Academic literature on the topic 'Vortex compressor'
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Journal articles on the topic "Vortex compressor"
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Hong, Shuli, Guoping Huang, Yuxuan Yang, and Zepeng Liu. "Introduction of DMD Method to Study the Dynamic Structures of a Three-Dimensional Centrifugal Compressor with and without Flow Control." Energies 11, no. 11 (November 9, 2018): 3098. http://dx.doi.org/10.3390/en11113098.
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The flow structures around the blade tip, mainly large-scale leakage vortex, exert a great influence on compressor performance. By applying unsteady jet control technology at the blade tip in this study, the performance of the compressor can be greatly improved. A numerical simulation is conducted to study the flow characteristics of a centrifugal compressor with and without a flow control. The complex flow structures cause great difficulties in the analysis of the dynamic behavior and flow control mechanism. Thus, we introduced a dynamic flow field analysis technology called dynamic mode decomposition (DMD). The global spectrums with different global energy norms and the coherent structures with different scales can be obtained through the DMD analysis of the three-dimensional controlled and uncontrolled compressors. The results show that the coherent structures are homogeneous in the controlled compressor. The leakage vortex is weakened, and its influence range of unsteady fluctuation is reduced in the controlled compressor. The effective flow control created uniform vortex structures and improved the overall order of the flow field in the compressor. This research provides a feasible direction for future flow control applications, such as transferring the energy of the dominant vortices to small-scale vortices.
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Jiang, Bin, Xiangtong Shi, Qun Zheng, Qingfang Zhu, Zhongliang Chen, and Zhitao Tian. "The Relationship of Spike Stall and Hub Corner Separation in Axial Compressor." International Journal of Turbo & Jet-Engines 37, no. 1 (March 26, 2020): 1–16. http://dx.doi.org/10.1515/tjj-2016-0046.
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AbstractThe onset of spike stall induced by the interaction of hub corner separation flow with the tip leakage flow is investigated in detail by numerical method in this paper. The time resolved results indicate that the remarkable radial secondary flow from hub to tip near the trailing edge is formed when the compressor approaching rotating stall. The radial secondary flow is unstable and cross-passages propagates, which flows in and away out of the tip region periodically. The disturbance caused by radial secondary flow will influence the tip leakage flow directly by reforming the vortexes in blade tip region. A secondary vortex which comes from the radial migration of corner separation and is induced by the tip leakage vortex appears in the tip region. The simulation result demonstrates that the generation of the secondary vortex is an important symbol of blockage growth in the tip region at the stall inception phase. The disturbance produced by secondary vortex is an incentive of the leading edge overflow and the intensity of secondary vortex could be used as a criterion of rotating stall before leading edge spillage.
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Schrapp, H., U. Stark, and H. Saathoff. "Unsteady behaviour of the tip clearance vortex in a rotor equivalent compressor cascade." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 223, no. 6 (July 6, 2009): 635–43. http://dx.doi.org/10.1243/09576509jpe816.
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From earlier experimental investigations in a single-stage axial-flow pump and different numerical calculations of the flow in single-stage axial-flow compressors, it is known that vortex breakdown of the tip clearance vortex can take place in turbomachines, although an experimental proof for subsonic compressors is lacking. Vortex breakdown, if existent, is a source of high instability in the sensitive tip region of axial-flow pumps and compressors and will also play an important role in the stall inception process. Therefore, the flow in a linear compressor cascade with a 3 per cent tip clearance to one side has been investigated at different flow angles from the design point up to the stability limit of the cascade. The cascade resembles the tip section of a single-stage, axial-flow, low-speed compressor that is also in use at the Technical University of Braunschweig. The measuring techniques used were (a) a commercial particle image velocimetry (PIV) system and (b) a pressure measuring system with several flush mounted high-response pressure transducers at selected locations where the vortex was expected. As the cascade approaches its stall limit, the analysis of the pressure signals in the frequency domain revealed a bump of increased amplitude at a certain non-dimensional frequency for some of the measuring positions. The measuring positions that exhibited the bump correlated very well with a paraboloid-shaped region of high standard deviation enveloping an area of very low momentum fluid. It is shown that the frequency of the striking bump corresponds to the rotational frequency of the vortex calculated from the PIV measurements.
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Wu, T. T., and W. H. Hsieh. "Compression Processes and Performance Analysis of a High-Pressure Reciprocating Gas Compressor." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 210, no. 2 (March 1996): 153–65. http://dx.doi.org/10.1243/pime_proc_1996_210_182_02.
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Compression processes and compressor performance in a two-stage 41.34 MPa (6000 lb/in2) reciprocating gas compressor were investigated by transient multi-dimensional and transient global thermodynamic models. The transient multi-dimensional model was adopted to predict the two-dimensional compression processes in the second-stage cylinder of the high-pressure reciprocating gas compressor. Calculated results showed no significant temperature gradients anywhere in the compressor cylinder except near the wall, throughout one complete compressor cycle. On the other hand, the calculated velocity fields and streamline contours showed a convergent flow pattern during the process of compression with discharge and a large recirculation vortex during the process of expansion with suction. A parametric study based on the transient global thermodynamic model was conducted to investigate the effect of various parameters, that is clearance volume, wall temperature of cylinder head and stroke length, on the compressor performance. Among these parameters, it was found that the clearance volume had the strongest effect on the compressor performance. A reduced clearance volume increased volumetric efficiency. Also, it was found that decreasing the stroke length would not degrade the compressor performance, but it could reduce the compressor size and thereby the manufacturing cost significantly.
