In this work, we study the enhanced endwall heat transfer for flow past non conducting pin fin arrays. The aim is to resolve the controversy over the heat transfer that is taking place from the endwall and the pin surface.Various parameters were studied and results were obtained. Our results are found to be consistent with some of the results that have been previously published. The results were surprisingly found to be dependent on the height of the pin fin.

Engine designers continually push the combustor exit temperature higher to produce more power from gas turbine engines. These high turbine inlet temperatures, coupled with high turbulence levels and flow field non-uniformities, make turbine vane and endwall cooling a very critical issue in engine design. To appropriately cool these surfaces, knowledge of the passage flow field and endwall temperature distribution at representative engine conditions is necessary. A combustor test section was used to simulate realistic turbine inlet profiles of turbulence, normalized temperature, normalized total pressure, and normalized streamwise velocity to study the flow field in a turbine vane passage and the adiabatic temperature distribution on the endwall. The combustor liner film-cooling and exit slot flows were varied independently to determine their relative effect on endwall cooling in the downstream turbine vane. Flow field measurements revealed the presence of a previously unmeasured third vortex in the vane passage. The tertiary vortex was located above the passage vortex and had rotation opposite to the passage vortex. Increasing the amount of slot flow reduced the size and strength of the nearwall vortices, while increasing the size and strength of the tertiary vortex. Adiabatic endwall temperature measurements revealed higher temperatures surrounding the base of the vane. The endwall measurements also showed that the exit slot flow was effective at cooling only a region of the endwall near the vane leading edge on the suction side. Increasing slot flow was found to have a larger thermal benefit to the endwall relative to increasing combustor liner film-cooling. / Master of Science

This paper presents details of an experimental and computational investigation on the effect of misalignment between the combustor exit and nozzle guide vane endwall on the heat transfer distribution across an axisymmetric converging endwall. The axisymmetric converging endwall investigated was representative of that found on the shroud side of a first stage turbine nozzle section. The experiment was conducted at a nominal exit M of 0.85 and exit Re 1.5 x 10⁶ with an inlet turbulence intensity of 16%. The experiment was conducted in a blowdown transonic linear cascade wind tunnel. Two different inlet configurations were investigated. The first configuration, Case I, was representative of a combustor exit aligned to the nozzle platform, with a gap located at the interface of the tow components. The second configuration, Case II, the endwall platform was offset in the span-wise direction to create a backward facing step at the inlet. This step is representative of a misalignment between the combustor exit and the NGV platform. An infrared camera was used to capture the temperature history on the endwall, from which the endwall heat transfer distribution was determined. A numerical study was also conducted by solving RANS equations using ANSYS Fluent v.15. The numerical results provided insight into the passage flow field which explained the observed heat transfer characteristics. Case I showed the typical characteristics of transonic vane cascade flow, such as the separation line, saddle point, and horseshoe vortices. The presence of a gap at the combustor-nozzle interface facilitated the formation of a separated flow which propagated through the passage. This flow feature caused the passage vortex reattach to the SS vane at 0.44 x/C. The addition of the platform misalignment in Case II caused the flow reattachment region to occur near the vane LE plane. The separated flow which formed at the inlet step, merged with the recirculation region on the endwall platform, forming two counter-rotating auxiliary vortices. These vortices significantly delayed migration of the passage vortex, causing it to reattach on the SS vane at 0.85 x/C. These two flow features also had a significant effect on the endwall heat transfer characteristics. The heat transfer levels on the endwall platform, from -0.50 to +0.50 Cx relative to the vane LE, had an average increase of ~40%. However, downstream of the vane mid-passage, the heat transfer levels showed no appreciable heat transfer augmentation due to flow acceleration through the passage throat. / Master of Science

