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1
Song, T. W., T. S. Kim, J. H. Kim, and S. T. Ro. "Performance prediction of axial flow compressors using stage characteristics and simultaneous calculation of interstage parameters." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 215, no. 1 (February 1, 2001): 89–98. http://dx.doi.org/10.1243/0957650011536598.
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A new method for predicting performance of multistage axial flow compressors is proposed that utilizes stage performance curves. The method differs from the conventional sequential stage-stacking method in that it employs simultaneous calculation of all interstage variables (temperature, pressure and flow velocity). A consistent functional formulation of governing equations enables this simultaneous calculation. The method is found to be effective, i.e. fast and stable, in obtaining solutions for compressor inlet and outlet boundary conditions encountered in gas turbine analyses. Another advantage of the method is that the effect of changing the angles of movable stator vanes on the compressor's operating behaviour can be simulated easily. Accordingly, the proposed method is very suitable for complicated gas turbine system analysis. This paper presents the methodology and performance estimation results for various multistage compressors employing both fixed and variable vane setting angles. The effect of interstage air bleeding on compressor performance is also demonstrated.
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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.
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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.
3
Wisler, D. C., R. C. Bauer, and T. H. Okiishi. "Secondary Flow, Turbulent Diffusion, and Mixing in Axial-Flow Compressors." Journal of Turbomachinery 109, no. 4 (October 1, 1987): 455–69. http://dx.doi.org/10.1115/1.3262127.
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The relative importance of convection by secondary flows and diffusion by turbulence as mechanisms responsible for mixing in multistage, axial-flow compressors has been investigated by using the ethylene tracer-gas technique and hot-wire anemometry. The tests were conducted at two loading levels in a large, low-speed, four-stage compressor. The experimental results show that considerable cross-passage and spanwise fluid motion can occur and that both secondary flow and turbulent diffusion can play important roles in the mixing process, depending upon location in the compressor and loading level. In the so-called freestream region, turbulent diffusion appeared to be the dominant mixing mechanism. However, near the endwalls and along airfoil surfaces at both loading levels, the convective effects from secondary flow were of the same order of magnitude as, and in some cases greater than, the diffusive effects from turbulence. Calculations of the secondary flowfield and mixing coefficients support the experimental findings.
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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.
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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.
5
Leylek, J. H., and D. C. Wisler. "Mixing in Axial-Flow Compressors: Conclusions Drawn From Three-Dimensional Navier–Stokes Analyses and Experiments." Journal of Turbomachinery 113, no. 2 (April 1, 1991): 139–56. http://dx.doi.org/10.1115/1.2929069.
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Extensive numerical analyses and experiments have been conducted to understand mixing phenomena in multistage, axial-flow compressors. For the first time in the literature the following are documented: Detailed three-dimensional Navier–Stokes solutions, with high order turbulence modeling, are presented for flow through a compressor vane row at both design and off-design (increased) loading; comparison of these computations with detailed experimental data show excellent agreement at both loading levels; the results are then used to explain important aspects of mixing in compressors. The three-dimensional analyses show the development of spanwise (radial) and circumferential flows in the stator and the change in location and extent of separated flow regions as loading increases. The numerical solutions support previous interpretations of experimental data obtained on the same blading using the ethylene tracer-gas technique and hot-wire anemometry. These results, plus new tracer-gas data, show that both secondary flow and turbulent diffusion are mechanisms responsible for both spanwise and circumferential mixing in axial-flow compressors. The relative importance of the two mechanisms depends upon the configuration and loading levels. It appears that using the correct spanwise distributions of time-averaged inlet boundary conditions for three-dimensional Navier–Stokes computations enables one to explain much of the flow physics for this stator.
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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.
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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.
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Ehrich, F. "Rotor Whirl Forces Induced by the Tip Clearance Effect in Axial Flow Compressors." Journal of Vibration and Acoustics 115, no. 4 (October 1, 1993): 509–15. http://dx.doi.org/10.1115/1.2930379.
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It is now widely recognized that destabilizing forces, tending to generate forward rotor whirl, are generated in axial flow turbines as a result of the nonuniform torque induced by the nonuniform tip-clearance in a deflected rotor—the so called Thomas/Alford force (Thomas, 1958, and Alford, 1965). It is also recognized that there will be a similar effect in axial flow compressors, but qualitative considerations cannot definitively establish the magnitude or even the direction of the induced whirling forces—that is, if they will tend to forward or backward whirl. Applying a “parallel compressor” model to simulate the operation of a compressor rotor deflected radially in its clearance, it is possible to derive a quantitative estimate of the proportionality factor β which relates the Thomas/Alford force in axial flow compressors (i.e., the tangential force generated by a radial deflection of the rotor) to the torque level in the compressor. The analysis makes use of experimental data from the GE Aircraft Engines Low Speed Research Compressor facility comparing the performance of three different axial flow compressors, each with four stages (typical of a mid-block of an aircraft gas turbine compressor) at two different clearances (expressed as a percent of blade length)—CL/L = 1.4 percent and CL/L = 2.8 percent. It is found that the value of β is in the range of +0.27 to −0.71 in the vicinity of the stages’ nominal operating line and +0.08 to −1.25 in the vicinity of the stages’ operation at peak efficiency. The value of β reaches a level of between −1.16 and −3.36 as the compressor is operated near its stalled condition. The final result bears a very strong resemblance to the correlation obtained by improvising a normalization of the experimental data of Vance and Laudadio (1984) and a generic relationship to the analytic results of Colding-Jorgensen (1990).
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Cruz-Manzo, Samuel, Senthil Krishnababu, Vili Panov, and Chris Bingham. "Inter-Stage Dynamic Performance of an Axial Compressor of a Twin-Shaft Industrial Gas Turbine." Machines 8, no. 4 (December 9, 2020): 83. http://dx.doi.org/10.3390/machines8040083.
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In this study, the inter-stage dynamic performance of a multistage axial compressor is simulated through a semi-empirical model constructed in the Matlab Simulink environment. A semi-empirical 1-D compressor model developed in a previous study has been integrated with a 0-D twin-shaft gas turbine model developed in the Simulink environment. Inter-stage performance data generated through a high-fidelity design tool and based on throughflow analysis are considered for the development of the inter-stage modeling framework. Inter-stage performance data comprise pressure ratio at various speeds with nominal variable stator guide vane (VGV) positions and with hypothetical offsets to them with respect to the gas generator speed (GGS). Compressor discharge pressure, fuel flow demand, GGS and power turbine speed measured during the operation of a twin-shaft industrial gas turbine are considered for the dynamic model validation. The dynamic performance of the axial-compressor, simulated by the developed modeling framework, is represented on the overall compressor map and individual stage characteristic maps. The effect of extracting air through the bleed port in the engine center-casing on transient performance represented on overall compressor map and stage performance maps is also presented. In addition, the dynamic performance of the axial-compressor with an offset in VGV position is represented on the overall compressor map and individual stage characteristic maps. The study couples the fundamental principles of axial compressors and a semi-empirical modeling architecture in a complementary manner. The developed modeling framework can provide a deeper understanding of the factors that affect the dynamic performance of axial compressors.
