To see the other types of publications on this topic, follow the link: Axial flow compressors. Compressors. Aerodynamics.
Journal articles on the topic 'Axial flow compressors. Compressors. Aerodynamics'
Create a spot-on reference in APA, MLA, Chicago, Harvard, and other styles
Consult the top 50 journal articles for your research on the topic 'Axial flow compressors. Compressors. Aerodynamics.'
Next to every source in the list of references, there is an 'Add to bibliography' button. Press on it, and we will generate automatically the bibliographic reference to the chosen work in the citation style you need: APA, MLA, Harvard, Chicago, Vancouver, etc.
You can also download the full text of the academic publication as pdf and read online its abstract whenever available in the metadata.
Browse journal articles on a wide variety of disciplines and organise your bibliography correctly.
1
Zhang, Yingying, and Shijie Zhang. "Performance prediction of transonic axial multistage compressor based on one-dimensional meanline method." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 235, no. 6 (February 25, 2021): 1355–69. http://dx.doi.org/10.1177/0957650921998819.
Full textAbstract:
This study proposes a 1D meanline program for the modeling of modern transonic axial multistage compressors. In this method, an improved blockage factor model is proposed. Work-done factor that varies with the compressor performance conditions is added in this program, and at the same time a notional blockage factor is kept. The coefficient of deviation angle model is tuned according to experimental data. In addition, two surge methods that originated from different sources are chosen to add in and compare with the new method called mass flow separation method. The salient issues presented here deal first with the construction of the compressor program. Three well-documented National Aerodynamics and Space Administration (NASA) axial transonic compressors are calculated, and the speedlines and aerodynamic parameters are compared with the experimental data to verify the reliability and robustness of the proposed method. Results show that consistent agreement can be obtained with such a performance prediction program. It was also apparent that the two common methods of surge prediction, which rely upon either stage or overall characteristic gradients, gave less agreement than the method called mass flow separation method.
2
Hall, E. J. "Aerodynamic modelling of multistage compressor flow fields Part 1: Analysis of rotor-stator-rotor aerodynamic interaction." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 212, no. 2 (February 1, 1998): 77–89. http://dx.doi.org/10.1243/0954410981532153.
Full textAbstract:
The primary purpose of this study was to investigate improved numerical techniques for predicting flows through multistage compressors. The vehicle chosen for this study was the Pennsylvania State University Research Compressor (PSRC). The PSRC facility consists of a 3 1/2-stage axial flow compressor which shares design features which are consistent with embedded stages of modern gas turbine engine axial flow compressors. In Part 1 of this two-part paper, several computational fluid dynamics techniques were applied to predict both steady and unsteady flows through the PSRC facility. Interblade row coupling via a circumferentially averaged mixing-plane approach was employed for steady flow analysis. A mesh density sensitivity study was performed to define the minimum mesh requirements necessary to achieve reasonable agreement with the experimental data. Time-dependent flow predictions were performed using a time-dependent interblade row coupling technique. These calculations evaluated the aerodynamic interactions occurring between rotor 2, stator 2 and rotor 3 for the PSRC rig.
3
Gallimore, S. J., and N. A. Cumpsty. "Spanwise Mixing in Multistage Axial Flow Compressors: Part I—Experimental Investigation." Journal of Turbomachinery 108, no. 1 (July 1, 1986): 2–9. http://dx.doi.org/10.1115/1.3262019.
Full textAbstract:
Spanwise mixing has been shown to be an essential feature of multistage compressor aerodynamics. The cause of spanwise mixing in multistage axial flow compressors has been investigated directly by using an ethylene tracer gas technique in two low-speed, four-stage machines. The results show that the dominant mechanism is that of turbulent type diffusion and not the radial convection of flow properties as has been previously suggested. The mixing was also found to be substantially uniform in magnitude all the way across the span with levels similar to those found in two-dimensional turbulent wakes.
4
Kang, Qiang, Shuguang Zuo, and Kaijun Wei. "Study on the aerodynamic noise of internal flow of regenerative flow compressors for a fuel-cell car." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 228, no. 7 (August 30, 2013): 1155–74. http://dx.doi.org/10.1177/0954406213500746.
Full textAbstract:
The regenerative flow compressor used in fuel-cell cars generates high aerodynamic noise, which is the main source of noise. Compared with the research on centrifugal or axial turbomachinery, research on the noise of regenerative flow compressors is far from adequate. This paper presents the on-going work on it at Tongji University based on both experimental and computational works. In this study, a three-dimensional unsteady computational fluid dynamic model of the compressor was constructed with the large eddy approach. The pressure fluctuation, vortex noise source and Ffowcs William-Hawkings (FW-H) method were used to analyze the characteristics of the aerodynamic noise sources. Additionally, the far-field aerodynamic noise generated by the internal flow of the compressor was predicted using the aeroacoustic finite element method. The simulation results were validated with the experimental data. It was found that combining the fluid dynamic model and aeroacoustic finite element analysis promising results for aerodynamic noise prediction of compressors could be produced. The effects of the impeller parameters on the aerodynamic noise of the compressor were also studied.
5
Li, Chang Zheng, and Yong Lei. "Compressor Surge Detection Based on Support Vector Data Description." Applied Mechanics and Materials 152-154 (January 2012): 1545–49. http://dx.doi.org/10.4028/www.scientific.net/amm.152-154.1545.
Full textAbstract:
Axial flow compressors work as an indispensable device in industry fields. Surge is a phenomenon of aerodynamic instability, which characterized by disruption of flow. When a compressor works in surge state, the vibration is so intense that it may causes accidents. Detecting surge timely and accurately not only insure safety of compressors but also is a key of active control of aerodynamic instability. Support vector data description (SVDD) is a one-class classification method developed based on the theory of support vector machine (SVM). In this paper, we introduce SVDD into the field of compressor surge detection. It demonstrates that SVDD method can give a warning far ahead of surge.
6
Kulyk, Mykola, Ivan Lastivka, and Yuri Tereshchenko. "EFFECT OF HYSTERESIS IN AXIAL COMPRESSORS OF GAS-TURBINE ENGINES." Aviation 16, no. 4 (December 24, 2012): 97–102. http://dx.doi.org/10.3846/16487788.2012.753679.
Full textAbstract:
The phenomenon of separated flow hysteresis in the process of the streamlining the axial compressor of gas-turbine engines is considered. Generalised results of research on the occurrence of hysteresis in the aerodynamic performance of compressor grids and its influence on the performance of the bladed disks of compressors that operate in real conditions of periodic circular non-uniformity are demonstrated.
7
Storace, A. F., D. C. Wisler, H. W. Shin, B. F. Beacher, F. F. Ehrich, Z. S. Spakovszky, M. Martinez-Sanchez, and S. J. Song. "Unsteady Flow and Whirl-Inducing Forces in Axial-Flow Compressors: Part I—Experiment." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 433–45. http://dx.doi.org/10.1115/1.1378299.
Full textAbstract:
An experimental and theoretical investigation has been conducted to evaluate the effects seen in axial-flow compressors when the centerline of the rotor is displaced from the centerline of the static structure of the engine. This creates circumferentially nonuniform rotor-tip clearances, unsteady flow, and potentially increased clearances if the rotating and stationary parts come in contact. The result not only adversely affects compressor stall margin, pressure rise capability, and efficiency, but also generates an unsteady, destabilizing, aerodynamic force, called the Thomas/Alford force, which contributes significantly to rotor whirl instabilities in turbomachinery. Determining both the direction and magnitude of this force in compressors, relative to those in turbines, is especially important for the design of mechanically stable turbomachinery components. Part I of this two-part paper addresses these issues experimentally and Part II presents analyses from relevant computational models. Our results clearly show that the Thomas/Alford force can promote significant backward rotor whirl over much of the operating range of modern compressors, although some regions of zero and forward whirl were found near the design point. This is the first time that definitive measurements, coupled with compelling analyses, have been reported in the literature to resolve the long-standing disparity in findings concerning the direction and magnitude of whirl-inducing forces important in the design of modern axial-flow compressors.
