstream pressure gradients. Cascade tests allow for testing of the design in a more realistic setting with curved surfaces and mainstream pressure gradients, providing important information about the performance of the design on suction and pressure surfaces of airfoils. Additionally, the cascade tests allow for an investigation into the aerodynamic penalties associated with the injection hole designs at various flow rates. Through this procedure the current state of film cooling technology may be improved, with more effective surface coverage achieved with reduced coolant usage, and with reduced performance penalties for the engine as a whole. This research has developed a new film hole design that is manufacturable and durable, and provides a detailed analysis of its performance under a variety of flow conditions. This cooling hole design provides 40% higher cooling effectiveness while using 50% less coolant mass flow. The interior serpentine channel research provides comparisons between correlations and experiments for internal passages with realistic cross sections.Ph. D.
12
Sargison, Jane Elizabeth. "Development of a novel film cooling hole geometry." Thesis, University of Oxford, 2001. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.365427.
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Oguntade, Habeeb Idowu. "Modelling of gas turbine film and effusion cooling." Thesis, University of Leeds, 2012. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.581946.
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This thesis presents CFD predictions of gas turbine film and effusion cooling. The dearth of detailed experimental adiabatic effusion cooling data led to the validation of the computational procedures against the experimental adiabatic cooling effectiveness data for a single row of inclined round film cooling holes. This showed that the overall best agreement of the CFD predictions with experimental data was for the realizable k-e turbulence model with enhanced wall function. This was also shown to give good predictions of experimental results for trench outlet film cooling. This film cooling CFD work was extended .to demonstrate trench outlet lip geometries that could further improve the cooling effectiveness. The limitation of the CFD model was at higher blowing rates, M, when the film jet lifted off from the surface, where the CFD did not accurately predict the adiabatic cooling effectiveness close to the hole. For attached jets at lower M the agreement was good. The same CFD procedures were used for all the effusion cooling conjugate heat transfer (CHT) predictions. The hot metal wall effusion cooling experimental data base of Andrews and co-workers (1983-1995) was used to validate the CHT effusion cooling predictions. This database was for combustor flat wall cooling with mainly 90° injection holes. The overall effusion cooling effectiveness was measured and this required conjugate heat transfer CFD predictions. The adiabatic film cooling effectiveness was also predicted, by using a gas tracer in the cooling air and predicting its concentration at the effusion wall. For each effusion hole configuration, the coolant mass flow rate, G kg/srrr2bar, was varied from 0.1 to 1.5 and each G required a separate computation. The influence of the number of holes at a constant X!D of 4.6 and the hole size at fixed X were investigated. The agreement between the predictions and experimental data was good. Finally, the influence of the effusion coolant jets flow direction to the hot-gas crossflow on effusion cooling performance was investigated. This included 30° inclined opposed-flow jets effusion wall, which was predicted to be the best effusion jets flow pattern. The addition of the filleted shape trench outlet to effusion cooling was predicted to improve the cooling performance with reduced coolant mass flow rate, due to the improved adiabatic film cooling.
14
Yang, Xiaobo. "Numerical study of film cooling in hypersonic flows." Thesis, University of Glasgow, 2002. http://theses.gla.ac.uk/2063/.
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In this thesis, a numerical study of film cooling in hypersonic laminar and turbulent flows has been performed using an in-house Navier-Stokes solver. The aim of this computational work is to investigate the mechanism and effectiveness of film cooling in hypersonic laminar and turbulent flows. Hypersonic flow over a flat plate without film cooling was first studied to provide a reference datum to check the effectiveness of film cooling. For laminar film cooling (M¥ = 9.9), three different primary flow conditions were first used for validation. The inclusion of the development of the flow in the plenum chamber upstream of the slot was found to provide better heat prediction than a uniform boundary condition at the slot exit. Detailed information of the flow field including velocity profile, Mach contour, temperature contour and heat transfer rate was presented. The mechanism of film cooling has been revealed according to the plots of calculated velocity profiles, Mach contours and temperature contours downstream of the slot. The coolant fluid was found to affect the primary boundary layer in two ways: 1) initially a separate layer established by the coolant fluid itself in the near slot area, 2) later a mixing layer between the primary and coolant flow streams. Then five coolant injection rates between 2.95 x 10-4 and 7.33 x 10-4kg/s and three slot heights 0.8382, 1.2192, 1.6002 mm, were examined in hypersonic laminar film cooling. For turbulent film cooling (M¥ = 8.2), for the geometry used in the experiment, the injection at an angle of 20° was found to be appropriate. Different turbulence models including Wilcox's k - w model. Menter's baseline and SST model have been tested. It is concluded that the Wilcox's k - w turbulence model with dilatation-dissipation correction provides the best heat prediction. Again, five coolant injection rates varies from 5.07 x 10-4 to 30.69 x 10-4 kg/s and three slot heights (the same as studied in the laminar film cooling) were studied to check the influence on film cooling effectiveness. Both the coolant and the primary flow were air. Film cooling was found to be an effective way to protect wall surfaces that are exposed under a high heat transfer environment especially in hypersonic laminar flow. Increasing the coolant injection rate can obviously increase the film cooling effectiveness. Again, this works better in laminar flow than in turbulent flow. The coolant injection rate in turbulent flow should be considered to be high enough to give good heat protection. Slot height in both laminar and turbulent flows under the flow conditions in this study was found to be less important, which means other factors can be considered in priority when constructing film cooling systems. With the application of curve fitting, the cooling length was described using power laws according to curve fitting results. A two-equation film coating model has been presented to illustrate the relation between the film cooling effectiveness and the parameter x/(h/m). For film cooling effectiveness in log-log coordinates, a second-order polynomial curve can be used to fit the laminar flows, whilst a straight line is suitable for the turbulent flows.
15
Forth, C. J. Patrick. "An investigation of scaling parameters governing film-cooling." Thesis, University of Oxford, 1985. http://ora.ox.ac.uk/objects/uuid:aea5722e-89e5-4916-9bc4-2bc2631de9d9.
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Experiments were performed using an Isentropic Light Piston Tunnel, a transient facility which enables conditions representative of those in engines to be attained. The results were interpreted using a superposition model, which is shown to be a valuable and concise method of characterising the effects of injection.
16
Thomas, Mitra. "Optimization of endwall film-cooling in axial turbines." Thesis, University of Oxford, 2014. http://ora.ox.ac.uk/objects/uuid:e369eb63-0c99-4ded-ab05-6b050004ce4c.
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Considerable reductions in gas turbine weight and fuel consumption can be achieved by operating at a higher turbine entry temperature. The move to lean combustors with flatter outlet temperature profiles will increase temperatures on the turbine endwalls. This work will study methods to improve endwall film cooling, to allow these advances. Turbine secondary flows are caused by a deficit in near-wall momentum. These flow features redistribute near-wall flows and make it difficult to film-cool endwalls. In this work, endwall film cooling was studied by CFD and validated by experimental measurements in a linear cascade. This study will add to the growing body of evidence that injection of high momentum coolant into the upstream boundary layer can suppress secondary flows by increasing near-wall momentum. The reduction of secondary flows allows for effective cooling of the endwall. It is also noted that excess near-wall momentum is undesirable. This leads to upwash on the vane, driving coolant away from the endwall. A passive-scalar tracking method has been devised to isolate the contribution of individual film cooling holes to cooling effectiveness. This method was used to systematically optimize endwall cooling systems. Designs are presented which use half the coolant mass flow compared to a baseline design, while maintaining similar cooling effectiveness levels on the critical trailing endwall. By studying the effect of coolant injection on vane inlet total pressure profile, secondary flows were suppressed and upwash on the vane was reduced. The methods and insight obtained from this study were applied to a high pressure nozzle guide vane endwall from a current engine. The optimized cooling system developed offers significant improvement over the baseline.