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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.
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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.
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Sieverding, Frank, Beat Ribi, Michael Casey, and Michael Meyer. "Design of Industrial Axial Compressor Blade Sections for Optimal Range and Performance." Journal of Turbomachinery 126, no. 2 (April 1, 2004): 323–31. http://dx.doi.org/10.1115/1.1737782.
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Background: The blade sections of industrial axial flow compressors require a wider range from surge to choke than typical gas turbine compressors in order to meet the high volume flow range requirements of the plant in which they operate. While in the past conventional blade profiles (NACA65 or C4 profiles) at moderate Mach number have mostly been used, recent well-documented experience in axial compressor design for gas turbines suggests that peak efficiency improvements and considerable enlargement of volume flow range can be achieved by the use of so-called prescribed velocity distribution (PVD) or controlled diffusion (CD) airfoils. Method of approach: The method combines a parametric geometry definition method, a powerful blade-to-blade flow solver and an optimization technique (breeder genetic algorithm) with an appropriate fitness function. Particular effort has been devoted to the design of the fitness function for this application which includes non-dimensional terms related to the required performance at design and off-design operating points. It has been found that essential aspects of the design (such as the required flow turning, or mechanical constraints) should not be part of the fitness function, but need to be treated as so-called “killer” criteria in the genetic algorithm. Finally, it has been found worthwhile to examine the effect of the weighting factors of the fitness function to identify how these affect the performance of the sections. Results: The system has been tested on the design of a repeating stage for the middle stages of an industrial axial compressor. The resulting profiles show an increased operating range compared to an earlier design using NACA65 profiles. Conclusions: A design system for the blade sections of industrial axial compressors has been developed. Three-dimensional CFD simulations and experimental measurements demonstrate the effectiveness of the new profiles with respect to the operating range.
11
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.
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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.
12
Mo¨nig, Reinhard, Frank Mildner, and Ralf Ro¨per. "Viscous-Flow Two-Dimensional Analysis Including Secondary Flow Effects." Journal of Turbomachinery 123, no. 3 (February 1, 2000): 558–67. http://dx.doi.org/10.1115/1.1370167.
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During the last few decades extremely powerful Quasi-three-dimensional (3D) codes and fully 3D Navier-Stokes solvers have been developed and successfully utilized in the design process and optimization of multistage axial-flow compressors. However, most of these methods proved to be difficult in handling and extremely time consuming. Due to these disadvantages, the primary stage design and stage matching as well as the off-design analysis is nowadays still based on fast 2D methods incorporating loss-, deviation- and end wall modeling. Only the detailed 3D optimization is normally performed by means of advanced 3D methods. In this paper a fast and efficient 2D calculation method is presented, which already in the initial design phase of multistage axial flow compressors, considers the influence of hub leakage flows, tip clearance effects, and other end wall flow phenomena. The method is generally based on the fundamental approach by Howard and Gallimore (1992). In order to allow a more accurate prediction of skewed and nondeveloped boundary layers in turbomachines, an improved theoretical approach was implemented. Particularly the splitting of the boundary layers into an axial and tangential component proved to be necessary in order to account for the change between rotating and stationary end walls. Additionally, a new approach is used for the prediction of the viscous end wall zones including hub leakage effects and strongly skewed boundary layers. As a result, empirical correlations for secondary flow effects are no longer required. The results of the improved method are compared with conventional 2D results including 3D loss- and deviation-models, with experimental data of a three-stage research compressor of the Institute for Jet Propulsion and Turbomachinery of the Technical University of Aachen and with 3D Navier-Stokes solutions of the V84.3A compressor and of a multistage Siemens research compressor. The results obtained using the new method show a remarkable improvement in comparison with conventional 2D methods. Due to the high quality and the extremely short computation time, the new method allows an overall viscous design of multistage compressors for heavy duty gas turbines and aeroengine applications.
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Aligoodarz, MR, A. Mehrpanahi, M. Moshtaghzadeh, and A. Hashiehbaf. "Improved criteria for stall-free preliminary design of axial compressor of aero gas turbine engines." Proceedings of the Institution of Mechanical Engineers, Part G: Journal of Aerospace Engineering 233, no. 9 (August 24, 2018): 3286–97. http://dx.doi.org/10.1177/0954410018795538.
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A worldwide effort has been devoted to developing highly efficient and reliable gas turbine engines. There exist many prominent factors in the development of these engines. One of the most important features of the optimal design of axial flow compressors is satisfying the allowable range for various parameters such as flow coefficient, stage loading, the degree of reaction, De-Haller number, etc. But, there are some applicable cases that the mentioned criteria are exceeded. One of the most famous parameters is De-Haller number, which according to literature data should not be kept less than 0.72 in any stage of the axial compressor. A deep insight into the current small- or large-scale axial flow compressors shows that a discrepancy will occur among design criterion for De-Haller number and experimental measurements in which the De-Haller number is less than the design limit but no stall or surge is observed. In this paper, an improved formulation is derived based on one-dimensional modeling for predicting the stall-free design parameter ranges especially stage loading, flow coefficient, etc. for various combinations. It was found that the current criterion is much more accurate than the De-Haller criterion for design purposes.
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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.
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Tarabrin, A. P., V. A. Schurovsky, A. I. Bodrov, and J. P. Stalder. "An Analysis of Axial Compressor Fouling and a Blade Cleaning Method." Journal of Turbomachinery 120, no. 2 (April 1, 1998): 256–61. http://dx.doi.org/10.1115/1.2841400.
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The paper describes the phenomenon of axial compressor fouling due to aerosols contained in the air. Key parameters having effect on the level of fouling are determined. A mathematical model of a progressive compressor fouling using the stage-by-stage calculation method is developed. Calculation results on the influence of fouling on the compressor performance are presented. A new index of sensitivity of axial compressors to fouling is suggested. The paper gives information about Turbotect’s deposit cleaning method of compressor blading and the results of its application on an operating industrial gas turbine. Regular on-line and off-line washings of the compressor flow path make it possible to maintain a high level of engine efficiency and output.
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Dorney, D. J., and O. P. Sharma. "Evaluation of Flow Field Approximations for Transonic Compressor Stages." Journal of Turbomachinery 119, no. 3 (July 1, 1997): 445–51. http://dx.doi.org/10.1115/1.2841143.
