Academic literature on the topic 'Gas Turbine Nozzle'
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Journal articles on the topic "Gas Turbine Nozzle"
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Rodgers, C. "Impingement Starting and Power Boosting of Small Gas Turbines." Journal of Engineering for Gas Turbines and Power 107, no. 4 (October 1, 1985): 821–27. http://dx.doi.org/10.1115/1.3239817.
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The technology of high-pressure air or hot-gas impingement from stationary shroud supplementary nozzles onto radial outflow compressors and radial inflow turbines to permit rapid gas turbine starting or power boosting is discussed. Data are presented on the equivalent turbine component performance for convergent/divergent shroud impingement nozzles, which reveal the sensitivity of nozzle velocity coefficient with Mach number and turbine efficiency with impingement nozzle admission arc. Compressor and turbine matching is addressed in the transient turbine start mode with the possibility of operating these components in braking or reverse flow regimes when impingement flow rates exceed design.
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Karstensen, K. W., and J. O. Wiggins. "A Variable-Geometry Power Turbine for Marine Gas Turbines." Journal of Turbomachinery 112, no. 2 (April 1, 1990): 165–74. http://dx.doi.org/10.1115/1.2927629.
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Gas turbines have been accepted in naval surface ship applications, and considerable effort has been made to improve their fuel consumption, particularly at part-load operation. This is an important parameter for shipboard engines because both propulsion and electrical-generator engines spend most of their lives operating at off-design power. An effective way to improve part-load efficiency of recuperated gas turbines is by using a variable power turbine nozzle. This paper discusses the successful use of variable power turbine nozzles in several applications in a family of engines developed for vehicular, industrial, and marine use. These engines incorporate a variable power turbine nozzle and primary surface recuperator to yield specific fuel consumption that rivals that of medium speed diesels. The paper concentrates on the experience with the variable nozzle, tracing its derivation from an existing fixed vane nozzle and its use across a wide range of engine sizes and applications. Emphasis is placed on its potential in marine propulsion and auxiliary gas turbines.
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Chaker, Mustapha, Cyrus B. Meher-Homji, and Thomas Mee. "Inlet Fogging of Gas Turbine Engines—Part II: Fog Droplet Sizing Analysis, Nozzle Types, Measurement, and Testing." Journal of Engineering for Gas Turbines and Power 126, no. 3 (July 1, 2004): 559–70. http://dx.doi.org/10.1115/1.1712982.
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The inlet fogging of gas turbine engines for power augmentation has seen increasing application over the past decade yet not a single technical paper treating the physics and engineering of the fogging process, droplet size measurement, droplet kinetics, or the duct behavior of droplets, from a gas turbine perspective, is available. This paper provides the results of extensive experimental and theoretical studies conducted over several years, coupled with practical aspects learned in the implementation of nearly 500 inlet fogging systems on gas turbines ranging in power from 5 to 250 MW. Part II of the paper treats the practical aspects of fog nozzle droplet sizing, measurement and testing presenting the information from a gas turbine fogging perspective. This paper describes the different measurement techniques available, covers design aspects of nozzles, provides experimental data on different nozzles, and provides recommendations for a standardized nozzle testing method for gas turbine inlet air fogging.
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Kriukov, Aleksei Alekseevich. "Influence of nozzle inclination angle on velocity coefficient of inflow turbine with partial blading of runner." Vestnik of Astrakhan State Technical University. Series: Marine engineering and technologies 2023, no. 1 (February 28, 2023): 23–29. http://dx.doi.org/10.24143/2073-1574-2023-1-23-29.
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The paper considers a numerical experiment with three groups of low-consuming inflow turbines with partial blading of the runner. Geometrical models of the turbine stages with partial blading of the runner at different angles of the nozzle inclination have been developed. The inflow turbine stages with partial blading of the runner are investigated. The efficiency and speed coefficient of the inflow turbine nozzles are calculated. There has been carried out numerical modeling of the working fluid flow by using the computational gas dynamics elements. Pick values of the nozzle apparatus velocity coefficient and the turbine stage efficiency are defined. Geometric models of the turbine stages with different nozzle inclination angles are developed, the boundary conditions of the experiment are determined, and the experiment results are analyzed and shown. There are given the dependence graphs of the nozzle velocity coefficient and the efficiency of a low-consuming inflow turbine stage. Comparative analysis of the nozzle velocity coefficient and the efficiency of three turbine stages with different inclination angles of the nozzles in the nozzle apparatus has been carried out. Conclusions are drawn about further procedures on improving the flow part of the runner to increase the efficiency of the stage.