8
Ho¨nen, H., and H. E. Gallus. "Monitoring of Aerodynamic Load and Detection of Stall in Multistage Axial Compressors." Journal of Turbomachinery 117, no. 1 (January 1, 1995): 81–86. http://dx.doi.org/10.1115/1.2835645.
Full textAbstract:
The unsteady flow in a single-stage axial flow compressor at different operating conditions has been investigated with hot-wire and hot-film probes to find out the influence of the aerodynamic compressor load on the periodic fluctuations. These results are compared with measurements in the last stages of a multistage high-pressure compressor of a gas turbine for normal operation and under stall conditions. From the patterns of the frequency spectra of the measuring signals a parameter for the detection of the approach to the stability line of a compressor is derived. A method for the on-line monitoring of the aerodynamic load is presented. Based on these results a monitoring system has been developed. First experiences with this system, applied to two multistage compressors, are reported.
9
Wisler, D. C. "Loss Reduction in Axial-Flow Compressors Through Low-Speed Model Testing." Journal of Engineering for Gas Turbines and Power 107, no. 2 (April 1, 1985): 354–63. http://dx.doi.org/10.1115/1.3239730.
Full textAbstract:
A systematic procedure for reducing losses in axial-flow compressors is presented. In this procedure, a large, low-speed, aerodynamic model of a high-speed core compressor is designed and fabricated based on aerodynamic similarity principles. This model is then tested at low speed where high-loss regions associated with three-dimensional endwall boundary layers, flow separation, leakage, and secondary flows can be located, detailed measurements made, and loss mechanisms determined with much greater accuracy and much lower cost and risk than is possible in small, high-speed compressors. Design modifications are made by using custom-tailored airfoils and vector diagrams, airfoil endbends, and modified wall geometries in the high-loss regions. The design improvements resulting in reduced loss or increased stall margin are then scaled to high speed. This paper describes the procedure and presents experimental results to show that in some cases endwall loss has been reduced by as much as 10 percent, flow separation has been reduced or eliminated, and stall margin has been substantially improved by using these techniques.
10
Zhang, Botao, Xiaochen Mao, Xiaoxiong Wu, and Bo Liu. "Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage." Energies 14, no. 14 (July 10, 2021): 4168. http://dx.doi.org/10.3390/en14144168.
Full textAbstract:
To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.
11
Koyyalamudi, V. V. N. K. Satish, and Quamber H. Nagpurwala. "Stall Margin Improvement in a Centrifugal Compressor through Inducer Casing Treatment." International Journal of Rotating Machinery 2016 (2016): 1–19. http://dx.doi.org/10.1155/2016/2371524.
Full textAbstract:
The increasing trend of high stage pressure ratio with increased aerodynamic loading has led to reduction in stable operating range of centrifugal compressors with stall and surge initiating at relatively higher mass flow rates. The casing treatment technique of stall control is found to be effective in axial compressors, but very limited research work is published on the application of this technique in centrifugal compressors. Present research was aimed to investigate the effect of casing treatment on the performance and stall margin of a high speed, 4 : 1 pressure ratio centrifugal compressor through numerical simulations using ANSYS CFX software. Three casing treatment configurations were developed and incorporated in the shroud over the inducer of the impeller. The predicted performance of baseline compressor (without casing treatment) was in good agreement with published experimental data. The compressor with different inducer casing treatment geometries showed varying levels of stall margin improvement, up to a maximum of 18%. While the peak efficiency of the compressor with casing treatment dropped by 0.8%–1% compared to the baseline compressor, the choke mass flow rate was improved by 9.5%, thus enhancing the total stable operating range. The inlet configuration of the casing treatment was found to play an important role in stall margin improvement.
12
Gallimore, S. J. "Axial flow compressor design†." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 213, no. 5 (May 1, 1999): 437–49. http://dx.doi.org/10.1243/0954406991522680.
Full textAbstract:
The purpose of this paper is to set out some of the basic principles and rules associated with the design of axial flow compressors, principally for aero-engines, as well as the practical constraints that are inevitably present. The thrust is primarily on the aerodynamic design but this cannot be divorced from the mechanical aspects and so some of these are touched upon but are not gone into so deeply. The paper has been written from the point of view of the designer and tries to cover most of the points that need to be considered in order to produce a successful compressor. The emphasis has been on the theory behind the design process and on minimizing the reliance on empirical rules. However, because of the complexity of the flow, some empiricism still remains.
13
Hall, E. J. "Aerodynamic modelling of multistage compressor flow fields Part 2: Modelling deterministic stresses." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 212, no. 2 (February 1, 1998): 91–107. http://dx.doi.org/10.1243/0954410981532162.
Full textAbstract:
The primary purpose of this study was to investigate improved numerical techniques for predicting flows through multistage compressors. The vehicle chosen for this study was the Pennsylvania State University Research Compressor (PSRC). The PSRC facility consists of a 3 1/2-stage axial flow compressor which shares design features which are consistent with embedded stages of modern gas turbine engine axial flow compressors. In Part 2 of this two-part paper, time-dependent predictions of rotor- stator-rotor aerodynamic interactions are employed to quantify the levels and distribution of deterministic stresses resulting from the average-passage flow field description. Details of the spanwise and blade-to- blade distributions of the velocity correlations are examined and compared with results based on physical deterministic flow structures such as blade wakes and clearance flows. The predicted ‘apparent’ wake profile decay resulting from the interaction of the wake through a downstream blade row is presented and compared with test data. This ‘apparent’ wake profile decay is employed to define a simplified model for deterministic stress correlations in a steady state flow field prediction scheme which retains the ‘mixing- plane’ methodology. Calculations based on this proposed model are described and predicted results are compared with both time-dependent predictions and test data. The resulting prediction strategy is computationally efficient and also contains sufficient physical realism to permit its use in design studies.
14
Ehrich, F. F., Z. S. Spakovszky, M. Martinez-Sanchez, S. J. Song, D. C. Wisler, A. F. Storace, H. W. Shin, and B. F. Beacher. "Unsteady Flow and Whirl-Inducing Forces in Axial-Flow Compressors: Part II—Analysis." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 446–52. http://dx.doi.org/10.1115/1.1370165.
Full textAbstract:
An experimental and theoretical investigation was conducted to evaluate the effects seen in axial-flow compressors when the centerline of the rotor becomes displaced from the centerline of the static structure of the engine, thus creating circumferentially nonuniform rotor-tip clearances. This displacement produces unsteady flow and creates a system of destabilizing forces, which contribute significantly to rotor whirl instability in turbomachinery. These forces were first identified by Thomas (1958. Bull. AIM, 71, No. 11/12, pp. 1039–1063.) for turbines and by Alford (1965. J. Eng. Power, Oct., pp. 333–334) for jet engines. In Part I, the results from an experimental investigation of these phenomena were presented. In this Part II, three analytic models were used to predict both the magnitude and direction of the Thomas/Alford force in its normalized form, known as the β coefficient, and the unsteady effects for the compressors tested in Part I. In addition, the effects of a whirling shaft were simulated to evaluate differences between a rotor with static offset and an actual whirling eccentric rotor. The models were also used to assess the influence of the nonaxisymmetric static pressure distribution on the rotor spool, which was not measured in the experiment. The models evaluated were (1) the two-sector parallel compressor (2SPC) model, (2) the infinite-segment-parallel-compressor (ISPC) model, and (3) the two-coupled actuator disk (2CAD) model. The results of these analyses were found to be in agreement with the experimental data in both sign and trend. Thus, the validated models provide a general means to predict the aerodynamic destabilizing forces for axial flow compressors in turbine engines. These tools have the potential to improve the design of rotordynamically stable turbomachinery.
