Compressor Stage Efficiency Research Articles

Impact of the inducer meridional length on the tandem impeller performance within a transonic centrifugal compressor

Tandem impeller configuration has the potential to improve the performance of highly loaded transonic centrifugal compressors, while the related research is still in the early stages and general design guidelines, such as the selection of inducer meridional length, are limited so far. In this paper, a highly loaded transonic centrifugal compressor with various tandem impeller configurations is numerically investigated. The objective is to evaluate the impact of inducer meridional length [Formula: see text] on the tandem impeller flow structure and performance, and further to look into the underlying flow mechanisms and design principles of the tandem impeller. The results indicate that the inducer meridional length [Formula: see text] has significant effect on the tandem impeller and compressor stage performance. There exits an optimized tandem impeller [Formula: see text] value of Appr. 38% that achieves a compressor stage pressure ratio and efficiency gain (at the P.E. points) of 2.6% and 1.8%, respectively, and deviating from the optimized [Formula: see text] value will lead to deteriorated compressor performance. In terms of the impeller performance, the inducer trailing edge determined by the inducer meridional length [Formula: see text] should be located downstream of the passage shock to reduce the shock and inducer wake losses. However, the inducer blade should also be placed in the impeller axial part to avoid the exudcer incidence angle deviation. In terms of the diffuser performance which is partly determined by the impeller discharge flow uniformity, the inducer meridional length [Formula: see text] directly affects the tandem impeller discharge flow uniformity by rearranging and controlling the secondary flow patterns, and the flow uniformity is prone to be insensitive to the [Formula: see text] variation when the [Formula: see text] value is large enough (i.e. ≥30%). Additionally, the [Formula: see text] value should be large enough (i.e. ≥20%) to generate intense inducer shedding vortexes, which can suppress the development of Coriolis vortex and finally improve the impeller flow uniformity.

Algorithm and the examples of multiparametric automated computational optimization for turboexpander compressor stage 3D impellers

The study presents an algorithm and results of the multiparametric optimization for high-flow centrifugal compressor stages. We present the use of a CAD parameterized model of the compressor stage in the automated computational optimization procedure by the AMO – Adaptive Multiple-Objective method in the Ansys complex. To predict the compressor stage efficiency in the design mode, a viscous three-dimensional CFD calculation is used. As an objective function of the optimization problem, the stage total to total polytropic efficiency and the total to total polytropic head coefficient constraint are selected. The algorithm was used for the centrifugal compressor stages improvement. As a result, the efficiency was increased by 1.5-2.5%, and the stages energy–efficient operation zone was expanded by 30% - 60%. The results of the study showed that computational optimization can be successfully applied for the projected centrifugal compressor flow part fine-tuning to obtain the highest efficiency.

Methodology for estimating the specific fuel consumption of a two-circuit turbojet engine

A technique was developed for determining the minimum specific fuel consumption of a two-circuit turbojet engine using statistical data on the polytropic efficiency of individual compressor and turbine stages. The proposed method allows the presence of a technological advantage in specific fuel consumption compared to similar engines to be identified at the initial stage of designing a two-circuit turbojet engine.

