Modern Compressor Research Articles

Adaptive jet flow control for performance improvement via coupling stator-blade with rotor-tip injections in a transonic compressor

With the increase in the blade load of modern compressors for aircraft engines and gas turbines, flow separation is always generated in stator and rotor blade passages. This not only deteriorates the compressor efficiency but also blocks the flow path that may induce instability such as stall or surge. In order to tackle these issues, an adaptive jet flow control system coupling stator-blade with rotor-tip injections is proposed, and the goal of this paper is to investigate its effects on the aerodynamic performance of an in-house 1.5 stage transonic compressor. The developed flow control system integrates a jet configuration on the stator-blade surface and on the rotor casing surface respectively. The first step to construct this system is designing the jet configurations, and the Coanda effect is adopted to achieve wall-attachment injection. Second, the flow control system is intended to execute real-time monitoring on the operating condition of the compressor and can inject air with appropriate mass flow rate to eliminate flow separation and rotating stall. To do this, an optimal injection mass flow rate prediction model is established via a back propagation neural network algorithm. It aids to ensure an adaptive jet mass flow rate control during the compressor operation. Then, the effectiveness of the adaptive jet flow control system is evaluated using numerical simulations, and the influence mechanism of the injection mass flow on flow fields and aerodynamic performance is analyzed. Results indicate that compared to the prototype compressor, the compressor with adaptive stator-blade and rotor-tip injections can significantly increase efficiency and improve stability margin.

Study of nonequilibrium regenerative heat exchange in positive displacement compressors

BACKGROUND: An important task in the design of thermal engines is the evaluation of power losses in different processes to determine the installation efficiency. One of the processes that generates power losses in the compressor is the process of nonequilibrium regenerative heat exchange (NRHE) of the compressed gas with the working cavity walls. This heat exchange occurs at a significant temperature difference, which can lead to noticeable power losses. Modern compressor design methods do not consider losses from nonequilibrium regenerative heat exchange and do not describe them separately from other types of losses. AIM: This study investigates the mechanism of losses during regenerative heat exchange between the gaseous working flow out and the working cavity walls and evaluates the scale of these losses. METHODS AND RESULTS: This work provides a qualitative description of the mechanism for the formation of losses from NRHE. As a result of solving the problem of unsteady thermal conductivity, an analytical expression for calculating such losses is derived based on the Gouy-Stodola theorem. CONCLUSIONS: The study revealed that power losses from NRHE account for a significant share in the overall balance of machine losses. Parameters influencing the magnitude of these losses were also determined, and recommendations were developed to reduce them.

Influence mechanism of tip clearance on flutter stability with different aerodynamic coupling effects for transonic compressor

Flutter is a highly destructive aeroelastic problem in modern compressors, which hinders the improvements of aero-engine performance and reliability. Because the aerodynamic work is mainly concentrated in the blade tip region, the tip clearance which is associated to the complex tip clearance flow has a considerable effect on the flutter stability. In this paper, in order to investigate the influence mechanism of the tip clearance on the flutter stability at different nodal diameters (ND), a series of compressor models were established with different tip clearances based on a transonic compressor rotor. The aerodynamic damping at each ND is obtained by influence coefficient method (ICM), while phase-shifted boundary method (PSB) is adopted to analyze the influence of the tip clearance on the flutter stability at different NDs. The results indicate the worst flutter stability for each tip clearance always appears at ND=1, where the aerodynamic damping exhibits a nonmonotonic trend of increasing first and decreasing thereafter along with the rising tip clearance. Two kinds of action are exerted by the tip clearance flow on the flow structures to alter the flutter stability: one is the tip clearance vortices which is generated on the suction side and impinge on the pressure side; the other is the interference of tip clearance vortices to the shock wave and to the flow separation. Besides, the influence of the changing aerodynamic coupling effect makes the above action bring about more drastic fluctuations to the unsteady pressure. At larger NDs, the unsteady pressure amplitude and phase fluctuate more sharply for larger tip clearances. Therefore, diverse change trends of the aerodynamic damping with the increasing tip clearance appear at different NDs, while the impingement of tip clearance vortices on the adjacent pressure side has a constantly stabilizing effect at different NDs.