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Mao, Xiaochen, Bo Liu, and Tianquan Tang. "Effect of casing aspiration on the tip leakage flow in the axial flow compressor cascade." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 232, no. 3 (August 3, 2017): 225–39. http://dx.doi.org/10.1177/0957650917724598.
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Tip leakage flow is usually responsible for the deterioration of compressor performance and stability. The current paper conducts numerical simulations on the impact of casing aspiration on the axial compressor cascade performance. Three aspiration schemes with different chordwise coverage are studied and analyzed. It is found that the cascade performance can be effectively improved by the appropriate casing aspiration, and the optimum aspiration scheme should cover the area including the onset point of tip leakage vortex and its vicinity. The control mechanisms are different for the aspiration schemes located at different blade chord ranges. For the aspiration scheme covering the onset point of tip leakage vortex, the improvement of the cascade performance is mainly due to that the starting point of the tip leakage vortex is shifted downstream. The original tip leakage vortex structure is divided into two parts if the aspiration scheme is located behind the onset point of tip leakage vortex and the final control effect is the combination of the influence from the two different parts of tip leakage vortex. Additionally, the casing aspiration redistributes the blade loading along the chord near blade tip. The results of these investigations may offer guidance for the appropriate design of aspiration scheme in the future updated compressors and the overall total pressure loss coefficient caused by aspiration slot should be considered in the design process.
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Khalid, S. A. "Factors Affecting Measured Axial Compressor Tip Clearance Vortex Circulation." Journal of Turbomachinery 117, no. 3 (July 1, 1995): 487–90. http://dx.doi.org/10.1115/1.2835685.
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The relationship between turbomachinery blade circulation and tip clearance vortex circulation measured experimentally is examined using three-dimensional viscous flow computations. It is shown that the clearance vortex circulation one would measure is dependent on the placement of the fluid contour around which the circulation measurement is taken. Radial transport of vorticity results in the magnitude of the measured clearance vortex circulation generally being less than the blade circulation. For compressors, radial transport of vorticity shed from the blade tip in proximity to the endwall is the principal contributor to the discrepancy between the measured vortex circulation and blade circulation. Further, diffusion of vorticity shed at the blade tip toward the endwall makes it impossible in most practical cases to construct a fluid contour around the vortex that encloses all, and only, the vorticity shed from the blade tip. One should thus not expect agreement between measured tip clearance vortex circulation and circulation around the blade.
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Mailach, R., I. Lehmann, and K. Vogeler. "Rotating Instabilities in an Axial Compressor Originating From the Fluctuating Blade Tip Vortex." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 453–60. http://dx.doi.org/10.1115/1.1370160.
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Rotating instabilities (RIs) have been observed in axial flow fans and centrifugal compressors as well as in low-speed and high-speed axial compressors. They are responsible for the excitation of high amplitude rotor blade vibrations and noise generation. This flow phenomenon moves relative to the rotor blades and causes periodic vortex separations at the blade tips and an axial reversed flow through the tip clearance of the rotor blades. The paper describes experimental investigations of RIs in the Dresden Low-Speed Research Compressor (LSRC). The objective is to show that the fluctuation of the blade tip vortex is responsible for the origination of this flow phenomenon. RIs have been found at operating points near the stability limit of the compressor with relatively large tip clearance of the rotor blades. The application of time-resolving sensors in both fixed and rotating frame of reference enables a detailed description of the circumferential structure and the spatial development of this unsteady flow phenomenon, which is limited to the blade tip region. Laser-Doppler-anemometry (LDA) within the rotor blade passages and within the tip clearance as well as unsteady pressure measurements on the rotor blades show the structure of the blade tip vortex. It will be shown that the periodical interaction of the blade tip vortex of one blade with the flow at the adjacent blade is responsible for the generation of a rotating structure with high mode orders, termed a rotating instability.
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Hah, C., and J. Loellbach. "Development of Hub Corner Stall and Its Influence on the Performance of Axial Compressor Blade Rows." Journal of Turbomachinery 121, no. 1 (January 1, 1999): 67–77. http://dx.doi.org/10.1115/1.2841235.
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A detailed investigation has been performed to study hub corner stall phenomena in compressor blade rows. Three-dimensional flows in a subsonic annular compressor stator and in a transonic compressor rotor have been analyzed numerically by solving the Reynolds-averaged Navier–Stokes equations. The numerical results and the existing experimental data are interrogated to understand the mechanism of compressor hub corner stall. Both the measurements and the numerical solutions for the stator indicate that a strong twisterlike vortex is formed near the rear part of the blade suction surface. Low-momentum fluid inside the hub boundary layer is transported toward the suction side of the blade by this vortex. On the blade suction surface near the hub, this vortex forces fluid to move against the main flow direction and a limiting stream surface is formed near the hub. The formation of this vortex is the main mechanism of hub corner stall. When the aerodynamic loading is increased, the vortex initiates further upstream, which results in a larger corner stall region. For the transonic compressor rotor studied in this paper, the numerical solution indicates that a mild hub corner stall exists at 100 percent rotor speed. The hub corner stall, however, disappears at the reduced blade loading, which occurs at 60 percent rotor design speed. The present study demonstrates that hub corner stall is caused by a three-dimensional vortex system and that it does not seem to be correlated with a simple diffusion factor for the blade row.
