Performance Of Compressor Stage Research Articles

A New Global Gas Dynamics Performance Optimization Method for Refrigeration Centrifugal Compressor Stage Based on the Immune Algorithm

Abstract This study proposes a new global gas dynamics optimization method, which was applied to a multi-objective optimization task of centrifugal compressor performance with the aim of determining the improvement probability for achieving high efficiency across a wide operating range. Initially, the original nondominated neighbor immune algorithm (NNIA) was extended to solve constrained multi-objective optimization problems for the first time, which mainly incorporated a procedure for handling inequality and equality constraints without additional parameters. Subsequently, an adaptive topological back-propagation multilayer feed-forward artificial neural network (BP-MLFANN) was trained using the modified NNIA to quickly evaluate the fitness value of the centrifugal compressor stage performance during the optimization. The feasibility of the method was validated using the first stage of a refrigeration centrifugal compressor. The results indicated a substantial enhancement in the stage efficiency of the optimized impeller at the Near-stall, Design, and Near-choke operating points, with increasement of 1.8%, 1.9%, and 4%, respectively, as compared to the baseline stage. The flow field analysis shows that the impact loss at impeller leading edge and flow separation in the passage reduced greatly, the mixing process between the leakage flow and mainstream in the channel is significantly weakened, thus the flow field becomes more uniform after optimization. The new global gas dynamics optimization method provides a reference for the development of efficient and rapid optimization techniques for centrifugal compressor.

Effect of Tip Gap Size on the Performance of an Axial Compressor Stage with and without Active Flow Control

The tip gap region of an axial compressor rotor is a source of complex flows, inducing losses and stability issues. Recent works have proven the ability of blowing high-speed jets in the tip region to improve the surge margin of an axial compressor stage with a narrow tip gap configuration. However, the tip gap size can evolve during the compressor lifetime, possibly affecting its performance and operability. The objective is to evaluate the performance of an active flow control system on a compressor with different tip gap sizes. The present work is based on the single-stage compressor CME2 located at the Laboratory of Fluid Mechanics of Lille and equipped with actuators blowing at the rotor tip leading edge. Configurations with two different values of the tip gap to chord ratio (0.6% and 2.4%) are experimentally tested. RANS simulations are also performed. The effect of tip gap sizes and tip blowing on the flow topology and compressor performance is evaluated (surge margin improvement of the order of 200% for the larger tip gap size).

Cfd Analysis to Investigate the Effect of Leaned Rotor On the Performance of Transonic Axial Flow Compressor Stage

This paper describes the numerical study carried out to understand the impact of combination of forward and backward leaned rotor on the overall performance of a transonic axial flow research compressor of CSIR National Aerospace Laboratories (NAL), Bangalore. The analyses were carried out using commercial Computational Fluid Dynamics (CFD) solver Ansys-CFX. Initially CFD analysis was carried out for the baseline rotor configuration and validated the results with the experimental data. Then, three new configurations of combined Leaned rotors were modeled from the radially stacked baseline rotor by changing the circumferential curvature of original stacking line using three control points located at 0%, 50% and 100% of the blade span. Analyses were carried out for all the three configurations of modified geometries and the results were compared with the baseline configuration. The results revealed many interesting aspects which will be very useful for better understanding of the blade curvature effects on the shock and secondary losses within the transonic rotor. The discussions on the effect of combination of forward and backward leaned rotor and the consequent developments on the overall performance of the compressor are presented in detail in this paper.

The Effect of Impeller Tip Clearance on the Performance of a MGT Mixed Flow Compressor Stage Fitted with a Crossover Diffuser

Mixed flow compressor stage designs are becoming increasingly popular for use in Micro Gas Turbine (MGT) engines. Due to manufacturing constraints and structural considerations, such compressors typically feature unshrouded impeller designs. A numeric investigation of the effect of impeller-shroud tip clearance on the performance of a MGT mixed flow compressor stage fitted with a crossover diffuser is presented. Three baseline test compressors are designed using an in-house-developed application based on one-dimensional mean line theory. These cover a wide range of design mass flow rates, design speeds, and impeller meridional exit (mixed flow) angles. The performance of each of the test compressors are evaluated at an impeller-shroud tip clearance of 0.1 mm, 0.3 mm, and 0.6 mm. Commercial CFD software is used to verify the performance results of the various test compressor configurations. It is found that an increase in impeller tip clearance negatively impacts both compressor performance (total-to-total pressure ratio and efficiency) and choke margin. Conversely, stall margin increases with an increase in impeller tip clearance.