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The flow through gas turbine compressors is often characterized by unsteady, transonic, and viscous phenomena. Accurately predicting the behavior of these complex multi-blade-row flows with unsteady rotor–stator interacting Navier–Stokes analyses can require enormous computer resources. In this investigation, several methods for predicting the flow field, losses, and performance quantities associated with axial compressor stages are presented. The methods studied include: (1) the unsteady fully coupled blade row technique, (2) the steady coupled blade row method, (3) the steady single blade row technique, and (4) the loosely coupled blade row method. The analyses have been evaluated in terms of accuracy and efficiency.
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Kalinkevych, M., V. Ihnatenko, O. Bolotnikova, and O. Obukhov. "Design of high efficiency centrifugal compressors stages." Refrigeration Engineering and Technology 54, no. 5 (October 31, 2018): 4–9. http://dx.doi.org/10.15673/ret.v54i5.1239.
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The modern trend in compressor industry is an extension of the use of multi-shaft centrifugal compressors. Multi-shaft compressors have a number of advantages over single-shaft. The design of such compressors gives opportunity to use an axial inlet for all stages and select the optimum rotational speed for each pair of impellers, which, along with the cooling of the gas after each stage, makes possible to achieve high levels of efficiency. The design of high-efficiency centrifugal compressor stages can be performed on the basis of highly effective stage elements. Such elements are: impellers with spatial blades, vaned and channel diffusers with given velocity distribution. In this paper, impellers with axial-radial blades are considered. The blade profile is determined by the specified pressure distribution along the blade. Such design improves the structure of the gas flow in the interblade channels of the impeller, which leads to an increase in its efficiency. Characteristics of loss coefficients from attack angles for impellers were obtained experimentally. Vaned and channel diffusers, the characteristics of which are given in this article, are designed with the given velocity distribution along the vane. Compared to the classic type of diffuser, such diffusers have lower losses and a wider range of economical operation. For diffusers as well as for impellers, characteristics of loss coefficients from attack angles were obtained. High efficient impellers and diffusers and obtained gas-dynamic characteristics were used in the design of a multi-shaft compressor unit for the production of liquefied natural gas. The initial pressure of the unit is 3bar. The obtained characteristics of loss coefficients from attack angles for the considered impellers and diffusers make it possible to calculate the gas-dynamic characteristics of high-efficient centrifugal compressors stages. The high-efficient centrifugal compressors stages can be designed using high-efficient elements, such as: impeller with spatial blades and vaned diffuser with given velocity distribution.
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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.
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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.
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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.
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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.
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Colding-Jorgensen, J. "Prediction of Rotor Dynamic Destabilizing Forces in Axial Flow Compressors." Journal of Fluids Engineering 114, no. 4 (December 1, 1992): 621–25. http://dx.doi.org/10.1115/1.2910076.
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It has been shown by Thomas (1958) and Alford (1965), that axial flow turbo-machinery is subject to rotor dynamic destabilizing gas forces produced by the circumferential variation of blade-tip clearance when the rotor is whirling. However, the magnitude and direction of these forces have yet to be clarified. For example, it is still uncertain, under which circumstances the rotor whirl direction will be forward, and when it will be backward, with respect to the rotation. In the present paper, a simple analysis of the perturbed flow in an axial compressor stage with whirling rotor is presented, based on the actuator disc analysis of Horlock and Greitzer (1983), and the gas force on the rotor is calculated on this basis. It appears that in the normal operation range of an axial compressor, the whirl direction is predicted to be forward always. Backward whirl is predicted to take place only at very low flow rates, well below the normally expected stall limit. Experimentally, forces were indeed found in direction of backward whirl for low flow rates, and in direction of forward whirl for high flow rates, in the results reported by Vance and Laudadio (1984), as analyzed by Ehrich (1989). While this experimental evidence supports the present theory qualitatively, a direct comparison of the measured and predicted destabilizing force has yet to be carried out.
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Zhao, Wenfeng, Qun Zheng, Bin Jiang, and Aqiang Lin. "A Passive Control Method of Hub Corner Stall in a 1.5-Stage Axial Compressor under Low-Speed Conditions." Energies 13, no. 11 (May 27, 2020): 2691. http://dx.doi.org/10.3390/en13112691.
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Since the use of the compressor of a ship gas turbine is unavoidable at a low-speed operation, the flow field characteristics and stall mechanism at off-design speeds are important aspects for compressor designers. In this study, the first 1.5 stages of an eight-stage compressor are numerically simulated. The mechanism of compressor rotor instability at lower speeds is identified. The characteristic lines of compressors with various partial clearance are calculated at low speed (0.6 N). The flow field of the same outlet pressure (near stall point of the original compressor without clearance) is compared and analyzed. The results show that, at the near stall point, the suction surface separation and backflow occur in the main flow of the rotor top. It develops along the blade span and finally blocks the flow passage of the rotor, which results in the compressor stall. At the same time, the stall also occurs at the corner of the stator hub. In this paper, the characteristics of partial clearance in four different positions of the stator hub are analyzed. The near stall point and the working point are selected for the flow field analysis. It is concluded that the radial development of the stall vortex on the suction surface of the stator can be restrained by the partial clearance at the stator. In this paper, a passive control method by partial clearance is used in the real compressors, which is different from previous studies on cascades. The margin increases at low speeds.
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Бойко, Людмила Георгиевна, Александр Евгеньевич Демин, and Наталия Владимировна Пижанкова. "МЕТОД РАСЧЁТА ТЕРМОГАЗОДИНАМИЧЕСКИХ ПАРАМЕТРОВ ТУРБОВАЛЬНОГО ГАЗОТУРБИННОГО ДВИГАТЕЛЯ НА ОСНОВЕ ПОВЕНЦОВОГО ОПИСАНИЯ ЛОПАТОЧНЫХ МАШИН. ЧАСТЬ II. ОПРЕДЕЛЕНИЕ ПАРАМЕТРОВ СТУПЕНЕЙ И МНОГОСТУПЕНЧАТЫХ КОМПРЕССОРОВ." Aerospace technic and technology, no. 1 (March 7, 2019): 18–28. http://dx.doi.org/10.32620/aktt.2019.1.02.
