Dissertations / Theses on the topic 'Nozzle guide vane'

The continuous demand from the airlines for reduced jet engine fuel consumption results in increasingly challenging high pressure turbine nozzle guide vane (NGV) working conditions. The capability to reproduce representative boundary conditions in a rig at the combustor-turbine interaction plane is a key feature when testing NGVs in an engine-representative environment. A large scale linear cascade rig to investigate NGV leading edge cooling systems has been designed with particular attention being paid to creating engine representative conditions at the NGV inlet plane. The combustor simulator replicates the main features of a rich-burn design including large dilution jets and extensive endwall film cooling. CFD simulations have been used to develop the design which matches Reynolds number and mainstream-to-dilution jet momentum flux ratio. Detailed measurements of velocity, turbulence and temperature have been acquired at the NGV inlet plane. A thermo-couple was manufactured from 12.7 Î1⁄4m diameter wire and carefully calibrated to obtain its time constant in the velocity range of interest. The results are compared to CFD predictions and data in the literature. The time-averaged measurements show that the flow field conditions are dominated by the endwall cooling flows. The time-resolved data show that the measured turbulence length scale reflects the scale of the relevant upstream jets while the spectrum of temperature fluctuations reports a thermal cascade independent of any geometrical features. Attention was also focused on the flow field downstream of different endwall film cooling holes configurations: three arrangements of a double row of staggered cylindrical holes (lateral pitch-to-diameter ratio of 2 - 3 - 6) and one with intersecting holes (intersecting angle of 90o) were experimentally and numerically analyzed. The research quantified the extent by which closer spaced hole configurations provide more effective film coverage. It was found that the turbulent integral length scales are strongly connected to the hole diameter and spacing. It was also found that intersecting holes can potentially reduce the amount of required coolant at a fixed pressure ratio, but offer worst film performance than cylindrical holes. RANS simulations proved successful at predicting the main trends shown by the measurements. A new concept to increase the pressure margin across the film cooling holes in a specific region of vane LE coolant passage was introduced and developed: an insert was used to cover the area with the highest risk of ingestion, slowing down the flow and increasing the local static pressure. Numerical simulations were initially used to compare different designs and to analyse the impact of the insert on the overall coolant flow distribution. In particular, the effect on the static pressure downstream of the insert was identified as a critical factor that needs to be taken into account during the design process in order to avoid hot gas ingestion in other areas. The experimental campaign proved the ability of this new design to significantly increase the pressure margin in the covered region.

The entropy noise in modern engines is mainly originating from two types of mechanisms.First, chemical reactions in the combustion chamber lead to unsteady heat releasewhich is responsible of the direct combustion noise. Second, hot and cold blobsof air coming from the combustion chamber are advected and accelerated throughturbine stages, giving rise to the so-called entropy noise (or indirect combustionnoise). In the present work, numerical characterization of indirect combustion noiseof a Nozzle Guide Vane passage was assessed using three-dimensional Large EddySimulations. The study was conducted on a simplified topology of a real turbinestator passage, for which experimental data were available in transonic operatingconditions. First, a baseline case was reproduced to validate a numerical finite volumesolver against the experimental measurements. Then, the same solver is used toreproduce the effects of incoming entropy waves from the combustion chamber andto characterize the additional generated acoustic power. Periodic temperature fluctuationsare imposed at the inlet, permitting to simulate hot and cold packets of aircoming from the unsteady combustion. The incoming waves are characterized bytheir characteristic wavelength; therefore, a parametric study has been conductedvarying the inlet temperature of the passage, generating entropy waves of greaterwavelengths. The study proves that the generated indirect combustion noise canbe significant. Moreover, the generated indirect combustion noise increases as thewavelength of the incoming disturbances increases. Finally, the present work suggeststhat, in transonic conditions, there might be flow features which enhance theindirect combustion noise generation mechanism.

Despite many new investigations over the last years, there is no indication that alternative energy conversion technologies will overtake the place of turbomachinery. Hence thermal turbines are still the most dominant movers for electricity generation.Although this leadership in the energy production does not seem to be in danger, the current drivers in turbomachinery industry are to work towards both less fuel consumption and less pollution. In order to meet the future economic and environmental goals, researchers press towards highly loaded vanes and blades. This has to be performed at maintained or improved aerodynamic performances. Increased performances and blade loading lead in turn to increased velocities and larger regions of supersonic fluid velocities and consequently general increasing of shock intensities. The biggest problem dealing with supersonic flow and high shock intensities is that the boundary layer, when walking through these regions, experiences strong pressure gradients and intense shock-boundary layer interaction. This may lead the blade to stall meaning detachment of both boundary layer and cooling-film from the wall. These effects can evidently lead to catastrophic consequences since nowadays the materials used in turbomachinery applications have temperature strengths much lower than those coming from the combustion chamber. This thanks to very complex blade and vane cooling systems.There are even other features that may take benefit from increased velocities such as an attenuation in the boundary layer growth and the static pressure distribution on the blade surface. For helping researchers studying these new geometries, a cold air annular test rig designed by “Siemens Industrial Turbomachinery AB”, it has been built and placed at “Division of Heat and Power Technology” at KTH.The present thesis has the goal to provide a numerical model for CFD calculations, optimized for boundary layer studies, able to give a good prediction of detachment of the boundary layer and losses for different working cases. A previous model was provided with a commercial software for both ideal vane and real test rig. Recover of results and adaptions of the model were performed with a new version of the same software starting from the previous model. A comparison between numerical and experimental results have shown a good match for the subsonic and transonic case. Instead, problems were met for the supersonic case. Many attempts of different boundary condition at the inlet have been run. No reliable solution has been reached with realistic pressure profile at inlet while realistic results have been found using the mass flow rate as Inlet boundary condition. At the end, an analysis of shock and detachment is provided in terms of density gradient and static entropy distribution through the blade passage. Future works may aim to solve the “supersonic problem”.