Aerodynamics of inter-spool duct under the influence of an upstream transonic compressor stage

The compressor inter-spool duct (ISD), with its offset and profiled curvature, is an inherent part of the turbofan system due to large density gradients and substantial area changes across each spool. The aerodynamic performance of the ISD is highly dependent on the upstream flow conditions, and hence its design process must account for the upstream effects. In the present investigation, the inter-spool duct design is carried out, suiting an existing high-speed transonic compressor stage, forming an integrated system. The emphasis of this investigation is to assess the effect of upstream secondary flow structure, i.e., shock losses, tip leakage vortices, blade wakes, and corner-separated flow, on the aerodynamic performance of the duct. The compressor stage performance is presented with the help of experimental and numerical data at design and off-design conditions. The inter-spool duct is numerically analyzed considering uniform inlet flow and the flow coming from the upstream high-speed compressor stage. This investigation reveals that the duct with uniform inlet flow is free from any severe flow separation; consequently, minimum loss. However, the losses are observed to increase significantly with the upstream compressor stage. The majority of the losses are incurred at the duct inner wall due to large wake structures emanating from the stator hub section. The stage exit swirl throughout the duct in the outer wall region, in combination with the hub-separated flow, results in the low-pressure region at the duct-inner wall, causing flow reversal at the exit. This creates flow blockage resulting in a decrease in the choke margin of the integrated system. The overall efficiency of the integrated system is also reduced owing to the losses encountered within the duct.

Leakage Flow Impact on Shrouded Stator Cavity Flow Topology and Associated High-Speed Axial Compressor Stage Performance

AbstractThis paper presents the investigation of realistic cavity leakage flow for the von Karman Institute for Fluid Dynamics H25 axial compressor stage, equipped with a shroud cavity of average Reynolds number Rer = 2.2 * 106 and average aspect ratio G = 0.06. On top of overall performance measurements, time averaged data from unprecedented experimental dataset are used to characterize the stator hub shroud cavity flow field and its interaction with the power stream. The sensitivity of the stage performance and stability as well as the cavity flow topology is evaluated against injection conditions, compressor throttling, and compressor speed. Parametric steady RANS simulations of the stage under injection are used to support the findings and state on the relevance of such approach in the preliminary design phase of axial compressor components. As an example, a stage efficiency reduction of 0.97% for a leakage fraction of 0.56% at design point is retrieved. This performance drop is attributed to the effect of cavity flow on: the blockage ratio, the boundary layer skewness, and total temperature increase at stator row inlet. It is also presented that there is a mutual interaction between the cavity geometry and the cavity flow field organization. The injection homogenizes the cavity flow field and the associated pressure gradient. The various operating conditions presented also demonstrate that the cavity flow studied is, on average (i.e., time), sensitive to changes in compressor rotation speed and changes in stage loading.

Experimental evaluation of heat transfer effect on turbocompressor performance operating with helium-neon gas mixtures

Within the framework of the Future Circular Collider (FCC) currently being investigated at CERN, the entire cryogenic cycle had to be revised with respect to the existing Large Hadron Collider (LHC). In particular, a novel pre-cooling cycle had to be developed for this purpose. This led to a closed-loop cryogenic cycle operating with a mixture of helium and neon, also called Nelium. To better understand the challenges and opportunities associated with the design and operation of radial compressors with such light gases, a closed loop test facility has been designed, built and commissioned at the ITSM (University of Stuttgart). The test facility has been developed to operate with air as well as with helium-neon gas mixtures of varying mixing ratios ranging from pure neon to pure helium. In this paper, the test facility architecture and operation procedure are briefly introduced together with a description of the newly installed compressor stage. Experimental performance measurements are then compared to adiabatic and diabatic numerical simulation validating respectively the pressure rise and diabatic stage efficiency for various gases. The heat transfer effect on compressor stage performance is then described and the respective contribution of the influencing factors are quantified.