17
Sundaram, Narayan. "Effects of Surface Conditions on Endwall Film-Cooling." Diss., Virginia Tech, 2007. http://hdl.handle.net/10919/27066.
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A higher demand in power output from modern land based gas turbines has resulted in an increase in combustor exit temperatures. High temperatures in turn have resulted in flatter profiles at the combustor exit warranting the need for sufficient cooling of the endwall region. Endwall cooling is affected by the coolant flow through certain design features. A typical endwall design includes a leakage slot at the interface between the combustor and the vane, a leakage slot at the vane-to-vane interface and film-cooling holes. In addition, with the increase in energy demands and depletion of natural gas resources, alternate fuels such as coal derived synthetic gas are being used in gas turbines. Coal derived fuels, however, contain traces of ash and other contaminants that deposit on endwall surfaces, thereby altering its surface conditions.
The purpose of this study was to investigate the effects of realistic endwall features and surface conditions on leading edge endwall cooling. Endwall designs like placing film-cooling holes in a trench, which provide an effective means of improving cooling were also studied at the leading edge. An infrared camera was used to obtain measurements of adiabatic effectiveness levels and a laser Doppler velocimeter was used for flowfield measurements.
This study was done on a large scale, low-speed, recirculating wind tunnel operating at a Reynolds number of 2.1e+5 and an inlet mainstream turbulence level of 1%. Endwall measurements were taken for coolant flow through varying slot width at the combustor-vane interface. A constant coolant mass flow and a narrower combustor-turbine interface slot caused the coolant to exit uniformly whereas increasing the slot width had an opposite effect. Measurements were also taken with hole blockage and spallation, which showed a 10-25% decrease in the effectiveness levels whereas near hole deposition showed a 20% increase in effectiveness levels.
A comparison of the cooling effectiveness due to placement of film-cooling holes in a trench was made to film-cooling holes not placed in a trench. Measurements indicated a superior performance of trenched holes to holes without a trench. Trenched holes showed a 60% increase in effectiveness levels due to decreased coolant jet separation.Ph. D.
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Ramesh, Sridharan. "Analysis of film cooling performance of tripod hole." Diss., Virginia Tech, 2016. http://hdl.handle.net/10919/72912.
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The thermal efficiency of a gas turbine directly depends on the rotor inlet temperature. The ever increasing demand for more power and advances in the field of engineering enabled this temperature to be pushed higher. But the material strength of the blades and vanes can often impose restrictions on the thermal load it can bear. This is where gas turbine cooling becomes very critical and a better cooling design has the potential to extend the blade life span, enables higher rotor inlet temperatures, conserves compressor bleed air. Among various kinds of cooling involved in gas turbines, film cooling will be the subject of this study.
A novel concept for film cooling holes referred to as anti-vortex design proposed in 2007 is explored in this study. Coolant exits through two bifurcated cylindrical holes that branched out on either side of the central hole resulting in a tripod-like arrangement. Coolant from the side holes interacted with the mainstream and produced vortices that countered the main central rotating vortex pairs, weakening it and pushing the coolant jet towards the surface. In order to understand the performance of this anti-vortex tripod film cooling, a flat plate test setup and a low speed subsonic wind tunnel linear cascade were built.
Transient heat transfer experiments were carried out in the flat plate test setup using Infrared thermography. Film cooling performance was quantified by measuring adiabatic effectiveness and heat transfer coefficient ratio. In order to gauge the performance, other standard hole geometries were also tested and compared with. Following the results from the flat plate test rig, film cooling performance was also evaluated on the surface of an airfoil. Adiabatic effectiveness was measured at different coolant mass flow rates. The tripod hole consistently provided better cooling compared to the standard cylindrical hole in both the flat plate and cascade experiments.
In order to understand the anti-vortex concept which is one of the primary reason behind better performance of the tripod film cooling hole geometry, numerical simulations (CFD) were carried out at steady state using RANS turbulence models. The interaction of the coolant from the side holes with the mainstream forms vortices that tries to suppress the vortex formed by the central hole. This causes the coolant jet from the central to stay close to the surface and increases its coverage. Additionally, the coolant getting distributed into three individual units reduces the exit momentum ratio. Tripod holes were found to be capable of providing better effectiveness even while consuming almost half the coolant used by the standard cylindrical holes.Ph. D.
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Jessen, Wilhelm. "Particle image velocimetry measurements of film cooling flows /." Aachen : Mainz, 2008. http://bvbr.bib-bvb.de:8991/F?func=service&doc_library=BVB01&doc_number=017075640&line_number=0001&func_code=DB_RECORDS&service_type=MEDIA.
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Modlin, James Michael. "Hypersonic aerospace vehicle leading edge cooling using heat pipe, transpiration and film cooling techniques." Diss., Georgia Institute of Technology, 1991. http://hdl.handle.net/1853/16347.
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Huang, Xi [Verfasser], and X. [Akademischer Betreuer] Cheng. "Study on Water Film Cooling for PWR's Passive Containment Cooling System / Xi Huang. Betreuer: X. Cheng." Karlsruhe : KIT-Bibliothek, 2015. http://d-nb.info/1077821875/34.
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Lim, Chia Hui. "The influence of film cooling on turbine aerodynamic performance." Thesis, University of Cambridge, 2011. https://www.repository.cam.ac.uk/handle/1810/283872.
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Narayanan, Shankar. "Gas assisted thin-film evaporation from confined spaces." Diss., Georgia Institute of Technology, 2011. http://hdl.handle.net/1853/42780.
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A novel cooling mechanism based on evaporation of thin liquid films is presented for thermal management of confined heat sources, such as microprocessor hotspots. The underlying idea involves utilization of thin nanoporous membranes for maintaining microscopically thin liquid films by capillary action, while providing a pathway for the vapor generated due to evaporation at the liquid-vapor interface. The vapor generated by evaporation is continuously removed by using a dry sweeping gas keeping the membrane outlet dry. This thesis presents a detailed theoretical, computational and experimental investigation of the heat and mass transfer mechanisms that result in dissipating heat.
Performance analysis of this cooling mechanism demonstrates heat fluxes over 600W/cm2 for sufficiently thin membrane and film thicknesses (~1-5µm) and by using air jet impingement for advection of vapor from the membrane surface. Based on the results from this performance analysis, a monolithic micro-fluidic device is designed and fabricated incorporating micro and nanoscale features. This MEMS/NEMS device serves multiple functionalities of hotspot simulation, temperature sensing, and evaporative cooling. Subsequent experimental investigations using this microfluidic device demonstrate heat fluxes in excess of 600W/cm2 at 90 C using water as the evaporating coolant.
