With the continued economic and socio-political pressure on aircraft manufacturers to produce more profitable and environmentally-friendly aircraft, the drive towards increasingly more efficient aircraft engines remains of prime importance to aircraft engine manufacturers. While the majority of axial flow turbomachines use cylindrically shaped endwalls between the blades on the hub or shroud, non-axisymmetric endwall contouring is a reasonably recent technique which relaxes this constraint, and allows the geometry of the endwalls to depart from that of a plain cylinder. Although a number of studies have shown non-axisymmetric endwall contouring to be an effective mechanism for the reduction of secondary flows (and the losses associated with them), within the open literature there still remains a general lack of detailed information relating to the optimal design of these devices. Among some of the most important issues which remain unresolved, are uncertainties such as: “What is the best way to identify and thereafter quantify the strength of turbine secondary flows?”, and thereafter, as a natural progression from this, “Of the metrics which are currently found within the literature, which are best for use in the design of secondary loss mitigating endwall contours for a real turbine?”. Some of the reasons for the lack of information as described above, result from the undertaking of many of the investigations into the design of endwall contours by or on behalf of the major engine manufacturers, and therefore, a general inability or perhaps even unwillingness to divulge many of the specific details related to the methodologies and quantities used as a result of the commercial sensitivity of these investigations. In addition to this, as a result of the relatively large number and diverse nature of groups involved in non-axisymmetric endwall contouring research, within the literature which has been made available, there exists a wide variety of different test geometries as well as conditions which have been used, making a neutral determination of the most successful approach to endwall contouring considerably more difficult. This thesis documents the design and testing of a number of different non-axisymmetric endwall configurations intended to produce flow conditions optimized using a selection of the metrics commonly found in the literature, for the rotor of a low speed, research turbine, whose baseline as well as performance using contoured endwalls has been reported on previously, in order to establish which of these metrics is the most effective. As part of this process, a fully validated computational fluid dynamics model of the turbine downstream of the first nozzle was developed and incorporated into an automated non-axisymmetric end- wall design routine, capable of producing endwall contours optimized for various objective functions. Numerical testing showed that, in order to distinguish accurately between the various endwall configurations, relatively fine computational meshes were required and therefore, as a result of corresponding computational expense associated with these meshes, the implementation of a surrogate modelling procedure in which part of this computational cost is offset by mathematical modelling, was necessary. Altogether, a total of 8 endwall designs were produced - 6 using a single metric each as the basis of their objective functions (the ‘simple’ designs) and a further 2 so-called ‘compound’ designs. Of the simple designs, the best performing endwalls in terms of improvements to the rotor exit efficiency were the ηtt-, Cske- & βdev-based designs, which were based in turn on the rotor total-total efficiency (ηtt), coefficient of secondary kinetic energy (Cske) and flow deviation from design angle (βdev) respectively. All three of these designs were predicted to result in very similar changes to the secondary flow characteristics although the increasing bias towards flow correction was found to have an inverse correlation with the overall efficiencies predicted for each rotor. Of these designs, the numerical predictions for both the ηtt- & Cske-based designs (which were included in the experimental subset), were found to be validated, at both the rotor exit as well as downstream measurement planes. Further to this (with the exception of the Cp0,rel-based case), although the remainder of the simple designs (i.e. the SKEH & ηde-based designs) were also predicted to improve the overall rotor efficiency, either the form or the performance of these endwalls resulted in the final corresponding designs for these metrics being considered unsatisfactory. Finally, the two ‘compound’ metrics were both formulated to to include a term designed to target the secondary flow within the target blade row, as well as an additional term which was designed to promote improvement in the flow into the downstream blade row. While both designs produced using the compound design objective functions were predicted to improve both the conditions for the target blade row, as well as the flow quality at the exit of the blade row, flow separations at the exit of the contoured regions for both designs resulted in only partial validation of each design when tested experimentally. Finally, although both designs were once again predicted to perform very well at the ‘mixed-out’ measurement plane, these predictions were found to be only partially validated by the experiment.

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.