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Zhang, Mingming, and Anping Hou. "Numerical Investigation on Unsteady Separation Flow Control in an Axial Compressor Using Detached-Eddy Simulation." Applied Sciences 9, no. 16 (August 12, 2019): 3298. http://dx.doi.org/10.3390/app9163298.
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Unsteady excitation has proved its effectiveness in separation flow control and has been extensively studied. It is observed that disordered shedding vortices in compressors can be controlled by unsteady excitation, especially when the excitation frequency coincides with the frequency of the shedding vortex. Furthermore, former experimental results indicated that unsteady excitation at other frequencies also had an impact on the structure of shedding vortices. To investigate the impact of excitation frequency on vortex shedding structure, the Detached-Eddy Simulation (DES) method was applied in the simulation of shedding vortex structure under unsteady excitations at different frequencies in an axial compressor. Effectiveness of the DES method was proved by comparison with URANS results. The simulation results showed a good agreement with the former experiment. The numerical results indicated that the separation flow can be partly controlled when the excitation frequency coincided with the unsteady flow inherent frequency. It showed an increase in stage performance under the less-studied separation flow control by excitation at a certain frequency of pressure side shedding vortex. Compared with other frequencies of shedding vortices, the frequency of pressure side shedding vortex was less sensitive to mass-flow variation. Therefore, it has potential for easier application on flow control in industrial compressors.
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Rattanongphisat, Waraporn. "Efficiency of Vortex Tube Enclosure Cooling." Applied Mechanics and Materials 666 (October 2014): 154–58. http://dx.doi.org/10.4028/www.scientific.net/amm.666.154.
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A vortex tube offers an alternative cooling with advantages of simplicity and compact. Using a natural refrigerant, a vortex tube enclosure cooling is environmentally benign. In this paper, the performance of a vortex tube enclosure cooling, VTEC, is investigated experimentally. The VTEC system comprises of the vortex tube cooling, an enclosure with a volume space of 0.045 m3, an air compressor, a compressed air storage tank and a compressed air line. The VTEC system is tested for its efficiency and cooling potential in the laboratory. An operating condition is controlled by a pressure regulator for an inlet air pressure of 3 bars, for energy saving, and a cold flow rate is adjusted by a needle valve near the hot exit of a vortex tube for the cold fraction between 0 and 1. Accordingly, the analysis of experimental data shows the maximum isentropic efficiency of the vortex tube enclosure cooling is 0.37 at the cold mass fraction of 0.45. Air temperature in the enclosure is about 13°C in average.
Dissertations / Theses on the topic "Vortex compressor"
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Lim, Choon Peng. "Experimental investigation of vortex shedding in high Reynolds number flow over compressor blades in cascade." Thesis, Monterey, Calif. : Springfield, Va. : Naval Postgraduate School ; Available from National Technical Information Service, 2003. http://library.nps.navy.mil/uhtbin/hyperion-image/03Mar%5FLim.pdf.
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Thesis (M.S. in Aeronautical Engineering)--Naval Postgraduate School, March 2003.Thesis advisor(s): Garth V. Hobson, Raymond P. Shreeve. Includes bibliographical references (p. 81-82). Also available online.
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Intaratep, Nanyaporn. "Formation and Development of the Tip Leakage Vortex in a Simulated Axial Compressor with Unsteady Inflow." Diss., Virginia Tech, 2006. http://hdl.handle.net/10919/26771.
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The interaction between rotor blade tip leakage vortex and inflow disturbances, such as encountered in shrouded marine propulsors, was simulated in the Virginia Tech Linear Cascade Wind Tunnel equipped with a moving endwall system. Upstream of the blade row, idealized periodic inflow unsteadiness was generated using vortex generator pairs attached to the endwall at the same spacing as the blade spacing. At three tip gap settings, 1.7%c, 3.3%c and 5.7%c, the flow near the lower endwall of the center blade passage was investigated through three-component mean velocity and turbulence distributions measured by four-sensor hotwires. Besides time-averaged data, the measurements were processed for phase-locked analysis, with respect to pitchwise locations of the vortex generators relative to the blade passage. Moreover, surface pressure distributions at the blade tip were acquired at eight tip gaps from 0.87%c to 12.9%c. Measurements of pressure-velocity correlation were also performed with wall motion but without inflow disturbances.
Achieved in this study is an understanding of the characteristics and structures of the tip leakage vortex at its initial formation. The mechanism of the tip leakage vortex formation seems to be independent of the tip gap setting. The tip leakage vortex consists of a vortical structure and a region of low streamwise-momentum fluid next to the endwall. The vortical structure is initially attached to the blade tip that creates it. This structure picks up circulation shed from that blade tip, as well as those from the endwall boundary layer, and becomes stronger with downstream distance. Partially induced by the mirror images in the endwall, the vortical structure starts to move across the passage resulting in a reduction in its rotational strength as the cross sectional area of the vortex increases but little circulation is added. The larger the tip gap, the longer the vortical structure stays attached to the blade tip, and the stronger the structure when it reaches downstream of the passage.
