Статті в журналах: "Nozzle cascade of axial turbine"

1

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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2

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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3

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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4

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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5

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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6

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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7

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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8

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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9

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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10

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.

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11

Harasgama, S. P., and C. D. Burton. "Film Cooling Research on the Endwall of a Turbine Nozzle Guide Vane in a Short Duration Annular Cascade: Part 2—Analysis and Correlation of Results." Journal of Turbomachinery 114, no. 4 (October 1, 1992): 741–46. http://dx.doi.org/10.1115/1.2928027.

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Results have been presented on the heat transfer characteristics of the film cooled endwall (platform) of a turbine nozzle guide vane in an annular cascade at engine representative conditions in a companion paper by Harasgama and Burton (1992). The present paper reports on the analysis of these measurements. The experimental results are well represented by the superposition theory of film cooling. It is shown that high cooling effectiveness can be achieved when the data are corrected for axial pressure gradients. The data are correlated against both the slot-wall jet parameter and the discrete hole injection function for flat-plate, zero pressure gradient cases. The pressure gradient correction brings the present data to within ± 11 percent of the discrete hole correlation. Preliminary predictions of heat transfer reduction have been carried out using the STANCOOL program. These indicate that the code can predict the magnitude of heat transfer reduction correctly, although the absolute values are not in good agreement. This is attributed to the three-dimensional nature of the flow at the endwall.

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12

Gribin, V. G., A. A. Tishchenko, R. A. Alekseev, V. A. Tishchenko, I. Yu Gavrilov, and V. V. Popov. "Application of the Parametric Method for Profiling the Interblade Channels in the Nozzle Cascades of Axial-Flow Turbine Machines." Thermal Engineering 67, no. 8 (July 31, 2020): 536–42. http://dx.doi.org/10.1134/s0040601520080029.

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13

Tiedemann, Maik, and Friedrich Kost. "Some Aspects of Wake-Wake Interactions Regarding Turbine Stator Clocking." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 526–33. http://dx.doi.org/10.1115/1.1370158.

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This investigation is aimed at an experimental determination of the unsteady flowfield downstream of a transonic high pressure turbine stage. The single stage measurements, which were part of a joined European project, were conducted in the windtunnel for rotating cascades of the DLR Go¨ttingen. Laser-2-focus (L2F) measurements were carried out in order to determine the Mach number, flow angle, and turbulence distributions. Furthermore, a fast response pitot probe was utilized to determine the total pressure distribution. The measurement position for both systems was 0.5 axial rotor chord downstream of the rotor trailing edge at midspan. While the measurement position remained fixed, the nozzle guide vane (NGV) was “clocked” to 12 positions covering one NGV pitch. The periodic fluctuations of the total pressure downstream of the turbine stage indicate that the NGV wake damps the total pressure fluctuations caused by the rotor wakes. Furthermore, the random fluctuations are significantly lower in the NGV wake affected region. Similar conclusions were drawn from the L2F turbulence data. Since the location of the interaction between NGV wake and rotor wake is determined by the NGV position, the described effects are potential causes for the benefits of “stator clocking” which have been observed by many researchers.

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14

Leibsohn Martins, Guilherme. "Axial Turbine Cascade Correlation." Applied Sciences 6, no. 12 (December 10, 2016): 420. http://dx.doi.org/10.3390/app6120420.

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15

Moustapha, S. H., W. E. Carscallen, and J. D. McGeachy. "Aerodynamic Performance of a Transonic Low Aspect Ratio Turbine Nozzle." Journal of Turbomachinery 115, no. 3 (July 1, 1993): 400–408. http://dx.doi.org/10.1115/1.2929267.

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This paper presents detailed information on the three-dimensional flow field in a realistic turbine nozzle with an aspect ratio of 0.65 and a turning angle of 76 deg. The nozzle has been tested in a large-scale planar cascade over a range of exit Mach numbers from 0.3 to 1.3. The experimental results are presented in the form of nozzle passage Mach number distributions and spanwise distribution of losses and exit flow angle. Details of the flow field inside the nozzle passage are examined by means of surface flow visualization and Schlieren pictures. The performance of the nozzle is compared to the data obtained for the same nozzle tested in an annular cascade and a stage environment. Excellent agreement is found between the measured pressure distribution and the prediction of a three-dimensional Euler flow solver.

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16

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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17

Tsuchiya, T., Y. Furuse, S. Yoshino, R. Chikami, Y. Tsukuda, and M. Mori. "Development of Air-Cooled Ceramic Nozzles for a Power-Generating Gas Turbine." Journal of Engineering for Gas Turbines and Power 118, no. 4 (October 1, 1996): 717–23. http://dx.doi.org/10.1115/1.2816986.