This thesis presents a computational fluid dynamic (CFD) study of high pressure gas turbine blade design with different realistic inlet temperature and velocity boundary conditions. The effects of blade shaping and inlet conditions can only be fully understood by considering the aerodynamics and heat transfer concurrently; this is in contrast to the sequential method of blade design for aerodynamics followed by cooling. The inlet boundary conditions to the NGV simulations are governed by the existence of discrete fuel injectors in the combustion chamber. An appreciation of NGV shaping design under engine realistic inflow conditions will allow for an identification of the correct three dimensional shaping parameters that should be considered for design optimisation. The Rolls-Royce efficient Navier-Stokes solver, HYDRA, was employed in all computational results for a transonic turbine stage. The single passage unsteady method based on the Fourier Shape Correction is adopted. The solver is validated under both rich burn (hot steak only) and the case with swirl inlet profiles for aerothermal characteristics; good agreement is noted with the validation data. Post processing methods were used in order to obtain time-averaged results and blade visualisations. Subsequently, a surrogate design optimisation methodology using machine learning combined with a Genetic Algorithm is implemented and validated. A study of the effect of NGV compound lean on stage performance is carried out and contrasted for uniform and rich burn inlets, and subsequently for lean burn. Compound lean is shown to produce a tip uploading at the rotor inlet, which is beneficial for rich burn, but detrimental for lean burn. It is also found that for rich burn, fluid driving temperature is more dominant than HTC in determining rotor blade heat transfer, the opposite sense to the uniform inlet. Also, for a lean burn inlet, there is another role reversal, with HTC dominating fluid driving temperature in determining heat transfer. A novel NGV design methodology is proposed that seeks to mitigate the combined effects of inlet hot streak and swirling flow. In essence, the concept two NGVs in a pair are shaped independently of each other, thus allowing the inlet flow non uniformity to be suitably accommodated. Finally, two numerical NGV optimisation studies are undertaken for the combined hot streak and swirl inlet for two clocking positions; vane impinging and passage aligned. Due to the prohibitive cost of unsteady CFD simulations for an optimisation strategy, a suitable objective function at the NGV exit plane is used to minimise rotor tip heat flux. The optimised shape for the passage case resulted in the lowest tip heat flux distribution, however the optimum shape for the impinging case led to the highest gain in stage efficiency. This therefore suggests that NGV lean and clocking position should be a consideration for future optimisation and design of the HP stage.

Thèse sciences techniques, EPF Lausanne, no 2602 (2002), Faculté Sciences et techniques de l'ingénieur, Domaine du génie mécanique. Directeur: A. Bölcs ; rapporteurs: M. Gritsch, J. Thome, B. Weigand.

The possibility of misalignments at interfaces would be increased due to individual parts' assembly and external factors during its operation. In actual engine representative conditions, the upstream misalignments have effects on turbines performance through the nozzle guide vane passages. The current experimental aerodynamic investigation over the nozzle guide vane passage was concentrated on the backward-facing step of upstream misalignments. The tests were performed using two types of vane endwall platforms in a 2D linear cascade: flat endwall and axisymmetric converging endwall. The test conditions were a Mach number of 0.85, Re_ex 1.5*10^6 based on exit condition and axial chord, and a high freestream turbulence intensity (16%), at the Virginia tech transonic cascade wind tunnel. The experimental results from the surface flow visualization and the five-hole probe measurements at the vane-passage exit were compared with the two cases with and without the backward-facing step for both types of endwall platforms. As a main source of secondary flow, a horseshoe vortex at stagnation region of the leading edge of the vane directly influences other secondary flows. The intensity of the vortex is associated with boundary layer thickness of inlet flow. In this regard, the upstream backward-facing step as a misalignment induces the separation and attachment of the inlet flow sequentially, and these cause the boundary layer of the inlet flow to reform and become thinner locally. The upstream-step positively affects loss reduction in aerodynamics due to the thinner inlet boundary layer, which attenuates a horseshoe vortex ahead of the vane cascade despite the development of the additional vortices. And converging endwall results in an increase of the effect of the upstream misalignment in aerodynamics, since the inlet boundary layer becomes thinner near the vane's leading edge due to local flow acceleration caused by steep contraction of the converging endwall. These results show good correlation with many previous studies presented herein.
Master of Science

Heat transfer study on gas turbine blades is very important due to the resultant increase in cycle thermal efficiency. This study is focused on the heat transfer effects on a reference nozzle guide vane and test rig component preparation in heat and power technology division at KTH. In order to prepare the current test rig for heat transfer experiments, some feature should be changed in the current layout to give a nearly instant temperature rise for heat transfer measurement. The heater mesh component is the main component to be added to the current test rig. Some preliminary design parameters were set and the necessary power for the heater mesh to achieve required step temperature rise was calculated. For the next step, it is needed to estimate the heat transfer coefficient and the other parameters for study on the reference blade using numerical methods. Boundary layer analysis is very important in heat transfer modeling so to model the reference blade heat transfer and boundary layer properties, a 2D boundary layer code TEXSTAN is used and the velocity distribution around the vane was set to an input dataset file. After elements refinement to ensure the numerical accuracy of TEXSTAN code, various turbulence modeling was check to predict the heat transfer coefficient and boundary layer assessments. It was concluded from TEXTAN calculations that both suction and pressure side have transition flow while for the suction side it was predicted that the flow regime at trailing edge is fully turbulent. Based on the Abu-Ghannam –Shaw Transition model and by the aid of shape factor data, momentum Reynolds number and various boundary layer properties, it was concluded that the pressure side remains in transient region.