In order to further enhance the device performance, a comprehensive theoretical and computational analysis of heat and mass transfer at micro and nanoscales is carried out. Since the coolant is confined using a nanoporous membrane, a detailed study of evaporation inside a nanoscale cylindrical pore is performed. The continuum analysis of water confined within a cylindrical nanopore determines the effect of electrostatic interaction and Van der Waals forces in addition to capillarity on the interfacial transport characteristics during evaporation. The detailed analysis demonstrates that the effective thermal resistance offered by the interface is negligible in comparison to the thermal resistance due to the thin film and vapor advection. In order to determine the factors limiting the performance of the MEMS device on a micro-scale, a device-level detailed computational analysis of heat and mass transfer is carried out, which is supported by experimental investigation. Identifying the contribution of various simultaneously occurring cooling mechanisms at different operating conditions, this analysis proposes utilization of hydrophilic membranes for maintaining very thin liquid films and further enhancement in vapor advection at the membrane outlet to achieve higher heat fluxes.
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Yusop, Nadiahnor Md. "CFD Predictions of Gas Turbine Full-Coverage Film Cooling." Thesis, University of Leeds, 2007. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.490967.
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The present study aims at conducting a numerical investigation of the classic film cooling scheme of transpiration film cooling and effusion film cooling for validation through computational methods. Steady-state simulations were performed and the flow was considered incompressible with low turbulence. The CFD package FLUENT 6.2.16 was used to solve the Navier-Stokes equations numerically, and the pre-processor, Gambit 2.2.30, was used to generate the required grid. The research aims at perfonning computational predictions on the film cooling performance and the aerodynamics aspect of flat plate film cooling ·on the transpiration and effusion film cooling. It was determined that the proposed scheme and type of geometry, coupled with the hybrid mesh generation, can validate the classic experimental results. with reasonable agreement. Computational predictions on the transpiration film cooling have shown that different boundary conditions used for the porous media may lead to different results, whether it is over-prediction or under-prediction results in comparison with the experimental data. It has been observed for the effusion film cooling, on the case of co-flow coolant , . ejection into the mainstream, that the adiabatic film cooling effectiveness continuously increases with the axial distance towards the leading edge where the flow of the coolant is fully-developed. Furthermore, the streamwise cooling uniformity was better than in the upstream region at the middle region of the test wall. In contrast, the adiabatic film cooling effectiveness for the opposed flow coolant ejection into the mainstream flow was gradually decreasing with the axial distance. Coflow coolant ejection into the mainstream has provide better cooling effectiveness but the oppose flow coolant ejection from the cooling holes has proved to be good aerodynamics in protecting the adjacent wall due to the large area of the film cooling coverage of the combustor wall. The present study was concerned only with the downstream effectiveness aspect on the performance of the coolant mass flow on the geometrical parameters effects; for transpiration film cooling - the pore size, and effusion film cooling - hole diameter, film cooling hole arrangement, number of holes, inclination and orientation of cooling hole with respect to the mainstream flow. The performance related to the heat transfer coefficient and conjugate heat transfer is a prospective topic for future studies. Advanced and innovative cooling techniques are essential in order to improve the efficiency and output power of the gas turbines. The CFD predictions performed have utilised a scalar tracer gas in the coolant flow and has been very effective at visualizing the coolant to the mainstream mixing phenomenon, determining the boundary layer development and directly predicting the adiabatic film cooling effectiveness. Current methods in determining the film cooling effectiveness using the scalar tracer gas concentration facilitate the future study on the conjugate heat transfer pred,iction where the temperature profiles cannot be used because conjugate heat transfer is highly affected by the effect ofthe temperature in the system. The technique· of providing an alternative method using the heat and mass transfer analogy in quantify the cooling effectiveness combines the advantages of using a scalar tracer gas in determining the cooling effectiveness and also provide clear insight into the film cooling structure in the cooling hole and coolant interaction in the mainstream when the experimental method is at 'off-limit'. The results of the present investigations performed were used to validate the computation model. Therefore, this study is of value for those interested in gas turbine cooling.
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Zhou, Jian Ming. "A multi-grid method for computation of film cooling." Thesis, University of British Columbia, 1990. http://hdl.handle.net/2429/29414.
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This thesis presents a multi-grid scheme applied to the solution of transport equations in turbulent flow associated with heat transfer. The multi-grid scheme is then applied to flow which occurs in the film cooling of turbine blades. The governing equations are discretized on a staggered grid with the hybrid differencing scheme. The momentum and continuity equations are solved by a nonlinear full multi-grid scheme with the SIMPLE algorithm as a relaxation smoother. The turbulence k — Є equations and the thermal energy equation are solved on each grid without multi-grid correction.
Observation shows that the multi-grid scheme has a faster convergence rate in solving the Navier-Stokes equations and that the rate is not sensitive to the number of mesh points or the Reynolds number. A significant acceleration of convergence is also produced for the k — Є and the thermal energy equations, even though the multi-grid correction is not applied to these equations. The multi-grid method provides a stable and efficient means for local mesh refinement with only little additional computational and.memory costs.
Driven cavity flows at high Reynolds numbers are computed on a number of fine meshes for both the multi-grid scheme and the local mesh-refinement scheme. Two-dimensional film cooling flow is studied using multi-grid processing and significant improvements
in the results are obtained. The non-uniformity of the flow at the slot exit and its influence on the film cooling are investigated with the fine grid resolution. A near-wall turbulence model is used. Film cooling results are presented for slot injection with different mass flow ratios.Science, Faculty of
Mathematics, Department of
Graduate
26
Gotzen, Thomas [Verfasser]. "Numerical Investigation of Film and Transpiration Cooling / Thomas Gotzen." München : Verlag Dr. Hut, 2013. http://d-nb.info/1043892435/34.
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Rowbury, David. "Discharge coefficients of nozzle guide vane film cooling holes." Thesis, University of Oxford, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.365838.
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Natsui, Gregory. "Surface Measurements and Predictions of Full-Coverage Film Cooling." Master's thesis, University of Central Florida, 2012. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/5351.
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Full-coverage film cooling is investigated both experimentally and numerically. First,surface measurements local of adiabatic film cooling effectiveness and heat transfer augmentation for four different arrays are described. Reported next is a comparison between two very common turbulence models, Realizable k-epsilon and SST k-omega, and their ability to predict local film cooling effectiveness throughout a full-coverage array. The objective of the experimental study is the quantification of local heat transferaugmentation and adiabatic film cooling effectiveness for four surfaces cooled by large, both in hole count and in non-dimensional spacing, arrays of film cooling holes. The four arrays are of two different hole-to-hole spacings (P=D = X=D = 14.5; 19.8) and two different hole inclination angles (alpha = 30°; 45°), with cylindrical holes compounded relative to the flow(beta = 45°) and arranged in a staggered configuration. Arrays of up to 30 rows are tested so that the superposition effect of the coolant film can be studied. In addition, shortened arrays of up to 20 rows of coolant holes are also tested so that the decay of the coolant film following injection can be studied. Levels of laterally averaged effectiveness reach values as high as η = 0.5, and are not yet at the asymptotic limit even after 20 - 30 rows of injection for all cases studied. Levels of heat transfer augmentation asymptotically approach values of h=h0 ≈ 1.35 rather quickly, only after 10 rows. It is conjectured that the heat transfer augmentation levels off very quickly due to the boundary layer reaching an equilibrium in which the perturbation from additional film rows has reached a balance with the damping effect resulting from viscosity. The levels of laterally averaged adiabatic film cooling effectiveness far exceeding eta = 0.5 are much higher than expected.