15
SINGH, Prabhat, and Dharmahinder Singh CHAND. "VIGV Angle Optimization for Subsonic Axial Flow Compressor." INCAS BULLETIN 12, no. 2 (June 5, 2020): 217–28. http://dx.doi.org/10.13111/2066-8201.2020.12.2.18.
Full textAbstract:
Long-term performance of an axial-flow compressor by the scheduling of inlet guide vanes at off-design conditions has been studied in this paper. The compressors are used in various industries and in aviation sectors for different operations at different climatic conditions; due to diverse climatic conditions the compressor is unable to give good performances as expected. At the design conditions, the results show that the variations in total pressure loss coefficient, volume flow rate, pressure ratio, and others like thermodynamics, aerodynamic properties are different at different stagger angles of 15°, 30° and 45°. BladeGen tools were used to design the inlet guide vanes for the investigations. The performance parameters of the axial flow compressor were analyzed by using ANSYS CFX and were validated using the analytical method. The objective of this study is to minimize the total pressure loss coefficient and improve the aerodynamic characteristics of an inlet guide vanes at different stagger angles and hence reduce the overall fuel consumption. The outcomes of this work give an improved insight into the efficient use of a VIGV in axial flow compressor.
16
Storer, J. A., and N. A. Cumpsty. "Tip Leakage Flow in Axial Compressors." Journal of Turbomachinery 113, no. 2 (April 1, 1991): 252–59. http://dx.doi.org/10.1115/1.2929095.
Full textAbstract:
Experimental measurements in a linear cascade with tip clearance are complemented by numerical solutions of the three-dimensional Navier–Stokes equations in an investigation of tip leakage flow. Measurements reveal that the clearance flow, which separates near the entry of the tip gap, remains unattached for the majority of the blade chord when the tip clearance is similar to that typical of a machine. The numerical predictions of leakage flow rate agree very well with measurements, and detailed comparisons show that the mechanism of tip leakage is primarily inviscid. It is demonstrated by simple calculation that it is the static pressure field near the end of the blade that controls chordwise distribution of the flow across the tip. Although the presence of a vortex caused by the roll-up of the leakage flow may affect the local pressure field, the overall magnitude of the tip leakage flow remains strongly related to the aerodynamic loading of the blades.
17
Wennerstrom, A. J. "Low Aspect Ratio Axial Flow Compressors: Why and What It Means." Journal of Turbomachinery 111, no. 4 (October 1, 1989): 357–65. http://dx.doi.org/10.1115/1.3262280.
Full textAbstract:
One of the more visible changes that has occurred in fans and compressors for aircraft turbine engines that have entered development since about 1970 has been a significant reduction in the aspect ratio of the blading. This has brought with it a greatly reduced engine parts count and improved ruggedness and aeroelastic stability. This paper traces the evolution of thinking concerning appropriate aspect ratios for axial flow compressors since the early years of the aircraft turbine engine. In the 1950’s, moderate aspect ratios were favored for reasons of mechanical design. As mechanical design capability became more sophisticated, several attempts were made, primarily in the 1960s, to employ very high aspect ratios to reduce engine size and weight. Four of these programs are described that were largely unsuccessful for both mechanical and aerodynamic reasons. After 1970, the pendulum swung strongly in the other direction and designs of very low aspect ratio began to emerge. This has had a significant impact on compressor design systems, and a number of the ways in which design systems have been affected are discussed. Some concluding remarks are made concerning the author’s opinion of trends in the near future in aerodynamic design technology.
18
Mustaffa, Ahmad Fikri, and Vasudevan Kanjirakkad. "Stall margin improvement in a low-speed axial compressor rotor using a blockage-optimised single circumferential casing groove." Journal of the Global Power and Propulsion Society 5 (May 3, 2021): 79–89. http://dx.doi.org/10.33737/jgpps/133912.
Full textAbstract:
The stall margin of tip-critical axial compressors can be improved by using circumferential casing grooves. From previous studies, in the literature, the stall margin improvement due to the casing grooves can be attributed to the reduction of the near casing blockage. The pressure rise across the compressor as the compressor is throttled intensifies the tip leakage flow. This results in a stronger tip leakage vortex that is thought to be the main source of the blockage. In this paper, the near casing blockage due to the tip region aerodynamics in a low-speed axial compressor rotor is numerically studied and quantified using a mass flow-based blockage parameter. The peak blockage location at the last stable operating point for a rotor with smooth casing is found to be at about 10% of the tip chord aft of the tip leading edge. Based on this information, an optimised single casing groove design that minimises the peak blockage is found using a surrogate-based optimisation approach. The implementation of the optimised groove is shown to produce a stall margin improvement of about 5%.
19
Mailach, Ronald, and Konrad Vogeler. "Unsteady Aerodynamic Blade Excitation at the Stability Limit and During Rotating Stall in an Axial Compressor." Journal of Turbomachinery 129, no. 3 (July 25, 2006): 503–11. http://dx.doi.org/10.1115/1.2720486.
Full textAbstract:
The stable operating range of axial compressors is limited by the onset of rotating stall and surge. These flow conditions endanger the reliability of operation and definitely have to be avoided in compressors of gas turbines. However, there is still a need to improve the physical understanding of these flow phenomena to prevent them while utilizing the maximum available working potential of the compressor. This paper discusses detailed experimental investigations of the rotating stall onset with the main emphasis on the aerodynamic blade excitation in the Dresden four-stage low-speed research compressor. The stall inception, which is triggered by modal waves, as well as the main flow features during rotating stall operation are discussed. To investigate the unsteady pressure distributions, both the rotor and the stator blades of the first stage were equipped with piezoresistive pressure transducers. Based on these measurements the unsteady blade pressure forces are calculated. Time-resolved results at the stability limit as well as during rotating stall are presented. For all operating conditions rotor–stator interactions play an important role on the blade force excitation. Furthermore the role of the inertia driven momentum exchange at the stall cell boundaries on the aerodynamic blade force excitation is pointed out.
20
Leichtfuß, Sebastian, Johannes Bühler, Heinz-Peter Schiffer, Patrick Peters, and Michael Hanna. "A Casing Treatment with Axial Grooves for Centrifugal Compressors." International Journal of Turbomachinery, Propulsion and Power 4, no. 3 (August 16, 2019): 27. http://dx.doi.org/10.3390/ijtpp4030027.
Full textAbstract:
This paper provides an investigation of a casing treatment (CT) approach for pressure ratio improvements of centrifugal compressors between peak efficiency and surge. Results were experimentally verified for a variety of automotive turbocharger compressors and analyzed with 3D CFD. The CT design is an adaptation from an axial high-pressure compressor, which was successfully applied and intensively investigated in recent years. The aerodynamic working principle of the applied CT design and the achievable improvements are shown and described. The demand of operating range for automotive applications typically dictates high inlet shroud to outlet radius ratio (high trim) and past experiences indicate that a recirculation zone is formed in the inducer for those centrifugal compressors. This recirculation at the inlet shroud causes losses, a massive blockage and induces a co-rotating swirl at the inlet of the impeller. The result is a reduced pressure ratio, often leading to flat speed lines between the onset of recirculation and surge. This paper provides an understanding of inducer recirculation, its impacts and suggests countermeasures. The CT design for centrifugal compressors only influences flow locally at the inducer and prevents recirculation. It differs substantially in design and functionality from the classical bleed slot system commonly used to increase operating range. An experimental and CFD comparison between these designs is presented. While the classical bleed slot system often provides a massive increase in operating range, it often fails to increase the pressure ratio between onset of inducer recirculation and surge. In contrast, the CT design achieves a gain in pressure ratio near surge.