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The development of air-cooled ceramic nozzle vanes for a power-generating gas turbine has been reported. To make up the limited temperature resistance of present ceramic materials, the utilization of a small amount of cooling air has been studied for the first-stage nozzle vanes of a 1500°C class gas turbine. A series of cascade tests were carried out for the designed air-cooled Si3N4 nozzle vanes under 6 atm and 1500°C conditions. It was confirmed that the maximum ceramic temperature can be maintained below 1300°C by a small amount of cooling air. In spite of the increased thermal stresses from local cooling, all Si3N4 nozzle vanes survived the cascade tests, including both steady-state and transients of emergency shutdown. The potential for an air-cooled ceramic nozzle was demonstrated for a 1500°C class gas turbine application.

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18

Luxa, Martin, Rudolf Dvorak, David Simurda, and Jan Vimmr. "Pneumatic measurements downstream of a radial turbine nozzle cascade." Journal of Thermal Science 19, no. 1 (January 29, 2010): 42–46. http://dx.doi.org/10.1007/s11630-010-0042-4.

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19

Williamson, R. G., and S. H. Moustapha. "Annular Cascade Testing of Turbine Nozzles at High Exit Mach Numbers." Journal of Fluids Engineering 108, no. 3 (September 1, 1986): 313–20. http://dx.doi.org/10.1115/1.3242579.

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This paper presents detailed information on the three-dimensional flow field in a realistic low aspect ratio, high turning nozzle vane design which incorporates end-wall contouring and which has been tested over a range of exit Mach number from subsonic up to the design value at mean section of 1.15. The experimental results, in the form of nozzle surface pressure distributions as well as surveys of pressure losses and flow angles at exit, are compared with those calculated by a three-dimensional flow analysis. The effects of exit Mach number on the measured nozzle performance are also presented.

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20

Gostelow, J. P., A. Mahallati, W. E. Carscallen, and A. Rona. "Encounters with Vortices in a Turbine Nozzle Passage." International Journal of Rotating Machinery 2012 (2012): 1–10. http://dx.doi.org/10.1155/2012/928623.

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Experiments were conducted on the flow through a transonic turbine cascade. Secondary flows and a wide range of vortex types were encountered, including horseshoe vortices, shock-induced passage vortices, and streamwise vortices on the suction surface. In the separation region on the suction surface, a large rollup of passage vorticity occurred. The blunt leading edge gave rise to strong horseshoe vortices and secondary flows. The suction surface had a strong convex curvature over the forward portion and was quite flat further downstream. Surface flow visualization was performed and this convex surface displayed coherent streamwise vorticity. At subsonic speeds, strong von Kármán vortex shedding resulted in a substantial base pressure deficit. For these conditions, time-resolved measurements were made of the Eckert-Weise energy separation in the blade wake. At transonic speeds, exotic shedding modes were observed. These phenomena all occurred in experiments on the flow around one particular turbine nozzle vane in a linear cascade.

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21

Qi, Wenjiao, Qinghua Deng, Yu Jiang, Zhenping Feng, and Qi Yuan. "Aerodynamic performance and flow characteristics analysis of Tesla turbines with different nozzle and outlet geometries." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 233, no. 3 (July 3, 2018): 358–78. http://dx.doi.org/10.1177/0957650918785312.

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The aerodynamic performance and flow characteristics of a multichannel nozzled Tesla turbine were investigated numerically with different nozzle and outlet geometries at different rotational speeds. Two kinds of nozzle geometries were proposed: one nozzle channel to one disc channel (named as one-to-one turbine) and one nozzle channel to several disc channels (named as one-to-many turbine). Simplified radial outlet and real axial outlet geometries of the Tesla turbines were adopted to research the influence of outlet geometries. The results show that compared with the one-to-many turbine, the isentropic efficiency of the one-to-one turbine is much higher; while the flow coefficient is much lower. In addition, in the middle disc channels (DC1 and DC2) of which two walls are rotating disc walls, the flow fields are almost the same, but different from that in the side channel (DC3) of which one wall is a rotating wall and the other one is a stationary casing wall. DC1 and DC2 generate more torque with less working fluid, thus the disc spacing distance of DC3 should be narrower than that of DC1 and DC2. Compared to the one-to-many turbine, the working fluid flowing through DC1 and DC2 of the one-to-one turbine is much less, and the flow path lines are much longer. The results of different turbine outlet geometries show that compared with the turbines with radial outlet, the isentropic efficiency of the one-to-many turbine with axial outlet is a little higher, while that of the one-to-one turbine with axial outlet is lower. This is due to the larger torque on the disc hole walls, despite a lot more total pressure loss in the exhaust vent of the one-to-many turbine. Therefore, the contribution of disc hole walls to torque cannot be neglected in numerical simulations.

