Compressor Stage Research Articles

Understanding the Flow Field in a Highly Loaded Tandem Compressor Cascade

Abstract Large flow turning in compressor cascades with single airfoils requires an effective control of the boundary layer growth under the diffusing flow. An alternative approach consists of distributing the loading over two subsequent airfoils, using a so-called tandem blade, to, in some measure, restart the boundary layer before flow separation occurs. It is, however, not always clear whether the benefits of the two-blade setup justify the additional manufacturing complexity. The present work explores the tandem blade concept using a gradient-based optimization method to produce an efficient, highly loaded compressor cascade blade directly. A comparison between two-dimensional single and tandem configurations is first presented to clarify the benefits of one over the other. The geometry is optimized for each concept using a gradient-based optimization technique to improve the pressure loss coefficient at multiple operating points for a given flow-turning constraint. While the optimized single and tandem blade designs have similar performance, the lower solidity of the latter provides a lighter compressor stage for the considered operating range. A three-dimensional tandem compressor cascade based on the two-dimensional study is then optimized to account for secondary flows. The aerodynamic performance and the operating range are assessed and compared, along with a study of the physical phenomena surrounding the tandem configuration. The resulting geometry presents similar non-conventional features observed during the two-dimensional study that exploits the flow mechanism of two-airfoil configurations.

Analysis of The Flow-Field Distorsions Induced by Pneumatic Pressure Probes In Test Bench Experiments On Centrifugal Compressors: Are We Testing Right?

Abstract Experimental testing at the test rig still plays a key role in the development of new centrifugal compressor designs for industrial applications. Although a wide set of different probes is used as per common practice, the overall impact of the probes? dimensions on the stage performance and rangeability is often assumed to be negligible. In this study, we aimed at quantifying the intrusiveness effects of pneumatic pressure probes on the flow-field inside a vaned diffuser using a model stage of an industrial centrifugal compressor, for which different experimental setups were tested at the University of Florence test rig. Performance and rangeability were initially assessed using a conventional set of pneumatic probes inserted in the flow path. Then, comparative analyses were realized through dedicated tests by either removing all probes at the diffuser inlet or introducing only a single pressure probe in the upstream section and varying its angular position in the measuring section. Experiments revealed an impact of upstream probes much higher than expected, with a decrease in efficiency and a substantial reduction of stall margin. Moving from this experimental evidence, accurate 3D CFD simulations of the entire stage were carried out including the actual geometry of the probes for two probes configurations and three operating points. Simulations highlighted the effect probes? wakes on downstream vanes and the induced blockage effect, which both appear in line with experimental evidence. As a handout, a recommendation on the optimal probe positioning is suggested.

Upgraded one-dimensional code for the design of a micro gas turbine mixed flow compressor stage with various crossover diffuser configurations

Abstract Micro Gas Turbine (MGT) designers often have the need for a design tool, capable of rapidly providing feasible design options. The upgrade of such an existing MATLAB® based mixed flow compressor design application (1D App) is discussed. In MGT engines, diffuser design largely influences the compressor stage performance. Crossover diffusers provide superior efficiency and pressure ratio performance compared to legacy type diffusers, albeit at the expense of operating range. The reduced operating range attributed to single vane crossover diffusers can be alleviated by splitter vane, tandem vane, and combined configurations. Therefore, the 1D App is updated to allow the user the ability to design these additional crossover diffuser configurations. Performance results are verified using Numeca/FINE™ Turbo CFD software. It is shown that the upgraded 1D App provides predicted performance results within 1 % of CFD results for both total-to-total efficiency and pressure ratio.

Uncertainty quantification based on active subspace dimensionality-reduction method for high-dimensional geometric deviations of compressors

Compressors are inevitably exposed to diverse geometric deviations from manufacturing errors and in-service degradation. Consequently, the evaluation of performance uncertainties becomes of utmost importance for compressors in engineering application. However, the presence of high-dimensional and strongly nonlinear geometric deviations poses significant challenges in efficiently and accurately assessing the performance uncertainties of compressors. This study proposes an active subspace-based dimensionality-reduction method for high-dimensional uncertainty quantification (UQ) of compressors. Based on the active subspace (AS) method, a dimensionality-reduction high-precision artificial neural network is raised to solve the dimension disaster problem for high-dimensional UQ. Additionally, a data-driven approach is used to calculate the gradient of the quantity of interest, addressing the issue of high computational cost during the AS dimensionality reduction process. Furthermore, the Shapley method is applied to explore the influence mechanism of geometric uncertainties on performance deviations of compressors. The UQ of one transonic compressor stage at design point and near stall point is conducted by the proposed method. The findings show that the original 24-dimensional uncertainties are reduced to three-dimensional uncertainties by using this method. Consequently, the required sample size is reduced by 75% while maintaining almost unchanged model accuracy. The findings reveal that the sweep and stagger deviation of the rotor are key uncertainties on the performance of the compressor. The dispersion in efficiency is attributed to variations in shock wave position and intensity, while the dispersion in total pressure ratio is primarily affected by changes in rotor work capacity. Moreover, the dispersion at near stall is 50% higher than that at design point. Therefore, when studying UQ, it is important to pay closer attention to the performance dispersion at near stall conditions.

