Academic literature on the topic 'Axial turbine stage'
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Journal articles on the topic "Axial turbine stage"
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Goodisman, M. I., M. L. G. Oldfield, R. C. Kingcombe, T. V. Jones, R. W. Ainsworth, and A. J. Brooks. "An Axial Turbobrake." Journal of Turbomachinery 114, no. 2 (1992): 419–25. http://dx.doi.org/10.1115/1.2929160.
Full textAbstract:
The “Axial Turbobrake” (patent applied for) is a novel turbomachine that can be used to absorb power generated by test turbines. Unlike a compressor, there is no pressure recovery through the turbobrake. This simplifies the aerodynamic design and enables high-stage loadings to be achieved. The blades used have high-turning two-dimensional profiles. This paper describes a single-stage axial turbobrake, which is driven by the exhaust gas of the test turbine and is isolated from the turbine by a choked throat. In this configuration no fast-acting controls are necessary as the turbobrake operates automatically with the turbine flow. Tests on a 0.17 scale model show that the performance is close to that predicted by a simple two-dimensional theory, and demonstrate that the turbobrake power absorption can be controlled and hence matched to that typically produced by the first stage of a modern highly loaded transonic turbine. A full-size axial turbobrake will be used in a short-duration rotating turbine experiment in an Isentropic Light Piston Tunnel at RAE Pyestock.
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Klimko, Marek, Pavel Žitek, and Richard Lenhard. "Measurement on Axial Reaction Turbine Stage." MATEC Web of Conferences 328 (2020): 03013. http://dx.doi.org/10.1051/matecconf/202032803013.
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This article describes a measuring methods and evaluating measured data on a single-stage axial turbine with reaction (~ 50 %). One turbine operating mode was selected, in which the traversing behind the nozzle and bucket with two 5-hole pneumatic probes took place. The results are distributions of flow angles, reactions, or losses distribution/efficiencies along the blades.
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Touil, Kaddour, and Adel Ghenaiet. "Blade stacking and clocking effects in two-stage high-pressure axial turbine." Aircraft Engineering and Aerospace Technology 91, no. 8 (2019): 1133–46. http://dx.doi.org/10.1108/aeat-03-2018-0110.
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Purpose The purpose of this paper is to characterize the blade–row interaction and investigate the effects of axial spacing and clocking in a two-stage high-pressure axial turbine. Design/methodology/approach Flow simulations were performed by means of Ansys-CFX code. First, the effects of blade–row stacking on the expansion performance were investigated by considering the stage interface. Second the axial spacing and the clocking positions between successive blade–rows were varied, the flow field considering the frozen interface was solved, and the flow interaction was assessed. Findings The axial spacing seems affecting the turbine isentropic efficiency in both design and off-design operating conditions. Besides, there are differences in aerodynamic loading and isentropic efficiency between the maximum efficiency clocking positions where the wakes of the first-stage vanes impinge around the leading edge of the second-stage vanes, compared to the clocking position of minimum efficiency where the ingested wakes pass halfway of the second-stage vanes. Research limitations/implications Research implications include understanding the effects of stacking, axial spacing and clocking in axial turbine stages, improving the expansion properties by determining the adequate spacing and locating the leading edge of vanes and blades in both first and second stages with respect to the maximum efficiency clocking positions. Practical implications Practical implications include improving the aerodynamic design of high-pressure axial turbine stages. Originality/value The expansion process in a two-stage high-pressure axial turbine and the effects of blade–row spacing and clocking are elucidated thoroughly.
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Gregory, Brent A. "How Many Turbine Stages?" Mechanical Engineering 139, no. 05 (2017): 56–57. http://dx.doi.org/10.1115/1.2017-may-5.
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This article discusses various stages of turbines and the importance of having more stages in turbine design. The article also highlights reasons that determine the designer’s choice to select the number of turbine stages for a given design of gas turbine. The highest performance turbines are defined by lower work requirements and slower velocities in the gas path. The fundamental factors determining performance might be relegated to only two factors: demand on the turbine and axial velocity. Aircraft engine technologies drive new initiatives because of the need to increase firing temperature and dramatically improve efficiency for substantially less weight. Also, the expansion across each stage determined the annulus area so that the optimums implied by the Pearson chart were largely ignored in the article. Developments in aircraft engine gas turbines have forced heavy frame gas turbines’ original equipment manufacturers to rethink many historical paradigms.
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Koprowski, Arkadiusz, and Romuald Rządkowski. "Optimization of Curtis stage in 1 MW steam turbine." E3S Web of Conferences 137 (2019): 01039. http://dx.doi.org/10.1051/e3sconf/201913701039.
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When operating at 3000 rpm, small turbines do not require a gear box and the generator does not require complex electronic software. This paper analyses the various geometries of the Curtis stage, comprising two rotor and stator blades with and without an outlet, from the efficiency point of view. Presented are 3D steady viscous flows. The results were compared with the performance of an axial turbine.
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Agbadede, Roupa, Dennis Uwakwe, and Isaiah Allison. "Preliminary Re-design of an Axial Turbine in an Existing Engine to Meet the Increased Load Demand." European Journal of Engineering Research and Science 5, no. 11 (2020): 1360–64. http://dx.doi.org/10.24018/ejers.2020.5.11.2141.