21
Sehra, A., J. Bettner, and A. Cohn. "Design of a High-Performance Axial Compressor for Utility Gas Turbine." Journal of Turbomachinery 114, no. 2 (April 1, 1992): 277–86. http://dx.doi.org/10.1115/1.2929141.
Full textAbstract:
An aerodynamic design study to configure a high-efficiency industrial-size gas turbine compressor is presented. This study was conducted using an advanced aircraft engine compressor design system. Starting with an initial configuration based on conventional design practice, compressor design parameters were progressively optimized. To improve the efficiency potential of this design further, several advanced design concepts (such as stator ends bends and velocity controlled airfoils) were introduced. The projected poly tropic efficiency of the final advanced concept compressor design having 19 axial stages was estimated at 92.8 percent, which is 2 to 3 percent higher than the current high-efficiency aircraft turbine engine compressors. The influence of variable geometry on the flow and efficiency (at design speed) was also investigated. Operation at 77 percent design flow with inlet guide vanes and front five variable stators is predicted to increase the compressor efficiency by 6 points as compared to conventional designs having only the inlet guide vane as variable geometry.
22
Lu, Hanan, Qiushi Li, Tianyu Pan, and Ramesh Agarwal. "Analysis and application of shroud wall optimization to an axial compressor with upstream boundary layer to improve aerodynamic performance." International Journal of Numerical Methods for Heat & Fluid Flow 29, no. 11 (November 4, 2019): 4237–61. http://dx.doi.org/10.1108/hff-01-2019-0071.
Full textAbstract:
PurposeFor an axial-flow compressor rotor, the upstream inflow conditions will vary as the aircraft faces harsh flight conditions (such as taking off, landing or maneuvering) or the whole compressor operates at off-design conditions. With the increase of upstream boundary layer thickness, the rotor blade tip will be loaded and the increased blade load will deteriorate the shock/boundary layer interaction and tip leakage flows, resulting in high aerodynamic losses in the tip region. The purpose of this paper is to achieve a better flow control for tip secondary flows and provide a probable design strategy for high-load compressors to tolerate complex upstream inflow conditions.Design/methodology/approachThis paper presents an analysis and application of shroud wall optimization to a typical transonic axial-flow compressor rotor by considering the inlet boundary layer (IBL). The design variables are selected to shape the shroud wall profile at the tip region with the purpose of controlling the tip leakage loss and the shock/boundary layer interaction loss. The objectives are to improve the compressor efficiency at the inlet-boundary-layer condition while keeping its aerodynamic performance at the uniform condition.FindingsAfter the optimization of shroud wall contour, aerodynamic benefits are achieved mainly on two aspects. On the one hand, the shroud wall optimization has reduced the intensity of the tip leakage flow and the interaction between the leakage and main flows, thereby decreasing the leakage loss. On the other hand, the optimized shroud design changes the shock structure and redistributes the shock intensity in the spanwise direction, especially weakening the shock near the tip. In this situation, the shock/boundary layer interaction and the associated flow separations and wakes are also eliminated. On the whole, at the inlet-boundary-layer condition, the compressor with optimized shroud design has achieved a 0.8 per cent improvement of peak efficiency over that with baseline shroud design without sacrificing the total pressure ratio. Moreover, the re-designed compressor also maintains the aerodynamic performance at the uniform condition. The results indicate that the shroud wall profile has significant influences on the rotor tip losses and could be properly designed to enhance the compressor aerodynamic performance against the negative impacts of the IBL.Originality/valueThe originality of this paper lies in developing a shroud wall contour optimization design strategy to control the tip leakage loss and the shock/boundary layer interaction loss in a transonic compressor rotor. The obtained results could be beneficial for transonic compressors to tolerate the complex upstream inflow conditions.
23
Bozza, F., A. Senatore, and R. Tuccillo. "Thermal Cycle Analysis and Component Aerodesign for Gas Turbine Concept in Low-Range Cogenerating Systems." Journal of Engineering for Gas Turbines and Power 118, no. 4 (October 1, 1996): 792–802. http://dx.doi.org/10.1115/1.2816995.
Full textAbstract:
The authors link together their previous experiences in gas turbine plant analysis and aerodynamic design of radial flow compressors. In recent papers they have introduced a method for the performance estimation of gas turbine engines, based on the prediction of the matching conditions among the several components in the whole operating range. On the other hand they have expressly paid attention to the problem of optimal design of radial flow compressors for satisfactory operation within an assigned operating range. In this paper, the authors present an integrated method, which aims to define the optimal characteristics of a low-power gas turbine engine (i.e., in the range 500–2000 kW). In this case, the radial compressor performance plays an important role as regards gas turbine operation for both power generation and cogeneration applications. The analysis proceeds with the optimization of rotating components (i.e., radial compressor and axial flow turbine) for given thermal cycle parameters. The prescribed objectives of the optimizing procedure are related to performance levels not only at the reference design conditions but also throughout the operating field. A particular emphasis is given to the extension of the field of satisfactory performance for cogeneration applications, with best fitting of mechanical and thermal power requirements. The aerodynamic design of radial flow compressor utilizes a method based on genetic algorithms.
24
Song, Wei, Xi De Lai, Wei Zhang, Zhen Lu, and Xiao Ming Chen. "Research on Reverse Engineering for Blades of Axial Flow Compressors Based on Parameterization." Advanced Materials Research 813 (September 2013): 3–6. http://dx.doi.org/10.4028/www.scientific.net/amr.813.3.
Full textAbstract:
To acquire the digital model of axial compressors on the actual projects, a Reverse Engineering procedure of the blade which has high accuracy and efficiency based on Parametric Modeling was developed. To meet the requirements of geometric characteristics and aerodynamic optimization design and accuracy of overall blade, a method of blade-measuring path planning based on curves of cross sections and blade body surfaces was presented. For solving the problems of difficulty in the definition and fitting in primary parametric surfaces and connections between them, a parametric surface definition method of typical axial compressors blade based on blade geometry was presented and the connections were fit with a process of primary parametric surface boundary-splicing surface generation. Finally, an example of Reverse Engineering for blades of axial flow compressors based on parameterization was given.
25
Zheng, Xin-qian, Xiao-bo Zhou, and Sheng Zhou. "Investigation on a Type of Flow Control to Weaken Unsteady Separated Flows by Unsteady Excitation in Axial Flow Compressors." Journal of Turbomachinery 127, no. 3 (March 1, 2004): 489–96. http://dx.doi.org/10.1115/1.1860572.
Full textAbstract:
By solving unsteady Reynolds-averaged Navier-Stokes equations discretized by a high-order scheme, the results showed that the disordered unsteady separated flow could be effectively controlled by periodic suction and blowing in a wide range of incidences, resulting in enhancement of time-averaged aerodynamic performances of an axial compressor cascade. The effects of unsteady excitation frequency, amplitude, and excitation location were investigated in detail. The effective excitation frequency spans a wide spectrum, and there is an optimal excitation frequency that is nearly equal to the characteristic frequency of vortex shedding. Excitation amplitude exhibits a threshold value (nearly 10% in terms of the ratio of maximum velocity of periodic suction and blowing to the velocity of free flow) and an optimal value (nearly 35%). The optimal excitation location is just upstream of the separation point. We also explored feasible unsteady actuators by utilizing the upstream wake for constraining unsteady separation in axial flow compressors.
26
Dunham, J. "A New Endwall Model for Axial Compressor Throughflow Calculations." Journal of Turbomachinery 117, no. 4 (October 1, 1995): 533–40. http://dx.doi.org/10.1115/1.2836565.