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Shen, J. J. S., V. C. Ting, and E. H. Jones. "Application of Sonic Nozzles in Field Calibration of Natural Gas Flows." Journal of Energy Resources Technology 111, no. 4 (December 1, 1989): 205–13. http://dx.doi.org/10.1115/1.3231425.
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This paper presents Chevron Oil Field Research Company’s operating experience using the sonic nozzle as a proving device for measuring natural gas flows in field tests. The nozzle reference flow rate was used for calibrating orifice, turbine, and vortex meters in three tests with a pipeline quality gas and an unprocessed natural gas as the working fluid. For pipeline gas, the field calibration results show good agreement between the sonic nozzle reference and a turbine meter while the accuracy of orifice metering is size dependent. The 4-in. (102-mm) orifice meter flow rates agree well with the nozzle reference, but the 16-in. (406-mm) orifice flow measurements are up to 2 percent lower. Deviations between the test meters and the sonic nozzles are generally larger for the unprocessed gas. These field projects demonstrate that sonic nozzles can be operated successfully as a prover for processed natural gas, while more work is needed to study the critical flow in nozzles for unprocessed natural gas.
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Sanaye, Sepehr, and Salahadin Hosseini. "Off-design performance improvement of twin-shaft gas turbine by variable geometry turbine and compressor besides fuel control." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 234, no. 7 (December 3, 2019): 957–80. http://dx.doi.org/10.1177/0957650919887888.
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A novel procedure for finding the optimum values of design parameters of industrial twin-shaft gas turbines at various ambient temperatures is presented here. This paper focuses on being off design due to various ambient temperatures. The gas turbine modeling is performed by applying compressor and turbine characteristic maps and using thermodynamic matching method. The gas turbine power output is selected as an objective function in optimization procedure with genetic algorithm. Design parameters are compressor inlet guide vane angle, turbine exit temperature, and power turbine inlet nozzle guide vane angle. The novel constrains in optimization are compressor surge margin and turbine blade life cycle. A trained neural network is used for life cycle estimation of high pressure (gas generator) turbine blades. Results for optimum values for nozzle guide vane/inlet guide vane (23°/27°–27°/6°) in ambient temperature range of 25–45 ℃ provided higher net power output (3–4.3%) and more secured compressor surge margin in comparison with that for gas turbines control by turbine exit temperature. Gas turbines thermal efficiency also increased from 0.09 to 0.34% (while the gas generator turbine first rotor blade creep life cycle was kept almost constant about 40,000 h). Meanwhile, the averaged values for turbine exit temperature/turbine inlet temperature changed from 831.2/1475 to 823/1471°K, respectively, which shows about 1% decrease in turbine exit temperature and 0.3% decrease in turbine inlet temperature.
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Kravchuk, Yu, and Т. Tatarchuk. "Methods of increasing the reliability of turbine elements." Innovative Materials and Technologies in Metallurgy and Mechanical Engineering, no. 2 (March 18, 2021): 57–65. http://dx.doi.org/10.15588/1607-6885-2020-2-8.
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The purpose of the work. Statistical and experimental analysis of coating methods on the turbine nozzle apparatus to increase the temperature regime.
Research methods. Calculation method of finite elements, experimental.
The results obtained. Studies have shown that the use of thermally protective coatings TZP thickness of 250 мm with a thermal conductivity of 1 W / mK on the two steps of the turbine can implement one of two possibilities:
- at constant operating temperature of the blade material to increase the temperature of the gas in front of the turbine by about 100 °C, which will increase efficiency and fuel savings by more than 13 %;- without changing the temperature of the gas in front of the turbine - to increase the durability of the blades by about 4 times, due to a decrease in their operating temperature. The analysis of two methods of drawing TZP was carried out, in the work the estimation of a temperature condition of the nozzle device (CA) of the turbine of a high pressure of the engine, decrease in its temperature due to drawing TZP and increase of its resource is carried out. The problem was solved by applying TZP on the blades of the nozzle apparatus. The analysis of two methods of drawing TZP was carried out, the estimation of a temperature condition of the nozzle device (CA) of the turbine of high pressure of the engine, decrease in its temperature due to drawing TZP and increase of its resource is carried out.
Scientific novelty. The problem of creating efficient, economical and reliable gas turbines is the most difficult among the many problems that arise in the development of gas turbine construction. Important elements of turbines are working and nozzle blades, the material and design of which determine the allowable gas temperature in front of the turbine and thus directly affect the technical and economic performance of the gas turbine engine.
