Academic literature on the topic 'Nozzle cascade of axial turbine'
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Journal articles on the topic "Nozzle cascade of axial turbine"
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Morphis, G., and J. P. Bindon. "The Flow in a Second-Stage Nozzle of a Low-Speed Axial Turbine and Its Effect on Tip Clearance Loss Development." Journal of Turbomachinery 117, no. 4 (October 1, 1995): 571–77. http://dx.doi.org/10.1115/1.2836569.
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The flow field in a one-and-a-half-stage low-speed axial turbine with varying levels of rotor tip clearance was measured in order to compare the behavior of the second nozzle with the first and to identify the manner in which the second nozzle responds to the complex tip clearance dependent flow presented to it and completes the formation of tip clearance loss. The tangentially averaged flow relative to the rotor blade in the tip clearance region was found to differ radically from that found in cascade and is not underturned with a high axial velocity. There is evidence rather of overturning caused by secondary flow. The axial velocity follows an almost normal endwall boundary layer pattern with almost no leakage jet effect. The cascade tip clearance model is therefore not accurate. The reduction in second-stage nozzle loss was shown to occur near the hub and tip, which confirms that it is probably a reduction in secondary flow loss. The nozzle exit loss contours showed that leakage suppressed the formation of the classical secondary flow pattern and that a new tip clearance related loss phenomenon exists on the suction surface. The second-stage nozzle reduced the hub endwall boundary layer below that of both the first nozzle and that behind the rotor. It also rectified secondary and tip clearance flows to such a degree that a second-stage rotor would experience no greater flow distortion than the first-stage rotor. Radial flow angles behind the second-stage nozzle were much smaller than found in a previous study with low-aspect-ratio untwisted blades.
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Flaszynski, Pawel, Michal Piotrowicz, and Tommaso Bacci. "Clocking and Potential Effects in Combustor–Turbine Stator Interactions." Aerospace 8, no. 10 (October 2, 2021): 285. http://dx.doi.org/10.3390/aerospace8100285.
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Investigations of combustors and turbines separately have been carried out for years by research institutes and aircraft engine companies, but there are still many questions about the interaction effect. In this paper, a prediction of a turbine stator’s potential effect on flow in a combustor and the clocking effect on temperature distribution in a nozzle guide vane are discussed. Numerical simulation results for the combustor simulator and the nozzle guide vane (NGV) of the first turbine stage are presented. The geometry and flow conditions were defined according to measurements carried out on a test section within the framework of the EU FACTOR (full aerothermal combustor–turbine interactions research) project. The numerical model was validated by a comparison of results against experimental data in the plane at a combustor outlet. Two turbulence models were employed: the Spalart–Allmaras and Explicit Algebraic Reynolds Stress models. It was shown that the NGV potential effect on flow distribution at the combustor–turbine interface located at 42.5% of the axial chord is weak. The clocking effect due to the azimuthal position of guide vanes downstream of the swirlers strongly affects the temperature and flow conditions in a stator cascade.
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Song, Bo, Wing F. Ng, Joseph A. Cotroneo, Douglas C. Hofer, and Gunnar Siden. "Aerodynamic Design and Testing of Three Low Solidity Steam Turbine Nozzle Cascades." Journal of Turbomachinery 129, no. 1 (March 1, 2004): 62–71. http://dx.doi.org/10.1115/1.2372774.
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Three sets of low solidity steam turbine nozzle cascades were designed and tested. The objective was to reduce cost through a reduction in parts count while maintaining or improving performance. The primary application is for steam turbine high pressure sections where Mach numbers are subsonic and high levels of unguided turning can be tolerated. The base line design A has a ratio of pitch to axial chord of 1.2. This is the pitch diameter section of a 50% reaction stage that has been verified by multistage testing on steam to have a high level of efficiency. Designs B and C have ratios of pitch to axial chord of 1.5 and 1.8, respectively. All three designs satisfy the same inlet and exit vector diagrams. Analytical surface Mach number distributions and boundary layer transition predictions are presented. Extensive cascade test measurements were carried out for a broad incidence range from −60to+35deg. At each incidence, four outlet Mach numbers were tested, ranging from 0.2 to 0.8, with the corresponding Reynolds number variation from 1.8×105 to 9.0×105. Experimental results of loss coefficient and blade surface Mach number are presented and compared for the three cascades. The experimental results have demonstrated low losses over the tested Mach number range for a wide range of incidence from −45to15deg. Designs B and C have lower profile losses than design A. The associated flow physics is interpreted using the results of wake profile, blade surface Mach number distribution, and blade surface oil flow visualization, with the emphasis placed on the loss mechanisms for different flow conditions and the loss reduction mechanism with lower solidity. The effect of the higher profile loading of the lower solidity designs on increased end wall losses induced by increased secondary flow, especially on low aspect ratio designs, is the subject of ongoing studies.
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Subotovich, Subotovich, Alexander Lapuzin, and Yuriy Yudin. "New Methods Used for the Smoothing of the Three-Dimensional Flow Behind the Turbine Nozzle Cascade." NTU "KhPI" Bulletin: Power and heat engineering processes and equipment, no. 1 (October 28, 2021): 38–46. http://dx.doi.org/10.20998/2078-774x.2021.01.07.