Phased-averaged data show that the inflow disturbances cause small-scale responses and large-scale responses upstream and downstream of the vortex shedding location, respectively. This difference in scale is possibly dictated by a variation in the shedding location since the amount of circulation in the vortex is dependent on this location. The inflow disturbances possibly cause a variation in the shedding location by manipulating the separation of the tip leakage flow from the endwall and consequently the flowâ s roll-up process. Even though this manipulation only perturbs the leakage flow in a small scale, the shedding mechanism of the tip leakage vortex amplifies the outcome.Ph. D.
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Baiense, Jr Joao C. "Vortex Generator Jet Flow Control in Highly Loaded Compressors." Digital WPI, 2014. https://digitalcommons.wpi.edu/etd-theses/916.
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"A flow control method for minimizing losses in a highly loaded compressor blade was analyzed. Passive and active flow control experiments with vortex generator jets were conducted on a seven blade linear compressor cascade to demonstrate the potential application of passive flow control on a highly loaded blade. Passive flow control vortex generator jets use the pressure distribution generated by air flow over the blade profile to drive jets from the pressure side to the suction side. Active flow control was analyzed by pressuring the blade plenum with an auxiliary compressor unit. Active flow control decreased profile losses by approximately 37 % while passive flow control had negligible impact on the profile loss of a highly loaded blade. Passive flow control was able to achieve a jet velocity ratio, jet velocity to upstream velocity, of 0.525. The success of active flow control with a velocity ratio of 0.9 suggests there is potential for passive flow control to be effective. The research presented in this thesis is motivated by the potential savings in the applications of passive flow control in gas turbine axial compressors by increasing the aerodynamic load of each stage. Increased stage loading that is properly controlled can reduce the number of stages required to achieve the desired pressure compression ratio."
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Lenglin, Geoffroy (Geoffroy Philippe) 1976. "Characterization of wake- and tip-vortex-induced unsteady blade response in multistage compressor environment." Thesis, Massachusetts Institute of Technology, 2002. http://hdl.handle.net/1721.1/82227.
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Znidarčić, Matej. "Computational Validation of the Compressor Design Program Blade Layout Method." Case Western Reserve University School of Graduate Studies / OhioLINK, 2012. http://rave.ohiolink.edu/etdc/view?acc_num=case1323361068.
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Семенов, Ф. Д. "Розроблення вихрового компресора для системи газодинамічних ущільнень і дослідження впливу параметрів на його ефективність." Master's thesis, Сумський державний університет, 2020. https://essuir.sumdu.edu.ua/handle/123456789/82264.
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У роботі виконано розрахунок одноступеневого та двоступеневого вихрового компресора. Виконано порівняльний аналіз розрахунків для різниці діаметрів коліс (D = 250/255/318/391) компресорів на параметри стиснення Рвх = 0,4 МПа, Рвих = 0,45 МПа, Vе = 0,006 м3/с, Т0 = 288 К, n = 3500 об/хв, середовище – аміак. В результаті визначено, що на задані значення кращим варіантом є двоступеневий компресор з діаметрами коліс D2 = 250 мм та D2 = 255 мм.В работе выполнен расчет одноступенчатого и двухступенчатого вихревого компрессора. Выполнен сравнительный анализ расчетов для разницы диаметров колес (D = 250/255/318/391) компрессоров на параметры сжатия Рвх = 0,4 МПа, Рвых = 0,45 МПа, Vе = 0,006 м3 / с, Т0 = 288 К, n = 3500 об / мин, среда - аммиак. В результате определено, что на заданные значения лучшим вариантом является двухступенчатый компрессор с диаметрами колес D2 = 250 мм и D2 = 255 мм.
The calculation of a single-stage and two-stage vortex compressor is performed in this work. A comparative analysis of calculations for the difference in wheel diameters (D = 250/255/318/391) of compressors for compression parameters Pvx = 0.4 MPa, Pvih = 0.45 MPa, Ve = 0.006 m3 / s, T0 = 288 K, n = 3500 rpm, medium - ammonia. As a result, it was determined that the best option for the set values is a two-stage compressor with wheel diameters D2 = 250 mm and D2 = 255 mm.
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Brown, Peter J. "Experimental investigation on vortex shedding in flow over second-generation, controlled-diffusion, compressor blades in cascade." Thesis, Monterey, California. Naval Postgraduate School, 2002. http://hdl.handle.net/10945/6092.
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Approved for public release, distribution unlimitedAn investigation of vortex shedding downstream of a cascade of second-generation, controlled-diffusion, compressor stator blades, at off-design inlet-flow angles of 31, 33 and 35 degrees and Reynolds numbers, based on chord length, of 280,000, 380,000 and 640,000 is reported. The objective of the study was to characterize the flow and shedding through various complementary methods. Blade surface pressure measurements were taken from a fully instrumented blade, and distributions of pressure coefficients were determined. Five-hole probe wake surveys were performed at midspan, and the total pressure loss coefficients and axial velocity ratios were calculated. Upstream inlet-flow angle was set, and further characterized through two-component laser-Doppler velocimetry (LDV). Hot-wire anemometry measurements were performed at mid span, in the wake, and the reduced data were compared with two-component LDV surveys of the same regions. Plots of hot-wire vs. LDV turbulence data are reported in addition to power spectra documenting the shedding events. Vortex shedding was determined to be a leading edge phenomenon as periodic shedding was only detected on the pressure side of the wake. The frequency and magnitude of shedding were found to be independent of incidence angle, and to increase with Reynolds number at constant incidence angle. The Strouhal number, based on leading edge diameter, was found to be in the range of 0.23-0.26, which is comparable to that of vortex shedding behind a circular cylinder in the Reynolds number range tested.