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22

Zhu, Xiaofeng, Zhirong Lin, Xin Yuan, Tomohiro Tejima, Yoshiki Niizeki, and Naoki Shibukawa. "Non-equilibrium Condensing Flow Modeling in Nozzle and Turbine Cascade." International Journal of Gas Turbine, Propulsion and Power Systems 4, no. 3 (2012): 9–16. http://dx.doi.org/10.38036/jgpp.4.3_9.

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23

Klimko, Marek, Richard Lenhard, Pavel Žitek, and Katarína Kaduchová. "Experimental Evaluation of Axial Reaction Turbine Stage Bucket Losses." Processes 9, no. 10 (October 13, 2021): 1816. http://dx.doi.org/10.3390/pr9101816.

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The article describes the measurement methods and data evaluation from a single-stage axial turbine with high reaction (50%). Four operating modes of the turbine were selected, in which the wake traversing behind nozzle and bucket with five-hole pneumatic probes took place. The article further focuses on the evaluation of bucket losses for all four measured operating modes, including the analysis of measurement uncertainties.

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24

Klimko, Marek, Richard Lenhard, Pavel Žitek, and Katarína Kaduchová. "Experimental Evaluation of Axial Reaction Turbine Stage Bucket Losses." Processes 9, no. 10 (October 13, 2021): 1816. http://dx.doi.org/10.3390/pr9101816.

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Анотація:

The article describes the measurement methods and data evaluation from a single-stage axial turbine with high reaction (50%). Four operating modes of the turbine were selected, in which the wake traversing behind nozzle and bucket with five-hole pneumatic probes took place. The article further focuses on the evaluation of bucket losses for all four measured operating modes, including the analysis of measurement uncertainties.

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25

Wedlake, E. T., A. J. Brooks, and S. P. Harasgama. "Aerodynamic and Heat Transfer Measurements on a Transonic Nozzle Guide Vane." Journal of Turbomachinery 111, no. 1 (January 1, 1989): 36–42. http://dx.doi.org/10.1115/1.3262234.

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Experimental determination of heat transfer rates to gas turbine blading plays an important part in the improvement of both the validation of existing design methods and the development of improved design codes. This paper describes a series of tests on an annular cascade of nozzle guide vanes designed for a high-work-capacity single-stage transonic turbine. The tests were carried out in the Isentropic Light Piston Cascade at the Royal Aerospace Establishment, Pyestock, and a brief description of this new test facility is included. Measurements of local heat transfer rates and aerodynamic data around the blade surface and on the end walls are described.

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Griffin, Lisa W., and Daniel J. Dorney. "Simulations of the Unsteady Flow Through the Fastrac Supersonic Turbine." Journal of Turbomachinery 122, no. 2 (February 1, 1999): 225–33. http://dx.doi.org/10.1115/1.555453.

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Analysis of the unsteady aerodynamic environment in the Fastrac supersonic turbine is presented. Modal analysis of the turbine blades indicated possible resonance in crucial operating ranges of the turbopump. Unsteady computational fluid dynamics (CFD) analysis was conducted to support the aerodynamic and structural dynamic assessments of the turbine. Before beginning the analysis, two major problems with current unsteady analytical capabilities had to be addressed: modeling a straight centerline nozzle with the turbine blades and exit guide vanes (EGVs), and reducing run times significantly while maintaining physical accuracy. Modifications were made to the CFD code used in this study to allow the coupled nozzle/blade/EGV analysis and to incorporate Message Passing Interface (MPI) software. Because unsteadiness is a key issue for the Fastrac turbine [and future rocket engine turbines such as the Reusable Launch Vehicle (RLV)], calculations were performed for two nozzle-to-blade axial gaps. Calculations were also performed for the nozzle alone, and the results were imposed as an inlet boundary condition for a blade/EGV calculation for the large gap case. These results are compared to the nozzle/blade/EGV results. [S0889-504X(00)02902-0]

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27

Mazur, Zdzislaw, Rafael Campos‐Amezcua, and Alfonso Campos‐Amezcua. "Shape modification of an axial flow turbine nozzle to reduce erosion." International Journal of Numerical Methods for Heat & Fluid Flow 19, no. 2 (March 27, 2009): 242–58. http://dx.doi.org/10.1108/09615530910930991.

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28

Aghaei tog, Reza, and Abolghasem Mesgarpoor Tousi. "Flow pattern improvement in nozzle-rotor axial gap in impulse turbine." Aircraft Engineering and Aerospace Technology 86, no. 2 (February 25, 2014): 108–16. http://dx.doi.org/10.1108/aeat-09-2012-0146.

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29

Matsunuma, Takayuki. "Unsteady Flow Field of an Axial-Flow Turbine Rotor at a Low Reynolds Number." Journal of Turbomachinery 129, no. 2 (July 18, 2006): 360–71. http://dx.doi.org/10.1115/1.2464143.