Full textAbstract:
It is well recognized that the endwall regions of a compressor—in which the annulus wall flow interacts with the mainstream flow—have a major influence on its efficiency and surge margin. Despite many attempts over the years to predict the very complex flow patterns in the endwall regions, current compressor design methods still rely largely on empirical estimates of the aerodynamic losses and flow angle deviations in these regions. This paper describes a new phenomenological model of the key endwall flow phenomena treated in a circumferentially averaged way. It starts from Hirsch and de Ruyck’s annulus wall boundary layer approach, but makes some important changes. The secondary vorticities arising from passage secondary flows and from tip clearance flows are calculated. Then the radial interchanges of momentum, energy, and entropy arising from both diffusion and convection are estimated. The model is incorporated into a streamline curvature program. The empirical blade force defect terms in the boundary layers are selected from cascade data. The effectiveness of the method is illustrated by comparing the predictions with experimental results on both low-speed and high-speed multistage compressors. It is found that the radial variation of flow parameters is quite well predicted, and so is the overall performance, except when significant endwall stall occurs.
27
Liang, Dong, Wenjie Wang, and Peter J. Thomas. "Aerodynamic analysis of nonuniform trailing edge blowing." Aircraft Engineering and Aerospace Technology 91, no. 1 (January 7, 2018): 134–44. http://dx.doi.org/10.1108/aeat-04-2018-0115.
Full textAbstract:
Purpose Numerical and experimental results for different oncoming base-flow conditions indicate that nonuniform trailing edge blowing (NTEB) can expand the performance range of compressors and reduce the thrust on the rotor, while the efficiency of the compressor can be improved by more than 2 per cent. Design/methodology/approach Relevant aerodynamic parameters, such as total pressure, ratio of efficiency and axial thrust, are calculated and analyzed under conditions with and without NTEB. Measurements are performed downstream of two adjacent stator blades, at seven equidistantly spaced reference locations. The experimental measurement of the interstage flow field used a dynamic four-hole probe with phase lock technique. Findings An axial low-speed single-stage compressor was established with flow field measurement system and nonuniform blowing system. NTEB was studied by means of numerical simulations and experiments, and it is found that the efficiency of the tested compressor can be improved by more than 2 per cent. Originality/value Unlike most of the previous research studies which mainly focused on the rotor/stator interaction and trailing edge uniform blowing, the research results summarized in the current paper on the stator/rotor interaction used inlet guide vanes for steady and unsteady calculations. An active control of the interstage flow field in a low-speed compressor was used to widen the working range and improve the performance of the compressor.
28
Calvert, W. J., and R. B. Ginder. "Transonic fan and compressor design." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 213, no. 5 (May 1, 1999): 419–36. http://dx.doi.org/10.1243/0954406991522671.
Full textAbstract:
Transonic fans and compressors are now widely used in gas turbine engines because of their benefits in terms of compactness and reduced weight and cost. However, careful and precise design is essential if high levels of performance are to be achieved. In this paper, the evolution of transonic compressor designs and methods is outlined, followed by more detailed descriptions of current compressor configurations and requirements and modern aerodynamic design methods and philosophies. Current procedures employ a range of methods to allow the designer to refine a new design progressively. Overall parameters, such as specific flow and mean stage loading, the axial matching between the stages at key operating conditions and the radial matching between the blade rows are set in turn, using one- and two-dimensional techniques. Finally, detailed quasi-three-dimensional and three-dimensional computational fluid dynamics (CFD) analyses are employed to refine the design. However, it is important to appreciate that the methods all have significant limitations and designers must take this into account if successful compressors are to be produced.
29
Shan, Peng. "Kinematic Analysis of 3-D Swept Shock Surfaces in Axial Flow Compressors." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 490–500. http://dx.doi.org/10.1115/1.1370159.
Full textAbstract:
This paper is part II of a comprehensive study on the blade leading edge sweep/bend of supersonic and transonic axial compressors. The paper explores and analyzes the kinematic characteristic variables of three-dimensional (3-D) swept shock surfaces. In the research field studying the sweep aerodynamics of axial flow compressors and fans, many types of high loading swept blades are under intensive study. So, in both direct and inverse design methods and experimental validations, an accurate grasp of the sweep characteristic of the blade’s 3-D swept shock surface becomes of more concern than before. Associated with relevant blading variables, this paper studies the forward and zero and backward sweeps of shock surfaces, defines and resolves every kind of useful sweep angle, obtains dimensionless sweep similarity factors, suggests a kind of method for the quantitative classification of 3-D shock structures, and proposes the principle of 3-D shock structure measurements. Two rotor blade leading edge shock surfaces from two high loading single stage fans are analyzed and contrasted. This study is the foundation of the kinematic design of swept shock surfaces.
30
Law, C. H., and A. J. Wennerstrom. "Performance of Two Transonic Axial Compressor Rotors Incorporating Inlet Counterswirl." Journal of Turbomachinery 109, no. 1 (January 1, 1987): 142–48. http://dx.doi.org/10.1115/1.3262059.
Full textAbstract:
A single-stage axial-flow compressor which incorporates rotor inlet counterswirl to improve stage performance is discussed. Results for two rotor configurations are presented, including design and experimental test data. In this compressor design, inlet guide vanes were used to add counterswirl to the inlet of the rotor. The magnitude of the counterswirl was radially distributed to maximize the overall stage efficiency by minimizing the rotor combined losses (diffusion losses and shock losses). The shock losses were minimized by simultaneously optimizing the rotor blade section geometry, through-blade static pressure distribution, and leading edge aerodynamic/geometric shock sweep angles. Results from both the design and experimental performance analyses are presented and comparisons are made between the experimental data and the analyses and between the performance of both rotor designs. The computation of the flow field for both rotor designs and for the analysis of both tests was performed in an identical fashion using an axisymmetric, streamline-curvature-type code. Results presented include tip section blade-to-blade static pressure distributions and rotor through-blade and exit distributions of various aerodynamic parameters. The performance of this compressor stage represents a significant improvement in axial compressor performance compared to previous attempts to use rotor inlet counterswirl and to current, more conventional, state-of-the-art axial compressors operating under similar conditions.
31
Aungier,, RH, and S. Farokhi,. "Axial-Flow Compressors: A Strategy for Aerodynamic Design and Analysis." Applied Mechanics Reviews 57, no. 4 (July 1, 2004): B22. http://dx.doi.org/10.1115/1.1786589.
Full text
32
Sayed Hassan, Ahmed. "ACTIVE SUPPRESSION OF AERODYNAMIC INSTABILITIES IN THE AXIAL FLOW COMPRESSORS." JES. Journal of Engineering Sciences 34, no. 6 (November 1, 2006): 1843–64. http://dx.doi.org/10.21608/jesaun.2006.111180.
Full text
33
Lastivka, Ivan. "RAISING GAS-DYNAMIC STABILITY MARGIN OF AXIAL AND CENTRIFUGAL COMPRESSOR STAGES BY MEANS OF VANED DIFFUSER BOUNDARY LAYER CONTROL / DUJŲ DINAMINIO STABILUMO RIBOS AŠINIO IR IŠCENTRINIO KOMPRESORIAUS PAKOPOSE DIDINIMAS ATLIEKANT STABILIZUOTO DIFUZORIAUS PARIBIO SLUOKSNIO KONTROLĘ." Aviation 15, no. 3 (October 4, 2011): 76–81. http://dx.doi.org/10.3846/16487788.2011.624262.