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Langford, Matthew David. "Experimental Investigation of the Effects of a Passing Shock on Compressor Stator Flow." Thesis, Virginia Tech, 2003. http://hdl.handle.net/10919/32020.
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A stator cascade was developed to simulate the flow conditions within a close-stage-spacing transonic axial compressor. Experiments were conducted in a linear transonic blowdown cascade wind tunnel with an inlet Mach number of 0.65. The bow shock from the downstream rotor was simulated by a single moving normal shock generated with a shock tube. First, steady pressure data were gathered to ensure that the stator cascade operated properly without the presence of the shock. Next, the effects of the passing shock on the stator flow field were investigated using shadowgraph photography and Digital Particle Image Velocimetry (DPIV). Measurements were taken for three different shock strengths. In every case studied, a vortex formed near the stator trailing edge as the shock impacted the blade. The size of this vortex was shown to be directly related to the shock strength, and the vortex remained present in the trailing edge flow field throughout the cycle duration. Analysis of the DPIV data showed that the vortex acts as a flow blockage, with the extent of this blockage ranging from 2.9% of the passage for the weakest shock, to 14.3% of the passage for the strongest shock. The vortex was also shown to cause flow deviation up to 75° for the case with the strongest shock. Further analysis estimated that the total pressure losses due to shock-induced vorticity ranged from 46% to 113% of the steady wake losses. Finally, the total pressure loss purely due to the upstream-propagating normal shock was estimated to be roughly 0.22%.Master of Science
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Shin, Sangmook. "Reynolds-Averaged Navier-Stokes Computation of Tip Clearance Flow in a Compressor Cascade Using an Unstructured Grid." Diss., Virginia Tech, 2001. http://hdl.handle.net/10919/28947.
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A three-dimensional unstructured incompressible RANS code has been developed using artificial compressibility and Spalart-Allmaras eddy viscosity model. A node-based finite volume method is used in which all flow variables are defined at the vertices of tetrahedrons in an unstructured grid. The inviscid fluxes are computed by using the Roe's flux difference splitting method, and higher order accuracy is attained by data reconstruction based on Taylor series expansion. Gauss theorem is used to formulate necessary gradients. For time integration, an implicit scheme based on linearized Euler backward method is used.
A tetrahedral unstructured grid generation code has been also developed and applied to the tip clearance flow in a highly staggered cascade. Surface grids are first generated in the flow passage and blade tip by using several triangulation methods including Delaunay triangulation, advancing front method and advancing layer method. Then the whole computational domain including tip gap region is filled with prisms using the surface grids. Each prism is divided into three tetrahedrons. To accomplish this division in a consistent manner, connectivity pattern is assigned to each triangle in the surface grids. A new algorithm is devised to assign the connectivity pattern without reference to the particular method of triangulation. This technique offers great flexibility in surface grid generation.
The code has been validated by comparisons with available computational and experimental results for several test cases: invisicd flow around NACA section, laminar and turbulent flow over a flat plate, turbulent flow through double-circular arc cascade and laminar flow through a square duct with 90° bend. For the laminar flat plate case, the velocity profile and skin friction coefficient are in excellent agreement with Blasius solution. For the turbulent flat plate case, velocity profiles are in full agreement with the law of the wall up to Reynolds number of 1.0E8, however, the skin friction coefficient is under-predicted by about 10% in comparison with empirical formula. Blade loading for the two-dimensional circular arc cascade is also compared with experiments. The results obtained with the experimental inflow angle (51.5° ) show some discrepancies at the trailing edge and severely under-predict the suction peak at the leading edge. These discrepancies are completely remedied if the inflow angle is increased to 53.5° . The code is also capable of predicting the secondary flow in the square duct with 90° bend, and the velocity profiles are in good agreement with measurements and published Navier-Stokes computations.
Finally the code is applied to a linear cascade that has GE rotor B section with tip clearance and a high stagger angle of 56.9° . The overall structure of the tip clearance flow is well predicted. Loss of loading due to tip leakage flow and reloading due to tip leakage vortex are presented. On the end wall, separation line of the tip leakage vortex and reattachment line of passage vortex are identified. The location of the tip leakage vortex in the passage agrees very well with oil flow visualization. Separation bubble on the blade tip is also predicted. Mean streamwise velocity contours and cross sectional velocity vectors are compared with experimental results in the near wake, and good agreements are observed. It is concluded that Spalart-Allmaras turbulence model is adequate for this type of flow field except at locations where the tip leakage vortex of one blade interacts with the wake of a following blade. This situation may prevail for blades with longer span and/or in the far wake. Prediction of such an interaction presents a challenge to RANS computations.
The effects of blade span on the flow structure have been also investigated. Two cascades with blades of aspect ratios of 0.5 and 1.0 are considered. By comparing pressure distributions on the blade, it is shown that the aspect ratio has strong effects on loading distribution on the blade although the tip gap height is very small (0.016 chord). Grid convergence study has been carried out with three different grids for pressure distributions and limiting streamlines on the end wall.Ph. D.