Performance Evaluation of Transonic Axial Flow Compressor under Distorted Conditions by Groove Casing Technique with Tip Injection and Surface Roughness Effects

The aerodynamic performance and pressure ratio of modern aircraft engine compressor are degrading due to unsteady flow distorted aerodynamic loads and complex phenomena in the blade tip. This paper attempts to improve the performance of modern axial flow transonic compressor stage under distorted flow conditions by incorporating the combination of groove casing, tip injection technique with surface roughness effects. This is performed for the Mach number ranging between 0.8 and 1.3. The obtained numerical results are compared with experimental studies and found to be in good agreement. Through the numerical analysis, the compressor stage flow interaction with the shock waves as well as vortex formation and boundary layer separation is studied in detail. While evaluating the performance of the axial flow compressor, more emphasis is given to the flow properties like pressure, density, temperature, velocity, etc. The performance improvement is observed at a particular Mach number for a specified aerodynamic flow property.

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.

CFD investigation of flow through a centrifugal compressor diffuser with splitter blades

The aerodynamic losses in centrifugal compressors are mainly associated with the separated flow on the suction sides of impeller and diffuser vanes. The overall performance of such compressors can be improved by adding splitter vanes. The present work examines the effect of varying the geometrical location of the splitter vanes in the diffuser on the overall performance of a high-speed centrifugal compressor stage of a small gas turbine. To increase the pressure recovery through the diffuser, two radial sets of vanes are used. The first set of vanes (diffuser-1) is equipped with splitter vanes, placed mid-distance between the main vanes, while the vanes of the second set (diffuser-2) are conventional vanes. Flow through the compressor was simulated using the ANSYS 19 workbench program. Flow characteristics and compressor performance were obtained and analyzed for different circumferential positions of the splitting vanes relative to the main vanes of diffuser-1. The study covered seven positions of the splitter vanes including the original design of the diffuser where the splitter vanes were located at mid-distance between the main vanes. The analysis shows that, at design conditions, selecting the position of the splitter vanes to be nearer to the pressure side of the main vanes improves the stage performance. In the present study, locating the splitters at 33% of the angular distance between the main vanes leads to the best performance, and a significant improvement in the overall stage performance is recorded. The pressure recovery coefficient is raised by about 17%, the pressure ratio is increased by about 1.13%, and the stage efficiency is increased by about 2.01%, compared to the original splitter position. Performance improvement is related to the suppression of the flow separation and the more uniformity of flow. On the contrary, further moving the splitter closer to the main blade, the pressure recovery coefficient is decreased by about 2% than the position of 33% of the angular distance, but still higher than the original position by about 15% and a limited improvement in the compressor performance is noticed. Moving the splitter far out the main blade annihilates the static pressure recovery of the diffuser by about 2:7% compared with the original position. So, for the investigated compressor, the best position of the splitter blade in the circumferential direction, which provides the best stage performance in our parametric analysis, is not necessary to be at the mid-angular distance between the diffuser’s main blades, but it is achieved by moving the splitter to about 33% of the angular distance where the diminished loss from the suppressed flow separation is more prevailing and the instigated friction losses from splitter surfaces are less critical.

Influence of Surface Roughness on a Highly Loaded Axial Compressor Stage Performance at Low Reynolds Number

In the present study, a numerical simulation was conducted to investigate the influence of surface roughness on the aerodynamic performance of a 1.5-stage highly loaded axial compressor at low Reynolds number. It was especially considered how the roughness Reynolds number ( k + ) affected the change of the inlet and outlet conditions, the growth of the separation bubble (LSB), the status of the limiting streamline, and the patterns of the wake. Regarding the roughness settings, five roughness magnitudes and four roughness locations were mainly studied. The results showed that at low Re, surface roughness mainly improved the stage performance by reducing the length and width of the LSB as well as the rotor tip vorticity, delaying the occurrence of three-dimensional flow separation and increasing the turbulence level near the wall. However, it also aggravated the incoordination between the subsequent stages to a certain extent, which limited further improvement of the overall aerodynamic performance. Generally, with k + increasing, the compressor aerodynamic performance improved and achieved the best at k + = 137.8 . The maximum increases in the total pressure ratio, peak efficiency, and chocked mass flow were approximately 4.01%, 5.34%, and 2.24%, respectively. In addition, for all the four roughness locations, the roughness covering from the leading edge to 50% of the axial chord length on the suction surface had a relatively evident advantage in improving the compressor peak efficiency because of the better control of the LSB and wall shear stress.