Practical value. The obtained results are important for the further development of aircraft engine construction, due to the application of TZP achieved an increase in the resource of CA from 40,000 hours to 67,000 hours.
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Myers, G. D., J. P. Armstrong, C. D. White, S. Clouser, and R. J. Harvey. "Development of an Innovative High-Temperature Gas Turbine Fuel Nozzle." Journal of Engineering for Gas Turbines and Power 114, no. 2 (April 1, 1992): 401–8. http://dx.doi.org/10.1115/1.2906605.
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The objective of the innovative high-temperature fuel nozzle program was to design, fabricate, and test propulsion engine fuel nozzles capable of performance despite extreme fuel and air inlet temperatures. Although a variety of both passive and active methods for reducing fuel wetted-surface temperatures were studied, simple thermal barriers were found to offer the best combination of operability, cycle flexibility, and performance. A separate nozzle material study examined several nonmetallics and coating schemes for evidence of passivating or catalytic tendencies. Two pilotless airblast nozzles were developed by employing finite-element modeling to optimize thermal barriers in the stem and tip. Operability of these prototypes was compared to a current state-of-the art piloted, prefliming airblast nozzle, both on the spray bench and through testing in a can-type combustor. The three nozzles were then equipped with internal thermocouples and operated at 1600°F air inlet temperature while injecting marine diesel fuel heated to 350°F. Measured and predicted internal temperatures as a function of fuel flow rate were compared. Results show that the thermal barrier systems dramatically reduced wetted-surface temperatures and the potential for coke fouling, even in an extreme environment.
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Richards, G. A., and M. C. Janus. "Characterization of Oscillations During Premix Gas Turbine Combustion." Journal of Engineering for Gas Turbines and Power 120, no. 2 (April 1, 1998): 294–302. http://dx.doi.org/10.1115/1.2818120.
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The use of premix combustion in stationary gas turbines can produce very low levels of Nox emissions. This benefit is widely recognized, but turbine developers routinely encounter problems with combustion oscillations during the testing of new premix combustors. Because of the associated pressure fluctuations, combustion oscillations must be eliminated in a final combustor design. Eliminating these oscillations is often time-consuming and costly because there is no single approach to solve an oscillation problem. Previous investigations of combustion stability have focused on rocket applications, industrial furnaces, and some aeroengine gas turbines. Comparatively little published data is available for premixed combustion at conditions typical of an industrial gas turbine. In this paper, we report experimental observations of oscillations produced by a fuel nozzle typical of industrial gas turbines. Tests are conducted in a specially designed combustor capable of providing the acoustic feedback needed to study oscillations. Tests results are presented for pressure up to 10 atmospheres, with inlet air temperatures up to 588 K (600 F) burning natural gas fuel. Based on theoretical considerations, it is expected that oscillations can be characterized by a nozzle reference velocity, with operating pressure playing a smaller role. This expectation is compared to observed data that shows both the benefits and limitations of characterizing the combustor oscillating behavior in terms of a reference velocity rather than other engine operating parameters. This approach to characterizing oscillations is then used to evaluate how geometric changes to the fuel nozzle will affect the boundary between stable and oscillating combustion.
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Придорожный, Роман Петрович, Александр Викторович Шереметьев, and Анатолий Павлович Зиньковский. "ВЛИЯНИЕ ПОЛЗУЧЕСТИ МАТЕРИАЛА НА РАБОТОСПОСОБНОСТЬ ЛОПАТОК СОПЛОВОГО АППАРАТА ТУРБИНЫ ВЫСОКОГО ДАВЛЕНИЯ." Aerospace technic and technology, no. 7 (August 31, 2020): 41–46. http://dx.doi.org/10.32620/aktt.2020.7.06.