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To smooth the parameters of the three-dimensional flow behind the nozzle cascade new methods were suggested that allow us to sustain the flow rate, stagnation enthalpy and the axial projection of the moment of momentum for initial-, nonuniform and averaged flows. It was shown that the choice of the fourth integral characteristic (the kinetic energy, the entropy and the quantity of motion) has no particular significance because it has no effect on the complex criterion of the cascade quality, i.e. the velocity coefficient-angle cosine product that characterizes the level of the radial component of velocity. The minimum values of the velocity coefficient and the cosine angle satisfy the method that allows us to sustain the quantity of motion during the smoothing and the maximum values of the specified nozzle characteristics satisfy method 2 that enables the entropy maintenance. To evaluate the aerodynamic efficiency of the nozzle cascade the preference should be given to method 1 that enables the kinetic energy conservation and the velocity coefficient allows for the precise determination of the degree of loss of the kinetic energy that is equal to 3.6 % as for the example given in the scientific paper. As for method 1, the kinematic losses in the cascade are defined by the angle cosine that characterizes the level of the radial component of the velocity behind the cascade. For the example in question, kinematic losses are equal to 1.9 % and the complex criterion of quality equal to 0.972 corresponds to the overall losses of 5.5 %. It was suggested to use the velocity coefficient and the two angles of flow as integral cascade characteristics. The use of these characteristics enables the correct computations of the efficiency factor for the stage within the one-dimensional computation. The incisive analysis was performed for different methods used for the averaging of the parameters of the axially asymmetric flow behind the nozzle cascade. It was suggested to neglect the flow rate factor in the case of thermal computations done for the turbine stage.
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Boletis, E. "Effects of Tip Endwall Contouring on the Three-Dimensional Flow Field in an Annular Turbine Nozzle Guide Vane: Part 1—Experimental Investigation." Journal of Engineering for Gas Turbines and Power 107, no. 4 (October 1, 1985): 983–90. http://dx.doi.org/10.1115/1.3239845.
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Tip endwall contouring is one of the most effective methods to improve the performance of low aspect ratio turbine vanes [1]. In view of the wide variety of geometric parameters, it appears that only the physical understanding of the three-dimensional flow field will allow us to evaluate the probable benefits of a particular endwall contouring. The paper describes the experimental investigation of the three-dimensional flow through a low-speed, low aspect ratio, high-turning annular turbine nozzle guide vane with meridional tip endwall contouring. The full impact of the effects of tip contouring is evaluated by comparison with the results of a previous study in an annular turbine nozzle guide vane of the same blade and cascade geometry with cylindrical endwalls [12]. In parallel, the present experimental study provides a fully three-dimensional test case for comparison with advanced theoretical calculation methods [15]. The flow is explored by means of double-head, four-hole pressure probes in five axial planes from far upstream to downstream of the blade row. The results are presented in the form of contour plots and spanwise pitch-averaged distributions.
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Rona, Aldo, Renato Paciorri, and Marco Geron. "Design and Testing of a Transonic Linear Cascade Tunnel With Optimized Slotted Walls." Journal of Turbomachinery 128, no. 1 (June 23, 2005): 23–34. http://dx.doi.org/10.1115/1.2101856.
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In linear cascade wind tunnel tests, a high level of pitchwise periodicity is desirable to reproduce the azimuthal periodicity in the stage of an axial compressor or turbine. Transonic tests in a cascade wind tunnel with open jet boundaries have been shown to suffer from spurious waves, reflected at the jet boundary, that compromise the flow periodicity in pitch. This problem can be tackled by placing at this boundary a slotted tailboard with a specific wall void ratio s and pitch angle α. The optimal value of the s-α pair depends on the test section geometry and on the tunnel running conditions. An inviscid two-dimensional numerical method has been developed to predict transonic linear cascade flows, with and without a tailboard, and quantify the nonperiodicity in the discharge. This method includes a new computational boundary condition to model the effects of the tailboard slots on the cascade interior flow. This method has been applied to a six-blade turbine nozzle cascade, transonically tested at the University of Leicester. The numerical results identified a specific slotted tailboard geometry, able to minimize the spurious reflected waves and regain some pitchwise flow periodicity. The wind tunnel open jet test section was redesigned accordingly. Pressure measurements at the cascade outlet and synchronous spark schlieren visualization of the test section, with and without the optimized slotted tailboard, have confirmed the gain in pitchwise periodicity predicted by the numerical model.
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Harasgama, S. P., and E. T. Wedlake. "Heat Transfer and Aerodynamics of a High Rim Speed Turbine Nozzle Guide Vane Tested in the RAE Isentropic Light Piston Cascade (ILPC)." Journal of Turbomachinery 113, no. 3 (July 1, 1991): 384–91. http://dx.doi.org/10.1115/1.2927887.
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Detailed heat transfer and aerodynamic measurements have been made on an annular cascade of highly loaded nozzle guide vanes. The tests were carried out in an Isentropic Light Piston test facility at engine representative Reynolds number, Mach number, and gas-to-wall temperature ratio. The aerodynamics indicate that the vane has a weak shock at 65–70 percent axial chord (midspan) with a peak Mach number of 1.14. The influence of Reynolds number and Mach number on the Nusselt number distributions on the vane and endwall surfaces are shown to be significant. Computational techniques are used for the interpretation of test data.
