Compressor Stage Research Articles

Numerical Analysis of Flow Characteristics of the Low-Speed Centrifugal Compressor Stage Comprising Wedge Diffuser

Wedge type diffusers are used generally in the highly loaded stages of smaller jet engines as it is compact in size. Low speed centrifugal compressor (LSCC) is selected for this present study as experimental details are available. A centrifugal compressor stage comprising wedge type diffuser is used in this numerical investigation in which studies are carried out at design mass flow rate (30kg/s) and off-design mass flow rates (23.64kg/s and 36.36kg/s) at constant rotational speed of 1920 rpm. Single passage approach is chosen to model the computational domain which is meshed with unstructured grid. Turbulence model chosen is kω-SST. The investigation revealed the jet-wake structure along the pressure side and suction side of the impeller and its subsequently mixing at impeller exit vaneless diffuser region. Diffusion process in the LSCC is observed to be effective as the outlet values of absolute velocity are lesser compared to its inlet values. Highest static pressure rise is observed for design mass flow rate and followed by below and above off-design mass flow rates.

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.

Numerical Investigation of Transient Flow Characteristics in a Centrifugal Compressor Stage with Variable Inlet Guide Vanes at Low Mass Flow Rates

This study numerically investigates the beneficial effects of positive pre-swirl on the aerodynamic performance and internal flow field in a centrifugal compressor stage with variable inlet guide vanes (VIGVs) at low mass flow rates. Four positions of VIGV are considered, including 0°, 30°, 45°, and 60° angle. The latter three positions of VIGV induce positive pre-swirl. Numerical results show that as positive pre-swirl increases, the aerodynamic performance curve of the stage moves in the low mass flow rate direction. In the three cases of positive pre-swirl, there was an improvement of approximately 9.95% of stall/surge margin greater than in conditions with no pre-swirl. The regulation of IGV can effectively improve the unstable flow of the compressor stage at low mass flow rates. A low frequency that has a great influence on the internal flow of the compressor stage is found, and the unstable flow caused by low frequency is analyzed by the combination of streamline distribution, spectrum analysis, vector, entropy increase, and modal decomposition method. Meanwhile, the modal decomposition method and flow field reconstruction techniques are used to investigate the coherent flow structures caused by low frequency under different guide vane openings.

Experimental and Numerical Analysis of a Compressor Stage under Flow Distortion

On many occasions, fan or compressor stages have to face azimuthal flow distortion at inlet, which affects their performance and stability. These flow distortions can be caused by external events or by some particular geometrical features. The aim of this work is to propose a joined numerical and experimental analysis of the flow behavior in a single axial compressor stage under flow distortion. The distortions are generated by different grids that are placed upstream to the rotor. Experimentally, the flow analysis is based on the measurements obtained by a series of unsteady pressure sensors flush-mounted at the casing of the machine rotor. URANS computations are conducted using the elsA software. The flow distortion is simulated by a drop of stagnation pressure ratio at the inlet boundary condition. The study is focusing first on the ability of a pressure drop, imposed as an inlet boundary condition in CFD, to reproduce accurately the effect of a flow distortion. The analysis is conducted using singular value decomposition (SVD) and dynamic mode decomposition (DMD). A special attention is then paid, on the experimental level, to the arising of rotating stall, from the onset of the instability up to completely developed stall cells.

Blade Roughness Effects on Compressor and Engine Performance—A CFD and Thermodynamic Study