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This work presents a preliminary design of an axial turbine section in an industrial gas turbine. The design was necessitated following the need to provide a gas turbine of a power output in the range of 48 to 60MW for a mini-city harbouring an oil rig, which was not possible with the old engine. The turbine section is designed to produce a power capable of driving the compressor as well as produce a useful power for electricity. Using proprietary gas turbine performance simulation software called TURBOMATCH and a computer program written in Microsoft Excel, a redesign of the axial turbine component was achieved. Consequent upon redesigning the axial turbine, a preliminary analysis was carried out to ascertain the new turbine stages introduced. The analysis revealed that when one or two turbine stage(s) was used for new engine, it proved unsatisfactory as the blade loading coefficient and the flow efficiency were both beyond the limit acceptable for an optimum performance. A three stage turbine was finally employed having provided a loading coefficient of 2.1, 1.9 and 1.7 for the first, second and the last stages respectively.
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Agbadede, Roupa, Dennis Uwakwe, and Isaiah Allison. "Preliminary Re-design of an Axial Turbine in an Existing Engine to Meet the Increased Load Demand." European Journal of Engineering and Technology Research 5, no. 11 (2020): 1360–64. http://dx.doi.org/10.24018/ejeng.2020.5.11.2141.
Full textAbstract:
This work presents a preliminary design of an axial turbine section in an industrial gas turbine. The design was necessitated following the need to provide a gas turbine of a power output in the range of 48 to 60MW for a mini-city harbouring an oil rig, which was not possible with the old engine. The turbine section is designed to produce a power capable of driving the compressor as well as produce a useful power for electricity. Using proprietary gas turbine performance simulation software called TURBOMATCH and a computer program written in Microsoft Excel, a redesign of the axial turbine component was achieved. Consequent upon redesigning the axial turbine, a preliminary analysis was carried out to ascertain the new turbine stages introduced. The analysis revealed that when one or two turbine stage(s) was used for new engine, it proved unsatisfactory as the blade loading coefficient and the flow efficiency were both beyond the limit acceptable for an optimum performance. A three stage turbine was finally employed having provided a loading coefficient of 2.1, 1.9 and 1.7 for the first, second and the last stages respectively.
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Salah, Salma I., Mahmoud A. Khader, Martin T. White, and Abdulnaser I. Sayma. "Mean-Line Design of a Supercritical CO2 Micro Axial Turbine." Applied Sciences 10, no. 15 (2020): 5069. http://dx.doi.org/10.3390/app10155069.
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Supercritical carbon dioxide (sCO2) power cycles are promising candidates for concentrated-solar power and waste-heat recovery applications, having advantages of compact turbomachinery and high cycle efficiencies at heat-source temperature in the range of 400 to 800 ∘C. However, for distributed-scale systems (0.1–1.0 MW) the choice of turbomachinery type is unclear. Radial turbines are known to be an effective machine for micro-scale applications. Alternatively, feasible single-stage axial turbine designs could be achieved allowing for better heat transfer control and improved bearing life. Thus, the aim of this study is to investigate the design of a single-stage 100 kW sCO2 axial turbine through the identification of optimal turbine design parameters from both mechanical and aerodynamic performance perspectives. For this purpose, a preliminary design tool has been developed and refined by accounting for passage losses using loss models that are widely used for the design of turbomachinery operating with fluids such as air or steam. The designs were assessed for a turbine that runs at inlet conditions of 923 K, 170 bar, expansion ratio of 3 and shaft speeds of 150k, 200k and 250k RPM respectively. It was found that feasible single-stage designs could be achieved if the turbine is designed with a high loading coefficient and low flow coefficient. Moreover, a turbine with the lowest degree of reaction, over a specified range from 0 to 0.5, was found to achieve the highest efficiency and highest inlet rotor angles.
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Němec, Martin, and Tomáš Jelínek. "Adaptable test rig for two-stage axial turbine." MATEC Web of Conferences 345 (2021): 00022. http://dx.doi.org/10.1051/matecconf/202134500022.
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This contribution describes a new test rig for a two-stage axial turbine built in the VZLÚ. The test rig has replaced an original facility used for a full stage aerodynamics investigation. The motivation for the design of the new test facility was the limitations of the original one. The design is briefly discussed, and then the first measurement results are presented. The first operation was performed with a turbine stage already measured in the original facility. This allows the comparison of the most important quantities.
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Jelínek, Tomáš, and Martin Němec. "Investigation of unsteady flow in axial turbine stage." EPJ Web of Conferences 25 (2012): 01035. http://dx.doi.org/10.1051/epjconf/20122501035.
Full textDissertations / Theses on the topic "Axial turbine stage"
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Shannon, Kevin R. (Kevin Robert). "Loss mechanisms in a highly loaded transonic axial turbine stage." Thesis, Massachusetts Institute of Technology, 2018. http://hdl.handle.net/1721.1/120440.