Full textAbstract:
Generalised research results that consider the upgradability of axial and centrifugal gas turbine engine compressors by means of gas-dynamic boundary layer control on bladed disks are demonstrated. Active and passive methods are used. Comparative analysis of the results has been carried out. The analysis is purposed to determine the influence of the flow circulation around the aerofoils on the performance of compressor single-row bladed disks with smooth blades and rough blades and under the condition that vortex generators are installed. An increase in the efficiency of aviation gas-turbine engines and in their gas-dynamic stability margin support leads to the enhancement of the parameters and performance of compressors: increase in loading of aerodynamic bladed disks, improvement of their economical efficiency, improvement of margin of the continuous flow around the compressor grids, etc. Airflow in the compressor grid is characterised by the flow region in the flow core and also by the flow regions in the wall boundary layers on the grid blades where shock waves, vortices, air swirls, and flow separation phenomena take place. The principle objective of the work is to research the possibilities of influence on the parameters of the elements of compressors and overall performance of gas-turbine engines via the methods of active and passive flow regulation. Active flow regulation is realised either by rendering the auxiliary gas mass to the blades boundary layer, or by suction (withdrawal) of the boundary layer (its part) from the surfaces of blades. Passive flow around regulation is characterised by influence on the boundary layer by means of energy redistribution in the flow itself. Santrauka Šiuo tyrimu siekiama nustatyti sparno profilio aptekejimo įtaką vienos eiles menčių kompresoriaus su lygiomis ir šiurkščiomis mentemis darbui, esant įdiegtiems sūkurio generatoriams. Pagrindinis darbo tikslas – ištirti kompresoriaus elementų ir bendro dujų turbininių variklių darbo įtaką parametrams, taikant pasyvų ir aktyvų srauto reguliavimo metodus. Padidinus dujų turbininių variklių našumą ir jų dujų dinamikos stabilumo ribas, pagereja kompresorių darbas ir parametrai: padideja aerodinaminių diskų su mentemis apkrova, jie tampa ekonomiškai našesni, padideja nepertraukiamo srauto riba aplink kompresoriaus plokšteles.
34
Vahdati, M., A. I. Sayma, M. Imregun, and G. Simpson. "Multibladerow Forced Response Modeling in Axial-Flow Core Compressors." Journal of Turbomachinery 129, no. 2 (June 3, 2005): 412–20. http://dx.doi.org/10.1115/1.2436892.
Full textAbstract:
This paper describes the formulation and application of an advanced numerical model for the simulation of blade-passing and low-engine order forced response in turbomachinery core compressors. The Reynolds averaged Navier–Stokes equations are used to represent the flow in a nonlinear time-accurate fashion on unstructured meshes of mixed elements. The structural model is based on a standard finite-element representation. The fluid mesh is moved at each time step according to the structural motion so that changes in blade aerodynamic damping and flow unsteadiness can be accommodated automatically. A whole-annulus 5-bladerow forced response calculation, where three upstream and one downstream bladerows were considered in addition to the rotor bladerow of interest, was undertaken using over 20 million grid points. The results showed not only some potential shortcomings of equivalent 2-bladerow computations for the determination of the main blade-passing forced response, but also revealed the potential importance of low engine-order harmonics. Such harmonics, due to stator blade number differences, or arising from common symmetric sectors, can only be taken into account by including all stator bladerows of interest. The low engine-order excitation that could arise from a blocked passage was investigated next. It was shown that high vibration response could arise in such cases.
35
Benichou, Emmanuel, Nicolas Binder, Yannick Bousquet, and Xavier Carbonneau. "Improvement of the Parallel Compressor Model and Application to Inlet Flow Distortion." International Journal of Turbomachinery, Propulsion and Power 6, no. 3 (August 25, 2021): 34. http://dx.doi.org/10.3390/ijtpp6030034.
Full textAbstract:
This paper introduces a semi-analytical approach which enables one to deal with distorted inflow in axial fans or compressors. It is inspired by the classical parallel compressor (PC) theory but relies on a local flow-loading coefficient formalism. It is applied to non-uniform flow conditions to study the aerodynamic behavior of a low-speed fan in response to upstream flow distortion. Experimental measurements and 3D RANS simulations are used to evaluate the prediction of fan performance obtained with the local PC method. The comparison proves that, despite its simplicity, the present approach enables to correctly capture first order phenomena, offering interesting perspectives for an early design phase if different fan geometries are to be tested and if the upstream distortion maps are available.
36
Niccolini Marmont Du Haut Champ, Carlo Alberto, Aristide Fausto Massardo, Mario Luigi Ferrari, and Paolo Silvestri. "Surge prevention in gas turbines: an overview over historical solutions and perspectives about the future." E3S Web of Conferences 113 (2019): 02003. http://dx.doi.org/10.1051/e3sconf/201911302003.
Full textAbstract:
The aim of the present work is to retrace experimental, analytical and numerical analyses which deal with compressor instability phenomena, such as rotating stall and surge. While the first affects only the machine itself, the second involves the whole energy system. Surge onset is characterized by strong pressure and mass flow rate fluctuations which can even lead to reverse flow through the compressor. Experimental studies on prevention of axial compressor fluid dynamic instabilities, which can be propagated and eventually damage the solid structure, have been carried out by many authors. The first important studies on this topic tried to underline the main aspects of the complex detailed mechanism of surge, by replacing the compression system with an equivalent conceptual lumped parameter model. This is specially meant to capture the unsteady behaviour and the transient response of the compression system itself, particularly its dependence on variations in the volume of discharge downstream and in the settings of the throttle valve at its outlet (which simulates the actual load coupled to the compressor). Greitzer’s model is still regarded as the milestone for new investigations about active control and stabilization of surge and, more generally, about active suppression of aerodynamic instabilities in turbomachinery. During the last years, a lot of simulations and experimental studies about surge have been conducted on multistage centrifugal compressors with different architectures (e.g. equipped with vaneless or vaned diffusers). Moreover, further kinds of analysis try to extend the stable working zone of compressors, identifying stall and surge precursors extractable from information contained in the vibro-acoustical and rotodynamic response of the system.
37
Cheng, Jinxin, Shengfeng Zhao, Zhaohui Dong, and Chengwu Yang. "Design Optimization of a Multistage Axial Flow Compressor Based on Full-Blade Surface Parameterization and Phased Strategy." International Journal of Aerospace Engineering 2021 (July 1, 2021): 1–25. http://dx.doi.org/10.1155/2021/5583681.
Full textAbstract:
A new design optimization method is proposed for the problem of high-precision aerodynamic design of multistage axial compressors. The method mainly contains three aspects: full-blade surface parametrization can significantly reduce the number of control variables per blade row and increase the degrees of freedom of the leading edge blade angle compared with the traditional semiblade parametric method; secondly, the artificial bee colony algorithm improved initialization and food source exploration and exploitation mechanism to enhance the global optimization ability and convergence speed, and a distributed optimization system is built on the supercomputing platform based on this method; finally, a phased optimization strategy based on the “synchronous change in multirow blades” is proposed, and expert experience is introduced to achieve a better balance between exploration and exploitation. The optimization method is applied to the AL-31F four-stage low-pressure compressor. As a result, the adiabatic efficiency is improved by 0.67% and the surge margin is improved by 3.1% under the premise that the total pressure ratio and mass flow rate satisfy the constraints, which verifies the effectiveness and engineering practicality of the proposed optimization method in the field of multistage axial flow compressor aerodynamic optimization.
38
Spakovszky, Z. S. "Analysis of Aerodynamically Induced Whirling Forces in Axial Flow Compressors." Journal of Turbomachinery 122, no. 4 (February 1, 2000): 761–68. http://dx.doi.org/10.1115/1.1312801.
Full textAbstract:
A new analytical model to predict the aerodynamic forces in axial flow compressors due to asymmetric tip-clearance is introduced. The model captures the effects of tip-clearance induced distortion (i.e., forced shaft whirl), unsteady momentum-induced tangential blade forces, and pressure-induced forces on the spool. Pressure forces are shown to lag the tip-clearance asymmetry, resulting in a tangential (i.e., whirl-inducing) force due to spool pressure. This force can be of comparable magnitude to the classical Alford force. Prediction and elucidation of the Alford force is also presented. In particular, a new parameter denoted as the blade loading indicator is deduced. This parameter depends only on stage geometry and mean flow and determines the direction of whirl tendency due to tangential blade loading forces in both compressors and turbines. All findings are suitable for incorporation into an overall dynamic system analysis and integration into existing engine design tools. [S0889-504X(00)01604-4]
39
He, Chen, Yunfei Ma, Xiaohua Liu, Dakun Sun, and Xiaofeng Sun. "Aerodynamic Instabilities of Swept Airfoil Design in Transonic Axial-Flow Compressors." AIAA Journal 56, no. 5 (May 2018): 1878–93. http://dx.doi.org/10.2514/1.j056053.