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Sieverding, C. H., T. Arts, R. De´nos, and F. Martelli. "Investigation of the Flow Field Downstream of a Turbine Trailing Edge Cooled Nozzle Guide Vane." Journal of Turbomachinery 118, no. 2 (April 1, 1996): 291–300. http://dx.doi.org/10.1115/1.2836639.
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A trailing edge cooled low aspect ratio transonic turbine guide vane is investigated in the VKI Compression Tube Cascade Facility at an outlet Mach number M2, is = 1.05 and a coolant flow rate m˙c/m˙g = 3 percent. The outlet flow field is surveyed by combined total-directional pressure probes and temperature probes. Special emphasis is put on the development of low blockage probes. Additional information is provided by oil flow visualizations and numerical flow visualizations with a three-dimensional Navier–Stokes code. The test results describe the strong differences in the axial evolution of the hub and tip endwall and secondary flows and demonstrate the self-similarity of the midspan wake profiles. According to the total pressure and temperature profiles, the wake mixing appears to be very fast in the near-wake but very slow in the far-wake region. The total pressure wake profile appears to be little affected by the coolant flow ejection.
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Martinez-Botas, R. F., G. D. Lock, and T. V. Jones. "Heat Transfer Measurements in an Annular Cascade of Transonic Gas Turbine Blades Using the Transient Liquid Crystal Technique." Journal of Turbomachinery 117, no. 3 (July 1, 1995): 425–31. http://dx.doi.org/10.1115/1.2835678.
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Heat transfer measurements have been made in the Oxford University Cold Heat Transfer Tunnel employing the transient liquid crystal technique. Complete contours of the heat transfer coefficient have been obtained on the aerofoil surfaces of a large annular cascade of high-pressure nozzle guide vanes (mean blade diameter of 1.11 m and axial chord of 0.0664 m). The measurements are made at engine representative Mach and Reynolds numbers (exit Mach number 0.96 and Reynolds number 2.0 × 106). A novel mechanism is used to isolate five preheated blades in the annulus before an unheated flow of air passes over the vanes, creating a step change in heat transfer. The surfaces of interest are coated with narrow-band thermochromic liquid crystals and the color crystal change is recorded during the run with a miniature CCD video camera. The heat transfer coefficient is obtained by solving the one-dimensional heat transfer equation for all the points of interest. This paper will describe the experimental technique and present results of heat transfer and flow visualization.
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Fan, S., and B. Lakshminarayana. "Time-Accurate Euler Simulation of Interaction of Nozzle Wake and Secondary Flow With Rotor Blade in an Axial Turbine Stage Using Nonreflecting Boundary Conditions." Journal of Turbomachinery 118, no. 4 (October 1, 1996): 663–78. http://dx.doi.org/10.1115/1.2840922.
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The objective of this paper is to investigate the three-dimensional unsteady flow interactions in a turbomachine stage. A three-dimensional time-accurate Euler code has been developed using an explicit four-stage Runge–Kutta scheme. Three-dimensional unsteady nonreflecting boundary conditions are formulated at the inlet and the outlet of the computational domain to remove the spurious numerical reflections. The three-dimensional code is first validated for two-dimensional and three-dimensional cascades with harmonic vortical inlet distortions. The effectiveness of the nonreflecting boundary conditions is demonstrated. The unsteady Euler solver is then used to simulate the propagation of nozzle wake and secondary flow through the rotor and the resulting unsteady pressure field in an axial turbine stage. The three-dimensional and time-dependent propagation of nozzle wakes in the rotor blade row and the effects of nozzle secondary flow on the rotor unsteady surface pressure and passage flow field are studied. It was found that the unsteady flow field in the rotor is highly three dimensional and the nozzle secondary flow has significant contribution to the unsteady pressure on the blade surfaces. Even though the steady flow at the midspan is nearly two dimensional, the unsteady flow is three dimensional and the unsteady pressure distribution cannot be predicted by a two-dimensional analysis.
Dissertations / Theses on the topic "Nozzle cascade of axial turbine"
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Баранник, Валентин Сергеевич. "Пространственная аэродинамическая оптимизация направляющей решетки осевой турбины." Thesis, НТУ "ХПИ", 2016. http://repository.kpi.kharkov.ua/handle/KhPI-Press/22677.