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The unsteady flow field of an annular turbine rotor was investigated experimentally using a laser Doppler velocimetry (LDV) system. Detailed measurements of the time-averaged and time-resolved distributions of the velocity, flow angle, turbulence intensity, etc., were carried out at a very low Reynolds number condition, Reout=3.5×104. The data obtained were analyzed from the viewpoints of both an absolute (stationary) frame of reference and a relative (rotating) frame of reference. The effect of the turbine nozzle wake and secondary vortices on the flow field inside the rotor passage was clearly captured. It was found that the nozzle wake and secondary vortices are suddenly distorted at the rotor inlet, because of the rotating potential field of the rotor. The nozzle flow (wake and passage vortices) and the rotor flow (boundary layer, wake, tip leakage vortex, and passage vortices) interact intensively inside the rotor passage.

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30

Joseph, Joby, Sathyanarayanan Subramanian, K. Vigney, B. V. S. S. S. Prasad, and D. Biswas. "Thermodynamic wetness loss calculation in nozzle and turbine cascade: nucleating steam flow." Heat and Mass Transfer 54, no. 8 (November 28, 2017): 2521–31. http://dx.doi.org/10.1007/s00231-017-2171-8.

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31

Stareček, Jakub, Pavel Čupr, and Miloslav Haluza. "Design of a high-specific speed turbine with non-uniform blade cascade." EPJ Web of Conferences 213 (2019): 02078. http://dx.doi.org/10.1051/epjconf/201921302078.

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This paper presents the hydraulic design of Kaplan type turbine in fully axial configuration. The turbine model consists of a straight pipe intake, a rib which covers the shaft, five axial guide vanes, a six blade runner and a 10,6°-full angle draft tube. Due to the presence of non-uniform velocity field behind the rib and interaction between stator and rotor, the pressure pulsations occur and cause some excitation forces acting on the runner. To reduce these negative phenomena the six non-uniform blade runner has been designed using computational fluid dynamics modelling (CFD). The runner blades are placed in various axial distances from the guide vanes. This new design should reduce the pressure pulsations caused by interaction between the rib and the runner, which is the main negative effect by the current design. The forces and torque which are acting on the runner blades as well as the pressure inside the turbine are observed in time domain. The amplitudes and frequencies are analysed using Fast Fourier Transform (FFT). The frequencies are compared to the modal analysis results of the runner submerged in water. This hydraulic design is based on previous axial turbine, which was designed for small hydro power plant located at Lužnice River. The design point is defined by head H = 1,5 m, volumetric flow rate Q = 2,1 m3/s and RPM n = 260 min-1. This turbine was designed using six runner blades to reduce the pressure, force and torque pulsations caused mainly by wakes, which were observed behind the rib.

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32

Karakoc, T. Hikmet, and Onder Turan. "Exergetic Destruction Effects of Operating Conditions on the Turbojet Engine Components." Applied Mechanics and Materials 110-116 (October 2011): 2390–94. http://dx.doi.org/10.4028/www.scientific.net/amm.110-116.2390.

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The minimization of exergy destruction brings the design as closely as permissible to the theoretical limit. This study presents exergy destruction analysis of a turbojet engine for different flight Mach number and altitudes. Turbojet engine being considered consists of an inlet, a centrifugal compressor, reverse flow combustion chamber, axial-flow turbine and exhaust nozzle. The flight Mach number and altitude are examined on the exergetic destructions of compressor, combustion chamber, turbine and exhaust nozzle. The results of component-based destruction analysis are given as three dimensional exergetic-destruction response surface plots related to altitude and flight Mach number.

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33

Pullan, Graham, and Neil W. Harvey. "Influence of Sweep on Axial Flow Turbine Aerodynamics at Midspan." Journal of Turbomachinery 129, no. 3 (July 14, 2006): 591–98. http://dx.doi.org/10.1115/1.2472397.

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Sweep, when the stacking axis of the blade is not perpendicular to the axisymmetric streamsurface in the meridional view, is often an unavoidable feature of turbine design. Although a high aspect ratio swept blade can be designed to achieve the same pressure distribution as an unswept design, this paper shows that the swept blade will inevitably have a higher profile loss. A modified Zweifel loading parameter, taking sweep into account, is first derived. If this loading coefficient is held constant, it is shown that sweep reduces the required pitch-to-chord ratio and thus increases the wetted area of the blades. Assuming fully turbulent boundary layers and a constant dissipation coefficient, the effect of sweep on profile loss is then estimated. A combination of increased blade area and a raised pressure surface velocity means that the profile loss rises with increasing sweep. The theory is then validated using experimental results from two linear cascade tests of highly loaded blade profiles of the type found in low-pressure aeroengine turbines: one cascade is unswept, the other has 45deg of sweep. The swept cascade is designed to perform the same duty with the same loading coefficient and pressure distribution as the unswept case. The measurements show that the simple method used to estimate the change in profile loss due to sweep is sufficiently accurate to be a useful aid in turbine design.