The heat transfer augmentation levels off quickly as opposed to the film effectiveness which continues to rise (although asymptotically) at large row numbers. This ensures that an increased row count represents coolant well spent. The numerical predictions are carried out in order to test the ability of the two most common turbulence models to properly predict full-coverage film cooling. The two models chosen, Realizable k-epsilon (RKE) and Shear Stress Transport k-omega (SSTKW), are both two-equation models coupled with Reynolds Averaged governing equations which makeseveral gross physical assumptions and require several empirical values. Hence, the models are not expected to provide perfect results. However, very good average values are seen tobe obtained through these simple models. Using RKE in order to model full-coverage film cooling will yield results with 30% less error than selecting SSTKW.M.S.A.E.
Masters
Mechanical and Aerospace Engineering
Engineering and Computer Science
Aerospace Engineering; Thermofluid Aerodynamic Systems
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Tang, Brian M. T. "Unshrouded turbine blade tip heat transfer and film cooling." Thesis, University of Oxford, 2011. http://ora.ox.ac.uk/objects/uuid:f8479e89-9cd1-4aa7-b5c8-8068ad80de54.
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This thesis presents a joint computational and experimental investigation into the heat transfer to unshrouded turbine blade tips suitable for use in high bypass ratio, large civil aviation turbofan engines. Both the heat transfer to the blade tip and the over-tip leakage flow over the blade tip are characterised, as each has a profound influence on overall engine efficiency. The study is divided into two sections; in the first, computational simulations of a very large scale, low speed linear cascade with a flat blade tip were conducted. These simulations were validated against experimental data collected by Palafox (2006). A thorough assessment of turbulence models and minimum meshing requirements was performed. The standard k-ω and standard k-ϵ turbulence models significantly overpredicted the turbulence levels within the tip gap. The other models were very similar in performance; the SST k-ω and realisable k-ϵ models were found to be the most suitable for the flow environment. The second section documents the development and testing of a novel hybrid blade tip design, the squealet tip, which seeks to combine the known benefits of winglet and double squealer tips. The development of the external geometry was performed primarily through engine-representative CFD simulations at a range of tip gaps from 0.45% to 1.34% blade chord. The squealet tip was found to have a similar aerodynamic sensitivity to tip clearance as a baseline double squealer tip, with a tip gap efficiency exchange rate of 2.03, although this was 18% greater than the alternative winglet tip. The squealet tip displayed higher predicted stage efficiency than the winglet tip over the majority of the range of tip clearances investigated, however. The overall heat load was reduced by 14% compared with the winglet tip but increased by 28% over the double squealer tip, primarily due to the change in wetted surface area. The predicted local heat transfer coefficients were similar across all geometries. A realistic internal cooling plenum and an array of blade tip cooling holes were subsequently added to the squealet tip geometry and the cooling configuration refined by the selective sealing of cooling holes. Film cooling performance was largely assessed by the predicted adiabatic wall temperature distributions. A viable cooling scheme which reduced the cooling air requirement by 38% was achieved, compared to the initial case which had all cooling holes open. This was associated with just a 7% increase in blade tip heat flux and no penalty in peak temperature on the blade tip. Film cooling air ejected from holes on the blade suction side was swept away from the blade tip region, making the squealet rim at the crown of the blade particularly challenging to cool. It was demonstrated that this region could be cooled effectively by ballistic cooling from holes located on the blade tip cavity floor, although this was expensive in terms of the mass flow rate of cooling air required. The computational results were reinforced with experimental data collected in a transonic linear cascade. Downstream aerodynamic loss measurements were taken for a linearised version of the squealet tip design without cooling at nominal tip gaps of 0.45%, 0.89% and 1.34% blade chord, which was compared to similar data taken by O’Dowd (2010) for flat and winglet tips. The squealet was seen to have a similar aerodynamic loss to the flat tip and a reduced loss compared with the winglet tip. Full surface heat transfer measurements were taken for the uncooled squealet tip, at tip gaps of 0.89% and 1.34% blade chord, and for two configurations of the cooled squealet tip, at a tip clearance of 0.89% blade chord. The qualitative similarity between the measured heat transfer distributions and the those predicted by the engine-representative CFD simulations was good. A CFD simulation of the uncooled linear cascade environment at the 1.34% blade chord tip clearance was performed using a single blade with translationally periodic boundary conditions. The predicted size of the over-tip leakage vortex was smaller than had been measured, resulting in a large underprediction in the magnitude of the downstream area-averaged aerodynamic loss. The magnitudes of the predicted blade tip Nusselt number distribution were similar to those produced by the engine-representative CFD simulations and lower than that measured experimentally. Differences in the shape of the Nusselt number distribution were observed in the vicinity of regions of separated and reattaching flow, but other salient features were replicated in the computational data. The squealet tip has been shown to be a promising, viable unshrouded blade tip design with an aerodynamic performance similar to the double squealer tip but is more amenable to film cooling. It is significantly lighter than a winglet tip and incurs a reduced thermal load. The squealet tip design can now be developed into a blade tip geometry for use in real engines to provide an alternative to shrouded turbine blades and current unshrouded blade tip designs. A commercial CFD solver, Fluent 6.3, was shown to capture blade tip heat transfer and over-tip leakage flow sufficiently well to be a useful design guide. However, the sensitivity of the flow structure (and hence, heat transfer) in the forward part of the blade tip cavity suggests that physical testing cannot be eliminated from the design process entirely.
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Cruz, Carlos A. "Experimental and numerical characterization of turbulent slot film cooling." College Park, Md. : University of Maryland, 2008. http://hdl.handle.net/1903/8145.
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Thesis (Ph. D.) -- University of Maryland, College Park, 2008.Thesis research directed by: Dept. of Aerospace Engineering. Title from t.p. of PDF. Includes bibliographical references. Published by UMI Dissertation Services, Ann Arbor, Mich. Also available in paper.
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Prenter, Robin Michael Patrick. "Investigating the Physics and Performance of Reverse-Oriented Film Cooling." The Ohio State University, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=osu1500505248644198.
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Rodriguez, Sylvette. "EFFECT OF PRESSURE GRADIENT AND WAKE ON ENDWALL FILM COOLING EFFECTIVENESS." Doctoral diss., University of Central Florida, 2008. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/2940.