Effects of geometric and operational uncertainties on aerodynamic performance of centrifugal compressor stage

No real centrifugal compressor can exactly conform to its design geometry and expected operating conditions due to the uncertainties existing in the manufacturing and operational processes. Such uncertainties have been increasingly recognised to be detrimental to compressor performance. However, quite few studies have focused on the combined effects of geometric and operational uncertainties on compressor performance, and the underlying flow mechanism even remains unclear. In this context, we here present an uncertainty analysis of a centrifugal compressor stage, with both geometric and operational uncertainties taken into account. With the combination of CFD simulation and nonintrusive sparse grid based stochastic collocation methods, the combined and individual effects of total inlet temperature, total inlet pressure, outlet mass flow, impeller tip clearance and hub fillet radius on the stage/impeller performance are quantified and analysed. Particular attention is paid to elucidating the compressor performance variations through flow field and energy decomposition analyses. Results show that the considered uncertainties exert more influence on the compressor stage performance rather than on the impeller performance. Amongst the examined uncertainties, the impeller tip clearance contributes the most to the stage performance. The underlying mechanism lies in that the wake of impeller tip clearance produces distorted flow downstream towards the diffuser, which causes complicated vortex structures and less conversion of kinetic energy to pressure rise in the diffuser passage. The present study lays a theoretical foundation for the further uncertainty quantification and robust design of centrifugal compressors against various sources of uncertainties.

Some effects of non-axisymmetric tip clearance layout on axial compressor aerodynamics

A steady and unsteady numerical research is carried out to explore some effects of a specific non-axisymmetric tip clearance layout on the overall performance and stability of an axial compressor stage. For a 4-stage low-speed research compressor (LSRC) in Shanghai Jiao Tong University (SJTU), one-eighth annulus of the inlet guide vane and the first stage rotor was modeled for this study. After the validation for the uniform tip clearance case, a specific non-axisymmetric tip clearance layout is chosen from several random cases generated by the Gaussian Probabilistic Density Function method. Unsteady time-averaged results at the near stall condition show that the chosen non-axisymmetric layout can improve the isentropic efficiency by 1.3% and extend the stall margin by 4%. Detailed analyses on flow fields are carried out to interpret the performance improvement. Due to the circumferential layout of clearance sizes, the inlet mass flow and incidence are redistributed in both the radial and circumferential directions. It leads to blade loading and tip leakage flow varying with the tip clearance size. The quantification of blockage manifests that the blockage arising from the tip leakage flow is significantly alleviated in the non-axisymmetric layout, which leads to improvements in overall performance and stall margin. Transient flow fields at the rotor tip are also analyzed at the near stall condition. For the non-axisymmetric layout, low-momentum regions originating from larger clearance sizes oscillate and develop downstream in one blade passage period.

Numerical Study of the Improvement in Stability and Performance by Use of a Partial Vaned Diffuser for a Centrifugal Compressor Stage

The influence of different diffuser configurations on the flow stability and aerodynamic performance of a centrifugal compressor stage with a mass flow coefficient of 0.196 is numerically investigated. Research results show that the performance of a traditional full-height vaned diffuser (TVD) deteriorates rapidly, and a shroud-side partial vaned diffuser (SVD) displays better adaptability in off-design conditions. SVD can suppress the development of vortices generating at the diffuser leading-edge. Therefore, it can reduce the flow loss inside the stage and improve the flow stability of the stage at low mass flow rates. The unsteady analysis for TVD and SVD shows that the stall cell propagates at about 35.7% of impeller rotational speed in the semi-vaneless space and diffuser passages. Furthermore, the internal flow in TVD and SVD is studied by employing the proper orthogonal decomposition (POD) and dynamic mode decomposition (DMD) methods. The flow loss and instability mechanism in the stage are consequently revealed more comprehensively.