Insights Into Turbocharger Centrifugal Compressor Stability Using High-Fidelity CFD and a Novel Impeller Diffusion Factor

Abstract Turbocharger technologies are widely applied in the modern automotive industry to meet increasingly stringent carbon emission standards. To support the continuous enhancement of these tactics, delivering improvement to the performance and stability of turbocharger compressors at low mass flow rate regimes is of paramount importance. The present paper investigated the stability of a turbocharger compressor stage equipped with a ported shroud, using high-fidelity CFD simulations. It was found that the recirculated flow passing through the ported shroud device had primary influence on the flow condition in the vicinity of the impeller tip region (80% to 100% span). This discovery facilitated the definition of a modified impeller diffusion factor (DF), which was applied to the CFD results for three different compressor geometry configurations to lead to the provision of quantitative metrics for impeller stability. Furthermore, the observed change in the limiting impeller DF along the compressor surge line indicated that the dominant components governing the instability of the entire stage had switched, i.e., the stalling behavior of the compressor stage had transitioned from impeller dominated stall to vaneless diffuser dominated stall. Additionally, the limiting impeller DF clearly showed that one of the compressor configurations with a change in the hub endwall recess around the rear of the impeller improved the impeller stability by 3%. By comparison, another configuration with a change to strut geometry in the ported shroud recirculation channel reduced the impeller stability by 10%.

Improving near-stall performance of axial flow compressors using variable rotor tandem stage. A steady analysis

This paper investigates the idea of substituting the first stage of a conventional axial-flow compressor and its inlet guide vane (IGV) row with a variable rotor tandem (VRT) stage to reduce the size and weight of axial-flow compressors. It is expected that a tandem stage gains higher pressure ratio than a conventional stage, and simultaneously, its variable stagger rotor provides more flexibility in handling unsteady phenomena at low mass flow rates. A three-dimensional numerical model based on Reynolds-Averaged Navier–Stokes equations is developed to conduct this investigation. To validate the numerical model, experimental tests are conducted on a tandem single-stage axial-flow compressor test bench in the Dana laboratory of Amirkabir University of Technology, to compare the performance map with numerical results. The validated numerical model is then employed to simulate the performance of both front and aft variable rotor blades. The proposed tandem stage effectively performs the IGV duty, suppresses instabilities at low mass flow rates, and eventually extends the compressor's stable operational range. This proves that a VRT stage is an efficient active surge-control system tool. The detailed survey of the three-dimensional flow field reveals that the aft tandem rotor blade stagger angle is more effective in increasing the stall margin and decreasing the minimum operational mass flow rate than the front row of blades. It is observed that when the aft blades rotate in the direction of decreasing the tandem blades' gap area, the flow momentum increases and causes the compressor stability at lower mass flow rates.

Investigation of mechanisms on leading edge vortex generation and suppression in vaned diffuser of a centrifugal compressor

This study numerically investigates the mechanisms for generating and suppressing the leading edge vortex (LEV) in a vaned diffuser of a centrifugal compressor, aiming to enhance the compressor's stable operating range. Detailed analyses focus on the mechanisms of the rotating stall and the initiation process of the LEV. Notably, the end wall contouring method is applied to vaned diffuser to suppress the LEV and improve diffuser stability. The suppression mechanisms of the LEV are examined to deepen the understanding of stability enhancement mechanisms. Results demonstrate that, under stall conditions, the vaned diffuser experiences two-cell rotating stalls moving opposite to the impeller rotation direction, each rotating at approximately 12% of impeller rotation speed. The formation of diffuser stall is attributed to a longitudinal vortex developing from the LEV and causing the passage blockage. The increasing spanwise distortion of leading edge (LE) incidence angles and circumferential distortion of static pressures at the diffuser inlet promote the evolution of the longitudinal vortex from LEV. The end wall contouring of vaned diffuser effectively enhances the compressor's stable operating range by 15.10% and 8.90% toward lower flow rate conditions for machine Mach numbers 1.3 and 1.2, respectively. At near-stall points, the end wall contoured diffuser reduces spanwise distortion of LE incidence angles and circumferential distortion of static pressures at the diffuser inlet, mitigates the LE flow separation and corner separation, suppresses the generation and development of the LEV, delays the onset of rotating stall in vaned diffuser, and thereby improves the stable operating range of centrifugal compressor stage.