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Gas Turbine Engine (GTE) operating characteristics such as thrust (or power), specific fuel consumption and other cycle parameters on different regimes, can be determined by engine modeling and applying correspondent calculation method. Its accuracy is the function of the engine’s element maps definition precision. So these maps representations influence for engines investigation results significantly. Main points and equation system for engine performances calculation method were represented in Part I of this article. The method gives an opportunity for the flow path thermodynamical parameters and engine integral values analyzing by using multistage axial blade machines blade-to-blade descriptions. The compressor and gas turbine and parameters are getting by special program modules, adding to the engine operating characteristics investigation program complex. These modules use the flow path and cascade middle radius geometrical parameters as the data for calculation. The goal of this article is the representation of the method for axial stages and multistage compressors performances definition. The calculation technique is based on one-dimensional (1D) multistage axial compressor flow description. Proposed 1D flow analysis method allows to get the multistage axial compressor maps taking into account the blade-to-blade gaps flow bleeding and by-pass. The method including is founded on the thermal and gas dynamic equations and turbomachinery theoretical dependences and empirical functions for losses and deviation angles determination. Besides, the representing method allows to calculate gas dynamic parameters, velocity triangles, angles of attack, evaluate their deviations from optimal values, hydraulic losses. Also, it can show accordance of stages working on different regimes, find the stage, which is a reason for compressor instability, and stall margin. This method can be used in GTE mathematic simulation, founded on blade-to-blade description multistage blade machines or also in multistage compressor designing. The proposed method gives the opportunity to control the stator variable vanes stagger angles control and to analyze its influence for stage and multistage compressor gas dynamic parameters and maps.
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Camp, T. R., and J. H. Horlock. "An Analytical Model of Axial Compressor Off-Design Performance." Journal of Turbomachinery 116, no. 3 (July 1, 1994): 425–34. http://dx.doi.org/10.1115/1.2929429.
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An analysis is presented of the off-design performance of multistage axial-flow compressors. It is based on an analytical solution, valid for small perturbations in operating conditions from the design point, and provides an insight into the effects of choices made during the compressor design process on performance and off-design stage matching. It is shown that the mean design value of stage loading coefficient (ψ = Δh0/U2) has a dominant effect on off-design performance, whereas the stage-wise distribution of stage loading coefficient and the design value of flow coefficient have little influence. The powerful effects of variable stator vanes on stage-matching are also demonstrated and these results are shown to agree well with previous work. The slope of the working line of a gas turbine engine, overlaid on overall compressor characteristics, is shown to have a strong effect on the off-design stage-matching through the compressor. The model is also used to analyze design changes to the compressor geometry and to show how errors in estimates of annulus blockage, decided during the design process, have less effect on compressor performance than has previously been thought.
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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.
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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.
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Perevoschikov, S. I. "DIFFERENTIAL DIAGNOSTICS OF GAS-TURBINE UNITSBY THEIR MAIN COMPONENTS." Oil and Gas Studies, no. 2 (May 1, 2016): 107–15. http://dx.doi.org/10.31660/0445-0108-2016-2-107-115.
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The dependences were obtained which enable to determine the values of the performance factors of gas- turbine units such as gas turbines and axial-flow compressors. The results of testing of the received relations applicability for practical calculations are presented. The test showed the dependences validity for real processes.
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Dong, Y., S. J. Gallimore, and H. P. Hodson. "Three-Dimensional Flows and Loss Reduction in Axial Compressors." Journal of Turbomachinery 109, no. 3 (July 1, 1987): 354–61. http://dx.doi.org/10.1115/1.3262113.
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Measurements have been performed in a low-speed high-reaction single-stage axial compressor. Data obtained within and downstream of the rotor, when correlated with the results of other investigations, provide a link between the existence of suction surface–hub corner separations, their associated loss mechanisms, and blade loading. Within the stator, it has been shown that introducing a small clearance between the stator blade and the stationary hub increases the efficiency of the stator compared to the case with no clearance. Oil flow visualizaton indicated that the leakage reduced the extensive suction surface–hub corner separation that would otherwise exist. A tracer gas experiment showed that the large radial shifts of the surface streamlines indicated by the oil flow technique were only present close to the blade. The investigation demonstrates the possible advantages of including hub clearance in axial flow compressor stator blade rows.
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Шкабура, Владимир Анатольевич. "РЕЗУЛЬТАТЫ ИССЛЕДОВАНИЙ КОМПРЕССОРНОЙ И ТУРБИННОЙ ЧАСТЕЙ ТУРБОКОМПРЕССОРОВ С ОБЩИМ РАБОЧИМ КОЛЕСОМ ДЛЯ ПРИМЕНЕНИЯ В МАЛОРАЗМЕРНЫХ ДВИГАТЕЛЯХ." Aerospace technic and technology, no. 4 (August 31, 2019): 39–43. http://dx.doi.org/10.32620/aktt.2019.4.07.
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It is considered the issues of improving small-sized engines through the application of a new type of turbomachines – turbo-compressors with general impeller (TCG) to develop engines and power plants. For example, it is shown a diagram of the simplest small-sized gas turbine engine using TCG. For the systematization of relatively efficient TCG schemes, a classification has been developed and is given in the article, of possible schemes for a turbocharger with a common impeller. The classification is based on 5 possible directions of movement of the working medium in the blade apparatus – axial (parallel to the axis of rotation of the machine), centrifugal, centripetal, diagonal and tangential. To implement one or another flowing pattern in the impeller, it is necessary to select the appropriate shape of the impeller blades and the location of the nozzle, exhaust, suction and discharge channels relative to each other. Depending on the direction of movement of the gas flows, turbo-compressors with a common impeller may have two flow patterns in interscapular impeller space – direct-flow and counter-flow. If the directions of the gas and airflow coincide concerning the axis of rotation of the impeller, then the flow pattern in the TCG is direct-flow, with opposite flow flows it is countercurrent. For carrying out the enlarged gas-dynamic calculation of TCG, formulas are given that make it possible to calculate the circumferential force arising on the blades of the impeller in the compressor and turbine working channels of the TCG. Also, formulas are given, with correction factors, for calculating the power factor of the compressor part and the load factor of the turbine part. In the process of computational and experimental studies, the characteristic of the compressor part of the TCG experimental model was obtained. The test results of the compressor part of the TCG experimental model showed good agreement between the calculated and experimental values. Studies have shown that a turbocharger with a common impeller can be used as part of small-size gas turbine engines and in a turbo-supercharging system of a small-capacity internal combustion engine with not high supercharging.
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Li, Y. S., and N. A. Cumpsty. "Mixing in Axial Flow Compressors: Part II—Measurements in a Single-Stage Compressor and a Duct." Journal of Turbomachinery 113, no. 2 (April 1, 1991): 166–72. http://dx.doi.org/10.1115/1.2929076.
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This paper follows directly from Part I, which contains not only the description of the facilities and the results for the C106 four-stage compressor, but also the background, list of nomenclature, acknowledgments, and references. The discussion and conclusions for Parts I and II are given here. The single-stage compressor results show the significant effects of inlet guide vane (IGV) wakes on mixing across the stage in the so-called “free-stream” region; in the casing region tip clearance flow is shown to play an important role in mixing. Explanations for these results are given. Investigations were also carried out in a two-dimensional rectangular duct flow to reveal the mixing mechanism in the corner region similar to those formed by blade surfaces and endwalls in a compressor. Turbulent diffusion has been found to be the dominant mechanism in spanwise mixing; anisotropic inhomogeneous turbulent diffusion is mainly responsible for the nonuniform mixing in the corner region. The larger spread of tracer gas in the tangential direction than in the radial direction is mainly caused by the wake dispersion and relative flow motions within the blade wakes as well as secondary flow contributions in the end-wall regions.