Increasing the combustion temperatures in a gas turbine engine to achieve higher efficiency and power output also results in high heat loads to turbine components downstream of the combustor. The challenge of adequately cooling the nozzle guide vane directly downstream of the combustor is compounded by a complex vortical secondary flow at the junction of the endwall and the airfoil. This flow tends to increase local heat transfer rates and sweep coolant away from component surfaces, as well as decrease the turbine aerodynamic efficiency. Past research has shown that a large fillet at the endwall-airfoil junction can reduce or eliminate the secondary flow. Also, leakage flow from the interface gap between the combustor and the turbine can provide some cooling to the endwall. This study examines the individual and combined effects of a large fillet and realistic combustor-turbine interface gap leakage flow for a nozzle guide vane. The first study focuses on the effect of leakage flow from the interface gap on the endwall upstream of the vane. The second study addresses the influence of large fillets at the endwall-airfoil junction, with and without upstream leakage flow. Both studies were performed in a large low-speed wind tunnel with the same vane geometry. Endwall shear stress measurements were obtained for various endwall-airfoil junction geometries without upstream leakage flow. Endwall heat transfer and cooling effectiveness were measured for various leakage flow rates and leakage gap widths, with a variety of endwall-airfoil junction geometries.

Results from these studies indicate that the secondary flow has a large influence on the coverage area of the leakage coolant. Increased leakage flow rates resulted in better cooling effectiveness and coverage, but also higher heat transfer rates. The two fillet geometries tested affected coolant coverage by displacing coolant around the base of the fillet, which could result in undesirably high gradients in endwall temperature. The addition of a large fillet to the endwall-airfoil junction, however, reduced heat transfer, even when upstream leakage flow was present.
Master of Science

In this study, the influence of inlet flow incidence on the aerodynamic and thermal performance of a film cooled linear nozzle vane cascade is fully assessed. Tests have been carried out on a solid and a cooled cascade. In the cooled cascade, coolant is ejected at the end wall through a slot located upstream of the leading edge plane. Moreover, a vane showerhead cooling system is also realized through 4 rows of cylindrical holes. The cascade was tested at a high inlet turbulence intensity level (Tu1 = 9%) and at a constant inlet Mach number of 0.12 and nominal cooling condition, varying the inlet flow angle. In addition to the reference incidence angle (0°), four other cases were investigated: +20°, +10°, -10° and -20°. The aero-thermal characterization of vane platform was obtained through 5-hole probe, endwall and vane showerhead adiabatic film cooling effectiveness measurements. Vane load distributions and surface flow visualizations supported the discussion of the results. On the vane, a significant movement in stagnation point happened when incidence angle varied, resulted in changing of the coolant distribution pattern between SS and PS of the cooled vane; which adversely affects the efficiency for both negative and positive inlet flow incidence angles. On the platform, however, a relevant negative impact of positive inlet flow incidence on the cooled cascade aerodynamic and endwall thermal performance was detected. A negligible influence was instead observed at negative incidence, even at the lowest tested value of -20°.

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

Over the last two decades, most developed countries have reached a consensus that greener energy production is necessary for the world, due to the climate changes and limited fossil fuel resources. More efficient turbine is desirable and can be archived by higher turbine-inlet temperature (TIT). However, it is difficult for nozzle guide vane (NGV), which is the first stage after combustion chamber, to withstand a very high temperature. Thus, cooling methods such as film cooling have to be implemented. Film-cooled NGV of an annular sector cascade (ASC) is studied in this thesis, for getting comprehensive calculation of vorticity, and analyzing applicability of existing loss models, namely Hartsel model and Young & Wilcock model. The flow-field calculation methods from previously published studies are reviewed. Literatures focusing on Hartsel model and Young & Wilcock model are studied. Measurement data from previously published studies are analyzed and compared with the loss models. In order to get experience of how measurements take place, participation of a test run experiment is involved. Calculation of flow vector has been evaluated and modified. Actual flow angle is introduced when calculating velocity components. Thus, more exact results are obtained from the new method. Calculation of vorticity has been evaluated and made more comprehensive. Vorticity components as well as magnitude of total streamwise vorticity are calculated and visualized. Vorticity is higher and more extensive for fully cooled case than uncooled case. Highest vorticity is found at regions near the hub, tip and TE. Axial and circumferential vorticities show similar patterns, while the radial vorticity is relatively simpler. Compressibility is introduced as a new method when calculating circumferential and radial vorticities, resulting more extensive and higher vorticities than results from incompressible solutions. Hartsel model and Young & Wilcock model have been evaluated and compared to the ASC to see the applicability of the models. In general, Hartsel model cannot agree with the ASC to a satisfactory level and thus cannot be applied. Coolant velocity is found to be the dominant factor of Hartsel model. Young & Wilcock model may match SS1 and SS2 cases, or even PS and SH4 cases, but cannot match TE case. The applicability of Young & Wilcock model is much dependent on the location of cooling rows.