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Thesis: S.M., Massachusetts Institute of Technology, Department of Aeronautics and Astronautics, 2018.<br>Cataloged from PDF version of thesis.<br>Includes bibliographical references (pages 129-130).<br>Flow in a one-and-a-half stage highly loaded transonic axial turbine representative of future generation turbine technology is assessed for its role in loss generation. Steady and unsteady two-dimensional and three-dimensional flow computations, complemented by simplistic control volume analyses as well as test data, provided results for establishing the quantitative level of loss from various sources. The test data has been acquired in a cascade and blowdown turbine research rig. Specifically, the overall loss determined from unsteady three-dimensional flow computations of a cooled one-and-a-half stage turbine is within 6% of that inferred from the blowdown turbine rig test data. The computed flows with different levels of flow and configuration complexities are post-processed and interrogated to allow an estimation of blade profile loss, trailing edge loss, shock loss, endwall loss, secondary flow loss, tip leakage loss, cooling injection loss, and unsteady flow loss. The dominant sources of loss are determined to be the trailing edge loss, profile loss, and tip leakage loss. The computed flows show that the flow deviation in a highly loaded transonic turbine airfoil with trailing edge shocks is negative (-2° to -4°); estimating the trailing edge loss by assuming zero flow deviation in a simple control volume approach would yield a significantly higher value. Loss arising from flow unsteadiness contributes an additional loss of about 1/6 of that in steady flow approximation; 3/4 of the flow unsteadiness induced loss occurs in the downstream vane where the flow is threaded with propagating shocks from the upstream blade and downstream shock reflections; and the remaining 1/4 is from unsteadiness in NGV wakes and shock oscillations from influence of the adjacent airfoil row. 1/5 of the overall loss in the one-and-a-half stage turbine is from the cooling and purge flows. A preliminary assessment of loss variation with turbine stage pressure ratio shows a non-monotonic trend.<br>by Kevin R. Shannon.<br>S.M.
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Miller, Robert John. "An investigation into the unsteady blade interaction in one and a half stage axial flow turbine." Thesis, University of Oxford, 1998. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.299161.
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Stefanis, Vassilis. "Investigation of flow and heat transfer in stator well cavities of a two-stage axial turbine." Thesis, University of Sussex, 2007. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.444347.
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Dominik, Dávid. "Návrh aeroderivátu pro využití v kompresních stanicích." Master's thesis, Vysoké učení technické v Brně. Fakulta strojního inženýrství, 2020. http://www.nusl.cz/ntk/nusl-417593.
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This thesis is concerned with the calculation of the power turbine. This turbine should be used in the automatic drive of the compressor used for compression of natural gas in compressor stations. Flight engine aeroderivate from the Rolls-Roye company, type RB211-22B, was used as gas generator. The main aim of the thesis is to summarize of the base atributes of the combustion turbines and aeroderivates. They are used for automatic engine, application a thermodynamic calculation of the power turbine, for reaction stage and basic strength calculations.
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Hauptmann, Thomas [Verfasser]. "Einfluss regenerationsbedingter Varianzen der Schaufelgeometrie auf erzwungene Schwingungen in einer mehrstufigen Turbine : The influence of regeneration-induced variances on forced response in a multi-stage axial turbine / Thomas Hauptmann." Hannover : Gottfried Wilhelm Leibniz Universität Hannover, 2020. http://d-nb.info/1216240930/34.
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Baker, Jonathan D. "Analysis of the sensitivity of multi-stage axial compressors to fouling at various stages." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2002. http://library.nps.navy.mil/uhtbin/hyperion-image/02Sep%5FBaker.pdf.
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Максюта, Дмитрий Игоревич. "Комбинированный метод аэродинамической оптимизации ступени осевой турбины". Thesis, НТУ "ХПИ", 2016. http://repository.kpi.kharkov.ua/handle/KhPI-Press/21648.
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Диссертация на соискание ученой степени кандидата технических наук по специальности 05.05.16 – турбомашины и турбоустановки. – Национальный технический университет "Харьковский политехнический институт", Харьков, 2016. Диссертация посвящена разработке комбинированного метода аэродинамической оптимизации ступени осевой турбины, который основываясь на поочередном решении одномерной и трехмерной задач, позволяет значительно повысить эффективность всей ступени при этом учитывая как характер течения рабочего тела в решетках, так и влияние на него протечек. На основании современной тенденций к использованию методов вычислительной аэродинамики (CFD) при оптимизации проточных частей осевых турбин и при этом задействуя как можно большее количество управляющих параметров в оптимизационном процессе, предложен комбинированный метод оптимизации. Предложенный метод использует одномерную и трехмерную оптимизацию, что позволяет существенно повышать аэродинамическую эффективность ступеней, при этом значительно экономя время, необходимое для проведения расчетов. С помощью предложенного метода оптимизации и методики расчета протечек в осерадиальном уплотнении выполнена оптимизация 3-й ступени ЦВД турбины К-540-23,5. Результаты проведенных расчетов показали, что повышение эффективности ступени на этапе одномерной оптимизации происходит за счет выбора на среднем радиусе оптимальных α1, β2, значений степени реактивности ρ и относительного шага решетки t/b. Повышение эффективности ступени на этапе трехмерной оптимизации происходит за счет: выбора оптимального значения входного геометрического угла β1г рабочего профиля, обеспечившего улучшение обтекания профиля; устранения локальных диффузорных участков в межлопаточном канале; нахождения оптимальных законов закрутки, обеспечивающих равномерное натекание потока по всей высоте рабочих лопаток. Суммарно абсолютный КПД новой ступени увеличился более чем на 1 %.<br>Thesis for degree of Candidate of Sciences in Technique for speciality 05.05.16 – turbomachinery and turbine-installations. – National Technical University "Kharkiv Polytechnical Institute", Kharkiv, 2016.