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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 методики пространственной аэродинамической оптимизации направляющих решеток осевых турбин путем поиска оптимальных формы профилей и меридиональных обводов межлопаточных каналов. Использование данной методики позволяет при решении оптимизационной задачи учесть дополнительные резервы повышения эффективности. Поиск оптимального варианта осуществлялся с использованием теории планирования эксперимента и ЛПτ – последовательности. Для описания полимодальных целевых функций исходная формальная макромодель в виде полного квадратичного полинома была уточнена путем замены суперпозиции параболы на суперпозицию кубического интерполяционного сплайна. На основе разработанной методики проведена оптимизация направляющей решетки третьей степени мощной паровой турбины с постоянным по высоте профилем при построении его различными типами кривых. Анализ результатов оптимизации показал, что наибольшее снижение интегральных потерь составило 7% в относительных величинах. Снижение потерь было достигнуто, как в ядре потока, так и в области вторичных течений. Существенно влиять на структуру течения в турбинных решетках, а следовательно получать дополнительных выигрыш при постановке оптимизационной задачи позволяет меридиональное профилирование поверхностей межлопаточного канала. Оптимизация периферийного меридионального обвода с помощью разработанного метода позволила дополнительно снизить интегральные потери 1,4%. в относительных величинах. Построение формы меридионального обвода осуществляется с использованием кривых Безье 4-го порядка для решеток без раскрытия и 3-го порядка – для решеток с раскрытием. Использование лопатки переменного по высоте профиля при постановке оптимизационной задачи также позволяет снизить интегральные потери.Thesis for degree of Candidate of Sciences in Technique for speciality 05.05.16 – turbomachine and turbo-installation. – National Technical University "Kharkiv Polytechnical Institute", Kharkiv, 2016. The thesis is devoted to development the methods of the three-dimensional aerodynamic optimization of axial turbine nozzle cascades by defining the optimal shape of profiles and nozzle channel meridional shape. The formulation of an optimization problem using this methods allows to consider the additional efficiency reserves.While implementing developed method design of the turbine profiles using different kinds of curves was carried out. For each of the curve types the control parameters that allow to widely vary the profile geometry were determined. The results reliability was confirmed by providing verification of the nozzle and blade cascade simulations with experimental data. Using developed methods the optimization of the third stage nozzle cascade with a constant height profile of the powerful steam turbine using different types of curves was conducted. As a result of optimization the largest reduction of the integral losses by 7% in relative values was shown. Further optimization of the shroud meridional shape using developed optimization method increased this value by 1.4%. Formulation optimization task Using variable nozzle height profile also reduces the integral loses.
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Бараннік, Валентин Сергійович. "Просторова аеродинамічна оптимізація направляючої решітки осьової турбіни." Thesis, НТУ "ХПІ", 2016. http://repository.kpi.kharkov.ua/handle/KhPI-Press/22676.
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Дисертація на здобуття наукового ступеня кандидата технічних наук за спеціальністю 05.05.16 – турбомашини та турбоустановки. – національний технічний університет "Харківський політехнічний інститут", Харків, 2016. Дисертація присвячена розробці методики просторової аеродинамічної оптимізації напрямних решіток осьових турбін шляхом пошуку оптимальних форми профілів та меридіональних обводів міжлопаткових каналів. Використання даної методики дозволяє при постановці оптимізаційної задачі врахувати додаткові резерви підвищення ефективності. При реалізації цієї методики було виконано проектування турбінних профілів з використанням різного роду кривих. Для кожного типу кривої визначені її параметри управління, що дозволяють в широких межах варіювати геометрію профілю. Достовірність отриманих результатів підтверджується проведеною верифікацією на направляючій та робочій решітці. На основі розробленої методики проведено оптимізацію направляючої решітки третього ступеня потужної парової турбіни з постійним по висоті профілем при побудові його різними типами кривих. Аналіз результатів оптимізації показав, що найбільше зниження інтегральних втрат склало 7% у відносних величинах. Подальша оптимізація периферійного меридіонального обводу за допомогою розробленого методу дозволила збільшити цю величину на 1,4%. Використання лопатки перемінного по висоті профілю при постановці оптимізаційної задачі також дозволяє знизити інтегральні втрати.Thesis for degree of Candidate of Sciences in Technique for speciality 05.05.16 – turbomachine and turbo-installation. – National Technical University "Kharkiv Polytechnical Institute", Kharkiv, 2016. The thesis is devoted to development the methods of the three-dimensional aerodynamic optimization of axial turbine nozzle cascades by defining the optimal shape of profiles and nozzle channel meridional shape. The formulation of an optimization problem using this methods allows to consider the additional efficiency reserves.While implementing developed method design of the turbine profiles using different kinds of curves was carried out. For each of the curve types the control parameters that allow to widely vary the profile geometry were determined. The results reliability was confirmed by providing verification of the nozzle and blade cascade simulations with experimental data. Using developed methods the optimization of the third stage nozzle cascade with a constant height profile of the powerful steam turbine using different types of curves was conducted. As a result of optimization the largest reduction of the integral losses by 7% in relative values was shown. Further optimization of the shroud meridional shape using developed optimization method increased this value by 1.4%. Formulation optimization task Using variable nozzle height profile also reduces the integral loses.
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Mamat, Zainul Asri. "The performance of a cascade of nozzle turbine blading in nucleating steam." Thesis, University of Birmingham, 1996. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.643566.
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The thesis describes an experimental investigation of nucleating flows of steam in a cascade of nozzle turbine blading. To obtain nucleation in subsonic flows of steam, a supercooled supply is necessary which has been achieved under blow-down conditions. The literature survey covers the early investigations of condensation in flowing steam, the development of nucleation theory, droplets growth laws and their simple applications in the study of phenomena associated with condensation in nozzles. Finally some of the problems associated with wetness in turbines are considered. Next the experimental facility is described. The facility is a blow-down steam tunnel constructed for the study of two-dimensional two-phase flows. The instrumentation caters for the measurements of surface pressure distributions, wake traverses downstream of the cascade as well as for the optical observations and droplet size measurements. The main experimental chapter includes the results of the surface pressure measurements, wake traverses downstream of the cascade, flow observations and droplet size measurements. Some measurements of the thermodynamic nucleation losses are also presented together with the calibration of the probe used for the wake traverses. Comparisons of results under steam superheated and nucleating flows reveal the blades to have different characteristics when the outlets have been supersonic or sonic. The most notable difference has been the pressure rise associated with the zone of rapid condensation, its location and its insensitivity to changes in inlet temperatures and overall pressure ratios. In contrast, when the outlet has been subsonic, the blades exhibit similar characteristics under steam superheated and nucleating conditions. In addition, comparisons made with the results obtained previously by other investigators using different blades suggest that the interaction between the zone of rapid condensation and the aerodynamic performance of the blades can vary depending on the blade shape.