Degradation of compressors is a common concern for operators of gas turbine engines (GTEs). Surface roughness, due to erosion or fouling, is considered one of the major factors of the degradation phenomenon in compressors that can negatively affect the designed pressure rise, efficiency, and, therefore, the engine aero/thermodynamic performance. The understanding of the aerodynamic implications of varying the blade surface roughness plays a significant role in establishing the magnitude of performance degradation. The present work investigates the implications due to the degradation of the compressor caused by the operation in eroding environments on the gas turbine cycle performance linking, thereby, the compressor aerodynamics with a thermodynamic cycle. At the core of the present study is the numerical assessment of the effect of surface roughness on compressor performance employing the Computational Fluid Dynamics (CFD) tools. The research engine test case employed in the study comprised a fan, bypass, and two stages of the low pressure compressor (booster). Three operating conditions on the 100% speed-line, including the design point, were investigated. Five roughness cases, in addition to the smooth case, with equivalent sand-grain roughness (ks) of 15, 30, 45, 60, and 150 µm were simulated. Turbomatch the Cranfield in-house gas turbine performance simulation software, was employed to model the degraded engine performance. The study showed that the increase in the uniform roughness is associated with sizable drops in efficiency, booster pressure ratio (PR), non-dimensional mass flow (NDMF), and overall engine pressure ratio (EPR) together with rises in turbine entry temperature (TET) and specific fuel consumption (SFC). The performance degradation evaluation employed variables such as isentropic efficiency (ηis), low pressure compressor (LPC) PR, NDMF, TET, SFC, andEPR. The variation in these quantities showed, for the maximum blade surface degradation case, drops of 7.68%, 2.62% and 3.53%, rises of 1.14% and 0.69%, and a drop of 0.86%, respectively.

Aerodynamic simulation of a high-pressure compressor stage using the lattice Boltzmann method

PurposeThe present paper aims at evaluating the lattice Boltzmann method (LBM) on a high-subsonic high-pressure compressor stage at nominal regime.Design/methodology/approachThe studied configuration corresponds to the H25 compressor operated in a closed-loop test rig at the von Karman Institute. Several operating points are simulated with LBM for two grids of successive refinements. A detailed analysis is performed on the time-averaged flow predicted by LBM, using a comparison with experimental and existing RANS data.FindingsThe finest grid is found to correctly predict the mean flow across the machine, as well as the influence of the rotor tip gap size. Going beyond time-averaged data, some flow analysis is performed to show the relevance of such a high-fidelity method applied to a compressor configuration. In particular, vortical structures and their evolution with the operating points are clearly highlighted. Spectral analyses finally hint at a proper prediction of tonal and broadband contents by LBM.Originality/valueThe application of LBM to high-speed turbomachinery flows is very recent. This paper validates one of the first LBM simulations of a high-subsonic high-pressure compressor stage.

Alternative positions of internal heat exchanger for CO2 booster refrigeration system: Thermodynamic analysis and annual thermal performance evaluation

• Three CO2 booster systems with different locations of IHX are proposed. • Thermodynamics models of booster systems are developed and annual performance is studied. • S1 is the optimal solution system to improve COP, which is enhanced by 0.6% - 6.36%. • By adding IHX, the APF of CO2 booster system can be improved significantly. • IHX can improve annual performance in subtropical and tropical climatic regions. Refrigeration systems running on transcritical CO 2 cycle are considered an alternative to phase-down the use of hydrofluorocarbons (HFCs) in response to the Kigali Amendment. Responding to the need to improve the efficiency of CO 2 booster systems and to identify optimum system designs, thermodynamic models of three booster systems are proposed with the use of an internal heat exchanger (IHX). Findings from this study suggest that placing the IHX with a low-temperature fluid side at the suction line of high pressure stage compressor and high-temperature fluid side at the outlet of gas cooler represents the most optimal approach in improving the coefficient of performance (COP) of the booster system. When operating in the transcritical conditions, the COP values can be improved by 6.35% at the IHX thermal effectiveness of 0.8 and by 6.48% at the ratio of medium temperature load to low temperature load of 6. Using IHX can significantly reduce the compressor discharge pressure, which can be reduced by 0.55 MPa at the ambient temperature of 40 °C. Furthermore, by adding IHX, the annual performance factor of CO 2 system can be improved significantly by 1.68% and the annual total power consumption can be decreased by 6.48% in the tropical climate. It can be concluded that IHX can improve the COP values of a booster system when operating in the subtropical and tropical regions.