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34

Dossena, V., A. Perdichizzi, and M. Savini. "The Influence of Endwall Contouring on the Performance of a Turbine Nozzle Guide Vane." Journal of Turbomachinery 121, no. 2 (April 1, 1999): 200–208. http://dx.doi.org/10.1115/1.2841302.

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The paper presents the results of a detailed investigation of the flow field in a gas turbine linear cascade. A comparison between a contoured and a planar configuration of the same cascade has been performed, and differences in the three-dimensional flow field are here analyzed and discussed. The flow evolution downstream of the trailing edge was surveyed by means of probe traversing while a three-dimensional Navier–Stokes solver was employed to obtain information on flow structures inside the vaned passages. The experimental measurements and the numerical simulation of the three-dimensional flow field have been performed for two cascades; one with planar endwalls, and the other with one planar and one profiled endwall, so as to present a reduction of the nozzle height. The investigation was carried out at an isentropic downstream Mach number of 0.6. Airfoils of both cascades were scaled from the same high-pressure gas turbine inlet guide vane. Measurements of the three-dimensional flow field have been performed on five planes downstream of the cascades by means of a miniaturized five-hole pressure probe. The presence of endwall contouring strongly influences the secondary effects; the vortex generation and their development are inhibited by the stronger acceleration taking place throughout the cascade. The results show that the secondary effects on the contoured side of the passage are confined in the endwall region, while on the flat side the secondary vortices display characteristics similar to the ones occurring downstream of the planar cascade. The spanwise outlet angle distribution presents a linear variation for most of the nozzle height, with quite low values approaching the contoured endwall. The analysis of mass-averaged losses shows a significant performance improvement in the contoured cascade. This can be ascribed not only to lower secondary losses but also to a reduction of the profile losses.

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35

Lapuzin, Alexander, Valery Subotovich, Yuriy Yudin, Svetlana Naumenko, and Ivan Malymon. "Flow Characteristics of the Nozzle Blade Cascade in the Mode of the Joint Operation with the Radial Diffuser." NTU "KhPI" Bulletin: Power and heat engineering processes and equipment, no. 3 (December 30, 2021): 5–11. http://dx.doi.org/10.20998/2078-774x.2021.03.01.

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The obtained research data are given for the nozzle cascade used by a small-size gas turbine of an average fanning in combination with the radial diffuser. Aerodynamic characteristics of the nozzle blade cascade were determined in a wide range of a change in the Reynolds number varying from 4∙105 to 106 and the reduced velocity varying in the range of 0.4 to 1.13. The flow rate coefficient of the nozzle cascade was derived for all modes using the integral methods and the drainages behind the cascade. The kinetic energy loss coefficient and the flow angles were calculated using the measurement data of flow parameters in three control modes that were obtained due to the use of orientable pneumometric probes. When the expansion degree of the convergent –divergent annular duct behind the cascade is equal to 1.43 the flow in the narrow section of this duct is “enlocked” in the mode when the reduced velocity behind the cascade is equal to 1.127. At such velocity the Reynolds number 106 is self-similar for the flow rate coefficient. At lower values of Reynolds number, the decrease of it is accompanied by an intensive decrease in the flow rate coefficient for all the values of the reduced velocity. For the Reynolds number lower than 7∙105 an increase in the velocity results in a decreased flow rate coefficient. When this number exceeds 8∙105 an increase in the velocity results in an increase of the flow coefficient up to the moment when the flow is “enlocked” in the nozzle cascade.

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36

Zhang, Fang Fang, Zhen Shan Zhang, and Rui Zhu. "Optimization Design Study on a New Type Underwater Turbine Engine." Advanced Materials Research 850-851 (December 2013): 292–95. http://dx.doi.org/10.4028/www.scientific.net/amr.850-851.292.

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In order to improve economic performance of a new type underwater turbine engine, objective function with maximizing inner efficiency in design condition is constructed based on establishing design calculation model of single stage impulse turbine engine. By using GA algorithm, optimization match study is finished among four parameters, such as working back pressure of turbine engine, mean diameter of turbine cascade, air-in inclination angle and diffuse angle of nozzle, and whose effect on inner efficiency is studied respectively. Results show that, for the engine studied, inner efficiency has sensitivity to working back pressure, while not to diffuse angle of nozzle. Which has been improved by 6.24% after optimization design in design condition, and working substance consumption per second has been decreased by 5.87% correspondingly, so economic performance of the studied engine has been improved obviously. The established model and its calculation results provide initial model and parameters for thermodynamic simulation for turbine engine in off-design condition.