This thesis deals with the development of the combined method of aerodynamic optimization of the axial turbine stage, based on the iterative usage of one-dimensional and three-dimensional theories, thereby can significantly improve the efficiency of the entire stage taking into account the nature of the flow around turbine profiles and the impact of leakage on it. Based on current trends of using computational fluid dynamic methods (CFD) while optimizing of the flow path of the axial turbines, with engaging the largest pos-sible number of control parameters in the optimization process, the combined optimization method is provided. Developed method uses one-dimensional and three-dimensional optimization theories and can noticeably improve aerodynamic efficiency of whole turbine stage, thus significantly saving the time required for the simulations. A three-step comprehensive comparison of the results of simulations with the experimental data confirmed the accuracy of CFD usage while developing the optimization method. To calculate amount of leakage in the radial clearance during one-dimensional optimization phase more accurate, the methodology of flow rate determining in axial-radial seals depending on geometrical, operational characteristics and considering rotor against stator displacement was developed using a series of CFD simulations. Advanced CFD study was conducted to compare the axial-radial seal with the straight-flow one and to identify the new more effective designs of seal. It was shown that creation of artificial roughness on the shaft of the straight-flow seal could reduce the leakage by 45 % compared to the axial-radial seal. Utilizing the developed optimization method and the methodology of leakage calculation in the axial-radial seal, the optimization of the 3rd stage of the high pressure turbine K-540-23,5 was made. As a result of the optimization a new stage with an absolute efficiency increase more than 1 % compared to the original design was obtained.
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Максюта, Дмитро Ігорович. "Комбінований метод аеродинамічної оптимізації ступеня осьової турбіни". Thesis, НТУ "ХПІ", 2016. http://repository.kpi.kharkov.ua/handle/KhPI-Press/21646.
Full textAbstract:
Дисертація присвячена розробці комбінованого методу аеродінамічної оптимізації ступеня осьової турбіни, який ґрунтуючись на почерговому вирішенні одновимірної та тривимірної задач, дозволяє значно підвищити ефективність всього ступеня враховуючи як характер течії робочого тіла в решітках, так і вплив на неї витоки. На підставі сучасної тенденцій до використання методів чисельної аеродинаміки (CFD) при оптимізації проточних частин осьових турбін і при цьому задіяючи якомога більшу кількість управляючих параметрів в оптимізаційному процесі, запропонований комбінований метод оптимізації. Запропонований метод використовує одновимірну та тривимірну оптимізації, що дозволяє істотно підвищувати аеродинамічну ефективність ступенів, при цьому значно заощаджуючи час, необхідний для проведення розрахунків. При розробці методу оптимізації достовірність застосування методів CFD підтверджена шляхом триетапного порівняння результатів розрахунків з результатами експериментальних досліджень. Для отримання більш точних даних кількості витоки робочого тіла в радіальний зазор при проведенні етапу одновимірної оптимізації, розроблена методика для визначення коефіцієнта витрати вісерадіального ущільнення в залежності від його геометричних і режимних характеристик, а також з урахуванням зсуву ротора відносно статора від теплового розширення. Дана методика розроблялася шляхом проведення серії CFD розрахунків. Додатково проведено CFD дослідження для порівняння вісерадіальних ущільнень з прямоточними та виявлення нових ефективних конструкцій ущільнень, яке показало, що шляхом створення штучної шорсткості на валу прямоточного ущільнення можна зменшити витрату через нього на 45 % в порівнянні з вісерадіальними ущільненнями. За допомогою запропонованого методу оптимізації та методики розрахунку витоки в вісерадіальному ущільненні виконана оптимізація 3-го ступеня ЦВТ турбіни К-540-23,5. Результати проведених розрахунків показали, що абсолютний ККД нового ступеня збільшився більш ніж на 1 %.<br>Thesis for degree of Candidate of Sciences in Technique for speciality 05.05.16 – turbomachinery and turbine-installations. – National Technical University "Kharkiv Polytechnical Institute", Kharkiv, 2016.
This thesis deals with the development of the combined method of aerodynamic optimization of the axial turbine stage, based on the iterative usage of one-dimensional and three-dimensional theories, thereby can significantly improve the efficiency of the entire stage taking into account the nature of the flow around turbine profiles and the impact of leakage on it. Based on current trends of using computational fluid dynamic methods (CFD) while optimizing of the flow path of the axial turbines, with engaging the largest pos-sible number of control parameters in the optimization process, the combined optimization method is provided. Developed method uses one-dimensional and three-dimensional optimization theories and can noticeably improve aerodynamic efficiency of whole turbine stage, thus significantly saving the time required for the simulations. A three-step comprehensive comparison of the results of simulations with the experimental data confirmed the accuracy of CFD usage while developing the optimization method. To calculate amount of leakage in the radial clearance during one-dimensional optimization phase more accurate, the methodology of flow rate determining in axial-radial seals depending on geometrical, operational characteristics and considering rotor against stator displacement was developed using a series of CFD simulations. Advanced CFD study was conducted to compare the axial-radial seal with the straight-flow one and to identify the new more effective designs of seal. It was shown that creation of artificial roughness on the shaft of the straight-flow seal could reduce the leakage by 45 % compared to the axial-radial seal. Utilizing the developed optimization method and the methodology of leakage calculation in the axial-radial seal, the optimization of the 3rd stage of the high pressure turbine K-540-23,5 was made. As a result of the optimization a new stage with an absolute efficiency increase more than 1 % compared to the original design was obtained.
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"Experimental Study of Main Gas Ingestion in a Subscale 1.5-stage Axial Flow Air Turbine." Master's thesis, 2015. http://hdl.handle.net/2286/R.I.36468.