Development of a deviation package method for low-cost robust optimization in compressor blade design

Manufacture variations can greatly increase the performance variability of compressor blades. Current robust design optimization methods have a critical role in reducing the adverse impact of the variations, but can be affected by errors if the assumptions of the deviation models and distribution parameters are inaccurate. A new approach for robust design optimization without the employment of the deviation models is proposed. The deviation package method and the interval estimation method are exploited in this new approach. Simultaneously, a stratified strategy is used to reduce the computational cost and assure the optimization accuracy. The test case employed for this study is a typical transonic compressor blade profile, which resembles most of the manufacture features of modern compressor blades. A set of 96 newly manufactured blades was measured using a coordinate measurement machine to obtain the manufacture variations and produce a deviation package. The optimization results show that the scatter of the aerodynamic performance for the optimal robust design is 20% less than the baseline value. By comparing the optimization results obtained from the deviation package method with those obtained from widely-used methods employing the deviation model, the efficiency and accuracy of the deviation package method are demonstrated. Finally, the physical mechanisms that control the robustness of different designs were further investigated, and some statistical laws of robust design were extracted.

Understanding the Effect of Three-Dimensional Design in Tandem Blade

Abstract In order to maximize the pressure ratio and efficiency, compressor designers have tried several unconventional design approaches. Tandem blading is one such unconventional design that promises a higher pressure ratio per stage through a higher diffusion factor. The nozzle shape created between the forward and aft blades of a tandem configuration acts as a passive boundary layer control mechanism. The boundary layer growth over the aft rotor is therefore effectively controlled with the help of this gap-nozzle flow. The flow complexity is likely to increase in the case of a tandem rotor due to the twin leakage vortices, twin wake regions, and their interaction with the hub and casing boundary layers. Modern compressor blades are often designed with three-dimensional blade techniques such as sweep, lean, dihedral, end bent, etc., to reduce the various losses and achieve optimum performance. However, to the best of the author’s knowledge, the effect of 3D blade designs on the performance of tandem rotors has not been fully explored so far. A comprehensive numerical investigation is undertaken to understand the effect of 3D designs on the performance of tandem blades. Axial sweep and dihedral failed to improve the performance of the tandem rotor. Significant improvement in the stall margin is observed for the forward chordwise-swept and negative lean tandem rotors and is largely attributed to lower tip incidence. The performance penalty of the forward-swept and negatively leaned cases can be reduced by integrating compound or variable lean and sweep into the design.

On Choosing the Optimal Impeller Exit Velocity Triangles in Preliminary Design

Abstract Impeller discharge flow plays an important role in centrifugal compressor performance and operability for two reasons. First, it determines the work factor and relative diffusion for the impeller. Second, it sets the flow into the downstream stationary diffusion system. The choice made in the preliminary design phase for the impeller exit velocity triangle is crucial for a successful design. The state-of-the-art design approach for determining the impeller exit velocity triangle in the preliminary design phase relies on several empirical guidelines, i.e., maximum work factor and diffusion ratio for an impeller, the optimal range of absolute flow angle, etc. However, as modern compressors continue pushing toward higher efficiency and higher work factor, this design approach falls short in providing exact guidance for choosing an optimal impeller exit velocity triangles due to its empirical nature as well as the competing mechanism of the two trends. In light of this challenge, this paper introduces a reduced-dimension, deterministic approach for the design of the impeller exit velocity triangle. The method gauges the design of the impeller exit velocity triangle from a different design philosophy using a relative diffusion effectiveness parameter and is validated using six impeller designs, representative of applications in both turbochargers and aero engines. Furthermore, with the deterministic method in place, optimal impeller exit velocity triangles are explored over a broad design space, and a one-to-one mapping from a selection of impeller total-to-total pressure ratios and backsweep angles to a unique optimal impeller exit velocity triangle is provided. This new approach is demonstrated, and discussions regarding the influences of impeller total-to-total pressure ratio, isentropic efficiency, and backsweep angle on the optimal impeller exit velocity triangle are presented.

Aeroelastic Stability of Combined Plunge-Pitch Mode Shapes in a Linear Compressor Cascade

Modern aeroengine designs strive for peak specific fuel and thermal efficiency. To achieve these goals, engines have more highly loaded compressor stages, thinner aerofoils, and blended titanium integrated disks (blisks) to reduce weight. These configurations promote the occurrence of aeroelastic phenomena such as flutter. Two important parameters known to influence flutter stability are the reduced frequency and the ratio of plunge and pitch components in a combined flap mode shape. These are used as design criteria in the engine development process. However, the limit of these criteria is not fully understood. The following research aims to bridge the gap between semi-analytical models and modern compressors by systematically investigating the flutter stability of a linear compressor cascade. This paper introduces the plunge-to-pitch incidence ratio, which is defined as a function of reduced frequency and pitch axis setback for a first flap (1F) mode shape. Using numerical simulations, in addition to experimental validation, aerodynamic damping is computed for many modes to build stability maps. The results confirm the importance of these two parameters in compressor aeroelastic stability as well as demonstrate the significance of the plunge-to-pitch incidence ratio for predicting the flutter limit.