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Cleak, James Gilbert Edwin. "Validation of viscous, three-dimensional flow calculations in an axial turbine cascade." Thesis, Durham University, 1989. http://etheses.dur.ac.uk/6429/.
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This thesis presents a detailed investigation of the capability of a modern three-dimensional Navier-Stokes solver to predict the secondary flows and losses in a linear cascade of high turning turbine rotor blades. Three codes were initially tested, to permit selection of the best of the available numerical solvers for this case. This program was then tested in more detail. Results showed that although very accurate prediction of the effects of inviscid fluid mechanics is now possible, the Reynolds stress modelling can have profound effects upon the quality of the solutions obtained. Solutions using two different calculation meshes, have shown that the results are not significantly grid dependent. The flowfield of the cascade was traversed with hot-wires to obtain measurements of the turbulent Reynolds stresses. A turbulence generating grid was placed upstream of the cascade, to produce a more realistic inlet turbulence intensity. Results showed that regions of high turbulent kinetic energy are associated with regions of high total pressure loss. Calculation of eddy viscosities from the Reynolds stresses showed that downstream of the -cascade the eddy viscosity is fairly isotropic. Evaluation of terms in the kinetic energy equation, also indicated that both the normal and shear Reynolds stresses are important as loss producing mechanisms in the downstream flow. The experimental Reynolds stresses have been compared with those calculated from the eddy viscosity and velocity fields of Navier-Stokes predictions using a mixing length turbulence model, a one equation model, and K - ϵ model. It was found that in the separated, shear flows, agreement was poor, although the K - ϵ model performed best. Further experimental work is suggested to obtain data with which to determine the accuracy of the models within the blade and endwall boundary layers.
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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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Rottmeier, Fabrice. "Experimental investigation of a vibrating axial turbine cascade in presence of upstream generated aerodynamic gusts /." Lausanne : EPFL, 2003. http://library.epfl.ch/theses/?display=detail&nr=2758.
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McQuilling, Mark. "EXPERIMENTAL STUDY OF ACTIVE SEPARATION FLOW CONTROL IN A LOW PRESSURE TURBINE BLADE CASCADE MODEL." UKnowledge, 2004. http://uknowledge.uky.edu/gradschool_theses/320.
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The flow field around a low pressure turbine (LPT) blade cascade model with and without flow control is examined using ejector nozzle (EN) and vortex generator jet (VGJ) geometries for separation control. The cascade model consists of 6 Pak-B Pratt andamp; Whitney low pressure turbine blades with Re = 30,000-50,000 at a free-stream turbulence intensity of 0.6%. The EN geometry consists of combined suction and blowing slots near the point of separation. The VGJs consist of a row of holes placed at an angle to the free-stream, and are tested at two locations of 69% and 10.5% of the suction surface length (SSL). Results are compared between flow control on and flow control off states, as well as between the EN, VGJs, and a baseline cascade with no flow control geometry for steady and pulsatile blowing. The EN geometry is shown to control separation with both steady and pulsatile blowing. The VGJs at 69% SSL are shown to be much more aggressive than the EN geometry, achieving the same level of separation control with lower energy input. Pulsed VGJs (PVGJ) have been shown to be just as effective as steady VGJs, and results show that a 10% duty cycle is almost as effective as a 50% duty cycle. The VGJs at 10.5% SSL are shown to be inefficient at controlling separation. No combination of duty cycle and pulsing frequency tested can eliminate the separation region, with only higher steady blowing rates achieving separation control. Thus, the VGJs at 69% SSL are shown to be the most effective in controlling separation.
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McGlumphy, Jonathan. "Numerical Investigation of Subsonic Axial-Flow Tandem Airfoils for a Core Compressor Rotor." Diss., Virginia Tech, 2008. http://hdl.handle.net/10919/26039.
Full textAbstract:
The tandem airfoil has potential to do more work as a compressor blade than a single airfoil without incurring significantly higher losses. Although tandem blades are sometimes employed as stators, they have not been used in any known commercial rotors. The goal of this work is to evaluate the aerodynamic feasibility of using a tandem rotor in the rear stages of a core compressor. As such, the results are constrained to shock-free, fully turbulent flow. The work is divided into 2-D and 3-D simulations. The 3-D results are subject to an additional constraint: thick endwall boundary layers at the inlet.
Existing literature data on tandem airfoils in 2-D rectilinear cascades have been compiled and presented in a Lieblein loss versus loading correlation. Large scatter in the data gave motivation to conduct an extensive 2-D CFD study evaluating the overall performance as a function of the relative positions of the forward and aft airfoils. CFD results were consistent with trends in the open literature, both of which indicate that a properly designed tandem airfoil can outperform a comparable single airfoil on- and off-design. The general agreement of the CFD and literature data serves as a validation for the computational approach.