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abstract: Gas turbine efficiency has improved over the years due to increases in compressor
pressure ratio and turbine entry temperature (TET) of main combustion gas, made viable
through advancements in material science and cooling techniques. Ingestion of main
combustion gas into the turbine rotor-stator disk cavities can cause major damage to the
gas turbine. To counter this ingestion, rim seals are installed at the periphery of turbine
disks, and purge air extracted from the compressor discharge is supplied to the disk
cavities. Optimum usage of purge air is essential as purge air extraction imparts a penalty on turbine efficiency and specific fuel consumption.
In the present work, experiments were conducted in a newly constructed 1.5-stage
axial flow air turbine featuring vanes and blades to study main gas ingestion. The disk
cavity upstream of the rotor, the 'front cavity', features a double seal with radial clearance
and axial overlap at its rim. The disk cavity downstream of the rotor, the 'aft cavity', features a double seal at its rim but with axial gap. Both cavities contain a labyrinth seal radially inboard; this divides each disk cavity into an 'inner cavity' and a 'rim cavity'.
Time-averaged static pressure at various locations in the main gas path and disk
cavities, and tracer gas (CO2) concentration at different locations in the cavities were
measured. Three sets of experiments were carried out; each set is defined by the main air flow rate and rotor speed. Each of the three sets comprises of four different purge air flow rates, low to high.
The mass flow rate of ingested main gas into the front and aft rim cavities is
reported at the different purge air flow rates, for the three experiment sets. For the present stage configuration, it appears that some ingestion persisted into both the front and aft rim cavities even at high purge air flow rates. On the other hand, the front and aft inner cavity were completely sealed at all purge flows.<br>Dissertation/Thesis<br>Masters Thesis Engineering 2015
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Abdelfattah, Sherif Alykadry. "Numerical and Experimental Analysis of Multi-Stage Axial Turbine Performance at Design and Off-Design Conditions." Thesis, 2013. http://hdl.handle.net/1969.1/151083.
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Computational fluid dynamics or CFD isan importanttool thatis used at various stages in the design of highly complex turbomachinery such as compressorand turbine stages that are used in land and air based power generation units. The ability of CFD to predict the performance characteristics of a specific blade design is challenged by the need to use various turbulence models to simulate turbulent flows as well as transition models to simulate laminar to turbulent transition that can be observed in various turbomachinery designs. Moreover, CFD is based on numerically solving highly complex differential equations, which through the use of a grid to discretize the geometry introduces numerical errors. Allthese factors combine to challenge CFD’s role as a predictor of blade performance. It has been generallyfound that CFD in its current state of the art is best used to compare between various design points and not as a pure predictor of performances.
In this study the capability of CFD, and turbulence modeling, in turbomachinery based geometry is assessed.Three different blade designs are tested, that include an advanced two-stage turbine blade design, a three stage 2D or cylindrical design and finally a three stage bowed stator and rotor design. Allcases were experimentally tested at the Texas A&Muniversity Turbomachinery Performance and Flow Research Laboratory (TPFL).In all cases CFD provided good insights into fundamental turbomachinery flow physics, showing the expected improvement from using 2D cylindrical blades to 3D bowed blade designs in abating the secondary flow effects which are dominant loss generators.However, comparing experimentally measured performance results to numerically predicted shows a clear deficiency, where the CFD overpredicts performance when compared to experimentallyobtained data, largely underestimating the various loss mechanisms. In a relative sense, CFD as a tool allows the user to calculate the impact a new feature or change can have on a baseline design. CFD will also provide insight into what are the dominant physics that explain why a change can provide an increase or decrease in performance.
Additionally,as part of this study, one of the main factors that affect the performance of modern turbomachinery is transition from laminar to turbulent flow.Transition is an influential phenomena especially in high pressure turbines, and is sensitive to factors such asupstream incidentwake frequency and turbulence intensity.A model experimentally developed, is implemented into a CFD solver and compared to various test results showing greater capability in modeling the effects of reduced frequency on the transition point and transitional flow physics. This model is compared to industry standard models showing favorable prediction performance due to its abilityto account for upstream wake effects which most current model are unable to account for.
Books on the topic "Axial turbine stage"
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Escudier, Marcel. Flow through axial-flow-turbomachinery blading. Oxford University Press, 2018. http://dx.doi.org/10.1093/oso/9780198719878.003.0014.
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This chapter is concerned primarily with the flow of a compressible fluid through stationary and moving blading, for the most part using the analysis introduced in Chapter 11. The principles of dimensional analysis are applied to determine the appropriate non-dimensional parameters to characterise the performance of a turbomachine. The analysis of incompressible flow through a linear cascade of aerofoil-like blades is followed by the analysis of compressible flow. Velocity triangles for flow relative to blades, and Euler’s turbomachinery equation, are introduced to analyse flow through a rotor. The concepts introduced are applied to the analysis of an axial-turbomachine stage comprising a stator and a rotor, which applies to either a compressor or a turbine.
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E, Steinthorsson, Rigby David L, and Lewis Research Center, eds. Effects of tip clearance and casing recess on heat transfer and stage efficiency in axial turbines. National Aeronautics and Space Administration, Lewis Research Center, 1998.
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Effects of tip clearance and casing recess on heat transfer and stage efficiency in axial turbines. National Aeronautics and Space Administration, Lewis Research Center, 1998.
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United States. National Aeronautics and Space Administration. and United States. Army Aviation Systems Command., eds. Unsteady flows in a single-stage transonic axial-flow fan stator row. National Aeronautics and Space Administration, 1986.