Trapped Acoustic Modes in an Axial Multi-Stage Compressor Leading to Non-Synchronous Blade Vibrations

Non-synchronous blade vibrations have been observed in an experimental multi-stage high-speed compressor setup at part-speed conditions. A detailed numerical study has been carried out to understand the observed phenomenon by performing unsteady full-annulus Reynolds-Averaged Navier–Stokes (RANS) simulations of the whole setup using the solver elsA. Several operating conditions have been simulated to observe this kind of phenomena along a speedline of interest. Based on the simulation results, the physical source of the non-synchronous blade vibration is identified: An aerodynamic disturbance appears in a highly loaded downstream rotor and excites a spinning acoustic mode. A “lock-in” phenomenon occurs between the blade boundary layer oscillations and the spinning acoustic mode. The establishment of axially propagating acoustic waves can lead to a complex coupling mechanism and this phenomenon is highly relevant in understanding the multi-physical interactions appearing in modern compressors. It is shown that aerodynamic disturbances occurring downstream can lead to critical excitation of rotor blades in upstream stages due to an axially propagating acoustic wave. The paper includes the analysis of a relevant transient test and a detailed analysis of the numerical results. The study shows the capability and necessity of a full-annulus multistage simulation to understand the phenomenon.

Influence of the inlet boundary layer on non-axisymmetric endwall contouring effects in a linear compressor cascade

The modern aero-engine compressors are commonly designed to be highly loaded which can introduce strong three-dimensional flow separation, and effective flow control is of vital importance in improving compressor performance. Non-axisymmetric endwall contouring nowadays is conducted on axial compressors to control the flow separation. In this paper, a contoured endwall is designed based on a linear CDA compressor cascade. The genetic algorithm and ANN surrogate model are used in the optimization process of the hub contouring, based on RANS simulations. Three inlet boundary layers which follow the fully turbulent boundary layer assumption are generated with different thickness, and the influence mechanism is numerically investigated. The optimal endwall with streamwise concave near the suction side provides great performance improvement at all inlet boundary layers at the aerodynamic design point. The endwall flow near the suction side is accelerated by the contoured endwall with cross-flow pressure gradient altered, suppressing the corner separation. The main pressure loss is closely related to one streamwise vortex that blocks the endwall flow. A thicker inlet boundary layer can induce early forming and increased strength on the streamwise vortex, enhancing the cross-flow movement and its interaction with the blade surface flow. In addition, the pressure loss near endwall in the inlet boundary layer is found more dominant on the endwall contouring effect than the boundary layer thickness.

METHODOLOGY FOR THE UTILIZATION OF WASTE HEAT BY AIR-COOLED COMPRESSORS

Compressed air is still a valid helper in many applications today, where it is necessary, for example, to move work equipment, pistons or it is used for cooling as a cooling medium. The producer of compressed air are air compressors, which need an external source for its production, usually an electric or internal combustion engine. Almost all the energy that is supplied to the compressor is always converted to heat during compression, regardless of the type of compressor. This carries the risk of overheating and therefore the cooling system must be optimally designed. Thus, during the compression of the air, a large part of the electrical energy supplied to the compressor is converted into heat, and only a small part of the supplied energy is in the compressed air. In the case of oil or water-cooled compressors, the exchangers can be used directly to obtain energy "for free". In the case of air cooling, a slight energy gain can only be achieved by modifying the exhaust hot air ducts. This energy can be used efficiently to heat water or heat buildings, instead of being uselessly ventilated. Modern compressors are already adapted for the use of waste heat, but most current companies still use older types of compressors that have not been directly adapted for the use of waste heat. In case of interest in obtaining waste heat, the reconstruction of the facility or development is inevitable.