Full text
40
Escuret, J. F., and V. Garnier. "Stall Inception Measurements in a High-Speed Multistage Compressor." Journal of Turbomachinery 118, no. 4 (October 1, 1996): 690–96. http://dx.doi.org/10.1115/1.2840924.
Full textAbstract:
This paper presents unsteady measurements taken in a high-speed four-stage aeroengine compressor prior to the onset of aerodynamic flow instabilities. In this experiment, 40 fast-response pressure transducers have been located at various axial and circumferential positions throughout the machine in order to give a very detailed picture of stall inception. At all the compressor speeds investigated, the stall pattern observed is initiated by a very short length-scale finite-amplitude disturbance, which propagates at a fast rate around the annulus. This initial stall cell leads to a large-amplitude system instability in less than five rotor revolutions. Varying the IGV setting angle is found to have a strong influence on the axial location of the first disturbance detected. In particular, transferring the aerodynamic loading from front to downstream stages moves the first disturbance detected from the first to the last stage of the compressor. Other repeatable features of the stall inception pattern in this compressor have been identified using a simple analysis technique particularly appropriate to the study of short length-scale disturbances. It is found that the origins of instabilities are tied to particular tangential positions in both the stationary and rotating frames of reference. These measurements lead to the conclusion that the stall inception process in high-speed multistage compressors can be characterized by some very local and organized flow phenomena. Moreover, there is no evidence of prestall waves in this compressor.
41
Taylor, Derek, and John Longley. "Effects of Stator Platform Geometry Features on Blade Row Performance." Journal of the Global Power and Propulsion Society 3 (September 17, 2019): 609–29. http://dx.doi.org/10.33737/jgpps/111508.
Full textAbstract:
This paper details an experimental investigation, using a linear cascade, into the effects of real geometry features on the aerodynamic performance of stator blade rows within axial flow compressors. The specific geometric features investigated include shroud cavities, inter-platform gaps, vane-pack gaps and the effects of misalignment of the platform endwalls due to manufacturing tolerances. A computational investigation into these effects is also included. To ensure that the linear cascade measurements are representative of a multi-stage compressor environment a novel experimental technique was developed to generate a hub endwall boundary layer which had skew. The boundary layer skew generation method involves injecting flow along the cascade endwall in such a manner as to control both the displacement thickness and tangential momentum thickness of the resulting boundary layer. Without the presence of the endwall boundary layer skew the linear cascade could not reproduce the flow features typically observed in a multi-stage compressor. The investigation reveals that real geometry features can have a significant impact on the flowfield within a blade passage. For a shrouded stator, increasing the leakage flow rate increases the stagnation pressure loss coefficient. However, high levels of whirl pickup of the leakage flow as it passes through the stator-shroud cavity can offset the natural secondary flow within the stator passage and thus reduce the stagnation pressure loss. All of the steps and gaps that were observed to be present in real compressors were found to increase the stagnation pressure loss relative to that of a smooth endwall. It is also shown that the computational method is capable of capturing the trends observed in the experiments.
42
Halstead, D. E., D. C. Wisler, T. H. Okiishi, G. J. Walker, H. P. Hodson, and H. W. Shin. "Boundary Layer Development in Axial Compressors and Turbines: Part 1 of 4—Composite Picture." Journal of Turbomachinery 119, no. 1 (January 1, 1997): 114–27. http://dx.doi.org/10.1115/1.2841000.
Full textAbstract:
Comprehensive experiments and computational analyses were conducted to understand boundary layer development on airfoil surfaces in multistage, axial-flow compressors and LP turbines. The tests were run over a broad range of Reynolds numbers and loading levels in large, low-speed research facilities which simulate the relevant aerodynamic features of modern engine components. Measurements of boundary layer characteristics were obtained by using arrays of densely packed, hot-film gauges mounted on airfoil surfaces and by making boundary layer surveys with hot wire probes. Computational predictions were made using both steady flow codes and an unsteady flow code. This is the first time that time-resolved boundary layer measurements and detailed comparisons of measured data with predictions of boundary layer codes have been reported for multistage compressor and turbine blading. Part 1 of this paper summarizes all of our experimental findings by using sketches to show how boundary layers develop on compressor and turbine blading. Parts 2 and 3 present the detailed experimental results for the compressor and turbine, respectively. Part 4 presents computational analyses and discusses comparisons with experimental data. Readers not interested in experimental detail can go directly from Part 1 to Part 4. For both compressor and turbine blading, the experimental results show large extents of laminar and transitional flow on the suction surface of embedded stages, with the boundary layer generally developing along two distinct but coupled paths. One path lies approximately under the wake trajectory while the other lies between wakes. Along both paths the boundary layer clearly goes from laminar to transitional to turbulent. The wake path and the non-wake path are coupled by a calmed region, which, being generated by turbulent spots produced in the wake path, is effective in suppressing flow separation and delaying transition in the non-wake path. The location and strength of the various regions within the paths, such as wake-induced transitional and turbulent strips, vary with Reynolds number, loading level, and turbulence intensity. On the pressure surface, transition takes place near the leading edge for the blading tested. For both surfaces, bypass transition and separated-flow transition were observed. Classical Tollmien–Schlichting transition did not play a significant role. Comparisons of embedded and first-stage results were also made to assess the relevance of applying single-stage and cascade studies to the multistage environment. Although doing well under certain conditions, the codes in general could not adequately predict the onset and extent of transition in regions affected by calming. However, assessments are made to guide designers in using current predictive schemes to compute boundary layer features and obtain reasonable loss predictions.
43
Huang, Song, Jinxin Cheng, Chengwu Yang, Chuangxin Zhou, Shengfeng Zhao, and Xingen Lu. "Optimization Design of a 2.5 Stage Highly Loaded Axial Compressor with a Bezier Surface Modeling Method." Applied Sciences 10, no. 11 (June 1, 2020): 3860. http://dx.doi.org/10.3390/app10113860.
Full textAbstract:
Due to the complexity of the internal flow field of compressors, the aerodynamic design and optimization of a highly loaded axial compressor with high performance still have three problems, which are rich engineering design experience, high dimensions, and time-consuming calculations. To overcome these three problems, this paper takes an engineering-designed 2.5-stage highly loaded axial flow compressor as an example to introduce the design process and the adopted design philosophies. Then, this paper verifies the numerical method of computational fluid dynamics. A new Bezier surface modeling method for the entire suction surface and pressure surface of blades is developed, and the multi-island genetic algorithm is directly used for further optimization. Only 32 optimization variables are used to optimize the rotors and stators of the compressor, which greatly overcome the problem of high dimensions, time-consuming calculations, and smooth blade surfaces. After optimization, compared with the original compressor, the peak efficiency is still improved by 0.12%, and the stall margin is increased by 2.69%. The increase in peak efficiency is mainly due to the rotors. Compared with the original compressor, for the second-stage rotor, the adiabatic efficiency is improved by about 0.4%, which is mainly due to the decreases of total pressure losses in the range of above 30% of the span height and 10%–30% of the chord length. Besides, for the original compressor, due to deterioration of the flow field near the tip region of the second-stage stator, the large low-speed region eventually evolves from corner separation into corner stall with three-dimensional space spiral backflow. For the optimized compressor, the main reason for the increased stall margin is that the flow field of the second-stage stator with a span height above 50% is improved, and the separation area and three-dimensional space spiral backflow are reduced.