Find full textBook chapters on the topic "Axial turbine stage"
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Lei, Zongqi, Lei Zhao, Weitao Hou, Shiji Wang, and Jing Wang. "Clocking of Stators and Rotors in a Three-Stage Axial Turbine." In Lecture Notes in Electrical Engineering. Springer Singapore, 2021. http://dx.doi.org/10.1007/978-981-16-7423-5_23.
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Subbarao, Rayapati. "Flow Rate and Axial Gap Studies on a One-and-a-Half-Stage Axial Flow Turbine." In Lecture Notes in Mechanical Engineering. Springer Singapore, 2020. http://dx.doi.org/10.1007/978-981-15-1892-8_30.
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Gallus, H. E., C. A. Poensgen, and J. Zeschky. "Three-Dimensional Unsteady Flow in a Single Stage Axial-Flow Turbine and Compressor." In Unsteady Aerodynamics, Aeroacoustics, and Aeroelasticity of Turbomachines and Propellers. Springer New York, 1993. http://dx.doi.org/10.1007/978-1-4613-9341-2_24.
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Topalovic, Daniel, Rudibert King, Markus Herbig, Alexander Heinrich, and Dieter Peitsch. "Efficiency Increase and Start-Up Strategy of an Axial Turbine Stage Under Periodic Inflow Conditions Using Extremum Seeking Control." In Notes on Numerical Fluid Mechanics and Multidisciplinary Design. Springer International Publishing, 2021. http://dx.doi.org/10.1007/978-3-030-90727-3_18.
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Biollo, Roberto, and Ernesto Benini. "State-of-Art of Transonic Axial Compressors." In Advances in Gas Turbine Technology. InTech, 2011. http://dx.doi.org/10.5772/25257.
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"Preliminary Aerodynamic Design of Axial-Flow Turbine Stages." In Turbine Aerodynamics: Axial-Flow and Radial-Flow Turbine Design and Analysis. ASME Press, 2006. http://dx.doi.org/10.1115/1.802418.ch6.
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"Preliminary Aerodynamic Design of Radial-Inflow Turbine Stages." In Turbine Aerodynamics: Axial-Flow and Radial-Flow Turbine Design and Analysis. ASME Press, 2006. http://dx.doi.org/10.1115/1.802418.ch10.
Full textConference papers on the topic "Axial turbine stage"
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Johnson, Mark S. "One-Dimensional, Stage-by-Stage, Axial Compressor Performance Model." 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-192.
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This paper presents a description of a one-dimensional, constant-radius, stage-by-stage (blade-element) axial compressor model used for compressor map generation and gas turbine off-design performance prediction. This model is designed for investigators who are without access to the proprietary compressor performance information of the gas turbine manufacturers but who are nevertheless interested in predicting the off-design performance of large utility gas turbine power systems. Model performance results (compressor maps) are reported for simulation of a nineteen-stage axial compressor designed by Allison Gas Turbine for the Electric Power Research Institute. The model is further demonstrated by simulating the NACA Eight Stage compressor. The resulting compressor maps are in good qualitative agreement with published maps and are useful for gas turbine power system performance simulation studies. This general-purpose modeling procedure can be applied to any axial compressor for which sufficient airfoil geometry and design-point performance information is known.
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Kumar, S. Satish, Ranjan Ganguli, S. B. Kandagal, and Soumendu Jana. "Flow Behavior in a Transonic Axial Compressor Stage." In ASME 2015 Gas Turbine India Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gtindia2015-1231.
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The steady and unsteady flow characteristics typically vary along and across the axial compressor stage. This coupled with asymmetric rotor tip clearance that occurs in practice makes flow even more complex. Understanding the complex flow behavior inside the transonic compressor stage will aid in developing flow control devices that are meant for purposes such as improving the rotating stall margin, flutter margin, etc. Here, a detailed time averaged numerical analysis is performed on the single stage transonic axial compressor with averaged rotor tip clearance (1.75% of rotor tip axial chord). An attempt is made to study the compressor stall phenomenon. Computational Fluid Dynamics (CFD) helps in resolving the complex flow features involved in a turbomachinery component and at transonic Mach numbers fairly well. Commercial tool ANSYS CFX is used for solving the 3D compressible Reynolds Averaged Navier-Stokes (RANS) equation with Shear Stress Transport (SST) turbulence model. Grid independency is carried out for three different mesh size models. All mesh models chosen have fine mesh near wall boundary regions to capture the boundary layer effects. Overall performance maps of the compressor are generated for 50% to 100% rated design speeds in steps of 10% for the chosen optimum grid. Flow variations along the blade annulus are studied for three different operating conditions: choke/free flow, peak efficiency and near stall flow conditions and for different speeds. Flow parameters such as Mach number, static and total pressure variations, etc. are studied at the inlet to rotor, exit to rotor and exit to stator for the various flow conditions and speeds. The boundary layer growth is clearly captured when the flow is throttled from choke/free flow conditions to near stall condition for all the speeds investigated. Mach number variation along blade height clearly shows decrease in Mach number as stall is approached. Blade loading distribution of the rotor at hub, mean and tip sections are clearly captured. Shock motion from around mid-chord region at free flow condition to towards the leading edge at near stall condition is clearly highlighted. Velocity streamlines near the tip section show the complex interaction of the tip leakage and clearance flows. Velocity vectors near the blade tip shows, the backflow near the trailing edge and tendency for leading edge spillage as the back pressure is increased. The flow blockage region is captured in the meridional plot and the motion of vortex core region as stall is approached is demarcated in the r-θ plots. Tangential velocity variation across the annulus for the two flow conditions investigated shows stall initiating from the tip section of the blade as compressor is throttled. Flow compensation at near stall conditions is explained.