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Carr, M. I. "The excitation of acoustic resonances in an axial flow compressor stage by vortex shedding from aerofoil section blading." Thesis, Swansea University, 1986. http://ethos.bl.uk/OrderDetails.do?uin=uk.bl.ethos.636209.
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In recent years continuing development of the axial flow compressor for use in the aero-engine has increased its susceptibility to unsteady flow phenomena which can cause severe blade vibration. A source which has emerged and become of considerable importance is excitation by acoustic resonances. An experimental investigation in a single stage axial flow compressor rig has been performed to ascertain whether acoustic resonances can be excited by vortex shedding from loaded aerofoil section blades. A further experimental programme, to study further the effect of inter-row spacing, was peformed in both an open jet facility and a wind tunnel facility with a tandem plate arrangement. Results showed that acoustic resonances could be excited in a compressor stage in which there was severe blade loading. The speed range over which the resonances were excited was demonstrated to be not only a function of the degree of loading but also the inter-row spacing. Vortex shedding will drive a resonance when the shedding is correlated by the resonant acoustic field and interaction between the vortices and the acoustic field in the vicinity of the blades may result in a net positive input of acoustic energy. As a result the phase of the acoustic field as vortices pass over the trailing edge of the shedding blades and the leading and trailing edges of the downstream blades, control the energy generation. The inter-row spacing controls the phase of the downstream blade interaction and therefore is a major factor influencing the resonant acoustic amplitude. As well as the fundamental acoustic mode, a resonance can also drive significant blade vibration in two other consequential frequency bands which are: a) Sum and Difference frequency bands due to acoustic non-linearity and b) Sidebands of the fundamental modes due to spatial modulation effects caused by flow distortions.
Books on the topic "Vortex compressor"
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Tip Vortex and Crenulation Effects in a Compressor Cascade with Moving Endwall. Storming Media, 1999.
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Experimental Investigation of Vortex Shedding in High Reynolds Number Flow Over Compressor Blades in Cascade. Storming Media, 2003.
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Experimental Investigation of Vortex Shedding in Flow Over Second- Generation, Controlled-Diffusion, Compressor Blades in Cascade. Storming Media, 2002.
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Experimental and computational investigation of the tip clearance flow in a transonic axial compressor rotor. [Washington, DC]: National Aeronautics and Space Administration, 1995.
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L, Celestina Mark, and United States. National Aeronautics and Space Administration., eds. Experimental and computational investigation of the tip clearance flow in a transonic axial compressor rotor. [Washington, DC]: National Aeronautics and Space Administration, 1995.
Find full textBook chapters on the topic "Vortex compressor"
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Decher, Reiner. "The Compressor: Gas Turbine Engine Keystone." In The Vortex and The Jet, 109–19. Singapore: Springer Singapore, 2022. http://dx.doi.org/10.1007/978-981-16-8028-1_10.
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AbstractAll engines involve the raising of pressure in the engine’s medium, be it liquid compression of water, piston-cylinder compression of air or compression by aerodynamic forces. The last is the hardest because it is required to be efficient and stable. An aerodynamic air compressor works by slowing velocities to raise pressure. The airfoil has a difficult time doing this reliably and well and so does the compressor. When it does, it allows for amazing performance of the engine in terms of power and efficiency.
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Decher, Reiner. "Bypass and Other Engines." In The Vortex and The Jet, 121–24. Singapore: Springer Singapore, 2022. http://dx.doi.org/10.1007/978-981-16-8028-1_11.
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AbstractThe modern fan engine has a propeller-like component whose aerodynamic performance differs from that of a propeller at the blade level but serves the same function. The blading aerodynamics differs from that of a compressor in notable ways.
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Manisankar, C., S. B. Verma, and C. Raju. "Shock-Wave Boundary-Layer Interaction Control on a Compression Corner Using Mechanical Vortex Generators." In 28th International Symposium on Shock Waves, 409–15. Berlin, Heidelberg: Springer Berlin Heidelberg, 2012. http://dx.doi.org/10.1007/978-3-642-25685-1_62.
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Renganathan, Manimaran. "Prediction of In-Cylinder Swirl in a Compression Ignition Engine with Vortex Tube Using Artificial and Recurrent Neural Networks." In Artificial Intelligence and Technologies, 53–62. Singapore: Springer Singapore, 2021. http://dx.doi.org/10.1007/978-981-16-6448-9_6.
Full textConference papers on the topic "Vortex compressor"
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Zheng, Tan, Xiaoqing Qiang, and Jinfang Teng. "Effects of Vortex-Vortex Interaction in a Compressor Cascade With Vortex Generators." In ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gt2015-43042.
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This paper presents a numerical investigation to explore the effects of vortex generators on a high speed compressor cascade. Secondary flow effects like the corner separation vortex have an influence on the performance of a compressor cascade such as leading to increased losses. In order to control the corner separation vortex and reduce losses, an extensive study of vortex generators applied to a compressor cascade is conducted. A preliminary study by steady 3D RANS simulations is performed using the Spalart-Allmaras turbulence model. The aerodynamic performance as well as the behavior of the corner separation vortex is investigated in the compressor cascade without vortex generators. Then, a vortex generator is added to the cascade, which is numerically simulated. Various configurations are considered, which are decided by the height and installation angle of the vortex generator. Comparison of the performance attained by these configurations results in an optimum scheme that has minimum losses. Furthermore, unsteady 3D DES simulations are performed with the optimum configuration. This method that predicts the flow field more precisely could help verify the accuracy of the RANS results. Finally, by analyzing all the resulting aerodynamic performance and numerical flow phenomena, the mechanism of vortex-vortex interaction is presented and discussed, which could be a criterion to reduce the corner separation flow and enhance the performance of axial compressors.