The high turbine inlet temperature of modern gas turbines poses a challenge to the material used in the turbine blading of the primary stages. Mechanical failure mechanisms are more pronounced at these high temperatures, setting the lifetime of the blade. It is therefore crucial to obtain accurate local metal temperature predictions of the turbine blade. Accurately predicting the external heat transfer coefficient (HTC) distribution of the blade is therefore of uttermost importance. At present time, Siemens Energy uses the boundary layer code TEXSTAN for this purpose. The limitations coupled to such codes however make them less applicable for the complex flow physics involved in the hot gas path of turbine blading. The thesis therefore aims at introducing CFD for calculating the external HTC. This includes conducting an extensive literature study to find and validate a suitable methodology. The literature study was centered around RANS modeling, reviewing how the calculation of the HTC has evolved and the performance of some common turbulence and transition models. From the literature study, the SST k − ω model in conjunction with the γ − Reθ transition model, the v2 − f model and the Lag EB k − ε model were chosen for the investigation of a suitable methodology. The validation of the methodology was based on the extensively studied LS89 vane linear cascade of the von Karman Institute. In total 13 test cases of the cascade were chosen to represent a wide range of flow conditions. Both a periodic model and a model of the entire LS89 cascade were tested but there were great uncertainties whether or not the correct flow conditions were achieved with the model of the entire cascade. It was therefore abandoned and a periodic model was used instead. The decay of turbulence intensity is not known in the LS89 cascade. This made the case difficult to model since the turbulence boundary conditions then were incomplete. Two approaches were attempted to handle this deficiency, where one was ultimately found invalid. It was recognized that the Steelant-Dick postulation could be used in order to find a turbulent length scale which when specified at the inlet, lead to fairly good agreement with data of the HTC. The validation showed that the SST γ − Reθ model performs relatively well on the suction side and in transition onset predictions but worse on the pressure side for certain flow conditions. The v2 − f model performed better on the pressure side and on a small portion of the suction side. Literature emphasized the importance of obtaining proper turbulence characteristics around the vane for accurate HTC-predictions. It was found that the results of the validation step could be closely coupled to this statement and that further work is needed regarding this. Further research must also be done on the Steelant-Dick postulation to validate it as a reliable method in prescribing the inlet length scale.

The goal of a gas turbine engine designer is to reduce the amount of coolant used to cool the critical turbine surfaces, while at the same time extracting more benefit from the coolant flow that is used. Fan-shaped holes offer this opportunity, reducing the normal jet momentum and spreading the coolant in the lateral direction providing better surface coverage. The main drawback of fan-shaped cooling holes is the added manufacturing cost from the need for electrical discharge machining instead of the laser drilling used for cylindrical holes. This research focused on examining the performance of fan-shaped holes on two critical turbine surfaces; the vane and endwall. This research was the first to offer a complete characterization of film-cooling on a turbine vane surface, both in single and multiple row configurations. Infrared thermography was used to measure adiabatic wall temperatures, and a unique rigorous image transformation routine was developed to unwrap the surface images. Film-cooling computations were also done comparing the performance of two popular turbulence models, the RNG-kε and the v2-f model, in predicting film-cooling effectiveness. Results showed that the RNG-kε offered the closest prediction in terms of averaged effectiveness along the vane surface. The v2-f model more accurately predicted the separated flow at the leading edge and on the suction side, but did not predict the lateral jet spreading well, which led to an over-prediction in film-cooling effectiveness. The intent for the endwall surface was to directly compare the cooling and aerodynamic performance of cylindrical holes to fan-shaped holes. This was the first direct comparison of the two geometries on the endwall. The effect of upstream injection and elevated inlet freestream turbulence was also investigated for both hole geometries. Results indicated that fan-shaped film-cooling holes provided an increase in film-cooling effectiveness of 75% on average above cylindrical film-cooling holes, while at the same time producing less total pressure losses through the passage. The effect of upstream injection was to saturate the near wall flow with coolant, increasing effectiveness levels in the downstream passage, while high freestream turbulence generally lowered effectiveness levels on the endwall.
Ph. D.

The growing need for increased performance from gas turbines has fueled the drive to raise turbine inlet temperatures. This results in high thermal stresses especially along the first stage nozzle guide vane cascade as the hot combustion products exiting modern day gas turbine combustors generally reach temperatures that could endanger the structural stability of these vanes and greatly reduce the vane life. The highest heat transfer coefficients in the vane passage occurs near the endwall, particularly in the leading edge-endwall junction where vortical flows cause the flow of hotter fluid in the mainstream to mix with relatively lower temperature boundary layer fluid. This work documents the computational investigation of air injection at the end wall through a cylindrical hole placed upstream of the nozzle guide vane leading edge-end wall junction. The effect of the secondary jet on the formation of the leading edge horseshoe vortex and the consequent formation of the passage vortex has been studied. For the computations, the Reynolds averaged Navier–Stokes (RANS) equations were solved with the commercial software ANSYS Fluent using the SST k-ω model. Total pressure loss coefficient and kinetic energy loss Coefficient contour plots at the exit of the cascade to estimate the effect of the endwall fluid injection on loss profiles at the vane cascade exit. Swirling strength contours were plotted at several axial chord locations in order to track the path of the passage vortex in and downstream of the vane cascade. Two different hole-positions (located at 1 hole diameter and 2 hole diameters from the leading edge) along a plane parallel to the incident flow were considered in order to study the effect of the hole position with respect to the vane leading edge-endwall junction. Three different streamwise hole inclination angles with respect to the mainstream flow direction were studied to identify the best angle for the injection of fluid through the endwall. This angle was combined with five different compound angles (0°, 30°, 45°, 60° and 90°) in order to study the effect of varying the compound angle on the leading edge vortex and the passage vortex. Each of the above studies were conducted at two different injected fluid-to-mainstream mass flow ratios (0.5% and 1%) in order to study the effect of varying injected flow rate on the formation of the leading edge vortex and the vane passage vortex. From the results it was observed that suitable selection of the secondary injection mass flow rate, injection angle and hole-position caused an absence of the leading edge horseshoe vortex and delayed migration of the passage vortex across the guide vane passage. Heat Transfer studies were also conducted to observe the absence/weakening of the leading edge vortex and the delayed pitch-wise movement of the passage vortex.
Master of Science