Effects of Tip Leakage Flow on the Aerodynamic Performance and Stability of an Axial-Flow Transonic Compressor Stage

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

Simulation of the effect of nonuniform fouling thickness on an axial compressor stage performance

This paper focuses on the numerical simulation of the flow status in the compressor with the condition of fouling. NASA stage 35 was considered as the object, and the commercial code ANSYS CFX was used. The deposition rate of contaminants on the surface was considered to be different along the height of the blade. A data from related study shows that the deposition rate of the contaminant on the side close to the hub is higher than the side near the shroud part. Based on the deposition law, this paper simulated the fouling of the compressor blades by changing the thickness on the blade surface. This subject only changed the thickness of the stator blade surface because of a data showing that the fouling on the stator blade surface is almost double that on the rotor blade surface. In the condition that the roughness value of the blade surface is constant, only the stable working range of the compressor is effected by the change of the surface thickness of the stator blade. There is a positive relationship between the value of compressor minimum flow rate and the value of thickness increment. After fouling the total pressure ratio and isentropic efficiency degenerated 1.59% and 3.76%, respectively.

Parametric study of rotor tip squealer geometry on the aerodynamic performance of a high subsonic axial compressor stage

Rotor with tip squealer is often used in axial compressor of aircraft engine; however, its effect on the aerodynamic performance of axial compressor is not well understood. The purpose of this paper is to numerically investigate the effect of squealer on the performance of a 1.5-stage axial compressor. In comparison with flat tip, suction and pressure side squealers (SSS and PSS) with different height and thickness are implemented on the rotor blade of the compressor stage. Numerical result shows that pressure ratio is increased slightly under different working conditions with SSS. At the reference working point (RWP), SSS helps to increase mass flow while PSS reduces both mass flow and efficiency. Detailed analysis indicates that blade tip loading rises due to SSS while PSS introduces more losses in the tip region. When squealer is used, the amount of tip leakage flow stays almost unchanged, while vortex trajectory and distribution of leakage flow within tip gap are different. PSS enhances tip leakage flow near the leading edge, which makes the mixing of tip leakage flow and main flow faster and stronger.

Effects of vaneless region length on the performance of centrifugal compressor with vaned diffuser

Effects of vaneless region length on the performance of a centrifugal compressor with vaned diffuser are studied in this paper. The vaneless region is defined as the space between the impeller outlet and the vaned diffuser inlet. In this work, the studied ratio between the diffuser inlet and impeller outlet is from 1.06 to 1.20. The simulation results show that the variation in the vaneless region has a great impact on the performance of the compressor stage. It is found that the proper extension of vaneless region improves the performance of compressor stage by significantly reducing the separation near shroud and uniformly disturbing flow velocity at impeller outlet and diffuser inlet. Sufficient extension of the vaneless region is necessary to produce beneficial effects, while the excessive extension will instead deteriorate the compressor performance. The mechanism of performance improvement is that reasonable vaneless space length is able to constrain the flow separation and to make the flow more uniform in the diffuser.

Inter-Stage Dynamic Performance of an Axial Compressor of a Twin-Shaft Industrial Gas Turbine

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.

Experimental and numerical studies on the influence of inlet guide vanes of centrifugal compressor on the flow field characteristics of inlet chamber

Inlet chambers (IC) are the typical upstream component of centrifugal compressors, and inlet guide vanes in the IC have a great impact on its internal flow and aerodynamic loss, which will significantly influence the performance of the downstream compressor stages. In this paper, an experimental study was carried out on the flow characteristics inside a radial IC of an industrial centrifugal compressor, including five testing sections and 968 measuring points for two schemes with and without guide vanes. Detailed distributions of flow parameters on each section were obtained as well as the overall performance of the radial IC, and the causes of the flow loss inside the IC and the non-uniformity of flow parameters at the outlet section were investigated. Besides, numerical simulations were performed to further analyze the flow characteristics inside the radial IC. The experimental and numerical results indicate that, in the scheme without guide vanes, sudden expansions in the spiral channel and flow separations in the annular convergence channel are the major sources of flow loss and distortions generated in the radial IC; while in the scheme with guide vanes, the flow impacts, separations and wakes caused by the inappropriate design of guide vanes are the main reasons for the flow loss of the IC itself and the uneven flow distributions at the IC outlet.