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Bloch, Gregory S., and Walter F. O’Brien. "A Wide-Range Axial-Flow Compressor Stage Performance Model." In ASME 1992 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1992. http://dx.doi.org/10.1115/92-gt-058.
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Dynamic compression system response is a major concern in the operability of aircraft gas turbine engines. Multi-stage compression system computer models have been developed to predict compressor response to changing operating conditions. These models require a knowledge of the wide-range, steady-state operating characteristics as inputs, which has limited their use as predicting tools. The full range of dynamic axial-flow compressor operation spans forward and reversed flow conditions. A model for predicting the wide flow range characteristics of axial-flow compressor stages was developed and applied to a 3-stage, low-speed compressor with very favorable results and to a 10-stage, high-speed compressor with mixed results. Conclusions were made regarding the inception of stall and the effects associated with operating a stage in a multistage environment. It was also concluded that there are operating points of an isolated compressor stage that are not attainable when that stage is operated in a multi-stage environment.
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Sell, M., J. Schlienger, A. Pfau, M. Treiber, and R. S. Abhari. "The 2-Stage Axial Turbine Test Facility “LISA”." In ASME Turbo Expo 2001: Power for Land, Sea, and Air. American Society of Mechanical Engineers, 2001. http://dx.doi.org/10.1115/2001-gt-0492.
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This paper describes the design and construction of a new two stage axial turbine test facility, christened “Lisa”. The research objective of the rig is to study the impact (relevance) of unsteady flow phenomena upon the aerodynamic performance, this being achieved through the use of systematic studies of parametric changes in the stage geometry and operating point. Noteworthy in the design of the rig is the use of a twin shaft arrangement to decouple the stages. The inner shaft carries the load from the first stage whilst the outer is used with an integral torque-meter to measure the loading upon the second stage alone. This gives an accurate measurement of the loading upon the aerodynamically representative second stage, which possesses the correct stage inlet conditions in comparison to the full two stage machine which has an unrealistic axial inlet flow at the first stator. A calibrated Venturi nozzle measures the mass flow at an accuracy of below 1%, from which stage efficiencies can be derived. The rig is arranged in a closed loop system. The turbine has a vertical arrangement and is connected through a gear box to a generator system that works as a brake to maintain the desired operating speed. The turbine exit is open to ambient pressure. The rig runs at a low pressure ratio of 1.5. The maximum Mach number at stator exit is 0.3 at an inlet pressure of 1.5 bar. The maximum mass flow is 14 kg/sec. Nominal rotor design speed is 3000 RPM. The tip to hub blade ratio is 1.29, and the nominal axial chord is 50 mm. The rig is designed to accommodate a broad range of measurement techniques, but with a strong emphasis upon unsteady flow methods, for example fast response aerodynamic pressure probes for time-resolved flow measurements. The first section of this paper describes the overall test facility hardware. This is followed by a detailed focus on the torque measurement device including stage efficiency measurements at operating conditions in Lisa. Discussion of measurement techniques completes the paper.
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Louis, J. F. "Axial Flow Contra-Rotating Turbines." In ASME 1985 International Gas Turbine Conference and Exhibit. American Society of Mechanical Engineers, 1985. http://dx.doi.org/10.1115/85-gt-218.
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Two types of contra-rotating stages are considered; the first uses guide vanes and the second is vaneless. The wheels of the first type use bladings which are mirror images of each other and they operate with inlet and outlet swirl. The second type uses dissimilar bladings in each of the two wheels with axial inlet velocity to the first wheel and axial outlet velocity for the second wheel. An analysis of their performance indicates that both types can reach stage loading coefficients comparable or larger than conventional turbines with the same number of wheels. A comparison of the contra-rotating stages with conventional ones indicate a significant stage efficiency advantage of the contra-rotating over the conventional single rotation stages due mainly to the elimination of stationary vanes. The off-design performance indicates that relative wheel speed must be controlled. The attributes of contra-rotating turbines suggest their potential use in high performance aircraft engines, in dynamic space power systems and in low speed industrial gas turbines.
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Kumar, S. Satish, Lakshya Kumar, R. Senthil Kumaran, Veera Sesha Kumar, and M. T. Shobhavathy. "Design of High Transonic Axial Compressor Stage for Small Gas Turbine Applications." In ASME 2019 Gas Turbine India Conference. American Society of Mechanical Engineers, 2019. http://dx.doi.org/10.1115/gtindia2019-2690.
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Abstract In the quest for achieving high performance, gas turbine engines demand efficient design of various engine components, mainly the compressor stages. The compressor stages consume most of the energy produced by the engine to provide the required pressure ratio. CSIR-NAL is involved in the development of a small gas turbine engine for UAV applications. In this regard, a high transonic single stage axial flow compressor is designed with a mass flow of 4.6 kg/s and pressure ratio of 1.6, for technology demonstration. In this paper, the aerodynamic and structural design of a high transonic axial compressor stage is discussed along with its performance characteristics. Preliminary mean-line design of the compressor stage is carried out, followed by detailed 3D blade design. Aerodynamic performance of the compressor stage is investigated numerically. Grid independency study is carried out, and the flow un-altering grid is used for steady simulations. Steady 3D RANS CFD simulations with SST turbulence model are carried out for estimating the compressor stage performance. At the design speed, the compressor is able to produce the desired pressure ratio and efficiency. Detailed flow investigations across the compressor stage are studied from choke to near stall flow conditions for different speeds. The compressor rotor blisk made of titanium alloy (Ti6AL4V) is subjected to stress analysis. The von-Mises stress and radial deformation are observed to be well within the safe limits of the chosen material. Modal analysis is carried out to study the structural dynamics of the rotor.