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The creation of aircraft gas turbine engines that meet modern requirements for the resource, especially its hot part, requires not only more advanced design methods but also an analysis of operability and damageability during the resource to find reserves aimed at improving reliability. One of the most complex and heat-stressed components of a modern gas turbine engine is the nozzle vanes of a high-pressure turbine, directly perceiving the high temperature of the gas at the exit of the combustion chamber and having an advanced convection-film cooling system. The service life nozzle vanes of modern aviation gas turbine engines can be tens of thousands of hours. At the same time, the maximum operating mode can reach only a few hundred hours. It is believed that damage to nozzle vanes on an engine occurs mainly in hot climatic zones. Nevertheless, as the analysis of computational studies for modern aviation gas turbine engines shows, such statements are erroneous. The direct consequence of the action of elevated temperatures and high thermal stresses is the creep of the material. A computational study of the effect of creep of the material of the nozzle vanes of a high-pressure turbine under various operating conditions of the engine on their operability was carried. It is shown that with increasing flight altitude the working temperature of the nozzle vane begins to increase, and creep processes are accelerated for all climatic zones of operation. Since with increasing flight altitude, the temperature difference for different climatic zones gradually decreases, at high altitudes, where the temperature in different climatic zones differs slightly, stress relaxation processes proceed identically. In this case, with an increase in temperature, creep processes proceed faster, and the stress level to which stress relaxation occurs becomes lower. Thus, with increasing flight altitude, damage in cold conditions approaches that under normal and hot conditions, and at high altitudes, it can even be higher. The regularities of the influence of climatic conditions and flight altitude on the strength of the nozzle vanes of high-pressure turbines and their operability are established, based on which the need to take into account the operating time of the engine in various climatic conditions when determining the service life of nozzle vanes is shown.
Dissertations / Theses on the topic "Gas Turbine Nozzle"
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Harvey, Neil William. "Heat transfer on nozzle guide vane end walls." Thesis, University of Oxford, 1991. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.293454.
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Rahim, Amir. "Effect of nozzle guide vane shaping on high pressure turbine stage performance." Thesis, University of Oxford, 2017. https://ora.ox.ac.uk/objects/uuid:35274ff0-0ea7-47bc-adc3-388f136b9555.
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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.
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Johnson, A. B. "The aerodynamic effects of nozzle guide vane shock wave and wake on a transonic turbine rotor." Thesis, University of Oxford, 1988. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.329958.
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Khorsand, Khashayar. "Numerical heat transfer studies and test rig preparation on a gas turbine nozzle guide vane." Thesis, KTH, Kraft- och värmeteknologi, 2014. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-144412.
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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.
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Lee, Yeong Jin. "Aerodynamic Investigation of Upstream Misalignment over the Nozzle Guide Vane in a Transonic Cascade." Thesis, Virginia Tech, 2017. http://hdl.handle.net/10919/77924.
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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
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Rubensdörffer, Frank G. "Numerical and Experimental Investigations of Design Parameters Defining Gas Turbine Nozzle Guide Vane Endwall Heat Transfer." Doctoral thesis, KTH, Energiteknik, 2006. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3884.
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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
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Rubensdörffer, Frank G. "Numerical and experimental investigations of design parameters defining gas turbine nozzle guide vane endwall heat transfer /." Stockholm, 2006. http://urn.kb.se/resolve?urn=urn:nbn:se:kth:diva-3884.
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Abdeh, Hamed. "Incidence Effects on Aerodynamic and Thermal Performance of a Film-Cooled Gas Turbine Nozzle Guide Vane." Doctoral thesis, Università degli studi di Bergamo, 2018. http://hdl.handle.net/10446/105183.
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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°.
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Agricola, Lucas. "Nozzle Guide Vane Sweeping Jet Impingement Cooling." The Ohio State University, 2018. http://rave.ohiolink.edu/etdc/view?acc_num=osu1525436077557298.
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Colban, William F. IV. "A Detailed Study of Fan-Shaped Film-Cooling for a Nozzle Guide Vane for an Industrial Gas Turbine." Diss., Virginia Tech, 2005. http://hdl.handle.net/10919/29856.
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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.
Books on the topic "Gas Turbine Nozzle"
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Escudier, Marcel. Introduction to Engineering Fluid Mechanics. Oxford University Press, 2018. http://dx.doi.org/10.1093/oso/9780198719878.001.0001.
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Turbojet and turbofan engines, rocket motors, road vehicles, aircraft, pumps, compressors, and turbines are examples of machines which require a knowledge of fluid mechanics for their design. The aim of this undergraduate-level textbook is to introduce the physical concepts and conservation laws which underlie the subject of fluid mechanics and show how they can be applied to practical engineering problems. The first ten chapters are concerned with fluid properties, dimensional analysis, the pressure variation in a fluid at rest (hydrostatics) and the associated forces on submerged surfaces, the relationship between pressure and velocity in the absence of viscosity, and fluid flow through straight pipes and bends. The examples used to illustrate the application of this introductory material include the calculation of rocket-motor thrust, jet-engine thrust, the reaction force required to restrain a pipe bend or junction, and the power generated by a hydraulic turbine. Compressible-gas flow is then dealt with, including flow through nozzles, normal and oblique shock waves, centred expansion fans, pipe flow with friction or wall heating, and flow through axial-flow turbomachinery blading. The fundamental Navier-Stokes equations are then derived from first principles, and examples given of their application to pipe and channel flows and to boundary layers. The final chapter is concerned with turbulent flow. Throughout the book the importance of dimensions and dimensional analysis is stressed. A historical perspective is provided by an appendix which gives brief biographical information about those engineers and scientists whose names are associated with key developments in fluid mechanics.