ДО ОПТИМАЛЬНОГО ПРОЕКТУВАННЯ ОСЬОВОГО БАГАТОСТУПІНЧАСТОГО КОМПРЕСОРА

The work is devoted to the development of an integrated approach to the aerodynamic improvement of the blade vanes of multistage axial compressors of gas turbine engines. One of the features of the studies carried out to optimize the blade vanes was to take into account the mixing effects by various sources of losses. Thus, minimization of losses sources in compressor blade vanes during their profiling is achieved due to the account of the complex three-dimensional vortex nature of the flow and the use of modern software systems for solving the Navier-Stokes equations. Calculations of the three-dimensional viscous flow in the work are performed with ANSYS software. The profiling of the multi-stage compressor blades is carried out with “Kompress 2.0” software package, the feature of which is the use of parametric models of rotor and stator blade vanes, adapted for their aerodynamic optimization, as well as taking into account the radial unevenness of the incident flow when selecting the tangential blades shape of the stator vanes. As a result, a complex methodology for the optimal design of the blades of a multi-stage compressor with the use of "controlled" diffusivity, S-shaped profiles, lean, and chord variation is developed, with simultaneous variation of geometric parameters according to special plans and the use of the compressor efficiency as an integral efficiency criterion, “Kompress 2.0” for the blades profiling, ANSYS for the calculation of threedimensional viscous flow. Multi-stage axial compressors have been designed, numerically and experimentally investigated. Based on the results of experimental studies of multistage axial compressors, new generalizing relationships are established, that establish a relationship between the geometric parameters of the profiles and their gas dynamic characteristics to take into account the features of the spatial shape of the flow at the stage when solving the inverse design problem. The results of the research were used in the design and improvement of the elements of the flow paths of the ZM multistage compressors. Test aerodynamic optimization of the blade systems of experimentally studied multistage compressors with the use of the proposed approaches made it possible to obtain the efficiency of high-pressure compressor stages at the level of 0.92 ... 0.93 and to increase the efficiency of multistage axial compressors by 1.0-2.4%. The efficiency of the 10-stage compressor is 0.8% (to the level of 0.853), the efficiency of the 6-stage compressor is 1.1% (to the level of 0.825), the efficiency of the 3-stage compressor is 2.4% (to the level of 0.885), and the efficiency of an 8-stage axial compressor is 1.6% (to the level of 0.879).

Experimental assessment on effect of lower porosities of bend skewed casing treatment on the performance of high speed compressor stage with tip critical rotor characteristics

Development of casing treatment geometries which will enhance the stable operating envelope of the compressor stage without sacrificing the efficiency is a great challenge for the designers. Bend skewed casing treatment with relatively lower porosities in the presence of plenum chamber is proven to be capable of improving the stall margin and gain in the compressor stage efficiency compared to other conventional casing treatments. Experimental investigations presented in this paper show very encouraging results with substantial improvement in the stable margin of the compressor stage with increase in the stage efficiency. The three bend skewed casing treatments designed for the experimental evaluations have porosities of 21%, 33% and 45%. The steady and transient performances of these three casing treatment geometries are evaluated at two axial extensions of 20% and 40%. Flow survey at the rotor exit shows distinct behaviors of the aerodynamic parameters in the tip region at near stall flow condition. The unsteady velocity measurements are performed with hot wire probe in the tip region ahead of the rotor leading edge to capture the fluctuations in the inlet axial velocity and the results are presented at the stall flow condition. The presence of the casing treatment changes the stalling behavior of the compressor stage. Highest porosity casing treatment results in 40.62% improvement in the stall margin with 0.88% improvement in the compressor stage efficiency at 20% axial extension.

Hub and shroud fillets influence on the radial compressor stage efficiency

Article describes numerical simulations of an air flow in the radial compressor stage in the NUMECA CFD software. During the tasks evaluating the stepped and straight impeller seals and hub and shroud fillets influence on working characteristics and flow field was observed. Also the CFD results comparison with results from the empiric design tool and the measurement is described.

Flank Milling 공법적용을 위한 자동차용 터보차져 임펠러의 설계체험

The performance of small-size impellers with ruled surfaces was investigated for flank milling over a wide speed range, using computational fluid dynamics analyses and gas bench tests. An impeller with a ruled surface was designed, manufactured, and tested to evaluate the effects of blade loading, the backsweep angle, and the relative velocity distribution on the compressor performance. The simulations and tests were completed using the same compressor cover with identical inlet and outlet channels to accurately compare the performance of the abovementioned impeller with a commercial impeller containing sculptured blades. Both impellers have the same number of blades, number of splitters, and shroud meridional profiles. The backsweep angles of the blades on the ruled impeller were selected to work with the same pinched diffuser as for a sculptured impeller. The inlet-to-exit relative velocity diffusion ratio and the blade loading were provided to maximize the flow rate and to minimize the surge flow rate. The design flow rate, rpm, were selected same for both impellers. Test results showed that for the compressor stage with a ruled impeller, the efficiency was increased by 0.32% with an extended surge margin without a reduction in the pressure ratio as compared to the impeller with the sculptured design. It was concluded that an increased relative velocity diffusion coupled with a large backsweep angle was an effective way to improve the compressor stage efficiency. Additionally, an appropriate blade loading distribution was important for achieving a wide operating range and higher efficiency.