Modelling carbon capture from power plants with low energy and water consumption using a novel cryogenic technology

Global warming mitigation requires modern energy generation technology to integrate low-cost carbon capture techniques. Despite volumes of research on the latter, the significantly high unit energy (MWh) of capture and high water consumption, do not enable any current carbon capture techniques to be adopted at a scale. To address this problem, we theoretically explore a novel low-cost carbon capture technology (NLCCT). In this paper, a detailed thermodynamic analysis of the effects of the operating conditions and parameters of NLCCT, such as the initial feed gas concentration, compressor intercooler temperature, number of compressor stages, efficiencies of the compressors, and turbines, ambient temperatures on the energy consumption for CO2 capture (ECC) is carried out with a view to obtain costs much lower than current techniques. The study identifies the inlet feed concentration, turbine and compressor efficiencies, and compressor intercooler temperatures (heat transfer effectiveness) as the significant parameters. For a drop of 20 % efficiency, the ECC was found to increase by 40 %, and for an increase in inlet CO2 feed concentration from 4 % to 10 %, the ECC was found to decrease by 31 %. Considering the effect of intercooler temperatures, the ECC increases by 25 % when the intercooler temperature is raised by 6 K. The effect of ambient conditions was also analyzed and observed that for an increase of 27C in the ambient temperature, the ECC would rise by 15 %.

Research on Effect of Endwall Contouring of Vaned Diffuser on Stable Operating Range of Centrifugal Compressor

Abstract The study investigates the impact of diverse endwall contouring strategies on the operational stability of a high-pressure ratio centrifugal compressor stage. Employing numerical simulation techniques, this study examines the implications of contoured wall positions, specifically convex profiles at the pressure side and/or concave profiles at the suction side, as well as asymmetric endwall contouring, on both stage performance and internal flow field within the centrifugal compressor. And the stability enhancement mechanism of endwall contouring is clarified. The results demonstrate that employing full circular contoured vaned diffusers on the hub-side wall can achieve a maximum improvement in stall margin of 15.10%. Additionally, the diffuser featuring solely convex profiles at the pressure side on the hub-side wall exhibits an augmented stall margin by 8.39%, with a reduction in performance loss at low flow rate conditions. The asymmetric endwall contouring improves stall margin by 8.39%, and mitigates the performance loss across the entire operating range. At the near-stall point, the contoured endwalls redirect fluid towards the suction side, thereby mitigating the incidence at the vane leading edge and attenuating the interaction between the impeller trailing edge vortex and the diffuser leading edge vortex. Moreover, concave profiles serve to accelerate the fluid on the suction side along the flow direction, thereby suppressing the corner separation and weakening passage blockages. Simultaneously, the advantages of endwall contouring in vane diffusers extend to the improvement in the circumferential non-uniformity of flow fields and the enhancement in diffuser stability.

Investigation of vaned diffuser endwall contouring technology for improving the stable operating range of a centrifugal compressor with an asymmetric volute

AbstractIn this study, nonaxisymmetric endwall contouring technology is employed as a passive control strategy to enhance the stable operating range of a centrifugal compressor with an asymmetric volute. The endwall contouring method is applied to the hub‐side wall of the vaned diffuser within a centrifugal compressor stage. Instead of utilizing a flat hub as the baseline diffuser, various diffuser configurations featuring hub‐side contoured endwalls are explored, with adjustments in the peak radial position and peak height of the convex and concave profiles. Numerical simulations are conducted to evaluate the performance of the centrifugal compressor stage with both the baseline and contoured vaned diffusers. The investigation explores the underlying flow control mechanisms and establishes the effective endwall contouring guidelines. The findings highlight the effectiveness of endwall contouring in enhancing the stable operating range of centrifugal compressors. Notably, a substantial enhancement of 15.1% in the stable operating range is achieved by employing the contoured diffuser with the peak radial position near the vane leading edge (LE) and the peak height at 20% of the vane height. The endwall contoured diffusers reduce the positive LE incidence angles to direct the fluid toward the suction side (SS), and increase the radial velocity near the SS to suppress the hub‐suction corner separations in the diffuser passages near the volute tongue. These improvements collectively contribute to the enhancement of diffuser flow characteristics. Finally, a set of effective endwall contouring guidelines is proposed to guide the endwall contouring design for enhancing the stable operating range of centrifugal compressors.