Oxy-combustion flow field study under power-generating and CO2-capturing Matiant cycle conditions and its influence on the turbine inlet conditions

Matiant cycles are one of the oxy-combustion ways to capture CO2 in power generation. This article analyzes the generated flow field when performing combustion in similar conditions to those in Matiant combustion chambers. It is complementary to other studies in which oxy-combustion is studied from a chemical or cycle-performance perspective. The used methodology comprises CFD (Computational Fluid Dynamics) simulations. These simulations took into account the study of reacting flows. The detailed GRI-Mech 3.0 reaction mechanism was used along with the detailed Eddy Dissipation Concept model (EDC). Air and pure CO2 working fluids were also studied to provide reference cases. The pressure effect was studied, taking into account the range of different modern compressors. Some target results were the temperature field and the conditions at the turbine-stage inlet. It was found a different flow field evolution compared to having a working fluid composed of air. Numerical data of the Matiant temperature deviations on the symmetry axis, taking the air oxy-combustion case as reference, in a point-to-point comparison, are provided. At a 30 atm operating pressure, the Matiant case temperature at the outlet was 88.1% of the air coflow case oxy-combustion temperature. These temperature differences decrease when increasing the operating pressure.

PURPOSE AND DESCRIPTION OF THE COMPRESSOR STATION

All work on the construction of pumping and compressor stations is usually divided into two groups of zero cycle work and ground cycle work. The work of the zero cycle includes the preparation of the construction site, earthworks, work on the construction of foundations for buildings, pumping units and technological equipment, work on the construction of underground pipelines and utilities. The work of the ground cycle includes work on the construction of buildings for pumping and compressor shops and auxiliary buildings, installation work on installation and fixing on the foundations in the design position of pumping units. Compressor stations (CS) have been installed along the pipeline route to maintain a certain flow rate of the transported gas and to ensure optimal pressure in the pipeline. A modern compressor station is a complex engineering structure that provides the basic technological processes for the preparation and transportation of natural gas. Keywords: compressor stations, gas pipeline, building structure, Booster compressor stations.

Modernization of industrial compressors based on modern automation tools

Modern compressors are built mainly by their computer-based automatic control systems that meet modern energy-saving requirements. Computer controls ensuring the safety of the equipment operation creates the conditions for saving energy.

AN OVERVIEW OF TIME-DOMAIN COMPUTATIONAL METHODS FOR AEROELASTIC INSTABILITIES OF MULTI-STAGE COMPRESSOR

Modern gas turbine design continues to move towards improved performance, reduced weight and reduced cost. As turbomachinery blade aerofoils are thinned to improve performance and reduce weight, aeroelastic issues such as flutter, forced response and stall driven vibrations become more predominant. Moreover, as the use of blisks (blade-integrated-disks) with very low mechanical damping becomes more common in modern compressor designs, accurate prediction of compressor aeroelastic stability in a multi-row environment becomes vital. This paper presents a review of aeroelasticity research carried out at Rolls-Royce Vibration University Technology Centre (VUTC) at Imperial College over the past 20 years. The aim is to summarise the unusual aeroelastic issues observed in multi-stage compressors into one document so that it can be used by other researchers in the field. Blade passing forced response is not addressed here as their existence can be detected by a Campbell diagram. The results presented here are based on numerical methods but where possible data from experiments are used to verify the numerical findings.

Sequence Compression Benchmark (SCB) database-A comprehensive evaluation of reference-free compressors for FASTA-formatted sequences.

BackgroundNearly all molecular sequence databases currently use gzip for data compression. Ongoing rapid accumulation of stored data calls for a more efficient compression tool. Although numerous compressors exist, both specialized and general-purpose, choosing one of them was difficult because no comprehensive analysis of their comparative advantages for sequence compression was available.FindingsWe systematically benchmarked 430 settings of 48 compressors (including 29 specialized sequence compressors and 19 general-purpose compressors) on representative FASTA-formatted datasets of DNA, RNA, and protein sequences. Each compressor was evaluated on 17 performance measures, including compression strength, as well as time and memory required for compression and decompression. We used 27 test datasets including individual genomes of various sizes, DNA and RNA datasets, and standard protein datasets. We summarized the results as the Sequence Compression Benchmark database (SCB database, http://kirr.dyndns.org/sequence-compression-benchmark/), which allows custom visualizations to be built for selected subsets of benchmark results.ConclusionWe found that modern compressors offer a large improvement in compactness and speed compared to gzip. Our benchmark allows compressors and their settings to be compared using a variety of performance measures, offering the opportunity to select the optimal compressor on the basis of the data type and usage scenario specific to a particular application.