Book chapters on the topic "Gas Turbine Nozzle"
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Nita, Kozo, Yoji Okita, and Chiyuki Nakamata. "Experimental and Numerical Study on Application of a CMC Nozzle for High Temperature Gas Turbine." In Mechanical Properties and Performance of Engineering Ceramics and Composites VII, 315–24. Hoboken, NJ, USA: John Wiley & Sons, Inc., 2012. http://dx.doi.org/10.1002/9781118217467.ch29.
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Decher, Reiner. "More Components: Inlets, Mixers, and Nozzles." In The Vortex and The Jet, 137–54. Singapore: Springer Singapore, 2022. http://dx.doi.org/10.1007/978-981-16-8028-1_13.
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AbstractTheintegrationof a gas turbine engine into a functioning jet propulsion engine for an airplane requires more components: inlets and nozzles. For the inlet, the special care exercised to avoid ingestion of boundary layers air is described. The design features of nozzles are described and extended to include discussion of more extreme configurations such as those found on rocket engines.
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Neidel, A., T. Gädicke, S. Wallich, and E. Engert. "Hydrogen Induced Stress Corrosion Cracking of Fuel Oil Premix Burner Nozzles in a Heavy-duty Gas Turbine Engine | Wasserstoffinduzierte Spannungsrisskorrosion in Heizölvormischbrennerdüsen einer Großgasturbine." In Schadensfallanalysen metallischer Bauteile 2, 335–49. 2nd ed. München: Carl Hanser Verlag GmbH & Co. KG, 2021. http://dx.doi.org/10.3139/9783446470538.031.
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"Failure Analysis of Gas Turbine Engine Fuel Nozzle Heat Shields." In ASM Failure Analysis Case Histories: Improper Maintenance, Repair, and Operating Conditions. ASM International, 2019. http://dx.doi.org/10.31399/asm.fach.usage.c9001508.
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"Inlets, Nozzles, and Combustion Systems." In Elements of Propulsion: Gas Turbines and Rockets, 685–784. Reston ,VA: American Institute of Aeronautics and Astronautics, 2006. http://dx.doi.org/10.2514/5.9781600861789.0685.0784.
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Karadimas, George, Konstantinos Salonitis, and Konstantinos Georgarakis. "Oxide Ceramic Matrix Composite Materials for Aero-Engine Applications: A Literature Review." In Advances in Transdisciplinary Engineering. IOS Press, 2021. http://dx.doi.org/10.3233/atde210029.
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The development of aircraft gas turbine engines has extensively been required for the development of advanced materials. This complex development process is however justified by the system-level benefits in terms of reduced weight, higher temperature capability, and/or reduced cooling, each of which increases efficiency. This is where high-temperature ceramics have made considerable progress and ceramic matrix composites (CMCs) are in the foreground. CMCs are classified into non-oxide and oxide-based ones. Both families have material types that have a high potential for use in high-temperature propulsion applications. Typical oxide-based ones are based on an oxide fiber and oxide matrix (Ox-Ox). Some of the most common oxide subcategories, are alumina, beryllia, ceria, and zirconia ceramics. Such matrix composites are used for example in combustion liners of gas turbine engines and exhaust nozzles. However, until now a thorough study on the available oxide-based CMCs for such applications has not been presented. This paper focus on assessing a literature survey of the available oxide ceramic matrix composite materials in terms of mechanical and thermal properties.
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FUJIYAMA, K., Y. YOSHIOKA, N. OKABE, and K. KIMURA. "CRACK SIMULATION AND LIFE ASSESSMENT OF GAS TURBINE NOZZLES." In Mechanical Behaviour of Materials VI, 73–78. Elsevier, 1992. http://dx.doi.org/10.1016/b978-0-08-037890-9.50131-9.
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"Hot Corrosion of Stage 1 Nozzles in an Industrial Gas Turbine." In Handbook of Case Histories in Failure Analysis, 502–5. ASM International, 1993. http://dx.doi.org/10.31399/asm.fach.v02.c9001281.