A coupled density based solver in the framework of foam-extend is used to perform simulations of transonic flows. The solver is based on an explicit and time-accurate algorithm and is coupled to a compressible Unsteady Reynolds-Averaged Navier-Stokes (URANS) and a Large Eddy Simulation (LES) module. The solver is first attested on canonical compressible flow scenarios such as a 1-D shock tube and the transonic flow through a 2-D channel. Following this, a 2-D URANS simulation of the flow within the passages of a High Pressure Turbine Nozzle Guide Vane (HPT-NGV) is performed and compared against experimental data. Finally, preliminary results of a 3-D LES on a simplified geometry of the HPT-NGV are presented. In the future, this numerical setup will be used to study indirect combustion noise in aircraft engines.

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

The following study presents an experimental and computational investigation on the effects of implementing a dual cavity slashface configuration and varying slashface coolant leakage mass flow rate on the thermal performance for a 1st stage nozzle guide vane axisymmetric converging endwall. An upstream doublet staggered cylindrical hole jet cooling scheme provides additional purged coolant with consistent conditions throughout the investigation. The effects are measured in engine representative transonic mainstream and coolant flow conditions where Mexit = 0.85, Reexit = 1.5 × 106, freestream turbulence intensity of 16%, and a coolant density ratio of 1.95. Four combinations of slashface Fwd and Aft cavity mass flow rate are experimentally analyzed by comparing key convective heat transfer parameters. Data is collected and reduced using a combination of IR thermography and a linear regression technique to map endwall heat transfer performance throughout the passage. A flow visualization study is employed using 100 cSt oil-based paint to gather qualitative insights into the endwall flow field. A complimentary CFD study is carried out to gather additional understanding of the endwall flow ingestion and egression behavior as well as comparing performance against a conventional cavity configuration. Experimental comparisons indicate slashface mass flow rate variations have a minor effect on passage film cooling coverage. Instead, coolant coverage across the passage is primarily driven by upstream purge coolant. However, endwall heat transfer coefficient is reduced as much as 20% in mid-passage areas as leakage flow decreases. This suggests that changes in leakage flow maintains a first order correlation in altering passage aerodynamics that, despite relatively consistent film cooling coverage, also leads to significant changes in net heat flux reduction in the passage. Endwall flow behavior proves to be complex along the gap interface showing signs of ingestion, egression, and tangential flow varying spatially throughout the gap. CFD comparisons suggests that a dual cavity configuration varies the gap static pressure distribution closer to the mainstream pressure throughout the passage in high speed applications compared to a single cavity configuration. The resulting decelerating flow creates a more stable endwall flow profile and favorable coolant environment by reducing boundary layer thinning and shear interaction in near gap endwall tangential flow.
Master of Science
Gas turbines are often exposed to high temperatures as they convert hot, energetic gas streams into mechanical motion. As turbines receive higher temperature gases, their efficiency increases and reduces waste. However, these temperatures can get too hot for turbine parts. To survive these high temperatures, turbine components are often assembled with a gap in between to allow the part to expand and contrast when it heats and cools. Relatively cold air is also fed into the gap to help prevent hot gases from entering. This cold air can also feed into other pathways to flow onto the turbine component's surface and act as an insulating layer to the hot gas and protect the component from overheating. The study presented investigates an assembly gap, referred to as a slashface gap, found in the middle of a vane located immediately after gas combustion with cold air leaking through. One unique aspect of this study is that there are two pathways for cold air, or coolant, to leak through when, typically, there is only one. The slashface gap lies on a wall which the vanes are attached to, referred to as the endwall. Multiple small holes on the endwall in between the combustor and vanes jet out coolant to try and protect the endwall from hot gases. These holes, called jump cooling holes, point out towards the vanes and angled more shallowly so that the holes do not face directly up from the endwall. The holes are angled as they are meant to gracefully spray coolant to cover and insulate the endwall instead of mixing with the hot air above. The experiments found that changing how much coolant is leaked through the slashface has little effect on how much coolant from jump cooling holes covered the endwall. However, smaller slashface leaks better protect the endwall from the hot gas by forcing it to move smoother and give off less heat across the endwall rather than a tumbling like manner. The experiment is modeled on a computer simulation to determine the differences of a slashface gap with the typical one coolant pathway and the coolant dual pathway configuration that is tested in the experiments. This simulation discovered that having two coolant pathways significantly reduces how much hot gas and jump cooling coolant enters and leaves the slashface gap. This makes the surrounding airflow along the endwall travel more smoothly and does not give off as much heat as if a single coolant pathway configuration is used instead.