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Bae, Jinwoo, Kenneth S. Breuer, and Choon S. Tan. "Periodic Unsteadiness of Compressor Tip Clearance Vortex." In ASME Turbo Expo 2004: Power for Land, Sea, and Air. ASMEDC, 2004. http://dx.doi.org/10.1115/gt2004-53015.
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Natural and forced responses of the tip clearance vortex are measured in a linear compressor cascade to characterize periodic unsteadiness of the tip clearance vortex. There exists a natural frequency at which the tip clearance vortex is the most receptive to external forcing, thus resulting in mixing enhancement and flow blockage reduction. A physical explanation of the source of the observed periodic unsteadiness is suggested based on the trailing vortex instability theory. Observations of the time scale for the unsteadiness from different compressor geometries and flow conditions are shown to scale with a reduced frequency based on convective time through the blade passage.
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Cao, Zhiyuan, Cheng Song, Xi Gao, and Xiang Zhang. "Effect of Pulsed Endwall Injection on Flow Separation and Vortex Structure of a Compressor Cascade." In GPPS Xi'an21. GPPS, 2022. http://dx.doi.org/10.33737/gpps21-tc-46.
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Pulsed endwall air injection (PEAI) is an effective method to eliminate flow separation and improve blade loading of compressor. Most studies on PEAI have focused on improving the aerodynamic performance of compressor, and few studies investigated the influence on the flow field in detail, especially vortex structure in compressors. Aiming at revealing the effect mechanism of PEAI on flow separation and vortex structure, unsteady numerical study is carried out. The effect of PEAI parameters (frequency and amplitude) on flow field of compressor cascade is addressed. Results show that cascade loss coefficient and endwall loss coefficient of cascade are respectively reduced 13.92% and 28.24% under the optimal PEAI scheme. Higher injection amplitude leads to more decrease of corner separation and deterioration of flow field at mid-span, and double concentration shedding vortex (CSV) structure is found when injection amplitude is 1.40. As the injection frequency increasing, the regular injection leads to a series of “injection vortexes” distributing with equal interval, which suppress CSV and PV. Keywords: axial flow compressor, corner separation, flow control, pulsed endwall air injection, vortex structure
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Ballmann, J., and W. Hofmann. "Tip clearance vortex development and shock-vortex-interaction in a transonic axial compressor rotor." In 40th AIAA Aerospace Sciences Meeting & Exhibit. Reston, Virigina: American Institute of Aeronautics and Astronautics, 2002. http://dx.doi.org/10.2514/6.2002-83.
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Tang, Yan-Ping, and Mao-Zhang Chen. "Vortex Control Over End Wall Flow in Compressor Cascades." 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-228.
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Three methods of vortex control over the end wall flow in compressor cascades have been investigated experimentally. The total pressure loss at the exit of a linear compressor cascade is reduced 6.5%, 10.5% and 26.5% respectively by these methods for different incidences over a range of moderate-high values. The physics of these methods has been discussed and some new concepts of vortex control techniques in compressor cascades have been proposed.
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Hah, Chunill, Jo¨rg Bergner, and Heinz-Peter Schiffer. "Tip Clearance Vortex Oscillation, Vortex Shedding and Rotating Instabilities in an Axial Transonic Compressor Rotor." In ASME Turbo Expo 2008: Power for Land, Sea, and Air. ASMEDC, 2008. http://dx.doi.org/10.1115/gt2008-50105.
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Unsteady flow characteristics in a modern transonic axial compressor operating near stall are studied in detail. Measured data from high-response pressure probes show that the tip clearance vortex oscillates substantially near stall. Instantaneous flow structure varies substantially among different blade passages even with uniform inlet flow. Fast Fourier transformation of measured wall pressure shows a dominant frequency component that is between 30% and 40% of the rotor speed. To identify and analyze this phenomenon, computational studies based on a single passage and full annulus were carried out. The flow field in a transonic compressor near stall is heavily influenced by the unsteady motion of tip clearance vortices. Therefore, a Large Eddy Simulation (LES) was carried out to capture transient characteristics of the tip clearance vortex more realistically. The wall pressure spectrum from the current full annulus analysis also shows a dominant frequency when the rotor operates near stall. The calculated peak frequency is about 30% of the rotor frequency. The dominant frequency, which is non-synchronous with the rotor blade, is due to rotating flow instabilities. Flow interactions across blade passages due to synchronized tip clearance vortex oscillation seem to be the main cause.
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Tomita, Isao, Seiichi Ibaraki, Masato Furukawa, and Kazutoyo Yamada. "The Effect of Tip Leakage Vortex for Operating Range Enhancement of Centrifugal Compressor." In ASME Turbo Expo 2012: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2012. http://dx.doi.org/10.1115/gt2012-68947.