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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.
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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.
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Rakhmanina, L. A., A. V. Zuev, A. Yu Petrov, A. A. Aksenov, and Minh Hai Nguyen. "The investigation of absolute flow non-uniform velocity distributions influence at the centrifugal compressor axial radial impeller inlet using numerical calculation methods in ANSYS CFX." E3S Web of Conferences 140 (2019): 05008. http://dx.doi.org/10.1051/e3sconf/201914005008.
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Currently, methods of numerical modelling are widely used. They are especially widely used in the design of turbo compressors. For the specific task of designing new flowing parts of a centrifugal compressor, it is not recommended to deviate from the canonical design techniques, but it is preferable to supplement them with numerical methods. This article is devoted to the end two-element stage investigation of a centrifugal compressor with an axial radial impeller; the stage main dimensions were obtained using the method of V.F. Rice. In order to obtain the necessary pressure characteristics and determine the dependence for the absolute velocity non-uniform distribution at the inlet to the axial radial impeller, the flow path main dimensions were optimized using numerical calculation methods. The calculation was performed using the SST turbulence model using computational gas dynamics methods in the ANSYS CFX software environment. Based on the optimization results, five compressor designs and corresponding characteristics were obtained. The absolute velocity distribution nature at the inlet to the centrifugal compressor axial radial impeller for five flow path variants is investigated. Empirical dependences are obtained for the deviation of the absolute velocity at the inlet in the hub section axial radial impeller and the absolute velocity deviation at the shroud from the absolute velocity at the average diameter based on the results of a numerical experiment. Recommendations are made for further absolute velocity distributions investigating at the inlet to the compressor impeller.
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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.
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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.
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Бойко, Людмила Георгиевна, Вадим Анатольевич Даценко, and Наталия Владимировна Пижанкова. "ОПРЕДЕЛЕНИЕ ДРОССЕЛЬНОЙ ХАРАКТЕРИСТИКИ ТУРБОВАЛЬНОГО ГТД НА ОСНОВЕ МЕТОДА МАТЕМАТИЧЕСКОГО МОДЕЛИРОВАНИЯ С ИСПОЛЬЗОВАНИЕМ ОДНО- И ДВУМЕРНЫХ ПОДХОДОВ К РАСЧЕТУ ПАРАМЕТРОВ КОМПРЕССОРА." Aerospace technic and technology, no. 7 (August 31, 2019): 21–30. http://dx.doi.org/10.32620/aktt.2019.7.03.
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The results of mathematical modeling processes in the turboshaft gas turbine engine (GTE) are presented. The using calculation method based on a high-level GTE mathematical model, which is founded on a multi-stage axial compressor blade-to-blade description. The model was developed at the Aviation Theory Chair of National Aerospace University “KhAI”. The model is based on a multistage axial compressor thermodynamic parameters calculations using a 1D and 2D approaches to analyzing of the flow. The model named above allows one to take into account air intakes from of the compressor blade gaps, as well as adjusting the angles of installation of the rotary stator vanes depending on the rotational speed. The GTE model has a modular structure. To determine the compressor parameters the modules for 1D or 2D flow calculation can be connected. As the initial data, besides the data traditionally specified in the 1st level GTE models it is necessary to set the geometrical parameters of the compressor flow path and blades on the medium radius (for the 2nd level GTE model) or along with the blade height (for the 3rd level). Both calculating compressor parameters methods are verified and have a fairly wide experience of practical use. The article presents the results of calculating the maps of the GTE multi-stage compressor using one- and two-dimensional approaches. Comparison of the compressor performances achieved by using of these two methods among themselves and with the experimental data has shown their good agreement. The approach used to simulate the flow in compressors makes it possible to estimate, by calculation, the surge margin, to consider the incidence angles and other flow parameters in the blade gaps in a wide range of GTE operation modes. Such results, as well as a comparison with experimental data, are presented in the article. The article also demonstrates the results of applying the described above model to the gas turbine engine performances calculation. The engine has the 12-stage axial compressor with the stator blades position of the first stages regulation. The calculated line of joint operation modes of the gas generator units, the dependence of the power and specific fuel consumption on the rotational speed. Presented are the processes in GTE on stationary modes analyzing results given in the article showed the used model advantage, reliability and expediency of its practical application.
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Prata, A. T., J. R. S. Fernandes, and F. Fagotti. "PISTON LUBRICATION IN RECIPROCATING COMPRESSORS." Revista de Engenharia Térmica 1, no. 1 (June 30, 2001): 56. http://dx.doi.org/10.5380/reterm.v1i1.3501.
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Piston dynamics plays a fundamental role in two critical processes related to fluid flow in reciprocating compressors. The first is the refrigerant leakage through the radial clearance, which may cause considerable loss in the pumping efficiency of the compressor. The second process is the viscous friction associated with the lubricant film in the radial clearance; certainly a significant factor in the compressor energy consumption. In the present contribution a numerical simulation of the piston movement inside the cylinder of a reciprocating compressor is performed. The compressor considered here is a small hermetic compressor employed in domestic refrigerators. For the problem formulation both the axial and the radial piston motion is considered. In operation, the piston moves up and down along the axis of the cylinder, but the radial oscillatory motion in the cylinder bore, despite being usually small, plays a very important role on the compressor performance and reliability. The compromise between sealing of the gas leakage through the piston-cylinder clearance and the friction losses requires a detailed analysis of the oscillatory motion for a good design. The forces acting on the piston are the hydrodynamic force due to the pressure build up in the oil film (lubrication effects), the force due to the connecting rod, the viscous force associated with the relative motion between the piston and oil, and the force exerted by the gas on the top of the piston. All corresponding moments are also included in the problem formulation of the piston dynamics, in order to determine the piston trajectory, velocity and acceleration at each time step. The hydrodynamic force is obtained from the integration of the pressure distribution on the piston skirt, which, in turn, is determined from a finite volume solution of the time dependent equation that governs the oil flow. A Newton-Raphson procedure was employed in solving the equations of the piston dynamics. The results explored the effects of some design parameters and operating conditions on the stability of the piston, its sealing performance and friction losses. Emphasis was placed on investigating the influence of the pin location, radial clearance and oil viscosity on the piston dynamics. The complexity of the piston movement in reciprocating compressors was demonstrated and the detailed model presented can be employed as an useful tool for engineering design.