Efficiency improvement in turbomachines is an important aspect in reducing the use of fossil-based fuel and thereby reducing carbon dioxide emissions in order to achieve a sustainable future. Gas turbines are mainly fossil-based turbomachines powering aviation and land-based power plants. In line with the present situation and the vision for the future, gas turbine engines will retain their central importance in coming decades. Though the world has made significant advancements in gas turbine technology development over past few decades, there are yet many design features remaining unexplored or worth further improvement. These features might have a great potential to increase efficiency. The high pressure turbine (HPT) stage is one of the most important elements of the engine where the increased efficiency has a significant influence on the overall efficiency as downstream losses are substantially affected by the prehistory. The overall objective of the thesis is to contribute to the development of gas turbine efficiency improvements in relation to the HPT stage.   Hence, this study has been incorporated into a research project that investigates leading edge contouring near endwall by fillet and external cooling on a nozzle guide vane with a common goal to contribute to the development of the HPT stage. In the search for HPT stage efficiency gains, leading edge contouring near the endwall is one of the methods found in the published literature that showed a potential to increase the efficiency by decreasing the amount of secondary losses. However, more attention is necessary regarding the realistic use of the leading edge fillet. On the other hand, external cooling has a significant influence on the HPT stage efficiency and more attention is needed regarding the aerodynamic implication of the external cooling. Therefore, the aerodynamic influence of a leading edge fillet and external cooling, here film cooling at profile and endwall as well as TE cooling, on losses and flow field have been investigated in the present work. The keystone of this research project has been an experimental investigation of a modern nozzle guide vane using a transonic annular sector cascade. Detailed investigations of the annular sector cascade have been presented using a geometric replica of a three dimensional gas turbine nozzle guide vane. Results from this investigation have led to a number of new important findings and also confirmed some conclusions established in previous investigations to enhance the understanding of complex turbine flows and associated losses.   The experimental investigations of the leading edge contouring by fillet indicate a unique outcome which is that the leading edge fillet has no significant effect on the flow and secondary losses of the investigated nozzle guide vane. The reason why the leading edge fillet does not affect the losses is due to the use of a three-dimensional vane with an existing typical fillet over the full hub and tip profile. Findings also reveal that the complex secondary flow depends heavily on the incoming boundary layer. The investigation of the external cooling indicates that a coolant discharge leads to an increase of profile losses compared to the uncooled case. Discharges on the profile suction side and through the trailing edge slot are most prone to the increase in profile losses. Results also reveal that individual film cooling rows have a weak mutual effect. A superposition principle of these influences is followed in the midspan region. An important finding is that the discharge through the trailing edge leads to an increase in the exit flow angle in line with an increase of losses and a mixture mass flow. Results also indicate that secondary losses can be reduced by the influence of the coolant discharge. In general, the exit flow angle increases considerably in the secondary flow zone compared to the midspan zone in all cases. Regarding the cooling influence, the distinct change in exit flow angle in the area of secondary flows is not noticeable at any cooling configuration compared to the uncooled case. This interesting zone requires an additional, accurate study. The investigation of a cooled vane, using a tracer gas carbon dioxide (CO2), reveals that the upstream platform film coolant is concentrated along the suction surfaces and does not reach the pressure side of the hub surface, leaving it less protected from the hot gas. This indicates a strong interaction of the secondary flow and cooling showing that the influence of the secondary flow cannot be easily influenced.   The overall outcome enhances the understanding of complex turbine flows, loss behaviour of cooled blade, secondary flow and interaction of cooling and secondary flow and provides recommendations to the turbine designers regarding the leading edge contouring and external cooling. Additionally, this study has provided to a number of new significant results and a vast amount of data, especially on profile and secondary losses and exit flow angles, which are believed to be helpful for the gas turbine community and for the validation of analytical and numerical calculations.
Ökad verkningsgrad i turbomaskiner är en viktig del i strävan att minska användningen av fossila bränslen och därmed minska växthuseffekten för att uppnå en hållbar framtid. Gasturbinen är huvudsakligen fossilbränslebaserad, och driver luftfart samt landbaserad kraftproduktion. Enligt rådande läge och framtidsutsikter bibehåller gasturbinen denna centrala roll under kommande decennier. Trots betydande framsteg inom gasturbinteknik under de senaste årtionden finns fortfarande många designaspekter kvar att utforska och vidareutveckla. Dessa designaspekter kan ha stor potential till ökad verkningsgrad. Högtrycksturbinsteget är en av de viktigaste delarna av gasturbinen, där verkningsgraden har betydande inverkan på den totala verkningsgraden eftersom förluster kraftigt påverkas av tidigare förlopp. Huvudsyftet med denna studie är att bidra till verkningsgradsförbättringar i högtrycksturbinsteget.   Studien är del i ett forskningsprojekt som undersöker ledskenans framkantskontur vid ändväggarna samt extern kylning, i jakten på dessa förbättringar. Den aerodynamiska inverkan av en förändrad geometri vid ledskenans ändväggar har i tidigare studier visat potential för ökad verkningsgrad genom minskade sekundärförluster. Ytterligare fokus krävs dock, med användning av en rimlig hålkälsradie. Samtidigt har extern kylning i form av filmkylning stor inverkan på verkningsgraden hos högtrycksturbinsteget och forskning behövs med fokus på den aerodynamiska inverkan. Av denna anledning studeras här inverkan både av ändrad hålkälsradie vid ledskenans framkant samt extern kylning i form av filmkylning av skovel, ändvägg och bakkant på aerodynamiska förluster och strömningsfält. Huvudpelaren i detta forskningsprojekt har varit en experimentell undersökning av en geometrisk replika av en modern tredimensionell gasturbinstator i en transonisk annulärkaskad. Detaljerade undersökningar i annulärkaskaden har gett betydande resultat, och bekräftat vissa tidigare studier. Detta har lett till ökad förståelsen för de komplexa flöden och förluster som karakteriserar gasturbiner.   De experimentella undersökningarna av förändrad framkantsgeometri leder till den unika slutsatsen att den modifierade hålkälsradien inte har någon betydande inverkan på strömningsfältet eller sekundärförluster av den undersökta ledskenan. Anledningen till att förändringen inte påverkar förlusterna är i detta fall den tredimensionella karaktären hos ledskenan med en redan existerande typisk framkantsgeometri. Undersökningarna visar också att de komplexa sekundärströmningarna är kraftigt beroende av det inkommande gränsskiktet. Undersökning av extern kylning visar att kylflödet leder till en ökad profilförlust. Kylflöde på sugsidan samt bakkanten har störst inverkan på profilförlusten. Resultaten visar också att individuella filmkylningsrader har liten påverkan sinsemellan och kan behandlas genom en superpositionsprincip längs mittsnittet. En viktig slutsats är att kylflöde vid bakkanten leder till ökad utloppsvinkel tillsammans med ökade förluster och massflöde. Resultat tuder på att sekundärströmning kan minskas genom ökad kylning. Generellt ökar utloppsvinkeln markant i den sekundära flödeszonen jämfört med mittsnittet för alla undersökta fall. Den kraftiga förändringen i utloppsvinkel är dock inte märkbar i den sekundära flödeszonen i något av kylfallen jämfört med de okylda referensfallet. Denna zon fordrar ytterligare studier. Spårgasundersökning av ledskenan med koldioxid (CO2) visar att plattformskylning uppströms ledskenan koncentreras till skovelns sugsida, och når inte trycksidan som därmed lämnas mer utsatt för het gas. Detta påvisar den kraftiga interaktionen mellan sekundärströmning och kylflöden, och att inverkan från sekundärströmningen ej enkelt kan påverkas. De generella resultaten från undersökningen ökar förståelsen av komplexa turbinflöden, förlustbeteenden för kylda ledskenor, interaktionen mellan sekundärströmning och kylflöden, och ger rekommendationer för turbinkonstruktörer kring förändrad framkantsgeometri i kombination med extern kylning. Dessutom har studien gett betydande resultat och en stor mängd data, särskilt rörande profil- och sekundärförluster och utloppsvinkel, vilket tros kunna vara till stor hjälp för gasturbinssamfundet vid validering av analytiska och numeriska beräkningar.