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Endwall film cooling is a necessity in modern gas turbines for safe and reliable operation. Performance of endwall film cooling is strongly influenced by the hot gas flow field, among other factors. For example, aerodynamic design determines secondary flow vortices such as passage vortices and corner vortices in the endwall region. Moreover blockage presented by the leading edge of the airfoil subjects the incoming flow to a stagnating pressure gradient leading to roll-up of the approaching boundary layer and horseshoe vortices. In addition, for a number of heavy frame power generation gas turbines that use cannular combustors, the hot and turbulent gases exiting from the combustor are delivered to the first stage vane through transition ducts. Wakes induced by walls separating adjacent transition ducts located upstream of first row vanes also influence the entering main gas flow field. Furthermore, as hot gas enters vane passages, it accelerates around the vane airfoils. This flow acceleration causes significant streamline curvature and impacts lateral spreading endwall coolant films. Thus endwall flow field, especially those in utility gas turbines with cannular combustors, is quite complicated in the presence of vortices, wakes and strong favorable pressure gradient with resulting flow acceleration. These flow features can seriously impact film cooling performance and make difficult the prediction of film cooling in endwall. This study investigates endwall film cooling under the influence of pressure gradient effects due to stagnation region of an axisymmetric airfoil and in mainstream favorable pressure gradient. It also investigates the impact of wake on endwall film cooling near the stagnation region of an airfoil. The investigation consists of experimental testing and numerical simulation. Endwall film cooling effectiveness is investigated near the stagnation region on an airfoil by placing an axisymmetric airfoil downstream of a single row of inclined cylindrical holes. The holes are inclined at 35° with a length-to-diameter ratio of 7.5 and pitch-to-diameter ratio of 3. The ratio of leading edge radius to hole diameter and the ratio of maximum airfoil thickness to hole diameter are 6 and 20 respectively. The distance of the leading edge of the airfoil is varied along the streamwise direction to simulate the different film cooling rows preceding the leading edge of the airfoil. Wake effects are induced by placing a rectangular plate upstream of the injection point where the ratio of plate thickness to hole diameter is 6.4, and its distance is also varied to investigate the impact of strong and mild wake on endwall film cooling effectiveness. Blowing ratio ranged from 0.5 to 1.5. Film cooling effectiveness is also investigated under the presence of mainstream pressure gradient with converging main flow streamlines. The streamwise pressure distribution is attained by placing side inserts into the mainstream. The results are presented for five holes of staggered inclined cylindrical holes. The inclination angle is 30° and the tests were conducted at two Reynolds number, 5000 and 8000. Numerical analysis is employed to aid the understanding of the mainstream and coolant flow interaction. The solution of the computational domain is performed using FLUENT software package from Fluent, Inc. The use of second order schemes were used in this study to provide the highest accuracy available. This study employed the Realizable κ-ε model with enhance wall treatment for all its cases. Endwall temperature distribution is measured using Temperature Sensitive Paint (TSP) technique and film cooling effectiveness is calculated from the measurements and compared against numerical predictions. Results show that the characteristics of average film effectiveness near the stagnation region do not change drastically. However, as the blowing ratio is increased jet to jet interaction is enhanced due to higher jet spreading resulting in higher jet coverage. In the presence of wake, mixing of the jet with the mainstream is enhanced particularly for low M. The velocity deficit created by the wake forms a pair of vortices offset from the wake centerline. These vortices lift the jet off the wall promoting the interaction of the jet with the mainstream resulting in a lower effectiveness. The jet interaction with the mainstream causes the jet to lose its cooling capabilities more rapidly which leads to a more sudden decay in film effectiveness. When film is discharged into accelerating main flow with converging streamlines, row-to-row coolant flow rate is not uniform leading to varying blowing ratios and cooling performance. Jet to jet interaction is reduced and jet lift off is observed for rows with high blowing ratio resulting in lower effectiveness.Ph.D.
Department of Mechanical, Materials and Aerospace Engineering
Engineering and Computer Science
Mechanical Engineering PhD
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Wright, Lesley Mae. "Experimental investigation of turbine blade platform film cooling and rotational effect on trailing edge internal cooling." [College Station, Tex. : Texas A&M University, 2006. http://hdl.handle.net/1969.1/ETD-TAMU-1826.
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Al-Zurfi, Nabeel. "Large eddy simulation of cooling practices for improved film cooling performance of a gas turbine blade." Thesis, University of Manchester, 2017. https://www.research.manchester.ac.uk/portal/en/theses/large-eddy-simulation-of-cooling-practices-for-improved-film-cooling-performance-of-a-gas-turbine-blade(966b0252-4b0e-45de-9c1f-effe007261b0).html.
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The Large Eddy Simulation approach is employed to predict the flow physics and heat transfer characteristics of a film-cooling problem that is formed from the interaction of a coolant jet with a hot mainstream flow. The film-cooling technique is used to protect turbine blades from thermal failure, allowing the gas inlet temperature to be increased beyond the failure temperature of the turbine blade material in order to enhance the efficiency of gas turbine engines. A coolant fluid is injected into the hot mainstream through several rows of injection holes placed on the surface of a gas turbine blade in order to form a protective coolant film layer on the blade surface. However, due to the complex, unsteady and three-dimensional interactions between the coolant and the hot gases, it is difficult to achieve the desired cooling performance. Understanding of this complex flow and heat transfer process will be helpful in designing more efficiently cooled rotor blades. A comprehensive numerical investigation of a rotating film-cooling performance under different conditions is conducted in this thesis, including film-cooling on a flat surface and film-cooling on a rotating gas turbine blade. The flow-governing equations are discretised based on the finite-volumes method and then solved iteratively using the well-known SIMPLE and PISO algorithms. An in-house FORTRAN code has been developed to investigate the flat plate film-cooling configuration, while the gas turbine blade geometry has been simulated using the STAR-CCM+ CFD commercial code. The first goal of the present thesis is to investigate the physics of the flow and heat transfer, which occurs during film-cooling from a standard film hole configuration. Film-cooling performance is analysed by looking at the distribution of flow and thermal fields downstream of the film holes. The predicted mean velocity profiles and spanwise-averaged film-cooling effectiveness are compared with experimental data in order to validate the reliability of the LES technique. Comparison of adiabatic film-cooling effectiveness with experiments shows excellent agreement for the local and spanwise-averaged film-cooling effectiveness, confirming the correct prediction of the film-cooling behaviour. The film coverage and film-cooling effectiveness distributions are presented along with discussions of the influence of blowing ratio and rotation number. Overall, it was found that both rotation number and blowing ratio play significant roles in determining the film-cooling effectiveness distributions. The second goal is to investigate the impact of innovative anti-vortex holes on the film-cooling performance. The anti-vortex hole design counteracts the detrimental kidney vorticity associated with the main hole, allowing coolant to remain attached to the blade surface. Thus, the new design significantly improves the film-cooling performance compared to the standard hole arrangement, particularly at high blowing ratios. The anti-vortex hole technique is unique in that it requires only readily machinable round holes, unlike shaped film-cooling holes and other advanced concepts. The effects of blowing ratio and the positions of the anti-vortex side holes on the physics of the hot mainstream-coolant interaction in a film-cooled turbine blade are also investigated. The results also indicate that the side holes of the anti-vortex design promote the interaction between the vortical structures; therefore, the film coverage contours reveal an improvement in the lateral spreading of the coolant jet.
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Kaisuwan, Pisut. "Effect of vortex circulation on injectant from a single film-cooling hole and a row of film-cooling holes in a turbulent boundary layer." Thesis, Monterey, California. Naval Postgraduate School, 1989. http://hdl.handle.net/10945/27011.
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Marsh, Jan H. "Development of an experimental setup for the study of film pulsation effects on film cooling effectiveness." Honors in the Major Thesis, University of Central Florida, 2008. http://digital.library.ucf.edu/cdm/ref/collection/ETH/id/1111.
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This item is only available in print in the UCF Libraries. If this is your Honors Thesis, you can help us make it available online for use by researchers around the world by following the instructions on the distribution consent form at http://library.ucf.edu/Systems/DigitalInitiatives/DigitalCollections/InternetDistributionConsentAgreementForm.pdf You may also contact the project coordinator, Kerri Bottorff, at kerri.bottorff@ucf.edu for more information.Bachelors
Engineering and Computer Science
Aerospace Engineering
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Gao, Zhihong. "Experimental investigation of film cooling effectiveness on gas turbine blades." [College Station, Tex. : Texas A&M University, 2007. http://hdl.handle.net/1969.1/ETD-TAMU-1557.
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Hombsch, Maximilian [Verfasser]. "Film Cooling in Supersonic Flows : Filmkühlung in Überschallströmungen / Maximilian Hombsch." Aachen : Shaker, 2017. http://d-nb.info/1138178691/34.