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37

Matsunuma, Takayuki. "Effects of Reynolds Number and Freestream Turbulence on Turbine Tip Clearance Flow." Journal of Turbomachinery 128, no. 1 (February 1, 2005): 166–77. http://dx.doi.org/10.1115/1.2103091.

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Tip clearance losses represent a major efficiency penalty of turbine blades. This paper describes the effect of tip clearance on the aerodynamic characteristics of an unshrouded axial-flow turbine cascade under very low Reynolds number conditions. The Reynolds number based on the true chord length and exit velocity of the turbine cascade was varied from 4.4×104 to 26.6×104 by changing the velocity of fluid flow. The freestream turbulence intensity was varied between 0.5% and 4.1% by modifying turbulence generation sheet settings. Three-dimensional flow fields at the exit of the turbine cascade were measured both with and without tip clearance using a five-hole pressure probe. Tip leakage flow generated a large high total pressure loss region. Variations in the Reynolds number and freestream turbulence intensity changed the distributions of three-dimensional flow, but had no effect on the mass-averaged tip clearance loss of the turbine cascade.

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38

Zaccaria, M. A., and B. Lakshminarayana. "Unsteady Flow Field Due to Nozzle Wake Interaction With the Rotor in an Axial Flow Turbine: Part I—Rotor Passage Flow Field." Journal of Turbomachinery 119, no. 2 (April 1, 1997): 201–13. http://dx.doi.org/10.1115/1.2841103.

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The flow field in turbine rotor passages is complex with unsteadiness caused by the aerodynamic interaction of the nozzle and rotor flow fields. The two-dimensional steady and unsteady flow field at midspan in an axial flow turbine rotor has been investigated experimentally using an LDV with emphasis on the interaction of the nozzle wake with the rotor flow field. The flow field in the rotor passage is presented in Part I. while the flow field downstream of the rotor is presented in Part II. Measurements were acquired at 37 axial locations from just upstream of the rotor to one chord downstream of the rotor. The time-averaged flow field and the unsteadiness caused by the wake have been captured. As the nozzle wake travels through the rotor flow field, the nozzle wake becomes distorted with the region of the nozzle wake near the rotor suction surface moving faster than the region near the rotor pressure surface, resulting in a highly distorted wake. The wake is found to be spread out along the rotor pressure surface, as it convects downstream of midchord. The magnitude of the nozzle wake velocity defect grows until close to midchord, after which it decreases. High values of unresolved unsteadiness were observed at the rotor leading edge. This is due to the large flow gradients near the leading edge and the interaction of the nozzle wake with the rotor leading edge. High values of unresolved unsteadiness were also observed near the rotor pressure surface. This increase in unresolved unsteadiness is caused by the interaction of the nozzle wake with the flow near the rotor pressure surface.

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39

Gregory-Smith, D. G., C. P. Graves, and J. A. Walsh. "Growth of Secondary Losses and Vorticity in an Axial Turbine Cascade." Journal of Turbomachinery 110, no. 1 (January 1, 1988): 1–8. http://dx.doi.org/10.1115/1.3262163.

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The growth of losses, secondary kinetic energy, and streamwise vorticity have been studied in a high turning rotor cascade. Negative vorticity associated with the passage vortex agreed well with predictions of classical secondary flow theory in the early part of the blade passage. However, toward the exit, the distortion of the flow by the secondary velocities rendered the predictions inaccurate. Areas of positive vorticity were associated with the feeding of loss into the bulk flow and have been related to separation lines observed by surface flow visualization.

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40

Leto, Angelo. "Investigation of a Radial Turbines Compatibility for Small Rocket Engine." E3S Web of Conferences 197 (2020): 11009. http://dx.doi.org/10.1051/e3sconf/202019711009.

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In the radial turbine preliminary design for an expander rocket engine, a comparison was made with axial turbine used in Pratt & Whitney RL10 engine. One of the primary requirements of a liquid propellant rocket engine is the generation of a high thrust, which depends on both the mass flow rate of the propellant and the pressure in the thrust chamber. In expander-cycle engines, which are the subject of the present study, the liquid propellant is first compressed using centrifugal turbo-pumps, then it is used to cool the combustion chamber and the nozzle and, once vaporized, it flows through the turbines used to drive the turbo-pumps. The aim was to demonstrate the greater efficiency of the radial turbine with a reduction of the pressure ratio with respect to the axial turbine.

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41

Su, Gao, G. Y. Zhou, and Fei Du. "Numercial Simulation on Three-Dimensional Unsteady Flow in a Supercharged Boiler Gas Turbine." Applied Mechanics and Materials 271-272 (December 2012): 1039–43. http://dx.doi.org/10.4028/www.scientific.net/amm.271-272.1039.