44
Nakano, Tsuguji, and Andy Breeze-Stringfellow. "A Method for Evaluating the Aerodynamic Stability of Multistage Axial-Flow Compressors." Journal of Turbomachinery 133, no. 3 (2011): 031004. http://dx.doi.org/10.1115/1.4001225.
Full text
45
Dunham, J. "Modelling of spanwise mixing in compressor through-flow computations." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 211, no. 3 (May 1, 1997): 243–51. http://dx.doi.org/10.1243/0957650971537150.
Full textAbstract:
Although three-dimensional Navier-Stokes computations are coming into use more and more, streamline curvature through-flow computations are still needed, especially for multistage compressors, and where codes which run in minutes rather than hours are preferred. These methods have been made more realistic by taking account of end-wall effects and spanwise mixing by four aerodynamic mechanisms: turbulent diffusion, turbulent convection by secondary flow, spanwise migration of aerofoil boundary layer fluid and spanwise convection of fluid in blade wakes. This paper describes the models adopted in the DRA streamline curvature method for axial compressor design and analysis. Previous papers are summarized briefly before describing the new part of the model—that accounting for aerofoil boundary layers and wakes. Other changes to the previously published annulus wall boundary layer model have been made to enable it to cater for separations and end bends. The resulting code is evaluated against a range of experimental and computational results.
46
Xiang, Honghui, Ning Ge, Jie Gao, Rongfei Yang, and Minjie Hou. "Experimental Investigation of the Effect of the Probe Support Tail Structure on the Compressor Cascade Flow Field." International Journal of Turbo & Jet-Engines 36, no. 1 (March 26, 2019): 9–18. http://dx.doi.org/10.1515/tjj-2016-0071.
Full textAbstract:
Abstract Aiming at resolving the problem of measuring probe blockage effect in the performance experiments of high loaded axial flow compressors, an experimental investigation of the probe support disturbance effect on the compressor cascade flow field was conducted on a transonic plane cascade test facility. The influence characteristics of the probe support tail structure on the cascade downstream flow field under different operation conditions were revealed through the detailed analysis of the test data. The results show that the aerodynamic coupling effect between the upstream probe support wake and the downstream cascade flow field is very intense. Some factors, i. e. inlet Mach number, probe support tail structure, circumferential installing position of probe, and axial distance from the probe support trailing edge to the downstream cascade, are found to have the most impact on the probe disturbance intensity. Under high speed inlet flow condition, changing probe support tail structure can’t inhibit probe support disturbance intensity effectively. Whereas under low speed inlet flow condition, compared with the cylindrical probe, the elliptic probe can inhibit probe support wake loss and reduce disturbance effects on the downstream cascade flow field.
47
Methling, Frank-Oliver, Horst Stoff, and Frank Grauer. "The Pre-Stall Behavior of a 4-Stage Transonic Compressor and Stall Monitoring Based on Artificial Neural Networks." International Journal of Rotating Machinery 10, no. 5 (2004): 387–99. http://dx.doi.org/10.1155/s1023621x04000399.
Full textAbstract:
Current research concerned with the aerodynamic instability of compressors aims at an extension of the operating range of the compressor towards decreased massflow. In practice, a safety margin is maintained between operating point and stability limit to prevent the compressor from going into stall and surge. In this article, we analyze the behavior of a 4-stage transonic axial compressor before entering the unstable range and present an approach to identifying incipient surge and stall using artificial neural networks. This method is based on measurements of the unsteady static wall pressure in front of the first rotor.Analyzing the static pressure signals by using the Fast Fourier Transform shows that peripheral disturbances (modal waves) can only be identified in a small range close to nominal speed (at 95%). At lower speeds (60 to 80% of nominal speed), the investigated compressor flow enters instability by spike-type stall.Monitoring stability over the entire speed range of the compressor relies on artificial neural networks using the unsteady wall pressure signal. In the present case, artificial neural networks show to be the most useful tool to indicate approaching instability. The method works reliably for both types of instabilities, spike-type stall as well as modal waves.
48
Adamczyk, John J. "Aerodynamic Analysis of Multistage Turbomachinery Flows in Support of Aerodynamic Design." Journal of Turbomachinery 122, no. 2 (February 1, 1999): 189–217. http://dx.doi.org/10.1115/1.555439.
Full textAbstract:
This paper summarizes the state of 3D CFD based models of the time-averaged flow field within axial flow multistage turbomachines. Emphasis is placed on models that are compatible with the industrial design environment and those models that offer the potential of providing credible results at both design and off-design operating conditions. The need to develop models free of aerodynamic input from semiempirical design systems is stressed. The accuracy of such models is shown to be dependent upon their ability to account for the unsteady flow environment in multistage turbomachinery. The relevant flow physics associated with some of the unsteady flow processes present in axial flow multistage machinery are presented along with procedures that can be used to account for them in 3D CFD simulations. Sample results are presented for both axial flow compressors and axial flow turbines that help to illustrate the enhanced predictive capabilities afforded by including these procedures in 3D CFD simulations. Finally, suggestions are given for future work on the development of time-averaged flow models. [S0889-504X(00)02002-X]
49
Mostefa, Brihmat, Refassi Kaddour, Douroum Embarek, and Kouadri Amar. "Analysis and Optimization of the Performances of the Centrifugal Compressor Using the CFD." International Journal of Heat and Technology 39, no. 1 (February 28, 2021): 107–20. http://dx.doi.org/10.18280/ijht.390111.
Full textAbstract:
Centrifugal compressors have been used in many areas of the machinery. The centrifugal compressor design is very complex, and a unique design system needs to be developed. A centrifugal compressor design system should be easy to use in interface and also flexible for inputs and outputs. The design tool also needs to be able to predicate the compressor performance in a fairly accurate level. In this study, we have developed a general analyses and optimization approach in the design and performance analysis of centrifugal turbomachines. This approach is based on different methods starting from a 1D approach up to the 3D study of the internal flow. It presents itself as a robust procedure for predicting and understanding the phenomena associated with the operation of turbomachines, but also for predicting performance. Current design system includes initial parameter studies, throughflow calculation, impeller design. The main improvements of the design system are adding the interface to allow users easy to use, adding the input and output capabilities and modifying few correlations. Current design system can predict the blade loading and compressor performance better compared with original design system. To check the aerodynamic appearance of the centrifugal compressor impeller blades, we must change the impeller dimensions and focus on changing axial length, but when changing the blade numbers, the model that improved efficiency and power at the same time introduced a design with a 0.274% and 10.735% improvement in each respectively in comparison to the initial impeller at the design point.
50
Zhu, N. G., L. Xu, and M. Z. Chen. "Similarity Transformations for Compressor Blading." Journal of Turbomachinery 114, no. 3 (July 1, 1992): 561–68. http://dx.doi.org/10.1115/1.2929180.
Full textAbstract:
Improving the performance of high-speed axial compressors through low-speed model compressor testing has proved to be economical and effective (Wisler, 1985). The key to this technique is to design low-speed blade profiles that are aerodynamically similar to their high-speed counterparts. The conventional aerodynamic similarity transformation involves the small disturbance potential flow assumption; therefore, its application is severely limited and generally not used in practical design. In this paper, a set of higher order transformation rules are presented that can accommodate large disturbances at transonic speed and are therefore applicable to similar transformations between the high-speed high-pressure compressor and its low-speed model. Local linearization is used in the nonlinear equations and the transformation is obtained in an iterative process. The transformation gives the global blading parameters such as camber, incidence, and solidity as well as the blade profile. Both numerical and experimental validations of the transformation show that the nonlinear similarity transformations do retain satisfactory accuracy for highly loaded blades up to low transonic speeds. Further improvement can be made by only slightly modifying profiles numerically without altering the global similarity parameters.