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Wirkowski, Paweł. "Modelling the characteristics of axial compressor of variable flow passage geometry, working in the gas turbine engine system." Polish Maritime Research 14, no. 3 (July 1, 2007): 27–32. http://dx.doi.org/10.2478/v10012-007-0015-z.
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Modelling the characteristics of axial compressor of variable flow passage geometry, working in the gas turbine engine system This paper concerns application of mathematical modelling methods to analyzing gas-dynamic processes in marine gas turbines. Influence of geometry changes in axial compressor flow passage on kinematical air flow characteristics, are presented. The elaborated mathematical model will make it possible to realize - in the future - simulative investigations of gas-dynamic processes taking place in a compressor fitted with controllable guide vanes.
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Shah, P. N., and C. S. Tan. "Effect of Blade Passage Surface Heat Extraction on Axial Compressor Performance." Journal of Turbomachinery 129, no. 3 (February 1, 2005): 457–67. http://dx.doi.org/10.1115/1.2372776.
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Axial compressor performance with heat extraction via blade passage surfaces (compressor cooling) is compared to its adiabatic counterpart, using computational experiments and mean line modeling. For a multistage compressor with an adiabatic design point, results at fixed corrected rotor speed indicate that, if available, compressor cooling would (i) raise the overall pressure ratio (at a given corrected flow), (ii) raise the maximum mass flow capability, (iii) raise the efficiency, defined as the ratio of isentropic work for a given pressure ratio to actual shaft work, and (iv) provide rear stage choking relief at low corrected speed. In addition, it is found that, if available, cooling in the front stages is better than in the rear stages. This is primarily a thermodynamic effect that results from the fact that, for a given gas, the compression work required to achieve a given pressure ratio decreases as the gas becomes colder. Heat transfer considerations indicate that the engineering challenges lie in achieving high enough heat transfer rates to provide a significant impact to the compressor’s performance.
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Carpenter, Chris. "Subsurface Compressor System Improves Gas Production in Unconventional Reservoirs." Journal of Petroleum Technology 73, no. 07 (July 1, 2021): 62–63. http://dx.doi.org/10.2118/0721-0062-jpt.
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This article, written by JPT Technology Editor Chris Carpenter, contains highlights of paper SPE 201138, “Liquid Removal To Improve Gas Production and Recoverable Reserves in Unconventional Liquid-Rich Reservoirs by Subsurface Wet Gas Compression,” by Lukas Nader, SPE, David Biddick, SPE, and Herman Artinian, SPE, Upwing Energy, et al., prepared for the 2020 SPE Virtual Artificial Lift Conference and Exhibition—Americas, 10–12 November. The paper has not been peer reviewed. This paper describes an artificial lift technology, a subsurface compressor system (SCS), that simultaneously removes liquids, increases gas production, and improves recoverable reserves in gas wells. The subsurface compressor can reverse the vicious cycle of liquid loading, which decreases gas production from a gas well and leads to premature abandonment, by creating a virtuous cycle of increased gas and condensate production. The first field trial of the technology in an unconventional shale gas well supports the mechanism of subsurface gas compression and its benefit to unconventional gas production. The SCS This paper focuses on the latest deployed design. As with all SCS systems, this unit has three major components (Fig. 1). High-Speed Motor. The motor is a four-pole, high-speed, permanent-magnet (PM) synchronous topology. The motor maximum operating speed is 50,000 rev/min, with a 55,000-rev/min overspeed. Surface-mounted PMs are retained on the shaft surface. A sine filter is also used to minimize harmonic losses in the rotor, eliminating the need for active cooling flow in the rotor cavity. With the motor housing hermetically sealed from the environment and maintaining a low pressure within the housing, a minimum life of 20 years is expected from the electrical motor section. The motor rotor is levitated with passive magnetic bearings, requiring no lubrication or a pressurized air source, to support the high-speed rotating shafts. Magnetic Coupling. The magnetic coupling consists of three major components: the male and female ends of the magnetic coupling as well as the isolation can in between. The female end of the magnetic coupling is attached directly to the motor. The isolation can is used to seal the female magnetic coupling section hermetically within the body of the PM motor from the environment. Using a magnetic coupling to transmit torque through an isolation can is one of the key features of the protectorless, rotating, sealless motor system to ensure reliability of the motor. Hybrid Wet Gas Compressor. The compressor is a multistage hybrid axial flow wet compressor. The key advantage of this proprietary compressor design is its relatively straight flow path compared with those of centrifugal compressors. When the flow path is straight, with little change of direction, the heavier constituents, including liquids and solids, will follow the gas phase because there is little or no centrifugal force to separate the high-density phases from the low-density one. Also, erosion of the compressor parts is minimized by the straight flow pattern because of the lower probability of impingements of solid particles on the compressor internal surfaces compared with the torturous internal paths of centrifugal compressors. The remainder of the system, as well as the deployment, is very similar to an electrical submersible pump.
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Xu, Hong, Hua Dong Yang, and Guang Ru Hua. "The Effect of Inlet Conditions on Particle Deposition in Axial Flow Compressor." Advanced Materials Research 915-916 (April 2014): 1066–69. http://dx.doi.org/10.4028/www.scientific.net/amr.915-916.1066.
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Axial flow compressor is an important component, so the compressor performance is of crucial. Fouling changes blade geometry and blade surface roughness is increased, thus aerodynamic performance is affected. The flow of gas phase and gas-solid coupling phase are implemented to reveal the effect of inlet condition on particle deposition. Based on Euler-Lagrange model, this paper made numerical simulation of gas-solid two phase flow in the axial flow compressor rotor cascade. Simulation result shows that the increase of inlet temperature can result in the reduction of particle volume fraction. And particle mass concentration is affected by particle mass flow rate.
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Boyko, Ludmila, and Alexandr Dyomin. "Numerical study of flows in axial compressors of aircraft gas-turbine engines." Eastern-European Journal of Enterprise Technologies 4, no. 8 (94) (July 24, 2018): 40–49. http://dx.doi.org/10.15587/1729-4061.2018.139445.
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Yang, Hua Dong, and Hong Xu. "Fouling Influence Factor Analysis of Axial Flow Compressor." Advanced Materials Research 516-517 (May 2012): 619–22. http://dx.doi.org/10.4028/www.scientific.net/amr.516-517.619.
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Fouling is an important performance degradation factor of axial flow compressor. In order to reveal fouling mechanism, flow simulation of four cases of NASA rotor37 has been performed, such as clean compressor, compressor with roughness of 50μm, 100μm and 150μm. Thermodynamics performance parameters of compressor at different rotational speed with different roughness are discussed. And then fouling sources and its influence factor are analyzed. Research finds that gas contaminants can deposit in blade surface existed with internal oil and water. Simultaneously, humidity, temperature, flow velocity and contact area are the main factors for the formation of compressor fouling. Finally, research finds that fouling can easily formed in suction surface and fouling level of leading edge is more critical than other locations.