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Recently, the application of turbochargers is increasing because they are effective in improving fuel consumption of engines. One of the most important turbocharger characteristics is compressor operating range, since it has been used in various driving patterns with the advent of variable geometry turbochargers. Owing to the complicated phenomena such as rotating stall occurring at low flow rate condition, flow analysis is very difficult and details of flow structure have not been fully understood for a long time since the early 1970’s. In this study, two compressors with different operating range width were investigated with experimental and computational flow analysis. In the compressor with narrow operating range, the amplitude of blade passing pressure fluctuation decreases rapidly and rotating stall occurs near surging. On the other hand, in the compressor with wide operating range, the blockage by the tip leakage vortex breakdown play a role in stabilizing the flow filed and keeping up a high performance even at low flow rates.
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Hergt, Alexander, Robert Meyer, and Karl Engel. "Effects of Vortex Generator Application on the Performance of a Compressor Cascade." In ASME Turbo Expo 2010: Power for Land, Sea, and Air. ASMEDC, 2010. http://dx.doi.org/10.1115/gt2010-22464.
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The performance of a compressor cascade is considerably influenced by secondary flow effects, like the cross flow on the end wall as well as the corner separation between the wall and the vane. An extensive experimental study of vortex generator application in a highly loaded compressor cascade was performed, in order to control these effects and enhance the aerodynamic performance. The results of the study will be used in future projects as a basis for parameterization in the design and optimization process for compressors in order to develop novel non-axisymmetric endwall as well as for blade modifications. The study includes the investigation of two vortex generator types, with different geometrical forms and their application on several positions in the compressor cascade. The investigation includes a detailed description of the secondary flow effects in the compressor cascade which is based on numerical and experimental results. This gives the basis for a specific approach of influencing the cascade flow by means of vortex generators. Depending on the vortex generator type and position, there is an impact on the end wall cross flow, the development of the horse shoe vortex at the leading edge of the vane and the extent of the corner separation achieved by improved mixing within the boundary layer. The experiments were carried out on a compressor cascade at a high-speed test facility at the DLR in Berlin at minimum loss (design point) and off-design of the cascade at Reynolds numbers up to Re = 0.6 × 106 (based on 40 mm chord) and Mach numbers up to M = 0.7. The cascade consisted of five vanes and their profiles represent the cut near hub of the stator vanes of the single stage axial compressor of the Technical University of Darmstadt. At the cascade design point the total pressure losses could be reduced by up to 9 percent with vortex generator configuration whereas the static pressure rise was nearly unaffected. Furthermore, the cascade deflection could be influenced considerably by vortex generators and also an enhancement of the cascade stall range could be achieved. All these results will be presented and discussed with respect to secondary flow mechanisms.
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Agarwal, Ruchika, Sridharan R. Narayanan, Shraman N. Goswami, and Balamurugan Srinivasan. "Numerical Analysis on Axial Compressor Stage Performance With Vortex Generators." In ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gt2015-43897.
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The performance of axial flow compressor stage can be improved by minimizing the effects of secondary flow and by avoiding flow separation. At higher blade loading, interaction of tip secondary flow and separated flow on blade surface is more near the tip of the rotor. This results in stall and hence decreases compressor performance. A previous study [1] was carried out to improve the performance of a rotating row of blades with the help of Vortex Generators (VGs) and considerable effects were observed. The current investigation is carried out to find out the effect of Vortex Generator (VG) on the performance of a compressor stage. NASA Rotor 37 with NASA Stator 37 (stage) is used as a test case for the current numerical investigation. VGs are placed at different chord wise as well as span wise locations. A mesh sensitivity study has been done so that simulation result is mesh independent. The results of numerical simulation with different geometrical forms and locations of VGs are presented in this paper. The investigation includes a description of the secondary flow effect and separation zone in compressor stage based on numerical as well as experimental results of NASA Rotor 37 with Stator 37 (without VG). It is also observed that the shape and location of the VG impacts the end wall cross flow and flow deflection [1], which result in enhanced stall range.
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Li, Xiaosa, Zegang Qian, and Qichao Yang. "Research of Whirlpool Sound Radiation in Turbulence Coherent Structures." In ASME 2013 International Mechanical Engineering Congress and Exposition. American Society of Mechanical Engineers, 2013. http://dx.doi.org/10.1115/imece2013-62890.
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The compressor aerodynamic noise consists of suction, exhausting noise, gas power noise and so on, and the exhausting noise is dominating. Gas which is compressed released instantaneously to form vortex injection noise and exhausting pulsation noise, yet will cause pipe vibration. Large Eddy Simulation model in fluid software FLUENT have been adopted to analyze unsteady flow field and the acoustic field and research unsteady vortex shedding and its noise radiation characteristics in compressor pipes. The results show that: the trailing edge vortex shedding phenomenon, interaction between separation vortex and the trailing vortex of downstream lead to a large gas pulsation which makes noise radiation enhanced in the compressor pipe flow field in screw compressor exhausting orifice. Combination turbulence and vortex-pair phenomena in coherent structures, based on the vortex sound equation, a mathematical model of vortex-pair acoustic radiation is established. Finally unit length sound power of the whirlpool is draw to lρ0U3M4.
Reports on the topic "Vortex compressor"
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Hershcovitch, Ady. Vortex stabilized electron beam compressed fusion grade plasma. Office of Scientific and Technical Information (OSTI), March 2014. http://dx.doi.org/10.2172/1127069.
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