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Vashi, Hardik K., Dilipkumar Bhanudasji Alone, and Harish S. Choksi. "Numerically Understanding the Steady State Response of Single Stage Transonic Axial Flow Compressor to Axial Locations of Step for Stepped Tip Clearance." In ASME 2014 Gas Turbine India Conference. American Society of Mechanical Engineers, 2014. http://dx.doi.org/10.1115/gtindia2014-8147.
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This paper describes the steady state numerical work carried out to study the influence of providing stepped tip clearances at various axial locations along the rotor chord on the performance of the single stage transonic axial compressor. Uniform tip clearance study on compressor under consideration showed performance deterioration of compressor at higher tip clearance of 2mm [3.4% of rotor axial chord] therefore in order to improve performance of compressor, stepped tip clearance method was introduced. Commercially available Ansys Fluent 12.0 software was used to perform steady state RANS simulations with three dimensional implicit pressure based solver and SST K-ω as turbulence model. Stepped tip clearance concept is based on providing smaller tip clearance in front portion and providing higher tip clearance after step above the rotor. In present case, study was carried out for stepped tip clearance with steps at four different axial locations [i.e. 10, 20, 40 & 60 % of rotor chord from leading edge of rotor] and results were compared with baseline model of 0.5 mm [0.9% of rotor axial chord] uniform tip clearance at 100% speed. The stepped tip clearance combinations used for the analysis was 0.5–2mm. It was observed that there was increase in compressor peak efficiency & peak pressure ratio for all stepped clearance cases. A trend was noticed where there was relative increase in peak efficiency as well as peak pressure ratio when step was moved downstream along rotor chord [i.e. moving from 10% to 60% axial location]for one of the combinations stepped tip clearance. Stall margin improvement was observed for all cases of stepped tip clearance. Stall margin gain obtained was higher when step was provided in front portion above rotor [i.e. 10 & 20 % axial location] compared to stall margin improvement when step was provided in rear portion above rotor [i.e. 40 & 60% axial location]. It can be concluded that stepped tip clearance provided near leading edge shows potential in improving performance for compressor under consideration. Numerical analysis of single stage has been carried out but data has been presented for rotor only to study the flow changes occurring in rotor vicinity created by implementation of stepped tip clearance method.
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Cyrus, Václav. "The Turbine Regime of a Rear Axial Compressor Stage." In ASME 1990 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1990. http://dx.doi.org/10.1115/90-gt-074.
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A detailed investigation of three-dimensional flow has been carried out in a low speed rear axial compressor stage with aspect ratio of 1 at the extreme off-design condition-turbine regime. Measurements were performed by means of both stationery and rotating pressure probes. The mechanism of flow in the rotor and stator blade row in the turbine regime is analysed. Comparison is made with flow mechanism at the design condition.
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Goodisman, M. I., M. L. G. Oldfield, R. C. Kingcombe, T. V. Jones, R. W. Ainsworth, and A. J. Brooks. "An Axial Turbobrake." 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-001.
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The Axial Turbobrake (Patent applied for) is a novel turbomachine which can be used to absorb power generated by test turbines. Unlike a compressor there is no pressure recovery through the turbobrake. This simplifies the aerodynamic design and enables high stage loadings to be achieved. The blades used have high turning two dimensional profiles. This paper describes a single stage axial turbobrake, which is driven by the exhaust gas of the test turbine and is isolated from the turbine by a choked throat. In this configuration no fast acting controls are necessary as the turbobrake operates automatically with the turbine flow. Tests on a 0.17 scale model, show that the performance is close to that predicted by a simple two-dimensional theory, and demonstrate that the turbobrake power absorption can be controlled and hence matched to that typically produced by the first stage of a modern highly loaded transonic turbine. A full size axial turbobrake will be used in a short duration rotating turbine experiment in an Isentropic Light Piston Tunnel at RAE Pyestock.
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Chen, Yang, Zhuhai Zhong, Jun Li, et al. "Effect of Stage Axial Distances on the Aerodynamic Performance of Three-Stage Axial Turbine Using Experimental Measurements and Numerical Simulations." In ASME Turbo Expo 2017: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/gt2017-63790.
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The stage axial distance significantly influences the aerodynamic performance of turbines under some constraints. Experimental measurements and numerical simulations are used to analyze the effect of stage axial distances on the aerodynamic performance of three-stage axial turbine in this work. The aerodynamic performance of three-stage axial turbine with three different stage axial distances is experimentally measured at the air turbine test rig of Dongfang Steam Turbine Co. LTD. Experimental results show that efficiency increases when the stage axial distance decreases for the geometry under study with relative stage distance ranged from 0.14 to 0.35, and the effect of stage axial distance on the optimization velocity ratio here is very limited. In addition, unsteady Reynolds-Averaged Navier-Stokes (RANS) simulations were carried out with nonlinear harmonic method to analyze the detailed flow field of the experimental three-stage axial turbine. The numerical aerodynamic efficiency of three-stage axial turbine is in good agreement with the experimental data. Furthermore, the small stage axial distance is preferred for the higher efficiency. The detailed flow field and aerodynamic parameters of three-stage axial turbine are also illustrated and discussed.