Full textConference papers on the topic "Gas Turbine Nozzle"
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Karstensen, Karl W., and Jesse O. Wiggins. "A Variable-Geometry Power Turbine for Marine Gas Turbines." In ASME 1989 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1989. http://dx.doi.org/10.1115/89-gt-282.
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Gas turbines have been accepted in naval surface ship applications, and considerable effort has been made to improve their fuel consumption, particularly at part-load operation. This is an important parameter for shipboard engines because both propulsion and electrical-generator engines spend most of their lives operating at off-design power. An effective way to improve part-load efficiency of recuperated gas turbines is by using a variable power turbine nozzle. This paper discusses the successful use of variable power turbine nozzles in several applications in a family of engines developed for vehicular, industrial, and marine use. These engines incorporate a variable power turbine nozzle and primary surface recuperator to yield specific fuel consumption that rivals that of medium speed diesels. The paper concentrates on the experience with the variable nozzle, tracing its derivation from an existing fixed vane nozzle and its use across a wide range of engine sizes and applications. Emphasis is placed on its potential in marine propulsion and auxiliary gas turbines.
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Napier, James C., and J. P. Arnold. "Advancements in Application of Ceramics to the Gemini Radial-Flow Gas Turbine." In ASME 1985 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1985. http://dx.doi.org/10.1115/85-gt-183.
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A ceramic vane turbine nozzle design has completed over 2000 hours of engine endurance testing for production qualification in the Gemini radial-flow turbine manufactured by the TurboMach Division of Solar Turbines Incorporated. The nozzle offers the advantage of substantial erosion and corrosion durability over all-superalloy types. Cost effective production methods and tooling have been demonstrated on a production lot of nozzles and all preproduction tasks including establishment of ceramic vane acceptance criteria have been completed. Further application of ceramics to the Gemini turbine are described and include all static hot-section components, i.e., combustor liner, combustor scroll, nozzle assembly and exhaust scroll. These ceramic components allow temperature up-rating of the Gemini turbine’s power by 50 percent.
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Rizk, N. K., J. S. Chin, and M. K. Razdan. "Modeling of Gas Turbine Fuel Nozzle Spray." In ASME 1995 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1995. http://dx.doi.org/10.1115/95-gt-225.
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Satisfactory performance of the gas turbine combustor relies on the careful design of various components, particularly the fuel injector. It is, therefore, essential to establish a fundamental basis for fuel injection modeling that involves various atomization processes. A 2-D fuel injection model has been formulated to simulate the airflow within and downstream of the atomizer and address the formation and breakup of the liquid sheet formed at the atomizer exit. The sheet breakup under the effects of airblast, fuel pressure, or the combined atomization mode of the air-assist type is considered in the calculation. The model accounts for secondary breakup of drops and the stochastic Lagrangian treatment of spray. The calculation of spray evaporation addresses both droplet heat-up and steady-state mechanisms, and fuel vapor concentration is based on partial pressure concept. An enhanced evaporation model has been developed that accounts for multicomponent, finite mass diffusivity and conductivity effects, and addresses near critical evaporation. The present investigation involved predictions of flow and spray characteristics of two distinctively different fuel atomizers under both nonreacting and reacting conditions. The predictions of the continuous phase velocity components and the spray mean drop sizes agree well with the detailed measurements obtained for the two atomizers, which indicates the model accounts for key aspects of atomization. The model also provides insight into ligament formation and breakup at the atomizer exit and the initial drop sizes formed in the atomizer near field region where measurements are difficult to obtain. The calculations of the reacting spray show the fuel rich region occupied most of the spray volume with two-peak radial gas temperature profiles. The results also provided local concentrations of unburned hydrocarbon (UHC) and carbon monoxide (CO) in atomizer flowfield, information that could support the effort to reduce emission levels of gas turbine combustors.
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Sundararaj, Ramraj H., T. Chandra sekar, Rajat Arora, A. N. Rao, and Abhijit Kushari. "Performance Simulation of an Engine Retrofitted With Thrust Vectoring Capabilities." In ASME 2019 Gas Turbine India Conference. American Society of Mechanical Engineers, 2019. http://dx.doi.org/10.1115/gtindia2019-2448.