Impact Investigation of Stator Seal Leakage on Aerodynamic Performance of Multistage Compressor

In this paper, a numerical model based on the mass flow rate of seal leakage is presented, and a 3D numerical method of a multistage axial compressor with good engineering practicability is established. Validation consists of modeling a nine-stage axial compressor in all operating rotation speeds and calculating results of the performance characteristic curves in good agreement with test data. Comparisons are made against different cases of seal leakage mass flow rate for analyzing the impact of increasing seal leakage on the aerodynamic performance of the multistage axial compressor. The results indicate that the performance of the nine-stage axial compressor is degenerated faster and faster with seal leakage increasing in all operating working points, and the degeneration of performance of this compressor can be evaluated by the relationships of main performance parameters with the mass flow rate of seal leakage. Comparisons of flow distribution in the compressor for different cases of seal leakage also show that stators located in front stages of the multistage axial compressor are affected more seriously by the increasing seal leakage, and it can be confirmed that relatively larger flow losses in front stages bring significant impact on the decay of aerodynamic performance of a multistage axial compressor.

Study of pitting corrosion under actual operating conditions of a first stage compressor blade

In this study, pitting corrosion was investigated numerically and experimentally under actual operational conditions of a gas turbine compressor. First, the samples cut from a blade made of Custom 450 stainless steel were exposed to different corrosion times in a 3.5 wt% NaCl solution. After the creation of pits in the samples, the dimensions of the corrosion pits were measured using a 3D profilometer device. The stoichiometry equations governing the corrosion of the Custom 450 alloy in a chloride solution were extracted and applied to the models. The results of this simulation were validated by comparing the depth of the simulated pits and the experimental samples. Subsequently, simulation was conducted for 48 months on a compressor having pits. There is a good agreement between the simulation results and the results obtained from the Turbo machine laboratory of the Texas A&M University. In addition, pits grew under the influence of corrosion, joined together, and made an equivalent pit.

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.

Forced Response Excitation Due to the Stator Vanes of Two and Three Compressor Stages Away

Abstract State-of-the-art axial compressors of gas turbines employed in power generation plants and aero engines should have both high efficiency and small footprint. Thus, compressors are designed to have thin rotor blades and stator vanes with short axial distances. Recently, problems of high cycle fatigue (HCF) associated with forced response excitation have gradually increased as a result of these trends. Rotor blade fatigue can be caused not only by the wake and potential effect of the adjacent stator vane, but also by the stator vanes of two, three, or four compressor stages away. Thus, accurate prediction and suppression methods are necessary in the design process. In this study, the problem of rotor blade vibration caused by the stator vanes of two and three compressor stages away is studied. In the first part of the study, one-way fluid structure interaction (FSI) simulation is carried out. To validate the accuracy of the simulation, experiments are also conducted using a gas turbine test facility. It is found that one-way FSI simulation can accurately predict the order of the vibration level. In the second part of the study, a method of controlling the blade vibration is investigated by optimizing the clocking of the stator vanes. It is confirmed that the vibration amplitude can be effectively suppressed without reducing the performance. Through this study, ways to evaluate and control the rotor blade vibration are validated.

Aerodynamic Functional Diagnostics Based on Angular Momentum Transport Lines

AbstractIn analogy with the classical concept of mass-flux-based streamlines, we define angular momentum transport (AMT) lines as an aerodynamic functional diagnostic tool. The AMT lines are the ones whose tangents are given by the average angular momentum flux. The mathematical and physical properties of these AMT lines are exploited to study the generation, removal, and transport of angular momentum in turbomachinery components. We illustrate the concept by visualizing AMT lines in two relatively simple flows, namely, vaneless incompressible diffuser and von Karman flow (a model of centrifugal compressors). Next, we apply the proposed diagnostic tool to flow in a return channel. A return channel is a part of a multistage centrifugal compressor stage. Its principal function is to remove angular momentum. In this work, we apply the diagnostic tool of AMT lines to a Reynolds-averaged Navier Stokes (RANS) simulation and a wall-modeled large eddy simulation (LES) of flow in the return channel. We show that AMT lines give us insights into the AMT process that are otherwise not available with conventional visualization tools.