A high hub-to-tip ratio 3-D blade geometry was developed based upon the best-case tandem airfoil configuration from the 2-D study. The 3-D tandem rotor was simulated in isolation in order to scrutinize the fluid mechanisms of the rotor, which had not previously been well documented. A geometrically similar single blade rotor was also simulated under the same conditions for a baseline comparison. The tandem rotor was found to outperform its single blade counterpart by attaining a higher work coefficient, polytropic efficiency and numerical stall margin. An examination of the tandem rotor fluid mechanics revealed that the forward blade acts in a similar manner to a conventional rotor. The aft blade is strongly dependent upon the flow it receives from the forward blade, and tends to be more three-dimensional and non-uniform than the forward blade.Ph. D.
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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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Whitehouse, David Richard Carleton University Dissertation Engineering Aerospace. "The effect of axial velocity ratio, turbulence intensity, incidence and leading edge geometry on the off-design performance of a turbine blade cascade." Ottawa, 1993.
Find full textBooks on the topic "Nozzle cascade of axial turbine"
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Ristić, D. Three dimensional viscous flow field in an axial flow turbine nozzle passage. [Washington, D.C.]: National Aeronautics and Administration, Office of Management, Scientific and Technical Information Program, 1997.
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Mamat, Zainul Asri. The performance of a cascade of nozzle turbine blading in nucleating steam. Birmingham: University of Birmingham, 1996.
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B, Lakshminarayana, and United States. National Aeronautics and Space Administration. Scientific and Technical Information Program., eds. Three dimensional viscous flow field in an axial flow turbine nozzle passage. [Washington, D.C.]: National Aeronautics and Space Administration, Office of Management, Scientific and Technical Information Program, 1997.
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W, Giel P., and United States. National Aeronautics and Space Administration., eds. Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section. [Washington, D.C: National Aeronautics and Space Administration, 1996.
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Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section. [Washington, D.C: National Aeronautics and Space Administration, 1996.
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Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section. [Washington, D.C: National Aeronautics and Space Administration, 1996.
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W, Giel P., and United States. National Aeronautics and Space Administration., eds. Three-dimensional Navier-Stokes analysis and redesign of an imbedded bellmouth nozzle in a turbine cascade inlet section. [Washington, D.C: National Aeronautics and Space Administration, 1996.
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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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Escudier, Marcel. Engineering applications of the linear momentum equation. Oxford University Press, 2018. http://dx.doi.org/10.1093/oso/9780198719878.003.0010.
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In this chapter a method is shown for applying the linear momentum equation, together with the continuity equation and either Bernoulli’s equation or some other information about static pressure, to the analysis of a diverse range of practical problems. A key aim is to demonstrate that it is possible to establish a relatively simple theoretical basis which can give quite accurate and useful information about the performance of such complex machines as jet and rocket engines, the jet pump, and the Pelton turbine. Other examples include flow through a sudden enlargement, a convergent nozzle, a pipe bend, a pipe junction, and a cascade of guidevanes. For each example it is shown how to define a suitable control volume.
Conference papers on the topic "Nozzle cascade of axial turbine"
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Putra, Mohammed Alexin, and Franz Joos. "Investigation of Secondary Flow Behavior in a Radial Turbine Nozzle." In ASME Turbo Expo 2006: Power for Land, Sea, and Air. ASMEDC, 2006. http://dx.doi.org/10.1115/gt2006-90019.
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Fundamental investigation of secondary flow phenomena in a radial turbine nozzle are presented. L2F measurements have been used for validation of numerical CFD calculations. Having a good agreement by using the Reynolds stress turbulence model (RSM) the numerical results have been used further to analyse the structure of secondary vortices. Contour plots of the flow angle with typical isoline pattern as well as the vorticity have been evaluated. It is shown that the channel of the radial nozzle similar secondary vorticity systems generates as known from the axial turbine nozzles. The formation and the development of the horse-shoe vortex and the corner vortex are discussed. The well known passage vortex of the axial turbines could not been found because of the small curvature of the streamlines. Instead of these an additional single vortex can be observed, called the “inflow” vortex caused by the unsymmetrical flow into the radial cascade from the upstream scroll.
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Madasseri Payyappalli, Manas, and S. R. Shine. "Numerical Investigation on Tandem Compressor Cascades." In ASME 2015 Gas Turbine India Conference. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gtindia2015-1311.
Full textAbstract:
Tandem blade arrangement of axial compressors has been proposed to obtain high loading and turning compared to a single blade. The objectives of the current study is to investigate the effect of percent pitch, axial overlap and incidence angle for a low speed axial compressor stator cascade and to supplement the results with the flow structures observed. 2-D numerical study was conducted using a finite volume scheme which solves the RANS equations along with the Spalart-Allmaras turbulence model. Comparison offlow structures corresponding to different percent pitch, axial overlap and incidence angle has been made to highlight all prominent flow mechanisms. It is observed that the flow through the gap nozzle between the two blades has significant effects on losses. The incidence range of the tandem cascades is found to be superior to the corresponding single blade cases.
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Morphis, G., and J. P. Bindon. "The Flow in a Second Stage Nozzle of a Low Speed Axial Turbine and its Effect on Tip Clearance Loss Development." In ASME 1994 International Gas Turbine and Aeroengine Congress and Exposition. American Society of Mechanical Engineers, 1994. http://dx.doi.org/10.1115/94-gt-145.