Aerodynamic improvement of GTE compressors by space profiling of blade vanes

The optimum spatial profiling of blade vanes of GTE compressor stages is considered on the basis of calculation of a three-dimensional viscid stream. The simultaneous varying of geometrical parameters of blades and blade cascades in their profiling substantially accelerates the procedure of parametric optimization. As an integral criterion of minimization of different types of hydraulic losses the efficiency of compressor stages is used. The coefficient of hydraulic resistance is shown to decrease as compared to the initial vane, in a wide range of incidence angles, as a result of experimental research of the optimized output straight vane cascade of a stationary GTE compressor.

Numerical study on the Effect of Interaction Vaned Diffuser with Impeller on the Performance of a Modified Centrifugal Compressor

ABSTRACTThis paper is a numerical simulation that was made in the three-dimensional flow, carried out in a modified centrifugal compressor, having vaned diffuser stage, used as an auto-motive turbo charger. Moreover, the performance of the centrifugal compressor was dependent on the proper matching between compressor impeller and vaned diffuser, influencing significantly surge and the efficiency of centrifugal compressor stages. In addition, a modified compressor impeller, coupled with vane and vaneless diffuser, has been found to have similar internal flow patterns for both the vaneless and vaned diffuser design. The vaned diffuser effect has been paid particular attention in terms of better analysis where the diffuser was designed for high sub-sonic inlet conditions. Another aim of this research was to study and simulate the effect of vaned diffuser on the performance of a centrifugal compressor. The simulation was undertaken by using a commercial software, the so-called ANSYS CFX, to predict numerically the performance in terms of pressure ratio, poly tropic efficiency and mass flow rate for the centrifugal compressor stage. The results were generated from CFD and were analyzed for better understanding of the fluid flow through centrifugal compressor stage. Conclusively, it was observed that the effect of the vaned diffuser is to convert the kinetic energy into a high static pressure after analyzing the results of the simulation.

Numerical simulation of radial compressor stage

Article describes numerical simulations of air flow in radial compressor stage in NUMECA CFD software. In simulations geometry variants with and without seals are used. During tasks evaluating was observed seals influence on flow field and performance parameters of compressor stage. Also is described CFDresults comparison with results from design software based on experimental measurements and monitoring of influence of seals construction on compressor stage efficiency.

Effect of Impeller Blade Loading on Compressor Stage Performance in a High Specific Speed Range

The authors previously found that compressor stage efficiency in a high specific speed range was significantly improved by employing an increased relative velocity diffusion ratio coupled with a high backsweep angle (Shibata et al., “Performance Improvement of a Centrifugal Compressor Stage by Increasing Degree of Reaction Optimizing Blade Loading of a 3D-Impeller,” ASME Paper No. GT2009-59588). In spite of such a high relative velocity diffusion ratio, the same surge margin as with a conventional design was able to be achieved by using a special front loading distribution with a lightly loaded inducer. In the present study, the blade loading distribution was further optimized in order to achieve a larger surge margin than previously. Four types of fully shrouded impellers were designed, manufactured, and tested to evaluate the effects of blade loading, backsweep angle, and relative velocity diffusion ratio on compressor performance. The design suction flow coefficient was 0.125 and the machine Mach number was 0.87. Test results showed that the developed impeller achieved 3.8% higher stage efficiency and 11% larger surge margin than the conventional design without reducing the pressure coefficient and choke margin. It was concluded that aft loading coupled with a high degree of reaction was a very effective way to improve surge margin as well as stage efficiency. Stator matching was also investigated by changing the design incidence angle, which was shown to have a little influence on surge margin in the present test results.