Aerodynamic Influence of a Bleed on the Last Stage of a Low-Pressure Compressor and S-Duct

Abstract This paper uses computational fluid dynamics to investigate the effect of an engine handling bleed situated on the outer casing downstream of the last rotor stage of a low-pressure compressor and upstream of the outlet guide vane and S-shaped duct. The model, validated against existing experimental data, utilized an unsteady Reynolds-averaging Navier–Stokes (RANS) solver incorporating a Reynolds stress closure to examine the unsteady component interactions. The results showed that at bleed rates less than 25% of the mainstream flow, the bleed effects were negligible. However, at higher bleed rates, performance was significantly degraded. A uniform flow extraction hypothesis was employed to separate the positional bias effects of the bleed from the bulk flow diffusion. This revealed that the bleed-induced radial flow distortion can significantly affect the outlet guide vane (OGV) loading distribution, which thereby dictates the position and type of stall within the OGV passage. Extraction of the rotor tip leakage via the shroud bleed, combined with the radial flow distortion, contributed to a 28% reduction in duct loss at 10% bleed and up to 50% reduced loss at 25% bleed. The actual amount of flow required to be extracted for an OGV stall to develop was 30%. That was independent of the bleed location and the type of stall. For bleeds up to 20%, the S-duct displayed a remarkable resilience and consistency of flow variables at duct exit. However, a stalled OGV deteriorated the radial flow uniformity that was presented to the high-pressure compressor.

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.

High Temperature Magnetic Sensors for the Hot Section of Aeroengines

Magnetic sensors are widely used in aeroengines and their health management systems, but they are rarely installed in the engine hot section due to the loss of magnetic properties by permanent magnets with increasing temperature. The paper presents and verifies models and design solutions aimed at improving the performance of an inductive sensor for measuring the motion of blades operated at elevated temperatures (200–1000 °C) in high pressure compressors and turbines. The interaction of blades with the sensor was studied. A prototype of the sensor was made, and its tests were carried out on the RK-4 rotor rig for the speed of 7000 rpm, in which the temperature of the sensor head was gradually increased to 1100 °C. The sensor signal level was compared to that of an identical sensor operating at room temperature. The heated sensor works continuously producing the output signal whose level does not change significantly. Moreover, a set of six probes passed an initial engine test in an SO-3 turbojet. It was confirmed that the proposed design of the inductive sensor is suitable for blade health monitoring (BHM) of the last stages of compressors and gas turbines operating below 1000 °C, even without a dedicated cooling system. In real-engine applications, sensor performance will depend on how the sensor is installed and the available heat dissipation capability. The presented technology extends the operating temperature of permanent magnets and is not specific for blade vibration but can be adapted to other magnetic measurements in the hot section of the aircraft engine.

Effect of Multi-Walled Carbon Nanotubes-Based Nanofluids on Marine Gas Turbine Intercooler Performance.

Coolants play a major role in the performance of heat exchanging systems. In a marine gas turbine engine, an intercooler is used to reduce the compressed gas temperature between the compressor stages. The thermophysical properties of the coolant running within the intercooler directly influence the level of enhancement in the performance of the unit. Therefore, employing working fluids of exceptional thermal properties is beneficial for improving performance in such applications, compared to conventional fluids. This paper investigates the effect of utilizing nanofluids for enhancing the performance of a marine gas turbine intercooler. Multi-walled carbon nanotubes (MWCNTs)-water with nanofluids at 0.01–0.10 vol % concentration were produced using a two-step controlled-temperature approach ranging from 10 °C to 50 °C. Next, the thermophysical properties of the as-prepared suspensions, such as density, thermal conductivity, specific heat capacity, and viscosity, were characterized. The intercooler performance was then determined by employing the measured data of the MWCNTs-based nanofluids thermophysical properties in theoretical formulae. This includes determining the intercooler effectiveness, heat transfer rate, gas outlet temperature, coolant outlet temperature, and pumping power. Finally, a comparison between a copper-based nanofluid from the literature with the as-prepared MWCNTs-based nanofluid was performed to determine the influence of each of these suspensions on the intercooler performance.