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The flow field in a one and a half stage low speed axial turbine with varying levels of rotor tip clearance was measured in order to compare the behaviour of the second nozzle with the first and to identify the manner in which second nozzle responds to the complex tip clearance dependent flow presented to it and completes the formation of tip clearance loss. The tangentially averaged flow relative to the rotor blade in the tip clearance region was found to differ radically from that found in cascade and is not underturned with a high axial velocity. There is evidence rather of overturning caused by secondary flow. The axial velocity follows an almost normal endwall boundary layer pattern with almost no leakage jet effect. The cascade tip clearance model is therefore not accurate. The reduction in second stage nozzle loss was shown to occur near the hub and tip which confirms that it is probably a reduction in secondary flow loss. The nozzle exit loss contours showed that leakage suppressed the formation of the classical secondary flow pattern and that a new tip clearance related loss phenomena exists on the suction surface. The second stage nozzle reduced hub endwall boundary layer below that of both the first nozzle and that behind the rotor. It also rectified secondary and tip clearance flows to such a degree that a second stage rotor would experience no greater flow distortion than the first stage rotor. Radial flow angles behind the second stage nozzle were much smaller than found in a previous study with low aspect ratio un-twisted blades.
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Song, Bo, Wing F. Ng, Joseph A. Cotroneo, Douglas C. Hofer, and Gunnar Siden. "Aerodynamic Design and Testing of Three Low Solidity Steam Turbine Nozzle Cascades." In ASME Turbo Expo 2004: Power for Land, Sea, and Air. ASMEDC, 2004. http://dx.doi.org/10.1115/gt2004-53329.
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Three sets of low solidity steam turbine nozzle cascades were designed and tested. The objective was to reduce cost through a reduction in parts count while maintaining or improving performance. The primary application is for steam turbine high pressure sections where Mach numbers are subsonic and high levels of unguided turning can be tolerated. The baseline Design A has a ratio of pitch to axial chord of 1.2. This is the pitch diameter section of a 50% reaction stage that has been verified by multistage testing on steam to have a high level of efficiency. Designs B and C have ratios of pitch to axial chord of 1.5 and 1.8 respectively. All three designs satisfy the same inlet and exit vector diagrams. Analytical surface Mach number distributions and boundary layer transition predictions are presented. Extensive cascade test measurements were carried out for a broad incidence range from −60 to +35 degrees. At each incidence, four outlet Mach numbers were tested, ranging from 0.2 to 0.8, with the corresponding Reynolds number variation from 1.8×105 to 9.0×105. Experimental results of loss coefficient and blade surface Mach number are presented and compared for the three cascades. The experimental results have demonstrated low losses over the tested Mach number range for a wide range of incidence from −45 to 15 degrees. Designs B and C have lower profile losses than Design A. The associated flow physics is interpreted using the results of wake profile, blade surface Mach number distribution and blade surface oil flow visualization, with the emphasis placed on the loss mechanisms for different flow conditions and the loss reduction mechanism with lower solidity. The effect of the higher profile loading of the lower solidity designs on increased end wall losses induced by increased secondary flow, especially on low aspect ratio designs, is the subject of ongoing studies.
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Giller, Lucas, and Heinz-Peter Schiffer. "Interactions Between the Combustor Swirl and the High Pressure Stator of a Turbine." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-69157.
Full textAbstract:
The interaction between the strongly swirling combustor outflow and the high pressure turbine nozzle guide vanes were investigated at the cascade test rig at Technische Universität Darmstadt. The test section of the rig consists of six swirl generators and five cascade vanes. The three middle vanes are equipped with film cooling holes at the leading edges. The swirler nozzles are aligned with the center of the cascade passages. The operating settings are defined by the swirl number, the distance between the swirler nozzles and the vanes, the blowing ratio and the radial angle of the film cooling holes. Flow field measurements using PIV downstream of the swirlers and five hole probe measurements at the inlet and outlet plane of the cascade were accomplished. Measurements using the ammonia diazo technique to determine the adiabatic film cooling effectiveness on the surface of the center cascade vane were also carried out. It is shown that a swirling inflow leads to a strong alteration of the flow field and the losses in the passages in comparison to an axial inflow. Furthermore, the impact of the swirl on the formation of the cooling film and it’s adiabatic film cooling effectiveness is presented.
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El-Gabry, Lamyaa A., Ranjan Saha, Jens Fridh, and Torsten Fransson. "Measurements of Hub Flow Interaction on Film Cooled Nozzle Guide Vane in Transonic Annular Cascade." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-68088.
Full textAbstract:
An experimental study has been performed in a transonic annular sector cascade of nozzle guide vanes to investigate the aerodynamic performance and the interaction between hub film cooling and mainstream flow. The focus of the study is on the endwalls, specifically the interaction between the hub film cooling and the mainstream. Carbon dioxide (CO2) has been supplied to the coolant holes to serve as tracer gas. Measurements of CO2 concentration downstream of the vane trailing edge can be used to visualize the mixing of the coolant flow with the mainstream. Flow field measurements are performed in the downstream plane with a 5-hole probe to characterize the aerodynamics in the vane. Results are presented for the fully cooled and partially cooled vane (only hub cooling) configurations. Data presented at the downstream plane include concentration contour, axial vorticity, velocity vectors, and yaw and pitch angles. From these investigations, secondary flow structures such as the horseshoe vortex, passage vortex, can be identified and show the cooling flow significantly impacts the secondary flow and downstream flow field. The results suggest that there is a region on the pressure side of the vane trailing edge where the coolant concentrations are very low suggesting that the cooling air introduced at the platform upstream of the leading edge does not reach the pressure side endwall, potentially creating a local hotspot.