The study presented herein describes and analyzes a detailed experimental investigation of the effects of density ratio on endwall thermal performance at varying blowing rates for a typical nozzle guide vane platform purge jet cooling scheme. An axisymmetric converging endwall with an upstream doublet staggered cylindrical hole purge jet cooling scheme was employed. Nominal exit flow conditions were engine representative and as follows: {rm Ma}_{Exit} = 0.85, {rm Re}_{Exit,C_{ax}} = 1.5 times {10}^6, and large-scale freestream Tu = 16%. Two blowing ratios were investigated corresponding to the upper and lower engine extrema. Each blowing ratio was investigated amid two density ratios; one representing typical experimental neglect of density ratio, at DR = 1.2, and another engine representative density ratio achieved by mixing foreign gases, DR = 1.95. All tests were conducted on a linear cascade in the Virginia Tech Transonic Blowdown Wind Tunnel using IR thermography and transient data reduction techniques. Oil paint flow visualization techniques were used to gather quantitative information regarding the alteration of endwall flow physics due two different blowing rates of high-density coolant. High resolution endwall adiabatic film cooling effectiveness, Nusselt number, and Net Heat Flux Reduction contour plots were used to analyze the thermal effects. The effect of density is dependent on the coolant blowing rate and varies greatly from the high to low blowing condition. At the low blowing condition better near-hole film cooling performance and heat transfer reduction is facilitated with increasing density. However, high density coolant at low blowing rates isn't adequately equipped to penetrate and suppress secondary flows, leaving the SS and PS largely exposed to high velocity and temperature mainstream gases. Conversely, it is observed that density ratio only marginally affects the high blowing condition, as the momentum effects become increasingly dominant. Overall it is concluded density ratio has a first order impact on the secondary flow alterations and subsequent heat transfer distributions that occur as a result of coolant injection and should be accounted for in purge jet cooling scheme design and analysis. Additionally, the effect of increasing high density coolant blowing rate was analyzed. Oil paint flow visualization indicated that significant secondary flow suppression occurs as a result of increasing the blowing rate of high-density coolant. Endwall adiabatic film cooling effectiveness, Nusselt number, and NHFR comparisons confirm this. Low blowing rate coolant has a more favorable thermal impact in the upstream region of the passage, especially near injection. The low momentum of the coolant is eventually dominated and entrained by secondary flows, providing less effectiveness near PS, near SS, and into the throat of the passage. The high momentum present for the high blowing rate, high-density coolant suppresses these secondary flows and provides enhanced cooling in the throat and in high secondary flow regions. However, the increased turbulence impartation due to lift off has an adverse effect on the heat load in the upstream region of the passage. It is concluded that only marginal gains near the throat of the passage are observed with an increase in high density coolant blowing rate, but severe thermal penalty is observed near the passage onset. / Master of Science / Gas turbine technology is used frequently in the burning of natural gas for power production. Increases in engine efficiency are observed with increasing firing temperatures, however this leads to the potential of overheating in the stages following. To prevent failure or melting of components, cooler air is extracted from the upstream compressor section and used to cool these components through various highly complex cooling schemes. The design and operational adequacy of these schemes is highly subject to the mainstream and coolant flow conditions, which are hard to represent in a laboratory setting. This experimental study explores the effects of various coolant conditions, and their respective response, for a purge jet cooling scheme commonly found in engine. This scheme utilizes two rows of staggered cylindrical holes to inject air into the mainstream from platform, upstream of the nozzle guide vane. It is the hope that this air forms a protective layer, effectively shielding the platform from the hostile mainstream conditions. Currently, little research has been done to quantify these effects of purge flow cooling scheme while mimicking engine geometry, mainstream and coolant conditions. For this study, an endwall geometry like that found in engine with a purge jet cooling scheme is studied. Commonly, an upstream gap is formed between the combustor lining and first stage vane platform, which is accounted for in this testing. Mainstream and coolant flow conditions can have large impacts on the results gathered, so both were matched to engine conditions. Varying of coolant density and injection rate is studied and quantitative results are gathered. Results indicate coolant fluid density plays a large role in purge jet cooling, and with neglection of this, potential thermal failure points could be overlooked This is exacerbated with less coolant injection. Interestingly, increasing the amount of coolant injected decreases performance across much of the passage, with only marginal gains in regions of complex flow. These results help to better explain the impacts of experimental neglect of coolant density, and aid in the understanding of purge jet coolant injection.