Numerical investigation of the potential of multi-row optimization of an axial compressor stage with a tandem vane stator

Numerical investigation of the potential of multi-row optimization of an axial compressor stage with a tandem vane stator

Experimental and Full-Annulus Simulation Analysis of the Rotating Stall in a Centrifugal Compressor Stage with a Vaned Diffuser

Experimental and Full-Annulus Simulation Analysis of the Rotating Stall in a Centrifugal Compressor Stage with a Vaned Diffuser

Sensitivity analysis and robust optimization to high-dimensional uncertainties of compressors with active subspace method

The performance of compressors often deviates significantly from the design value due to the substantial geometric deviations arising from manufacturing errors and in-service degradation. Robust design optimization (RDO) plays a crucial role in ensuring the high performance and reliability of compressors. The study proposes an active subspace-based dimensionality-reduction RDO method for high-dimensional uncertainties of compressors. By employing the active subspace (AS) method, a dimensionality-reduction artificial neural network (ANN) is raised to solve the "dimension disaster" problem for high-dimensional uncertainty quantification (UQ) analysis. Additionally, to decrease the computation cost of the original AS method, a data-driven method is utilized to calculate the gradient of the quantity of interest (QoI) within the AS dimensionality-reduction framework. Furthermore, Shapley method is applied to study the sensitivity of geometric variables and their variation direction on the mean and variance of the performance, and the robustness improvement method by only modifying key geometric variables based on Shapley results is discussed. The RDO of high dimensional geometrical uncertainties of a transonic compressor stage is studied by using this method. The results demonstrate a reduction from the original 24-dimensional uncertainties to a 4-dimensional space, achieved through this method. Consequently, the required sample size is reduced by 75 % while maintaining nearly unchanged model accuracy. Furthermore, the findings reveal that after RDO, the mean efficiency increases by 1.04 %, with a 42 % reduction in variance. Compared to RDO, which requires simultaneous optimization of all 24 geometric variables, the robustness improvement based on Shapley only necessitates modifying 4 key geometric variables, resulting in a substantial increase in both the mean and variance of efficiency.

On setting boundary conditions for calculations of turbomachines in 3-d computer packages

Unfortunately, in a growing number of publications of the results of the study of turbomachines in ANSYS and other packages, insufficient attention is paid to the choice of the flow model (absolute or relative) and the setting of boundary conditions in the areas adjacent to the rotating impeller: shroud seal, diaphragm (end) seals, gap over shroudless blades, steam balance holes, etc.As a rule, these areas contain such elements of geometry as a “ledge”, “projection”, “crest”, the flow around which, and, consequently, their resistance depends on the angle of incoming flow relative to the barrier. In the stages of turbomachines in sections behind fixed blade units, angle α1 between the flow velocity vector (in absolute motion) and the circumferential velocity is 10°÷25° (for steam turbine stages), 20°÷35° (for gas turbine stages) and 40°÷75° ((for axial compressor stages). It is shown (using the example of viscous flow around the “ledge + projection” system) that a change in the angle in the range of 15°÷90° changes the resistance by almost 2 times.The choice of a particular type of flow model (absolute or relative) for the domain primarily determines the value of the angle of interaction of the flow lines with an obstacle: it remains close to α1, or close to β1 = α1 + 20°÷70°, which significantly changes the qualitative picture of the viscous flow around the characteristic elements of geometry (“ledge”, “projection”, “crest”), and the values of integral parameters (flow rate, power, efficiency). The lack of recommendations on the correct choice of boundary conditions will inevitably lead not only to the inability to compare the results of various published studies, but also to their objective value. Therefore, the need to substantiate proposals for the choice of flow models in the areas of labyrinth seals, steam balance holes, gaps over shroudless blades, etc. is very relevant.

The loss mechanisms and vortex dynamics of rotor platform step geometries in a low-speed research compressor stage

The construction and assembly of an axial compressor blade row introduces real geometric features, which leads to the deviation of flow characteristics from the intended design. According to the realistic assembly errors observed in the four-stage low-speed research compressor developed by Shanghai Jiao Tong University, two representative step geometries at the rotor platform were abstracted. Unsteady simulations were carried out in the stage environment to study the influence of hub discontinuities on compressor performance and loss mechanisms. The comparisons of smooth hub, elevated hub (EH), and a wedged hub (WH) demonstrated that all the step geometries resulted in increased loss inside both rotor and stator rows. The loss induced by WH was more sensitive to step heights than EH. Under the influence of forward facing step, rotor incidence was reduced and horseshoe vortex (HV) was exacerbated, leading to the accumulation and separation of low momentum fluids at the pressure side corner. This led to a redistribution of losses across the span. In the lower half span, the relative total pressure loss increased almost linearly with step heights, while in the tip region, it reduced slightly. The backward facing step generated the step corner vortices (SCV) that shed and propagated downstream. The SCV interacted with the HV and cavity leakage vortex in the stator passage. Coupled with the augmented stator inflow angle, the corner separation at the suction side was dramatically intensified and led to a significant loss increase.