Performance Improvement of a Centrifugal Compressor Stage by Increasing Degree of Reaction and Optimizing Blade Loading of a 3D Impeller

Performance improvement of 3D impellers in a high specific speed range was investigated using computational fluid dynamics analyses and experimental tests. In order to reduce the loss production within the stator passages, the backsweep angle of the impellers was increased. At the same time, the inlet-to-exit relative velocity diffusion ratio was also increased by increasing the impeller exit width to prevent the reduction in the pressure ratio. Moreover, the blade loading distribution at the impeller shroud side was optimized to suppress the surge margin reduction caused by the increased relative velocity diffusion ratio. Five types of unshrouded impellers were designed, manufactured, and tested to evaluate the effects of blade loading, backsweep angle, and relative velocity diffusion ratio on the compressor performance. The design suction flow coefficient was 0.125 and the machine Mach number was 0.87. Test results showed that the compressor stage efficiency was increased by 5% compared with the base design without reducing the pressure coefficient and surge margin. It was concluded that an increased relative velocity diffusion ratio coupled with large backsweep angle was a very effective way to improve the compressor stage efficiency. An appropriate blade loading distribution was also important in order to achieve a wide operating range as well as high efficiency.

Tailor-Made Coatings for Turbine Applications Using the Triplex Pro 200

In modern jet engines, the efficiency of compressor stages is highly dependent on the clearance between blade tip and casing. To improve efficiency of gas turbines, the gap between the turbine blades and casing has to be minimized. Abradable coatings permit a minimization of the clearance and control of the over-tip leakage by allowing the blade tips to cut into the coating. Thermal sprayed abradables aim at a well-balanced profile of properties relevant for abradable seals. Amongst others these include: abradability, aging resistance, corrosion and oxidation resistance, and bond strength to substrate. Here, abradable coatings consisting of a multiphase material, comprising a metal matrix and a solid lubricant as well as a defined porosity, were developed using the Triplex Pro 200 to increase reproducibility and deposition efficiency. In addition, the influence of the process parameters on coating characteristics (porosity, hardness) and, as a result, erosion properties and abradability was investigated.

막오염에 의한 압축기 성능 저하가 발전용 가스터빈 설계점 성능에 미치는 영향에 관한 연구

Operating performance of industrial gas turbines in combined cycle power plants depends upon atmospheric conditions. Compressor fouling caused by airborne particles in the atmosphere and their adhesions on compressor blades is one of critical phenomena related to the performance degradation of industrial gas turbines. Compressor fouling provokes increase of pressure loss in inlet duct, decrease of mass flow rate of intake air and decrease of compressor stage efficiency. In this study, impacts of compressor fouling on the performance of an industrial gas turbine operating at design point condition are investigated analytically. As results, it is found that the reduction of produced power with decreased mass flow rate of intake air caused by narrowed flow area by the adhesion of airborne particles on compressor blades is the most dominant impact on the gas turbine performance by the compressor fouling phenomena.

The influence of inlet flow distortion on the performance of a centrifugal compressor and the development of an improved inlet using numerical simulations

The performance of centrifugal compressors can be seriously affected by inlet flow distortions due to the unsatisfactory nature of the inlet configuration and the resulting inlet flow structure. Experimental tests have been carried out for the comparison of centrifugal compressor stage efficiency with two different inlet configurations, one of which is straight with constant cross-sectional area and the other a 90° curved pipe with nozzle shape. The comparative test results indicated significant compressor stage performance difference between the two different inlet configurations and the details are discussed to understand the performance behaviour of the compressor exposed to the distorted flow from the bend inlet configuration. The experimental investigation motivated the need for a new inlet design as well as a clear picture of the detailed flow field in the existing inlet design using numerical simulations. Two design approaches are reported in this paper, one of which is the location of vanes and the other the length of the curvature radius. For a more effective design method, a generalized formula is developed for the optimum position and number of vanes in such a way that each divided flow passage with vanes shares the same pressure gradient in radial direction. Numerical simulation results are presented and discussed in terms of mass-averaged parameters and flow structures, based on the comparison of flow properties at the pipe exit cross-sectional area of each design. Finally, new designs of different inlet systems are proposed to reduce the secondary flow and to provide flow as uniform as possible for a compressor.