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Abstract Thrust vectoring is a requirement for fifth generation fighters, giving them super-maneuverability capabilities, allowing them to execute tactical maneuvers that are not possible using conventional aerodynamic mechanisms. The most widely used and successful method for achieving this is by using gimbaled engines or nozzles. The complexities involved in this method, have encouraged future engine designers to explore different avenues for achieving thrust vectoring, one of which is fluidic thrust vectoring. In fluidic thrust vectoring, jet deflection is achieved by fluid injection at various locations on the nozzle. During thrust vectoring operations, the engine performance is affected. This is primarily due to the change in effective nozzle area. When a nozzle is gimbaled, as is the method used in currently operational thrust vectored engines, or during fluidic thrust vectoring operations, there is a change in effective nozzle area. This impacts the engine mass flow rate, thus affecting the engine operation. The change in performance is similar to that of an engine fitted with a variable area nozzle. In this study, we attempted to retrofit a thrust vectoring nozzle to an existing engine designed for a fourth-generation fighter aircraft, in order to give it fifth-generation fighter aircraft capabilities. A Twin spool mixed flow turbofan engine with a convergent nozzle is selected and its performance is simulated using Gasturb 13. The baseline engine consists of a low pressure spool, high pressure spool, combustion chamber and convergent-divergent nozzle. For the sake of simplicity, the convergent-divergent nozzle is replaced with a convergent nozzle, with no loss in thrust at design point. The design point is arrived at based on engine data available in open literature. Following this, offdesign performance is simulated, for studying the effect of thrust vectoring operations, which are modeled as a nozzle area change. Suitably scaled generic maps provided in Gasturb are used for off-design simulations. The effect of nozzle area change on engine performance is studied at sea level static conditions. The nozzle area is decreased by a maximum of 15%, in steps of 1%. During thrust vectoring operations, there is a significant change in bypass ratio and fan surge margin, with the other performance parameters being relatively constant. Following this, simulations are conducted at different flight conditions to understand the effect of nozzle area change for different flight regimes. A total of seven different flight conditions are selected to understand the operational envelope of thrust vectoring operation. It is found that at all flight conditions, thrust vectoring has a significant influence on bypass ratio and fan surge margin. While for most conditions, there is an improvement in fan surge margin, there are two conditions where fan surge margin decreases substantially.
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Zhu, J. Y., M. Y. Hou, and J. S. Chin. "Experimental Study on the Atomization of Plain Orifice Injector Under Uniform Cross Air Flow." In ASME 1986 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1986. http://dx.doi.org/10.1115/86-gt-43.
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The effects of air velocity, fuel nozzle pressure drop, nozzle orifice diameter, axial and radial distance from the nozzle on a plain orifice injector atomization have been studied. The effect of interaction of two nozzles on atomization has been examined. The empirical correlation of Mass Median Diameter (MMD) of a plain orifice injector under uniform cross air flow is presented.
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Shirmohamadi, Manuchehr, Shawn Bratt, Difei Wang, Ahmad Ganji, and Rob Neville. "Nozzle Cracking in Gas Turbines." In ASME Turbo Expo 2000: Power for Land, Sea, and Air. American Society of Mechanical Engineers, 2000. http://dx.doi.org/10.1115/2000-gt-0423.
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Cracking of 1st stage nozzles occurs routinely in industrial gas turbines. Such cracking, which sometimes is observed after only a few operating cycles, leads to short intervals between overhauls and expensive repairs or replacements of the nozzle sections. Despite the enormous economic costs associated with nozzle cracking; mechanisms, root causes, and optimum mitigation of such cracking has not been convincingly determined. In fact, some original equipment manufacturers (OEM) consider nozzle cracking as a “fact of life” and have only issued recommendations for their monitoring and replacements. The objective of this work was to identify the mechanisms and the root causes for the 1st stage nozzle cracking of a Frame-3 gas turbine. In this project, we measured, monitored, modeled, and analyzed the temperature and stress response of the nozzles during actual operating practices. This work concluded that the primary mechanism of such cracking is low cycle fatigue caused by high cyclic stresses. The major source of these high stresses was determined to be related to thermal and pressure shocks caused by transonic events which occur during the unit startups. This work offers recommendations to mitigate such cracking and a testing program to verify the root cause.
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Kukutla, Pol Reddy, and B. V. S. S. S. Prasad. "Fluid Thermal Network Studies on Cooled Nozzle Guide Vane." In ASME 2019 Gas Turbine India Conference. American Society of Mechanical Engineers, 2019. http://dx.doi.org/10.1115/gtindia2019-2651.