Flow Analysis of Centrifugal Compressor using ANSYS

Abstract: The aerodynamic and structural designs of the impeller are vital to the functioning of the entire compressor stage in centrifugal compressors. In contrast to previous years, there has been a greater demand for industrial centrifugal compressors to operate with greater efficiency and operational range. Although the efficiency of low-pressure-ratio centrifugal compressors has nearly achieved its maximum, intermediate and high-pressure ratio machines still lag behind the state-of-the-art in terms of efficiency. Maintaining maximum efficiency while extending the working range is a challenging task in centrifugal compressor design. Furthermore, a single design must be usable globally due to product globalisation, even in the face of technological disparities and variances in electricity frequency of up to 10 percent in some nations. Often operating at different rotational speeds, centrifugal compressors add another layer of complexity to the impeller structural design. A full-3D impeller design that is optimised for both structural integrity and aerodynamic performance is shown in this study .For air conditioning systems of the future, centrifugal compressor performance must be improved. This research focusses on systematic numerical simulation, a computationally efficient method of optimising centrifugal compressor performance. ANSYS simulations are used in conjunction with a thorough literature review on the aerodynamic behaviour of materials used in centrifugal compressors. For the purpose of performing flow simulations under subsonic circumstances with suitable boundary conditions and grid independence checks, geometric modelling was completed using CATIAV5R22 and imported into ANSYS. Tests were conducted using various aluminium materials and mass flow rates, and enhanced efficiency findings were confirmed by comparison with experimental data. Materials Processing and Characterisations is in charge of peer review.

Darmstadt Open Test Case: Experimental investigation of forced response phenomena in a transonic compressor stage

Darmstadt Open Test Case: Experimental investigation of forced response phenomena in a transonic compressor stage

Unsteady Analysis of Secondary Vortex Formations Within An Axial Compressor Stage with Tandem Rotor

Unsteady Analysis of Secondary Vortex Formations Within An Axial Compressor Stage with Tandem Rotor

Performance Analysis and Aerodynamic Design of Axial Flow Compressors

The main objective of the paper is to analyze theperformance of axial flow compressors and to generate asystematic design approach which enables to designsubsonic flow ones. In order to investigate the validity of this approach, the LP axial flow compressor of the RR SpeyMK511 turbofan engine is taken as an example. The designcalculations were based on thermodynamics, gas dynamics,fluid mechanics and empirical relations. The flow isassumed to be of two-dimensional compressible type withconstant axial and rotor blade velocities with a free-vortexswirl distribution. Design calculations includethermodynamic properties of the working fluid, number ofcompressor performance parameters such as, stagetemperature rise and number, flow and blade angles (bladetwist), velocity triangles and relative inlet Mach number atrotor blades tips as well as blades tip and hub diameters. Arepeated calculation is made to determine these parametersalong compressor stages. The variation of velocity whirlcomponents, air and blade angles, deflection and degree ofreaction from root to tip of the blades were also determined.The twist of the blades along the blade length is setaccording to the recommended values in order to obtainsmooth blade twist profile.

Investigation of an Inter-Compressor S-Duct Using the Lattice Boltzmann Method

Abstract This article presents the study of a subsonic inter-compressor S-duct. Numerical simulations are performed using large-eddy simulation (LES) based on a compressible hybrid thermal lattice Boltzmann method (LBM) implemented within the ProLB solver. Comparisons are made between the LES–LBM results, Reynolds-averaged Navier–Stokes (RANS) computations, and experimental measurements on a representative S-duct taken from the European project AIDA. Several cases with increasing complexity are addressed where the different rows surrounding the duct are gradually included in the computations. The effects of each row on the flow field development and loss levels are studied. The goal is to evaluate the ability of the LES–LBM to recover the aerodynamic behavior and the total pressure loss evolution within the duct. Results show that the LES–LBM retrieves the correct flow evolution inside the S-duct compared to the experiment and previous RANS results. The case where the upstream stator row or the low-pressure compressor stage is integrated shows an increase in total pressure loss, as previously observed in the literature, and a more developed flow field with complex flow features contributing to the loss generation. To further analyze the loss mechanism, an entropy-based approach is presented and highlights that most losses are generated close to the hub wall due to the migration of the upstream stator wakes.