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Tran, Nghia Trong Van. "Film cooling with wake passing applied to an annular endwall." Master's thesis, University of Central Florida, 2010. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/4582.
Full textAbstract:
Advancement in turbine technology has far reaching effects on today's society and environment. With more than 90% of electricity and 100% of commercial air transport being produced by the usage of gas turbine, any advancement in turbine technology can have an impact on fuel used, pollutants and carbon dioxide emitted to the environment. Within the turbine engine, fully understanding film cooling is critical to reliability of a turbine engine. Film cooling is an efficient way to protect the engine surface from the extremely hot incoming gas, which is at a temperature much higher than allowable temperature of even the most advanced super alloy used in turbine. Film cooling performance is affected by many factors: geometrical factors and as well as flow conditions. In most of the film cooling literature, film effectiveness has been used as criterion to judge and/or compare between film cooling designs. Film uniformity is also a critical factor, since it determines how well the coolant spread out downstream to protect the hot-gas-path surface of a gas turbine engine. Even after consideration of all geometrical factors and flow conditions, the film effectiveness is still affected by the stator-rotor interaction, in particular by the moving wakes produced by upstream airfoils. A complete analysis of end wall film cooling inside turbine is required to fully understand the phenomena. This full analysis is almost impossible in the academic arena. Therefore, a simplified but critical experimental rig and computational fluid model were designed to capture the effect of wake on film cooling inside an annular test section. The moving wakes are created by rotating a wheel with 12 spokes or rods with a variable speed motor. Thus changing the motor speed will alter the wake passing frequency. This design is an advancement over most previous studies in rectangular duct, which cannot simulate wakes in an annular passage as in an engine.; This rig also includes film injection that allows study of impact of moving wakes on film cooling. This wake is a simplified representation of the trailing edge created by an upstream airfoil. An annulus with 30?? pitch test section is considered in this study. This experimental rig is based on an existing flat plate film cooling (BFC) rig that has been validated in the past. Measurement of velocity profiles within the moving wake downstream from the wake generator is used to validate the CFD rotating wake model. The open literature on film cooling and past experiments performed in the laboratory validated the CFD film cooling model. With these validations completed, the full CFD model predicts the wake and film cooling interaction. Nine CFD cases were considered by varying the film cooling blowing ratio and the wake Strouhal number. The results indicated that wakes highly enhance film cooling effectiveness near film cooling holes and degrades the film blanket downstream of the film injection, at the moment of wake passing. However, the time-averaged film cooling effectiveness is more or less the same with or without wake.ID: 029049607; System requirements: World Wide Web browser and PDF reader.; Mode of access: World Wide Web.; Thesis (M.S.M.E.)--University of Central Florida, 2010.; Includes bibliographical references (p. 118-123).
M.S.M.E.
Masters
Department of Mechanical, Materials and Aerospace Engineering
Engineering and Computer Science
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Renze, Peter C. [Verfasser]. "Large-Eddy Simulation of Film Cooling Flows / Peter C Renze." Aachen : Shaker, 2008. http://d-nb.info/1162793457/34.
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Li, Ke. "Experimental Study of Heat Transfer Coefficient and Film Cooling Effectiveness." Thesis, KTH, Energiteknik, 2019. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-249061.
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This thesis investigates the possibility to evaluate the film cooling thermal performance on flat plate using Thermochromic Liquid Crystal. After an introduction of the basic concept and background of gas turbine blades film cooling and Thermochromic Liquid Crystal, a thorough explanation of four methods is presented. Dimensional or similarity analysis is implemented to build relationship between real engine and laboratory model. Also, the Reynolds number and Blowing ratio are the fundamental of test object design and TLC selection. This study illustrated the layout of the test rig and corresponding setups, and the following part explains the data collection system and image processing MATLAB script which is vital for the success of data extraction. The least square method is applied to figure time-series optimal solution in solver. All the experiments are conducted at near room temperature as opposed to the extremely high gas turbine exhausted gas, including two calibration test and one heat transfer experiment. The heat transfer coefficient and film cooling effectiveness are the target objective through the entire project. By comparison with a similar experiment in a literature, the outcomes partially validated the film cooling performance under the pre-set flow and thermal condition and the Liquid Crystal thermography technique is proved to be a trustworthy method to mapping heat transfer surface.Denna avhandling undersöker möjligheten att utvärdera filmkylningens termiska prestanda på plan platta med användning av Termokromisk Flytande Kristall (TLC). Efter en introduktion av grundkonceptet och bakgrunden till gasturbinbladens filmkylning och termokromisk flytande kristall presenteras en grundlig förklaring av fyra metoder. Dimensionell eller likhetsanalys implementeras för att bygga upp förhållandet mellan verklig motor och laboratoriemodell. Reynoldstalet och blåsningsförhållandet (blowing ratio) är också grunden för testobjektdesign och TLC-val. Denna studie illustrerade provriggens layout och tillhörande inställningar. I följande del förklaras datainsamlingssystemet och bildbehandling, MATLABTM-skriptet som är avgörande för framgång med datautvärdering. Den minsta kvadratiska metoden tillämpas för att hitta tidsseriens optimala lösning i lösaren. Alla experiment utförs vid nära rumstemperatur i motsats till den höga temperature på gasturbingasen, inklusive två kalibreringstest och ett värmeöverföringsexperiment. Värmeöverföringskoefficienten och filmkylningseffektiviteten är målmålet genom hela projektet. Resultaten validerade partiellt filmkylningens prestanda under det förinställda flödet och det termiska tillståndet. Liquid Crystal-termografitekniken har visat sig vara en pålitlig metod för att kartlägga värmeöverföringsytan jämfört med ett liknande experiment i den öppna litteraturen.
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O'Neil, Alanna R. "Chemiluminescence and High Speed Imaging of Reacting Film Cooling Layers." University of Dayton / OhioLINK, 2011. http://rave.ohiolink.edu/etdc/view?acc_num=dayton1324042434.
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Kandampalayam, Kandasamy Palaniappan Mouleeswaran. "Design, Development and Validation of UC Film Cooling Research Facility." University of Cincinnati / OhioLINK, 2017. http://rave.ohiolink.edu/etdc/view?acc_num=ucin1511863107613678.
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Juhany, Khalid A. Hunt Melany L. Hunt Melany L. "Supersonic film cooling including the effect of shock wave interaction /." Diss., Pasadena, Calif. : California Institute of Technology, 1994. http://resolver.caltech.edu/CaltechETD:etd-12112007-084103.
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Hinse, Mathieu. "Investigation of Transpiration Cooling Film Protection for Gas Turbine Engine Combustion Liner Application." Thesis, Université d'Ottawa / University of Ottawa, 2021. http://hdl.handle.net/10393/42425.
Full textAbstract:
Transpiration cooling as potential replacement of multi-hole effusion cooling for gas turbine engines combustion liner application is investigated by comparing their cooling film effectiveness based on the mass transfer analogy (CFEM). Pressure sensitive paint was used to measure CFEM over PM surfaces which was found to be on average 40% higher than multi-hole effusion cooling. High porosity PM with low resistance to flow movement were found to offer uneven distribution of exiting coolant, with large amounts leaving the trailing edge, leading to lopsided CFEM. Design of anisotropic PM based on PM properties (porosity, permeability, and inertia coefficient) were investigated using numerical models to obtain more uniform CFEM. Heat transfer analysis of different PM showed that anisotropic samples offered better thermal protection over isotropic PM for the same porosity. Comparison between cooling film effectiveness obtained from temperatures CFET against CFEM revealed large differences in the predicted protection. This is attributed to the assumptions made to apply CFEM, nonetheless, CFEM remains a good proxy to study and improve transpiration cooling. A method for creating a CAD model of designed PM is proposed based on critical characteristics of transpiration cooling for future use in 3D printing manufacturing.