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Редин, И. И., and М. А. Шевченко. "УЛУЧШЕНИЕ ТОПЛИВНОЙ ЭФФЕКТИВНОСТИ ГАЗОТУРБИННОГО ДВИГАТЕЛЯ УСТАНОВКОЙ В КОМПРЕССОРЕ НАДРОТОРНОГО УСТРОЙСТВА." Open Information and Computer Integrated Technologies, no. 81 (November 16, 2018): 72–85. http://dx.doi.org/10.32620/oikit.2018.81.08.
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The analysis of the influence of annular grooves on the flow in the compressor rotor air gas channel and the axial compressor characteristics as well as on the fuel efficiency of the gas turbine engine is presented. The hypothetical mechanism of the flow effect in the cavity of the annular groove on the main flow in the tip end of the blade air-foil of the axial compressor stage is outlined. The effectiveness of the casing treatment in the form of single annular groove, width is 20% of the axial projection of the chord of the tip end section of the blade is shown experimentally in a single-stage and multi-stage axial compressor system. The increase of the compressor efficiency with ten single annular grooves installed above the leading edges of the blades of each stage, has reduced the specific fuel consumption of the serial GTE in its main operating modes.
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Prata, A. T., J. R. S. Fernandes, and F. Fagotti. "Dynamic Analysis of Piston Secondary Motion for Small Reciprocating Compressors." Journal of Tribology 122, no. 4 (April 4, 2000): 752–60. http://dx.doi.org/10.1115/1.1314603.
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Piston dynamics plays a fundamental role in two critical processes related to fluid flow in reciprocating compressors. The first is the gas leakage through the radial clearance, which may cause considerable loss in the pumping efficiency of the compressor. The second process is the viscous friction associated with the lubricant film in the radial clearance. In the present contribution a numerical simulation is performed for a ringless piston inside the cylinder of a reciprocating compressor, including both the axial and the radial piston motion. The compressor considered here is a small hermetic compressor employed in domestic refrigerators, with the radial clearance between piston and cylinder filled with lubricant oil. In operation, the piston moves up and down along the axis of the cylinder, but the radial oscillatory motion in the cylinder bore, despite being usually small, plays a very important role on the compressor performance and reliability. The compromise between oil leakage through the piston-cylinder clearance and the friction losses requires a detailed analysis of the oscillatory motion for a good design. All corresponding forces and moments are included in the problem formulation of the piston dynamics in order to determine the piston trajectory, velocity and acceleration at each time step. The hydrodynamic force is obtained from the integration of the pressure distribution on the piston skirt, which, in turn, is determined from a finite volume solution of the time dependent equation that governs the oil flow. A Newton-Raphson procedure was employed in solving the equations of the piston dynamics. The results explored the effects of some design parameters and operating conditions on the stability of the piston, the oil leakage, and friction losses. Emphasis was placed on investigating the influence of the pin location, radial clearance and oil viscosity on the piston dynamics. [S0742-4787(11)00301-8]
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Potapov, V. A., and A. A. Sanko. "Performance simulation of multi-stage axial-flow compressor of turbo-shaft engine with account for erosive wear nonlinearity of its blades." Civil Aviation High Technologies 23, no. 5 (October 28, 2020): 39–53. http://dx.doi.org/10.26467/2079-0619-2020-23-5-39-53.
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The construction and useful practice of gas-turbine engine diagnosis systems depend largely on the availability of the engine mathematical models and its certain components in their structure. Utilization of multi-stage axial flow compressor performance with account for erosive wear of its parts during the operation fundamentally raises possibilities of such systems as erosive wear of flow channel, blade rings of impellers and vane rings of multi-stage compressor is a common cause of preschedule gas-turbine engine detaching from an aircraft. As evidenced by various contributions presented in the article, special emphasis on abrasive wear impact assessment on axial flow compressor performance is placed upon rotor-wing turbo-shaft engine due to their particular operating conditions. One of the main tasks in the process of mathematic simulation of an axial flow compressor blade ring is consideration of its wear type that again has a nonlinear distribution along the level of the blade. In addition, wear rate at entry and exit blade edges often have different principles. Detecting of these principles and their consideration when constructing the compressor mathematical model is a crucial task in diagnostic assessment and integrity monitoring of rotor-wing turbo-shaft engine in operation. The article represents a concept to an estimate nonlinear erosive wear effect of axial flow compressor blades on its performance based on the three-dimensional flow approach in the gas-air flow duct of compressor with a formulation of the blade rings. This approach renders possible to take into account the nonlinearity of the compressor blades wear during their operation. Through the example of the inlet compressor stage of a rotor-wing aircraft gas-turbine engine, the engine pump properties predictions with different kind of rotor blade wear have been presented.
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Yang, Xu, and Zongchang Qu. "Suction port design for a synchronal rotary multiphase pump." Proceedings of the Institution of Mechanical Engineers, Part E: Journal of Process Mechanical Engineering 232, no. 1 (November 10, 2016): 127–32. http://dx.doi.org/10.1177/0954408916678268.
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The synchronal rotary multiphase pump is a new type of positive-displacement pump for multiphase boosting. Given that the discharge valve is eliminated in the synchronal rotary multiphase pump, the radial suction port as used in traditional rotary compressors is no longer suitable for the gas–liquid working condition. Thus, a special design for the axial suction port is presented to ensure the synchronal rotary multiphase pump’s flexibility in handling work fluids with any inlet gas volume fraction. The design concept and working principle of the axial suction port of the synchronal rotary multiphase pump are introduced herein, and a geometrical model of the suction port is established to calculate the critical dimensions for the synchronal rotary multiphase pump manufacture. The results show that the axial suction port’s design avoids the inner compression of work fluids and the back flow from the outlet to the inlet of the synchronal rotary multiphase pump during operation, which meets the requirements for operating under the gas–liquid working condition. The matching between the variation in the sectional area of the suction port and the varying trend of suction volume results in an acceptable sectional velocity of suction flow. Although the designed axial suction port reduces the suction angle range in a revolution, the decrease in the volumetric efficiency is negligible.
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Manjunath S. Dalbanjan et al.,, Manjunath S. Dalbanjan et al ,. "An Effect of Tip Clearance on Aero Performance in Axial Flow Compressors for Aero Gas Turbine Engines." International Journal of Mechanical and Production Engineering Research and Development 9, no. 4 (2019): 769–76. http://dx.doi.org/10.24247/ijmperdaug201977.
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Schnoes, Markus, Christian Voß, and Eberhard Nicke. "Design optimization of a multi-stage axial compressor using throughflow and a database of optimal airfoils." Journal of the Global Power and Propulsion Society 2 (October 18, 2018): W5N91I. http://dx.doi.org/10.22261/jgpps.w5n91i.