QC 20140909

Turbopower, Sector rig

Efficiency improvement is one of the most important aspects of engineering and especially important in the field of energy production. In the past decades, energy was mostly produced by fossil based technologies involving turbomachines, and the efficiency of these machines nearly quadrupled since the introduction of the first economically viable gas turbines. The progress continues, as there are still areas where improvement can be made. Such area is the High Pressure Turbine stage (HPT), which influences the flow characteristics and losses downstream, which this thesis will examine in more detail. In the open literature it can be found that one of the areas with potential for progress is the external cooling of the nozzle guide vanes (NGV) of the HPT stage. However not many studies go towards supersonic exit velocities even though that is the most common trend followed by the industry these days. The external cooling allows the turbine entry temperature (TET) to go beyond the melting point of the blade material thus increase Carnot efficiency but in the meantime influences the flow characteristics and losses. To understand these influences of the cooling, experiments in an annular sector cascade (ASC) were conducted with exit velocities from Mach 0.95 to 1.2 without and with cooling applied. The findings of the experiments are believed to help the more detailed understanding of the flow behaviour at high exit velocities. When comparing the corresponding runs in the two cases it became obvious that with cooling applied the deviation of the exit flow angle is generally smaller than in the uncooled case. This might be a highly important design feature for designers to work with. From the available data it was concluded that the total pressure distribution across the span is not significantly affected with the introduction of cooling.

This thesis presents detailed investigations of the effect of end wall profiling on the secondary flow in a large scale, linear cascade with nozzle guide vanes. The purpose of this project is to look into the secondary flow structure in the linear cascade and the influence of the shaped end wall on the secondary flow. By applying the non-axisymmetric end wall, the secondary flow is reduced compared to the flat end wall data. The yaw angle variation at the exit of the blade passage is reduced. The cascade was designed according to the nozzle guide vane from ALSTOM Energy Ltd. It was manufactured and connected to the low speed wind tunnel in the Thermo- Fluids Lab in Durham. The data acquisition system was designed and commissioned. Five hole probes were designed and calibrated according to the cascade test condition. The flow field with the flat end wall in the cascade was investigated using five hole probes through different traverse slots. Flow visualisations were conducted as well. The secondary flow structure and the loss development in the cascade are understood. Transitional trips were put on the blade surfaces and their effects on the secondary flow were observed. The CFD code was modified to fit the cascade case. It was validated against the Durham standard case and the flat end wall results. Different numerical schemes and turbulence models were evaluated. Different shaped end walls were systematically tested by the CFD code. The best end wall profile was selected and manufactured. It was then tested in the cascade. Detailed investigations by five hole probes, flow visualisation and wall static pressure measurements were conducted. The results were compared to the flat end wall results and the CFD prediction. The secondary flow and the total pressure loss were reduced. The test data in the cascade will supply the evidence and data for the real turbine design. The chosen end wall profile will hopefully be tested in a test turbine.