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To the unsteady characteristic of three-dimensional flow in the gas turbine blade cascades, based on sliding mesh and a standard turbulent flow model, Fluent software was employed to solve the Reynolds averaged N-S equation. The numberical result of unsteady flow field is obtained in gas turbine cascade for supercharged marine boiler. This paper shows the axial distribution of Ma in the position of β=0 in a calculational period time, and the effect of trails to flow field characteristics. The result can provide guidelines for aerodynamic optimization design of gas turbine stage cascade.

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42

Kawai, Tatsuo, Hidetoshi Ohshige, Kazuto Nakai, and Tsutomu Adachi. "Effect of Boundary Layer Fences Attached to Casing of Annular Turbine Nozzle Cascade." Transactions of the Japan Society of Mechanical Engineers Series B 60, no. 569 (1994): 169–75. http://dx.doi.org/10.1299/kikaib.60.169.

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43

Dishart, P. T., and J. Moore. "Tip Leakage Losses in a Linear Turbine Cascade." Journal of Turbomachinery 112, no. 4 (October 1, 1990): 599–608. http://dx.doi.org/10.1115/1.2927700.

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An investigation of tip leakage flow and its effects on loss production was performed on a large-scale linear turbine cascade having a tip gap measuring 2.1 percent of the blade height. The flow exiting the tip gap was measured to determine the losses incurred within the tip gap and the secondary kinetic energy due to tip leakage. Additional measurements, 40 percent of an axial chord downstream of the blade trailing edges, showed the development of the leakage flow and the overall cascade losses. At the downstream location, the additional loss due to tip leakage was found to be the sum of the measured loss at the tip gap exit plane and the amount of tip gap secondary kinetic energy that had been dissipated by that downstream location.

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44

Sznajder, Janusz. "Simulations of Hot-Gas Flow in Internally Cooled Cascade of Turbine Vanes." Journal of KONES 26, no. 2 (June 1, 2019): 151–58. http://dx.doi.org/10.2478/kones-2019-0044.

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Abstract An experiment in cooling of gas turbine nozzle guide vanes was modelled numerically with a conjugate viscous-flow and solid-material heat conduction solver. The nozzle vanes were arranged in a cascade and operated in high-pressure, hot-temperature conditions, typical for first turbine stage in a flow of controlled-intensity, artificially-generated turbulence. The vane cooling was internal, accomplished by 10 channels in each vane with cooling-air flow. Numerical simulations of the experiment were conducted applying two turbulence models of the k-omega family: k-omega-SST and Transition SST implemented in the ANSYS Fluent solver. Boundary conditions for the simulations were set based on conditions of experiment: total pressures and total temperature on inlet to cascade, static pressure on the outlet of the cascade and heat flux on the surface of cooling channels. The values of heat flux on the surface of cooling channels were evaluated based on Nusselt numbers obtained from experiment and varied in time until steady-state conditions were obtained. Two test cases, one with subcritical outlet flow, and another one, with supercritical outlet flow were simulated. The result of experiment – distributions of pressure, surface temperature, and heat transfer coefficients on the vane external surface were compared to results of numerical simulations. Sensitivity of the vane surface temperatures and heat transfer coefficients to turbulence models and to boundary-condition values of parameters of turbulence models: turbulence energy and specific dissipation of turbulence energy was also studied.

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45

Murawski, Christopher G., and Kambiz Vafai. "Effect of Wake Disturbance Frequency on the Secondary Flow Vortex Structure in a Turbine Blade Cascade." Journal of Fluids Engineering 122, no. 3 (March 27, 2000): 606–13. http://dx.doi.org/10.1115/1.1287792.

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An experimental study of the effect of wake disturbance frequency on the secondary flow vortices in a two-dimensional linear cascade is presented. The flow Reynolds numbers, based on exit velocity and suction side surface length were 25,000, 50,000 and 85,000. Secondary flow was visualized by injecting smoke into the boundary layer and illuminating it with a laser light sheet located at the exit of the cascade. To simulate wakes from upstream blade rows, a set of spanwise cylinders were traversed across the front of the blade row. The flow visualization results with a single wake disturbance reveal that the recovery time of the secondary flow vortex structure decreases as the wake traverse velocity is increased. The results of flow visualization with multiple wakes showed that wake disturbance frequencies below the axial chord flow frequency allowed complete recovery of the secondary flow vortex structure before the next wake encounters the blade leading edge. Wake disturbance frequencies that exceeded the axial chord flow frequency resulted in no observable recovery of the secondary flow vortex structure. Axial chord flow frequency is defined as the axial velocity in the cascade divided by the axial chord length of the turbine blade. [S0098-2202(00)02203-3]

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46

Nono Suprayetno, Priyono Sutikno, Nathanael P. Tandian, and Firman Hartono. "Numerical Simulation of Cascade Flow: Vortex Element Method for Inviscid Flow Analysis and Axial Turbine Blade Design." Journal of Advanced Research in Fluid Mechanics and Thermal Sciences 85, no. 2 (August 5, 2021): 14–23. http://dx.doi.org/10.37934/arfmts.85.2.1423.