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The basic tool set to design multi-stage axial compressors consists of fast codes for throughflow and blade-to-blade analysis. Detailed blade row design is conducted with 3D CFD, mainly to control the end wall flow. This work focuses on the interaction between throughflow and blade-to-blade design and the transition to 3D CFD. A design strategy is presented that is based on a versatile airfoil family. The new class of airfoils is generated by optimizing a large number of airfoil shapes for varying design requirements. Each airfoil geometry satisfies the need for a wide working range as well as low losses. Based on this data, machine learning is applied to estimate optimal airfoil shape and performance. The performance prediction is incorporated into the throughflow code. Based on a throughflow design, the airfoils can be stacked automatically to generate 3D blades. On this basis, a 3D CFD setup can be derived. This strategy is applied to study upgrade options for a 15-stage stationary gas turbine compressor test rig. At first, the behavior of the new airfoils is studied in detail. Afterwards, the design is optimized for mass flow rate as well as efficiency. Selected configurations from the Pareto-front are evaluated with 3D CFD.
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Srinivas, G., K. Raghunandana, and B. Satish Shenoy. "Flow blockage in a transonic axial flow compressor: simulation analysis under distorted conditions." International Journal of Engineering & Technology 7, no. 2.21 (April 20, 2018): 43. http://dx.doi.org/10.14419/ijet.v7i2.21.11833.
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Today the aircraft industry is looking for faster and safer engines for both civil and military applications. The performance of all different types of air breathing engines depends on the amount of mass flow rate of air entering and hot gas ejecting out from the engine. Thrust is the key role for any engine performance. To achieve more thrust all the turbo machinery components like axial fan, axial flow compressor and axial flow turbine should function effectively. This paper is primarily dealing about one of the turbo machinery component, axial flow compressor performance where the study is more focused on flow blockage formation under distorted phenomena. The complete blade boundary layer formation and related flow numerical theory are discussed in detail, accordingly the boundary conditions were set to have better numerical simulations using ANSYS tool. To find the flow blockage formation suitable turbulence model was coded using the well know compressible equations. The flow blockage between the rotor and stator of the compressor stage was calculated and also validated with that of experimental data effectively. The flow simulation results also revealed that the performance parameters under the modern engine transonic speed from Mach 0.8 to 1.2 under the distorted conditions are better for aeromechanical features.
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Yushchenko, K. A., V. S. Savchenko, L. V. Chervyakova, V. I. Izbash, and G. Solyanik. "Assessment of causes of fracture in 14Kh17N2 steel blades of GTK-25I gas-pumping unit axial-flow compressors." Strength of Materials 40, no. 5 (September 2008): 597–600. http://dx.doi.org/10.1007/s11223-008-9071-6.
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Редин, И. И., and М. А. Шевченко. "СИСТЕМАТИЗАЦІЯ І УЗАГАЛЬНЕННЯ ТЕОРЕТИЧНИХ ТА ЕКСПЕРИМЕНТАЛЬНИХ ДАНИХ ПО ЕФЕКТИВНОСТІ НАДРОТОРНОГО ПРИСТРОЮ В ОСЬОВОМУ КОМПРЕСОРІ." Open Information and Computer Integrated Technologies, no. 87 (June 30, 2020): 181–99. http://dx.doi.org/10.32620/oikit.2020.87.11.
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The systematization of physical flow models at the peripheral region of the rotors of axial compressor is carried out. Based on the experimental and numerical studies, the flow features in subsonic and transonic rotors are analyzed. Similar features of the flow near the wall at the periphery of subsonic and transonic rotors are formulated. The characteristic areas and individual features of the near-wall flow in them, which are obtained in experimental studies of the flow structure at the periphery of the blade rows, are reflected. The analysis of the influence of annular grooves in the axial compressor case on the flow in the airfoil channel of the subsonic and transonic rotors is presented. The hypothetical mechanism of the flow effect in the cavity of the annular groove on the main flow at the tip region of the airfoil channel of axial compressor rotor is described. An approach to generalize the experimental data of the axial compressor stages with the casing treatment based on the selected fundamental system of dimensionless parameters characterizing the main features of the flow at the rotor tip region are proposed. Using the approach, the dependences of the casing treatment effect on the gas-dynamic stability limit and efficiency are obtained. It was found, when Reynolds numbers ReΔr> 400 increase, the efficiency of annular groove casing treatment in the axial compressor wall on the gas-dynamic stability limit of the compressor decreases. The existence of the region of aerodynamic efficiency modes of the annular groove casing treatment in the case is shown. In this area, there is an optimal mode when the maximum effect of efficiency from install annular groove casing treatment is achieved. The obtained generalization al-lows us, at the step of making design decisions, to evaluate the effectiveness of the annular groove casing treatment in the case when it is used in subsonic and transonic rotors of the axial compressor stages.
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., Nilesh P. Salunke. "DESIGN OPTIMIZATION OF AN AXIAL FLOW COMPRESSOR FOR INDUSTRIAL GAS TURBINE." International Journal of Research in Engineering and Technology 03, no. 05 (May 25, 2014): 458–64. http://dx.doi.org/10.15623/ijret.2014.0305084.
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Poljak, Igor, Josip Orović, Vedran Mrzljak, and Dean Bernečić. "Energy and Exergy Evaluation of a Two-Stage Axial Vapour Compressor on the LNG Carrier." Entropy 22, no. 1 (January 17, 2020): 115. http://dx.doi.org/10.3390/e22010115.
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Data from a two-stage axial vapor cryogenic compressor on the dual-fuel diesel–electric (DFDE) liquefied natural gas (LNG) carrier were measured and analyzed to investigate compressor energy and exergy efficiency in real exploitation conditions. The running parameters of the two-stage compressor were collected while changing the main propeller shafts rpm. As the compressor supply of vaporized gas to the main engines increases, so does the load and rpm in propulsion electric motors, and vice versa. The results show that when the main engine load varied from 46 to 56 rpm at main propulsion shafts increased mass flow rate of vaporized LNG at a two-stage compressor has an influence on compressor performance. Compressor average energy efficiency is around 50%, while the exergy efficiency of the compressor is significantly lower in all measured ranges and on average is around 34%. The change in the ambient temperature from 0 to 50 °C also influences the compressor’s exergy efficiency. Higher exergy efficiency is achieved at lower ambient temperatures. As temperature increases, overall compressor exergy efficiency decreases by about 7% on average over the whole analyzed range. The proposed new concept of energy-saving and increasing the compressor efficiency based on pre-cooling of the compressor second stage is also analyzed. The temperature at the second stage was varied in the range from 0 to −50 °C, which results in power savings up to 26 kW for optimal running regimes.