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

In a typical modern gas turbine engine, the nozzle guide vanes (NGVs) endure the highest operating temperatures. There exists a great drive in the turbine industry to increase the turbine inlet temperatures leading to higher thermal efficiency. This has led to a drive to increase turbine vane- and blade-cooling. Numerical modelling has to a large degree replaced empirical codes and models as the main research tool regarding simulation of blade-cooling. Outdated empirical solvers have made way for commercial CFD solvers such as FLUENT, a Reynolds-averaged Navier-Stokes (RANS) solver. One such empirical solver, TACT1, has until recently still proved to yield acceptable results. A comparative study has been done using the T56 NGV blade to establish the differences, advantages and disadvantages of these 2 codes. The engine and subsequent NGV blade were analysed using NREC, STAN5, LOSS3D and TACT1. RANS simulations were found to be computationally expensive. TACT1 yielded acceptable results compared with computational cost. For modern-day designers, RANS would be the preferred tool.
Dissertation (MEng)--University of Pretoria, 2009.
Mechanical and Aeronautical Engineering
unrestricted

A 2-D plain strain CFD/FEM model to simulate thermal shocks and stresses in a turbine blade has been set up using the commercially available software FLUENT and NASTRAN. The model was validated against the experimental data of Bohn et. al. and used to simulate real test cases. The steady state numerical model was set up for a single Mark II nozzle guide vane using the correct boundary conditions to resolve the flow field. A combined laminar and turbulent model was developed in FLUENT that was used to highly accurately predict the pressure, temperature and heat transfer coefficient distribution on the blade surface as well as the temperature distribution on the cooling holes inside the blade. The resulting temperature profiles on the blade and cooling holes were used as boundary conditions for the FEM analysis to resolve the internal temperature and stress profiles. The pressure, temperature and heat transfer distribution on the blade, from FLUENT, were compared to those from Bohn et. al. The predicted pressure distribution was exact with the experimental results and the predicted temperature distribution had an average overprediction of 1.4 % on both the pressure and suction side. The internal temperature profile predicted by NASTRAN was correctly predicted with an average over-prediction of 2 %. The stress contours were accurately predicted with the stress magnitude varying by 17 % to that of Bohn et. al. The reason for the difference between the MSC.NASTRAN and Bohn et. al. stress results is believed to be purely solver related. Bohn et al. used a FEM package called MSC.MARClMentat. With the steady state model validated, transient test cases were simulated that represent typical operational data. The mission profile was obtained for the T-56 engine found on the C130 cargo plane. The model was used to simulate the test case where the turbine inlet temperature (TIT) varied with time. The simulation results showed that stress was proportional to TIT, where changes in the TIT were seen later in the stress curve, due to conduction in the blade. Steep TIT changes, such as shock loads affected stress later than gentler TIT changes. Thus, the FLUENT / NASTRAN model was successfully validated, and used to simulate a flight mission profile. The goal to calculate quality unsteady stress profiles was achieved and forms the boundary conditions for thermal fatigue calculations.
Thesis (M.Sc.Eng.)-University of KwaZulu-Natal, 2005.

碩士
大葉大學
機械工程研究所
88
The reason for simultaneous cracking at the leading edge and annular casing endwall of the first stage guide vanes is not fully understood in the industry nowadays. Although this might be due to drawbacks in the design of the internal cooling and film cooling systems, it might possibly be that the external heat transfer was underestimated. The objective of his study is to discuss a factor which is frequently ignored in the estimate of external heat transfer for a guide vane, namely, the effect of thermal expansion displacement of the combustor transition nozzle on the flow field and the heat transfer coefficient distribution around the first stage guide vanes. Since the cracking of the first stage guide vanes is usually found to occur at the leading edge, the trailing edge near the endwall, and at the annual casing endwall, to study the external heat transfer, it is quite necessary to fully understand the flow field around these areas. Many researches have pointed out that the flow near the endwall of a vane is three-dimensional and complicated, including secondary flows such as horseshoe vortex, passage vortex, and corner vortex, and the interaction of these vortices. The formation of these vortices, however, is intimately tied to the development of the upstream boundary layer along the endwall. The matching of the exit of the transition nozzle and the annular casing endwall of the guide vanes will obviously affect the boundary layer along the endwall. It seems questionable to fully ignore the fact of the expansion displacement of the transition nozzle in the estimate of the heat transfer distribution around the endwall of a vane. Experiments were conducted in this work to investigate the effects of the displacement of transition nozzle on the heat transfer coefficient distribution near endwall region of a first stage turbine guide vane. Two-half-vane model was used in the test. The sizes of the side bleed slots of the test model were determined using a CFD package to simulate the flow situation between the two-half-vane model and a corresponding cascade. The displacement of transition nozzle relative to the annular casing wall due to thermal expansion was represented by a finite step in the test. The distribution of the heat transfer coefficient was determined by transient heat transfer experiment with liquid crystal thermography. Results of the experiments show that, when such a displacement occurs, the heat transfer coefficients at the endwall and on the suction wall near the base of the vane are enhanced. On the other hand ,the heat transfer coefficient on the pressure wall seems decreased slightly. The size of the triangular region on the suction side surface due to the three-dimensional flow pattern near the endwall is enlarged. These results indicate that the three-dimensional flow pattern has been changed and that the possibility of damage to a vane is increased.