This experiment investigated the effects of a non-axisymmetric endwall contour and upstream purge flow on the secondary flow of an inlet guide vane. Three cases were tested in a transonic wind tunnel with an exit Mach number of 0.93-a flat endwall with no upstream purge flow, the same flat endwall with upstream purge flow, and a 3D contoured endwall with upstream purge flow. All cases had a backward-facing step upstream of the vanes. Five-hole probe measurements were taken 0.2, 0.4, and 0.6 Cx downstream of the vane row trailing edge, and were used to calculate loss coefficient, secondary velocity, and secondary kinetic energy. Additionally, surface static pressure measurements were taken to determine the vane loading at 4% spanwise position. Surface oil flow visualizations were performed to analyze the flow qualitatively. No statistically significant differences were found between the three cases in mass averaged downstream measurements. The contoured endwall redistributed losses, rather than making an improvement distinguishable beyond experimental uncertainty. Flow visualization found that the passage vortex penetrated further in the spanwise direction into the passage for the contoured endwall (compared to the flat endwall), and stayed closer to the endwall with a blowing ratio of 1.5 with a flat endwall (compared to no blowing with flat endwall). This was corroborated by the five hole probe results. / Master of Science

In gas turbine development, the direction has been toward higher turbine inlet temperatures to increase the work output and thermal efficiency. This extreme environment can significantly impact component life. One means of preventing component burnout in the turbine is to effectively use film-cooling whereby coolant is extracted from the compressor and injected through component surfaces. One such surface is the endwall of the first stage nozzle guide vane. This thesis details the design, prediction, and testing of two endwall film-cooling hole patterns provided by leading gas turbine engine companies. In addition a flush, two-dimensional slot was included to simulate leakage flow from the combustor-turbine interface. The slot coolant was found to exit in a non-uniform manner leaving a large, uncooled ring around the vane. Film-cooling holes were effective at distributing coolant throughout much of the passage, but at low blowing rates were unable to provide any benefit to the critical vane-endwall junction both at the leading edge and along the pressure side. At high blowing ratios, the increased momentum of the jets induced separation at the leading edge and in the upstream portion of the passage along the pressure side, while the jets near the passage exit remained attached and penetrated completely to the vane surface. Computational fluid dynamics (CFD) was successful at predicting coolant trajectory, but tended to under-predict thermal spreading and jet separation. Superposition was shown to be inaccurate, over-predicting effectiveness levels and thus component life, because the flow field was altered by the coolant injection. / Master of Science

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

Engine development has been in the direction of increased turbine inlet temperatures to improve efficiency and power output. Secondary flows develop as a result of a near-wall pressure gradient in the stagnating flow approaching the inlet nozzle guide vane as well as a strong cross-passage gradient within the passage. These flow structures enhance heat transfer and convect hot core flow gases onto component surfaces. In modern engines it has become critical to cool component surfaces to extend part life. Bypass leakage flow emerging from the slot between the combustor and turbine endwalls can be utilized for cooling purposes if properly designed. This study examines a three-dimensional slot geometry, scalloped to manipulated leakage flow distribution. Statistical techniques are used to decouple the effects of four geometric parameters and quantify the relative influence of each on endwall cooling levels and near-wall total pressure losses. The slot geometry is also optimized for robustness across a range of inlet conditions. Average upstream distance to the slot is shown to dominate overall cooling levels with nominal slot width gaining influence at higher leakage flow rates. Scalloping amplitude is most influential to near-wall total pressure loss as formation of the horseshoe vortex and cross flow within the passage are affected. Scalloping phase alters local cooling levels as leakage injection is shifted laterally across the endwall. / Ph. D.

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.

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

This study describes a detailed investigation on the effects that upstream step misalignment and upstream purge film cooling have on the endwall heat transfer for nozzle guide vanes in a land based power generation gas turbine at transonic conditions. Endwall Nusselt Number and adiabatic film cooling effectiveness distributions were experimentally calculated and compared with qualitative data gathered via oil paint flow visualization which also depicts endwall flow physics. Tests were conducted in a transonic linear cascade blowdown facility. Data were gathered at an exit Mach number of 0.85 with a freestream turbulence intensity of 16% at a Re = 1.5 x 106 based on axial chord. Varied upstream purge blowing ratios and a no blowing case were tested for 3 different upstream step geometries, one of which was the baseline (no step). The other two geometries are a backward step geometry and a forward step geometry, which comprised of a span-wise upstream step of +4.86% span and -4.86% span respectively. Experimentation shows that the addition of upstream purge film cooling increases the Nusselt Number at injection upwards of 50% but lowers it in the throat of the passage by approximately 20%. The addition of a backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow, thus aiding to help keep the film attached to the endwall at higher blowing ratios. Increasing the blowing ratio increases film cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff. Near endwall streamlines without purge cooling generated by Li et al. [1] for the same geometries were compared to the experimental data. It was shown that even with the addition of upstream purge cooling, the near endwall streamlines as they moved downstream matched strikingly well with the experimental data. This discovery indicates that while the coolant flow will likely affect the flow streamlines three dimensionally, they are minimally effected by the coolant flow near the endwall as the flow moves downstream. / Master of Science

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