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This study aims to design an axial turbine rotor blade and predict the turbine performance at preliminary design stage. Quasi three dimensional method was applied to design including blade to blade flow analysis. The blade profile uses a NACA 0015 airfoil by varying the profile thickness from hub to tip. The profile is divided into eleven segments which has different parameters. The profile was analysed using blade to blade flow/cascade flow analysis called vortex panel method to obtain lift coefficient. The analysis of cascade flow was performed in potential flow and prediction of turbine perfomance is carried out involving common best practice to give drag effect on the blade. The design of the turbine was applied on three different rotors, which also have a different discharge, head, and design rotation. The outer diameter of turbine 1 is 0.65 m, while turbine 2 and turbine 3 have an outer diameter of 0,60 m. The calculation result show that the efficiency of turbines 1, 2, and 3 were 88,32%, 89,67%, and 89,04%, respectively.

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47

Addison, J. S., and H. P. Hodson. "Unsteady Transition in an Axial-Flow Turbine: Part 1—Measurements on the Turbine Rotor." Journal of Turbomachinery 112, no. 2 (April 1, 1990): 206–14. http://dx.doi.org/10.1115/1.2927634.

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Previously published measurements in a low-speed, single-stage, axial-flow turbine have been reanalyzed in the light of more recent understanding. The measurements include time-resolved hot-wire traverses and surface hot film gage measurements at the midspan of the rotor suction surface with three different rotor-stator spacings. Part 1 investigates the suction surface boundary layer transition process, using surface-distance time plots and boundary layer cross sections to demonstrate the unsteady and two-dimensional nature of the process. Part 2 of the paper will describe the results of supporting experiments carried out in a linear cascade together with a simple transition model, which explains the features seen in the turbine.

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48

Wang, S.-S., J.-R. Mao, G.-W. Liu, and Z.-P. Feng. "Reduction of Solid-Particle Erosion on the Control-Stage Nozzle of a Steam Turbine through Improved End-Wall Contouring." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 224, no. 10 (April 22, 2010): 2199–210. http://dx.doi.org/10.1243/09544062jmes1987.

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The iron oxide scales exfoliated from the inner wall of a boiler tube and a main steam pipe is known to cause solid-particle erosion on the control-stage nozzle. A combined experimental and numerical investigation was conducted to explore the optimization method of end-wall contouring for reducing the nozzle's erosion damage most effectively. The results indicate that increasing the end-wall contraction ratio and (or) decreasing the distance between the starting point of end-wall contouring and the trailing edge can significantly reduce the erosion-induced weight-loss of the nozzle, and can slightly improve the nozzle efficiency, irrespective of the variation in the particles size distribution and the aerodynamic parameters of a steam turbine. A main reason of erosion reduction is that the movement of loading towards the rear of the nozzle cascade caused by these contoured end walls has reduced the incident velocity of particles. In this study, the weight-loss of the nozzle was reduced by 40—50 per cent, and the nozzle efficiency was improved by 0.4—0.5 per cent by improving the end-wall contouring of the nozzle according to the methods mentioned above.

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49

Kirtley, K. R., and B. Lakshminarayana. "Computation of Three-Dimensional Turbulent Turbomachinery Flows Using a Coupled Parabolic-Marching Method." Journal of Turbomachinery 110, no. 4 (October 1, 1988): 549–56. http://dx.doi.org/10.1115/1.3262230.

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A new coupled parabolic-marching method was developed to compute the three-dimensional turbulent flow in a turbine endwall cascade, a compressor cascade wake, and an axial flow compressor rotor passage. The method solves the partially parabolized incompressible Navier–Stokes equation and continuity in a coupled fashion. The continuity equation was manipulated using pseudocompressibility theory to give a convergent algorithm for complex geometries. The computed end-wall boundary layers and secondary flow compared well with the experimental data for the turbine cascade as did the wake profiles for the compressor cascade using a k–ε turbulence model. Suction side boundary layers, pressure distributions, and exit stagnation pressure losses compared reasonably well with the data for the compressor rotor.

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50

Fershalov, Mikhail Yu, Andrey Yu Fershalov, and Juriy Yu Fershalov. "Calculation Reactivity Degree for Axial Low-Account Turbines with Small Emergence Angles of Nozzle Devices." Advanced Materials Research 915-916 (April 2014): 341–44. http://dx.doi.org/10.4028/www.scientific.net/amr.915-916.341.

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This paper presents the possibility of constructing a mathematical regression model for the reactivity degree of turbine stages (with the nozzle exit angles 5 ° ... 9 ° and a diameter of 0.17 m) depending on the constructed and model factors that have the greatest impact on the function value.

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