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Abstract The aerothermal analysis is reported with the help of one-dimensional network modeling for the impingement cum film cooled gas turbine vane. The purpose of this one-dimensional simulation is to obtain the optimized film hole diameters of each row by analyzing the coolant flow distribution and overall effectiveness variations. FlownexR2017 commercial code is used to determine the detailed steady-state performance of the cooling vane. The results show that it is a useful simulation tool to obtain improved effectiveness of film cooling rows in a relatively short turn around time.
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Kannan, Ramesh, Bhamidi Prasad, Sridhara Koppa, Libin George, and Kuppusamy Karuppanan. "The Study on Effect of the Number of Nozzle Vanes in a Radial Flow Turbine for the Turbocharger." In ASME 2019 Gas Turbine India Conference. American Society of Mechanical Engineers, 2019. http://dx.doi.org/10.1115/gtindia2019-2729.
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Abstract The effect of number of nozzle vanes in the turbine stage of a turbocharger is studied using computational fluid dynamics. The nozzle vane unit having 8, 9 and 10 numbers of nozzle vanes configuration is proposed for the radial flow turbine with 30 mm wheel tip diameter. At maximum opening position of the nozzle vanes and for the typical turbine expansion ratio of 2.5, the reduction in mass flow parameter with 10 numbers of nozzle vanes is about 1% lower compared to the 8 numbers of nozzle vanes. The maximum turbine flow range is not affected with higher number of nozzle vanes. The improvement in flow guidance is observed in nozzle vane unit having 10 numbers of nozzle vanes. The improvement in pressure distribution is observed in both the nozzle vane and turbine wheel with increase in number of nozzle vanes. The entropy generation in a turbine stage is found to decrease with increase in the number of nozzle vanes.
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Myers, G. D., J. P. Armstrong, C. D. White, S. Clouser, and R. J. Harvey. "Development of an Innovative High-Temperature Gas Turbine Fuel Nozzle." In ASME 1991 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1991. http://dx.doi.org/10.1115/91-gt-036.
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The objective of the Innovative High-Temperature Fuel Nozzle Program was to design, fabricate, and test propulsion engine fuel nozzles capable of performance despite extreme fuel and air inlet temperatures. Although a variety of both passive and active methods for reducing fuel wetted-surface temperatures were studied, simple thermal barriers were found to offer the best combination of operability, cycle flexibility, and performance. A separate nozzle material study examined several nonmetallics and coating schemes for evidence of passivating or catalytic tendencies. Two pilotless airblast nozzles were developed by employing finite-element modeling to optimize thermal barriers in the stem and tip. Operability of these prototypes was compared to a current state-of-the-art piloted, prefilming airblast nozzle, both on the spray bench and through testing in a can-type combustor. The three nozzles were then equipped with internal thermocouples and operated at 1600F air inlet temperature while injecting marine diesel fuel heated to 350F. Measured and predicted internal temperatures as a function of fuel flow rate were compared. Results show that the thermal barrier systems dramatically reduced wetted-surface temperatures and the potential for coke fouling, even in an extreme environment.
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McClung, R. Craig, Henry L. Bernstein, Janet P. Buckingham, James M. Allen, and George L. Touchton. "Probabilistic Analyses of Industrial Gas Turbine Durability." In ASME 1993 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1993. http://dx.doi.org/10.1115/93-gt-427.
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Probabilistic analyses of component durability have been developed for first stage buckets and nozzles in General Electric MS7001B and MS7001E industrial gas turbine engines. The analyses illustrate two different approaches to the development of probabilistic durability algorithms. The bucket algorithm is built around an existing thermal-mechanical fatigue life model which predicts average fatigue life as a function of local strains, dwell times, and constants in the fatigue model derived from laboratory testing. The fatigue model is linked to a fast probability integration method which calculates the uncertainty or variability in the response variable (life or damage) resulting from uncertainties or variabilities in the physical input variables. The nozzle cracking algorithm, in contrast, is built around a large data base of crack length sums for each vane on seven different nozzles from the field. The crack length sums on nozzles with similar numbers of fired starts were described with standard normal distributions. Extreme value distributions were then calculated analytically to describe the crack length sums on the single most heavily cracked vane per nozzle for different probability levels.
Reports on the topic "Gas Turbine Nozzle"
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Butcher, Thomas, and Michael Furey. Development and Validation of Plasma Fuel Nozzles for Gas Turbine and Boiler Applications. Office of Scientific and Technical Information (OSTI), July 2013. http://dx.doi.org/10.2172/1095289.
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