A Study of Influences of Inlet Total Pressure Distortions on Clearance Flow in an Axial Compressor

Abstract The rotating distortion generated by upstream wakes or low speed flow cells is a kind of phenomenon in the inlet of middle and rear stages of an axial compressor. Highly complex inflow can obviously affect the performance and the stability of these stages, and is needed to be considered during compressor design. In this paper, a series of unsteady computational fluid dynamics (CFD) simulations is conducted based on a model of a 1-1/2 stage axial compressor to investigate the effects of the distorted inflows near the casing on the compressor performance and the clearance flow. Detailed analysis of the flow field has been performed and interesting results are concluded. The distortions, such as total pressure distortion in circumferential and radial directions, can block the tip region so that the separation loss and the mixing loss in this area are increased, and the efficiency and the total pressure ratio are dropped correspondingly. Besides, the distortions can change the static pressure distribution near the leading edge of the rotor, and make the clearance flow spill out of the rotor edge more easily under near stall (NS) condition, especially in the cases with corotating distortions. This phenomenon can be used to explain why the stall margin is deteriorated with nonuniform inflows.

Implementing the principles of operating processes schematization and of performance loss distribution when designing long-stroke reciprocating compressor stages

To compress gases up to 10.0 MPa and higher, diaphragm and multistage reciprocating compressor units are used. They are heavy, massive and complicated. The results of analyzing laboratory examples of slow-speed long-stroke stages of reciprocating compressors while gas pressurizing up to such pressure have shown the possibility of solving the problems mentioned. Such stages differ from others, that is why we cannot use common design methods in this case. The performance design method for slow-speed long-stroke compressor stages with a discharge pressure of up to 10.0 MPa is based on the well-known principles of schematization of operating processes and distribution of performance losses. With regard to the experimental studies results (indicator and temperature diagrams, integral characteristics) we have found: the volumetric efficiency, which characterizes the ratio between real and theoretical compressor performance, its main components: the volume coefficient, the pressure coefficient, the temperature coefficient, and the leakage coefficient. Dependences are obtained for calculating these coefficients, which differ significantly from the known dependences for calculating high-speed compressor stages. The analysis of the influence of various factors on the performance losses of low-capacity long-stroke compressor stages of medium and high pressure revealed that the gaps in the working chamber have the maximum impact on the performance losses, the leakage coefficient significantly depending on the discharge pressure change. The volume coefficient has much less impact due to the increased stroke-to-diameter ratio of the piston, which provides a small amount of relative dead volume. The developed method can be used for preliminary calculation of compressor stages of the considered type.

Part load performance of single and two stage compressors — a comparative experimental study in a R410A chiller unit

The rising cooling needs for the residential and commercial air conditioning sectors and the requirement of higher SEER values increased the focus on energy efficient part load performance. Although several studies exist on the topic of capacity modulation, there is no comprehensive study that compares different modulation strategies in the same experimental facility. In a first step, this paper aims to experimentally compare two different capacity controls for scroll compressors (single speed vs two stage compressor) using the same R410A water chiller having a nominal cooling capacity of 8kW. The tests were conducted according to AHRI Standard 551/591 (2018). The Integrated Part Load Value (IPLV.SI) was used to fairly compare all the different capacity modulation strategies. Based on the experimental results, differences on the order of 8% in IPLV.SI are found.