The primary requirements for a modern industrial gas turbine consist of a continuous trend of an increasing efficiency combined with very low emissions in a robust, cost-effective manner. To fulfil these tasks a high turbine inlet temperature together with advanced dry low NOX combustion chambers are employed. These dry low NOX combustion chambers generate a rather flat temperature profile compared to previous generation gas turbines, which have a rather parabolic temperature profile before the nozzle guide vane. This means that the nozzle guide vane endwall heat load for modern gas turbines is much higher compared to previous generation gas turbines. Therefore the prediction of the nozzle guide vane flow field and endwall heat transfer is crucial for the engineering task of the design layout of the vane endwall cooling system. The present study is directed towards establishing new in-depth aerodynamic and endwall heat transfer knowledge for an advanced nozzle guide vane of a modern industrial gas turbine. To reach this objective the physical processes and effects which cause the different flow fields and the endwall heat transfer pattern in a baseline configuration, a combustion chamber variant, a heat shield variant without and with additional cooling air and a cavity variant without and with additional cooling air have been investigated. The variants, which differ from the simplified baseline configuration, apply design elements which are commonly used in real modern gas turbines. This research area is crucial for the nozzle guide vane endwall heat transfer, especially for the advanced design of the nozzle guide vane of a modern industrial gas turbine and has so far hardly been investigated in the open literature. For the experimental aerodynamic and endwall heat transfer research of the baseline configuration of the advanced nozzle guide vane geometry a new low pressure, low temperature test facility has been developed, designed and constructed, since no experimental heat transfer data exist in the open literature for this type of vane configuration. The new test rig consists of a linear cascade with the baseline configuration of the advanced nozzle guide vane geometry with four upscaled airfoils and three flow passages. For the aerodynamic tests the two middle airfoils and the hub and the tip endwall are instrumented with pressure taps to monitor the Mach number distribution. For the heat transfer tests the temperature distribution on the hub endwall is measured via thermography. The analysis of these measurements, including comparisons to research in the open literature shows that the new test rig generates accurate and reproducible results which give confidence that it is a reliable tool for the experimental aerodynamic and heat transfer research on the advanced nozzle guide vane of a modern industrial gas turbine. Previous own research work together with the numerical analysis performed in another part of the project as well as conclusions from a detailed literature study lead to the conclusion that advanced Navier-Stokes CFD tools with the v2-f turbulence model are most suitable for the calculation of the flow field and the endwall heat transfer of turbine vanes and blades. Therefore this numerical tool, validated against different vane and blade geometries and for different flow conditions, has been chosen for the numerical aerodynamic and endwall heat transfer research of the advanced nozzle guide vane of a modern industrial gas turbine. The evaluation of the numerical and experimental investigations of the baseline configuration of the advanced design of a nozzle guide vane shows the flow field of an advanced mid-loaded airfoil design with the features to reduce total airfoil losses. For the hub endwall of the baseline configuration of the advanced design of a nozzle guide vane the flow characteristics and heat transfer features of the classical vane endwall secondary flow model can be detected with a very weak intensity and geometric extension compared to the studies of less advanced vane geometries in the open literature. A detailed analysis of the numerical simulations and the experimental data showed very good qualitative and quantitative agreement for the three-dimensional flow field and the endwall heat transfer. These findings, together with the evaluations obtained from the open literature, lead to the conclusions that selected CFD software Fluent together with the applied v2-f turbulence model exhibits a high level of general applicability and is not tuned to a special vane or blade geometry. Therefore the CFD code Fluent with the v2-f turbulence model has been selected for the research of the influence of the several geometric variants of the baseline configuration on the flow field and the hub endwall heat transfer of the advanced nozzle guide vane of a modern industrial gas turbine. Most of the vane endwall heat transfer research in the open literature has been carried out only for baseline configurations of the flow path between combustion chamber and nozzle guide vane. Such a simplified geometry consists of a long, planar undisturbed approach length upstream of the nozzle guide vane. The design of real modern industrial gas turbines however requires often significant variations from this baseline configuration consisting of air-cooled heat shields and purged cavities between the combustion chamber and the nozzle guide vane. A detailed evaluation of the flow field and the endwall heat transfer shows major differences between the baseline and the heat shield configuration. The heat shield in front of the airfoil of the nozzle guide vane influences the secondary flow field and the endwall heat transfer pattern strongly. Additional cooling air, released under the heat shield has a distinctive influence as well. Also the cavity between the combustion chamber and the nozzle guide vane affects the secondary flow field and the endwall heat transfer pattern. Here the influence of additional cavity cooling air is more decisive. The results of the detailed studies of the geometric variants are applied to formulate guidelines for an optimized design of the flow path between the combustion chamber and the nozzle guide vane and the nozzle guide vane endwall cooling configuration of next-generation industrial gas turbines. / QC 20100917