The Influence of Clocking on Unsteady Forces of Compressor and Turbine Blades

An important part of the unsteady effects in turbomachinery is generated by the relative motion between rotor and stator airfoils. The main sources of unsteadiness due to the rotor-stator interaction are potential flow (or inviscid) interaction and viscous interaction. Recently, the concept of indexing, or clocking, has been introduced to increase the efficiency of turbine and compressor stages. Previous research on airfoil clocking has been confined to studies of efficiency variation. The present paper addresses the clocking effects on the variation of unsteady forces on compressor and turbine blades. Numerical simulation of compressor and turbine flow shows that the pressure amplitude on the airfoil surface increases when the clocking position corresponds to higher efficiencies. Based on the numerical results, it is hypothesized that, at the maximum efficiency clocking position, the inlet guide vane is intensifying the pressure response of the stator which then increases the potential flow interaction between the rotor and stator. This coupling resonance between the inlet guide vane and the stator, which occurs at maximum efficiency, energizes boundary layer, producing a reduction of airfoil losses. ∗Assistant Professor, Member AIAA †Associate Professor, Senior Member AIAA Copyright c ©1999 by Paul Cizmas and Daniel Dorney. Published by the American Institute of Aeronautics and Astronautics, Inc. with permission. NOMENCLATURE Cf Skin friction coefficient f Frequency P Pressure P.S. Pressure surface Pt Total pressure S Arc length S.S. Suction surface t∗ Non-dimensional time Utip Rotor tip velocity α Absolute frame flow angle (from axial) β Relative frame flow angle η Efficiency θ Momentum thickness

Closure to “Discussion of ‘The Effects of Reynolds Number on the Efficiency of Centrifugal Compressor Stages’” (1985, ASME J. Eng. Gas Turbines Power, 107, p. 1047)

Closure to “Discussion of ‘The Effects of Reynolds Number on the Efficiency of Centrifugal Compressor Stages’” (1985, ASME J. Eng. Gas Turbines Power, 107, p. 1047)

The Effects of Reynolds Number on the Efficiency of Centrifugal Compressor Stages

An analysis of the friction losses in a centrifugal compressor stage is used to suggest a new form of correction equation for the effect of Reynolds number on efficiency. This equation relates the effect of Reynolds number to the surface roughness, the impeller outlet width ratio and the work input coefficient. Systematic tests on a wide range of compressor stages are used to calibrate the single empirical coefficient in the equation. Despite its simplicity this equation provides more accurate predictions of the Reynolds number effects than existing empirical methods.

Investigation of Blade Characteristics: Performance and Efficiency of Turbine and Axial-Flow Compressor Stages

Abstract The investigation reported in this paper deals with the flow through blade rows of turbine or axial-flow compressor stages on the basis of airfoil theory. Correlation was established by means of this theory between the aerodynamic characteristics of blade elements and the performance characteristics of the blade row. Test results from a rotating grid of sufficiently low solidity to eliminate interaction of neighboring blades are presented. It was found that the coefficients of lift of blade sections in rotating grids, especially near the hub of the wheel, are higher than might be expected from wind-tunnel tests. Methods of experimental aerodynamic investigation were adapted to the rotating-blade grid for the determination of the profile drag, in an attempt to broaden the basis for a rational determination of the efficiency of turbine and compressor stages.

Closure to “Discussion of ‘Investigation of Blade Characteristics: Performance and Efficiency of Turbine and Axial-Flow Compressor Stages’” (1944, Trans. ASME, 66, p. 419)

Closure to “Discussion of ‘Investigation of Blade Characteristics: Performance and Efficiency of Turbine and Axial-Flow Compressor Stages’” (1944, Trans. ASME, 66, p. 419)