Aerodynamic optimization to tandem stators by using sweep and dihedral

In the modern aerodynamic design of turbomachinery blades, the geometries of blades often need to be reshaped to achieve better aerodynamic performance by introducing extra parametric design variables. A higher variable dimension will lead to a larger sampling range as well as a sparser sample distribution, which challenges the effectiveness and stability of optimization schemes based on surrogate model by making the model prediction quality even poorer. In this paper, a multi-objective optimization based on Gaussian process model was carried out for a high dimensional design space. Based on the previous two-dimensional optimization, tandem stators of a modern compressor were optimized by the design of sweep and dihedral. The purpose of the study is to improve the aerodynamic performance of the compressor tandem stators as well as to provide an effective optimization scheme for high dimensional multi-objective optimization problems. The design of sweep and dihedral for reshaping the tandem stators consists of a total of 18 design variables. An improvement in total pressure recovery coefficient of at least 0.7% at positive incidence and at least 0.3% at negative incidence was obtained, much larger than that in the previous two-dimensional optimization. The optimization process shows that, by using Gaussian process as the surrogate model and a special sampling strategy, this optimization scheme is effective and efficient to handle this high dimensional space. The aerodynamic influences of design parameters of tandem blades were analyzed in detail and the superiority of sweep and dihedral in reducing aerodynamic loss was confirmed.

Centrifugal Compressor Design for a Gasoline Engine Turbocharger

With increasing challenges in both fuel consumption and emission, the turbochargers have played a very important role for gasoline engines. Turbochargers can increase engine power density, reduce physical dimensions and reduce engine weight. Most of the turbochargers have a rotor system including a centrifugal compressor and a turbo turbine. Centrifugal compressors is a turbomachine which increases the gas pressure with the help of a turbine. Centrifugal compressors dominate the turbocharger applications. Centrifugal compressor performances are very critical for turbocharger performance. Gasoline engines need compressor not only to have high efficiency at whole operating range but also have a wide operating range. This paper discussed a centrifugal compressor design for gasoline engine turbocharger. The modern compressor design process developed recently was used for this new compressor design. The performance of the new design was compared with original compressor in CFD (Computational Fluid Dynamics) analysis, gas stand test and engine test. It demonstrated that the newly developed compressor has better performance and also meet the engine operating needs from both numerical analysis and test. The new centrifugal compressor development is successful.

Aeroelastic vibration analysis of a 1.5 stage compressor

Aeroelastic vibration problems are commonly found in modern compressors operating in off-design conditions. Large amplitude vibration could lead to high cycle fatigue (HCF) of blade and usually occurs in front stages of axial compressors. In this study, the influence of tip gap size on aeroelastic stability is analyzed in a 1.5 stage compressor with an in-house fluid-structure interaction code. A three-dimensional unstructured finite-volume compressible flow solver is applied in the fluid domain and a structure dynamic solver with the modal superimposition method for blade motion is used in the structure domain. Rotor tip clearances of 1%, 2% and 3% of tip axial chord at maximum rotor loading conditions at off-design speeds are analyzed for aeroelastic stability. The tip leakage flow and vortex structure can be seen near the blade tip region at a larger tip gap size. The aeroelastic stability of rotor blade at different tip gap sizes is mainly influenced by the 1st torsion mode, and the variation of aerodynamic damping is not monotonous. The intensity of the tip vortex and shock wave are the key factors affecting the aeroelastic stability of rotor when tip gap size increases.

Pseudo random binary signal for MIMO frequency response identification of a magnetically levitated rotor system

High-speed machine technology with magnetically levitated rotor systems has become interesting solution in modern compressor, pump and turbo applications. Often, the commissioning of such machines is system identification-based with injected artificially generated excitation. In this paper a statistically uncorrelated PRBS design is proposed for multi-input multi-output MIMO system that gives the possibility to utilize same PRBS signal for all inputs. The proposed routine is validated with an experimental high-speed generator with active magnetic bearings (AMBs). The obtained frequency responses are validated by comparing the results with the mathematical model.

Performance Characteristic Modeling for 2D Compressor Cascades

Abstract The flow field in a compressor cascade is very complex owing to the highly 3D, turbulent, and viscous properties. However, in the through-flow analysis method, the viscosity effects are taken into account using empirical models. These models were based on experimental results for early blades. However, the blade types in modern compressors are quite different from those in older compressors. Therefore, it is not possible to predict the performance of modern compressors using the old empirical models. In this study, several multiple circular-arc (MCA) blades commonly used in modern compressors were simulated. After the simulations, a database of the cascade performance was built. Based on this database, some new models were established to predict the performance of modern cascades using regression analysis methods and an artificial neural network (ANN) method. The accuracy of all these new models is high enough for use in engineering applications.