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Chua, Khim Heng. "Experimental characterisation of the coolant film generated by various gas turbine combustor liner geometries." Thesis, Loughborough University, 2005. https://dspace.lboro.ac.uk/2134/12704.
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In modern, low emission, gas turbine combustion systems the amount of air available for cooling of the flame tube liner is limited. This has led to the development of more complex cooling systems such as cooling tiles i.e. a double skin system, as opposed to the use of more conventional cooling slots i.e. a single skin system. An isothennal experimental facility has been constructed which can incorporate 10 times full size single and double skin (cooling tile) test specimens. The specimens can be tested with or without effusion cooling and measurements have been made to characterise the flow through each cooling system along with the velocity field and cooling effectiveness distributions that subsequently develop along the length of each test section. The velocity field of the coolant film has been defined using pneumatic probes, hot-wire anemometry and PIV instrumentation, whilst gas tracing technique is used to indicate (i) the adiabatic film cooling effectiveness and (ii) mixing of the coolant film with the mainstream flow. Tests have been undertaken both with a datum low turbulence mainstream flow passing over the test section, along with various configurations in which large magnitudes and scales of turbulence were present in the mainstream flow. These high turbulence test cases simulate some of the flow conditions found within a gas turbine combustor. Results are presented relating to a variety of operating conditions for both types of cooling system. The nominal operating condition for the double skin system was at a coolant to mainstream blowing ratio of approximately 1.0. At this condition, mixing of the mainstream and coolant film was relatively small with low mainstream turbulence. However, at high mainstream turbulence levels there was rapid penetration of the mainstream flow into the coolant film. This break up of the coolant film leads to a significant reduction in the cooling effectiveness. In addition to the time-averaged characteristics, the time dependent behaviour of the .:coolantfilm was. also investigated. In particular, unsteadiness associated with large scale structures in the mainstream flow was observed within the coolant film and adjacent to the tile surface. Relative to a double skin system the single skin geometry requires a higher coolant flow rate that, along with other geometrical changes, results in typically higher coolant to mainstream velocity ratios. At low mainstream turbulence levels this difference in velocity between the coolant and mainstream promotes the generation of turbulence and mixing between the streams so leading to some reduction in cooling effectiveness. However, this higher momentum coolant fluid is more resistant to high mainstream turbulence levels and scales so that the coolant film break up is not as significant under these conditions as that observed for the double skin system. For all the configurations tested the use of effusion cooling helped restore the coolant film along the rear of the test section. For the same total coolant flow, the minimum value of cooling effectiveness observed along the test section was increased relative to the no effusion case. In addition the effectiveness of the effusion patch depends on the amount of coolant injected and the axial location of the patch. The overall experimental data suggested the importance of the initial cooling film conditions together with better understanding of the possible mechanisms that results in the rapid cooling film break-up, such as high turbulence mainstream flow and scales, and this will lead to a more effective cooling system design. This experimental data is also thought to be ideal for the validation of numerical predictions.
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Yang, Huitao. "Investigations of flow and film cooling on turbine blade edge regions." Texas A&M University, 2006. http://hdl.handle.net/1969.1/4338.
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The inlet temperature of modern gas turbine engines has been increased to achieve higher thermal
efficiency and increased output. The blade edge regions, including the blade tip, the leading edge, and the
platform, are exposed to the most extreme heat loads, and therefore, must be adequately cooled to
maintain safety.
For the blade tip, there is tip leakage flow due to the pressure gradient across the tip. This leakage
flow not only reduces the blade aerodynamic performance, but also yields a high heat load due to the thin
boundary layer and high speed. Various tip configurations, such as plane tip, double side squealer tip, and
single suction side squealer tip, have been studied to find which one is the best configuration to reduce the
tip leakage flow and the heat load. In addition to the flow and heat transfer on the blade tip, film cooling
with various arrangements, including camber line, upstream, and two row configurations, have been
studied. Besides these cases of low inlet/outlet pressure ratio, low temperature, non-rotating, the high
inlet/outlet pressure ratio, high temperature, and rotating cases have been investigated, since they are
closer to real turbine working conditions.
The leading edge of the rotor blade experiences high heat transfer because of the stagnation flow.
Film cooling on the rotor leading edge in a 1-1/2 turbine stage has been numerically studied for the design
and off-design conditions. Simulations find that the increasing rotating speed shifts the stagnation line
from the pressure side, to the leading edge and the suction side, while film cooling protection moves in the
reverse direction with decreasing cooling effectiveness. Film cooling brings a high unsteady intensity of
the heat transfer coefficient, especially on the suction side. The unsteady intensity of film cooling
effectiveness is higher than that of the heat transfer coefficient.
The film cooling on the rotor platform has gained significant attention due to the usage of low-aspect
ratio and low-solidity turbine designs. Film cooling and its heat transfer are strongly influenced by the
secondary flow of the end-wall and the stator-rotor interaction. Numerical predictions have been
performed for the film cooling on the rotating platform of a whole turbine stage. The design conditions
yield a high cooling effectiveness and decrease the cooling effectiveness unsteady intensity, while the high rpm condition dramatically reduces the film cooling effectiveness. High purge flow rates provide a better
cooling protection. In addition, the impact of the turbine work process on film cooling effectiveness and
heat transfer coefficient has been investigated. The overall cooling effectiveness shows a higher value than
the adiabatic effectiveness does.
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Ahn, Jaeyong. "Film cooling effectiveness measurements on rotating and non-rotating turbine components." Texas A&M University, 2005. http://hdl.handle.net/1969.1/4664.
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Detailed film cooling effectiveness distributions were measured on the stationary
blade tip and on the leading edge region of a rotating blade using a Pressure Sensitive
Paint technique. Air and nitrogen gas were used as the film cooling gases and the
oxygen concentration distribution for each case was measured. The film cooling
effectiveness information was obtained from the difference of the oxygen concentration
between air and nitrogen gas cases by applying the mass transfer analogy. In the case of
the stationary blade tip, plane tip and squealer tip blades were used while the film
cooling holes were located (a) along the camber line on the tip or (b) along the span of
the pressure side. The average blowing ratio of the cooling gas was controlled to be 0.5,
1.0, and 2.0. Tests were conducted in a five-bladed linear cascade with a blow down
facility. The free stream Reynolds number, based on the axial chord length and the exit
velocity, was 1,100,000 and the inlet and the exit Mach number were 0.25 and 0.59,
respectively. Turbulence intensity level at the cascade inlet was 9.7%. All
measurements were made at three different tip gap clearances of 1%, 1.5%, and 2.5% of
blade span. Results show that the locations of the film cooling holes and the presence
of squealer have significant effects on surface static pressure and film-cooling effectiveness. Same technique was applied to the rotating turbine blade leading edge
region. Tests were conducted on the first stage rotor of a 3-stage axial turbine. The
Reynolds number based on the axial chord length and the exit velocity was 200,000 and
the total to exit pressure ratio was 1.12 for the first rotor. The effects of the rotational
speed and the blowing ratio were studied. The rotational speed was controlled to be
2400, 2550, and 3000 rpm and the blowing ratio was 0.5, 1.0, and 2.0. Two different
film cooling hole geometries were used; 2-row and 3-row film cooling holes. Results
show that the rotational speed changes the directions of the coolant flows. Blowing
ratio also changes the distributions of the coolant flows. The results of this study will
be helpful in understanding the physical phenomena regarding the film injection and
designing more efficient turbine blades.