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Liu, Kevin, Hongzhou Xu, and Michael Fox. "Turbine Nozzle Endwall Phantom Cooling With Compound Angled Pressure Side Injection." In ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2018. http://dx.doi.org/10.1115/gt2018-75881.
Full textAbstract:
Cooling of the turbine nozzle endwall is challenging due to its complex flow field involving strong secondary flows. Increasingly-effective cooling schemes are required to meet the higher turbine inlet temperatures required by today’s gas turbine applications. Therefore, in order to cool the endwall surface near the pressure side of the airfoil and the trailing edge extended area, the spent cooling air from the airfoil film cooling and pressure side discharge slots, referred to as “phantom cooling” is utilized. This paper studies the effect of compound angled pressure side injection on nozzle endwall surface. The measurements were conducted in a high speed linear cascade, which consists of three nozzle vanes and four flow passages. Two nozzle test models with a similar film cooling design were investigated, one with an axial pressure side film cooling row and trailing edge slots; the other with the same cooling features but with compound angled injection, aiming at the test endwall. Phantom cooling effectiveness on the endwall was measured using a Pressure Sensitive Paint (PSP) technique through the mass transfer analogy. Two-dimensional phantom cooling effectiveness distributions on the endwall surface are presented for four MFR (Mass Flow Ratio) values in each test case. Then the phantom cooling effectiveness distributions are pitchwise-averaged along the axial direction and comparisons were made to show the effect of the compound angled injection. The results indicated that the endwall phantom cooling effectiveness increases with the MFR significantly. A compound angle of the pressure side slots also enhanced the endwall phantom cooling significantly. For combined injections, the phantom cooling effectiveness is much higher than the pressure side slots injection only in the endwall downstream extended area.
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Harasgama, S. P., and E. T. Wedlake. "Heat Transfer and Aerodynamics of a High Rim Speed Turbine Nozzle Guide Vane Tested in the RAE Isentropic Light Piston Cascade (ILPC)." 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-041.
Full textAbstract:
Detailed heat transfer and aerodynamic measurements have been made on an annular cascade of highly loaded nozzle guide vanes. The tests were carried out in an Isentropic Light Piston test facility at engine representative Reynolds number, Mach number and gas-to-wall temperature ratio. The aerodynamics indicate that the vane has a weak shock at 65–70% axial chord (mid span) with a peak Mach number of 1.14. The influence of Reynolds number and Mach number on the Nusselt number distributions on the vane and endwall surfaces are shown to be significant. Computational techniques are used for the interpretation of test data.
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Martinez-Botas, R. F., A. J. Main, G. D. Lock, and T. V. Jones. "A Cold Heat Transfer Tunnel for Gas Turbine Research on an Annular Cascade." 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-248.
Full textAbstract:
The Oxford University Blowdown Tunnel has been substantially modified to test a large annular cascade of high pressure nozzle guide vanes (mean blade diameter of 1.11 m and axial chord of 0.0673 m). The new transonic facility has been constructed to obtain complete contours of heat transfer coefficient for both the end walls and blade surfaces using the transient liquid crystal technique, to measure pressure distributions and losses, and to study fundamental aspects of boundary layers and secondary flows. The facility allows an independent variation of Reynolds and Mach numbers, providing aerodynamic and heat transfer measurements in the region of interest for gas turbine design. The mass flow rate through the cascade at NGV design conditions (exit Mach number 0.96 and Reynolds number 2.0 × 106) is 38 kg/s and the pressure-regulated test duration exceeds 7 seconds.
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Aziz, Imran, Imran Akhtar, Usama Bin Perwez, and Auwais Ahmed. "Three Dimensional Flow Investigation in One and a Half Stage Axial Turbine." In ASME 2016 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2016. http://dx.doi.org/10.1115/imece2016-66703.
Full textAbstract:
In this study, three dimensional flow analysis of one and a half stage axial turbine is investigated. The objective of this study is to analyse the effect of rotor stator interaction and the resulting unsteadiness. This includes the effect of first row of Nozzle guide vane (NGV) wakes on rotor blades, secondary vortical flow prediction, influence of rotor wakes on the flow pattern of second stator, appreciation and application of techniques to model the exact blade counts across the rotor-stator interfaces. We employ a three-dimensional finite-volume based solver to simulate the flow in the turbine using SST model to account for turbulence effects. Sliding mesh technique is used to allow the transfer of flow parameters across the sliding rotor/stator interfaces. In order to model a single passage configuration, profile transformation and time transformation method is used. The flow physics for the visualization and understanding of flow behavior in a 3D turbine cascade is explained in detail and validated with the previous experimental and numerical studies. The study provides application of computationally efficient methods for simulating the fluid flow in a turbine which contain unequal number of rotor and stator blades.