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Zuniga, Humberto. "STUDY OF FILM COOLING EFFECTIVENESS: CONICAL, TRENCHED AND ASYMMETRICAL SHAPED HOLES." Doctoral diss., University of Central Florida, 2009. http://digital.library.ucf.edu/cdm/ref/collection/ETD/id/2239.
Full textAbstract:
Film cooling is a technique whereby air from the compressor stage of a gas turbine engine is diverted for cooling purposes to parts, such as the turbine stage, that operate at very high temperatures. Cooling arrangements include impingement jets, finned, ribbed and turbulated channels, and rows of film cooling holes, all of which over the years have become progressively more complex. This costly, but necessary complexity is a result of the industry's push to run engines at increasingly higher turbine inlet temperatures. Higher temperatures mean higher efficiency, but they also mean that the turbine first stage operates hundreds of degrees Kelvin above the melting point of the metal core of the vanes and blades. Existing cooling technology and materials make it possible to protect these parts and allow them to function for extended periods of time--but this comes at a price: the compressed air that is used for cooling represents a considerable penalty in overall turbine efficiency. The aim of current cooling research is threefold: to improve the protection of components from extreme fluxes in order to extend the life of the parts; to increase the inlet turbine operating temperature; and to reduce the amount of air that is diverted from the compressor for cooling. Current film cooling schemes consist of forcing air through carefully machined holes on a part and ejecting it at an angle with the intent of cooling that part by blanketing the surface downstream of the point of ejection. The last major development in the field has been the use of expanded hole exits, which reduce coolant momentum and allow for greater surface coverage. Researchers and designers are continuously looking for novel geometries and arrangements that would increase the level of protection or maintain it while using less coolant.
It was found that the performance of fan-shaped holes inside trenches is actually diminished by the presence of the trench. It is obvious, since the fan diffuses the flow, reducing the momentum of the coolant; the addition of the trench further slows the flow down. This, in turn, leads to the quicker ingestion of the main flow by the jets resulting in lower effectiveness. The next part of the study consisted of systematically increasing the depth of the trench for the fan-shaped holes. The purpose of this was to quantify the effect of the trench on the film cooling effectiveness. It was found that the presence of the trench significantly reduces the film effectiveness, especially for the deeper cases. At the higher blowing ratios, the overall performance of the fans collapses to the same value signifying insensitivity to the blowing ratio. A recent study suggests that having a compound angle could reduce the protective effect of the film due to the elevated interaction between the non-co-flowing coolant jet and the mainstream. Although it has been suggested that a non-symmetric lateral diffusion could mitigate the ill effects of having a compound angle, little has been understood on the effect this non-symmetry has on film cooling effectiveness. The last part of this study investigates the effect of non-symmetric lateral diffusion on film cooling effectiveness by systematically varying one side of a fan-shaped hole. For this part of the study, one of the lateral angles of diffusion of a fan-shaped hole was changed from 5° to 13°, while the other side was kept at 7°. It was found that a lower angle of diffusion hurts performance, while a larger diffusion angle improves it. However, the more significant result was that the jet seemed to be slightly turning. This dissertation investigates such novel methods which one day may include combinations of cylindrical and fan-shaped holes embedded inside trenches, conical holes, or even rows of asymmetric fan-shaped holes. The review of current literature reveals that very few investigations have been done on film cooling effectiveness for uniformly diffusing conical holes. They have been treated as a sort of side novelty since industrial partners often say they are hard to manufacture. To extend our understanding of effectiveness of conical holes, the present study investigates the effect of increasing diffusion angle, as well as the effect of adding a cylindrical entrance length to a conical hole. The measurements were made in the form of film cooling effectiveness and the technique used was temperature sensitive paint. Eight different conical geometries were tested in the form of coupons with rows of holes. The geometry of the holes changed from pure cylindrical holes, a 0° cylindrical baseline, to an 8° pure cone. The coupons were tested in a closed loop wind tunnel at blowing ratios varying from 0.5 to 1.5, and the coolant employed was nitrogen gas. Results indicate that the larger conical holes do, in fact offer appropriate protection and that the holes with the higher expansion angles perform similar to fan-shaped baseline holes, even at the higher blower ratios. The study was also extended to two other plates in which the conical hole was preceded by a cylindrical entry length. The performance of the conical holes improves as a result of the entry length and this is seen at the higher blowing ratios in the form of a delay in the onset of jet detachment. The results of this study show that conical expanding holes are a viable geometry and that their manufacturing can be made easier with a cylindrical entry length, at the same time improving the performance of these holes.
This suggests that the jets actually have two regions: one region with reduced momentum, ideal for protecting a large area downstream of the point of injection; and another region with more integrity which could withstand more aggressive main flow conditions. A further study should be conducted for this geometry at compound angles with the main flow to test this theory. The studies conducted show that the temperature sensitive paint technique can be used to study the performance of film cooling holes for various geometries. The studies also show the film cooling performance of novel geometries and explain why, in some cases, such new arrangements are desirable, and in others, how they can hurt performance. The studies also point in the direction of further investigations in order to advance cooling technology to more effective applications and reduced coolant consumption, the main goal of applied turbine cooling research.
Trench cooling consists of having film cooling holes embedded inside a gap, commonly called a trench. The walls of this gap are commonly vertical with respect to the direction of the main flow and are directly in the path of the coolant. The coolant hits the downstream trench wall which forces it to spread laterally, resulting in more even film coverage downstream than that of regular holes flush with the surface. Recent literature has focused on the effect that trenching has on cylindrical cooling holes only. While the results indicate that trenches are an exciting, promising new geometry derived from the refurbishing process of thermal barrier ceramic coatings, not all the parameters affecting film cooling have been investigated relating to trenched holes. For example, nothing has been said about how far apart holes inside the trench will need to be placed for them to stop interacting. Nothing has been said about shaped holes inside a trench, either. This dissertation explores the extent to which trenching is useful by expanding the PI/D from 4 to 12 for rows of round and fan holes. In addition the effect that trenching has on fan-shaped holes is studied by systematically increasing the trench depth. Values of local, laterally-averaged and spatially-averaged film cooling effectiveness are reported. It is found that placing the cylinders inside the trench and doubling the distance between the holes provides better performance than the cylindrical, non-trenched baseline, especially at the higher blowing ratios, M [greater than] 1.0. At these higher coolant flow rates, the regular cylindrical jets show detachment, while those in the trench do not. They, in fact perform very well. The importance of this finding implies that the number of holes, and coolant, can be cut in half while still improving performance over regular holes. The trenched cylindrical holes did not, however, perform like the fan shaped holes.Ph.D.
Department of Mechanical, Materials and Aerospace Engineering;
Engineering and Computer Science
Mechanical Engineering PhD
50
Khaldi, A. "Discharge coefficient of film cooling holes with rounded entries or exits." Thesis, University of Nottingham, 1987. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.378758.
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