High Pressure Ratio Compressor Research Articles

Numerical and experimental investigations of flow separation control through a linear compressor cascade

Modern large-scale gas turbines are equipped with high-pressure ratio compressors to increase engine work and its overall efficiency. Flow separation and energy losses are also two interrelated phenomenon associated with changes in compressor loading level and performance. This paper examines therefore the control of flow separation using a passive-control technique. An arced divergent-convergent slot grooved from the blade pressure side to its suction side was adopted to control flow separation, reducing the losses through a linear compressor cascade. The spanwise location of the slot was selected based on CFD simulations where the corner separation was predicted. The slot height in the spanwise direction was selected to be 8% of the blade height at the end-wall side. The present work was performed experimentally and numerically at an inlet Reynolds number,Rec=ρV∞C/μ=2.98×105\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$${R}_{ec}=\\rho {V}_{\\infty }C/\\mu =2.98\ imes {10}^{5}$$\\end{document}, covering a wide range of incidence angles from +6∘to-6∘\\documentclass[12pt]{minimal} \\usepackage{amsmath} \\usepackage{wasysym} \\usepackage{amsfonts} \\usepackage{amssymb} \\usepackage{amsbsy} \\usepackage{mathrsfs} \\usepackage{upgreek} \\setlength{\\oddsidemargin}{-69pt} \\begin{document}$$+ 6^{ \\circ } {\ ext{ to}} - 6^{ \\circ }$$\\end{document}. The experimental work was carried out using a linear cascade test section consisting of six NACA 65-009 blade profiles integrated into a low-speed wind tunnel. A five-hole pressure probe system was used to obtain main flow parameters. Numerically, four turbulence models, including Spalart–Allmaras (S–A) model, Realizable (R k-ε) model, Shear-Stress Transport (SST k-ω) model, and Reynolds Stress model (RSM) were tested to predict the velocity and pressure fields. Good agreement between the experimental measurements and the numerical results, which were obtained using the RSM turbulence model in terms of velocity profiles and total pressure downstream of blades. It was observed also that the use of the arced-slotted blades for positive incident angles was more effective in reducing the separation than the negative and zero incident angles, approaching a maximum value of 33% for 6° with enhanced blade loading reaching 17.6%. It is to be concluded that, the use of arced slotted blade improves the compressor performance specially for positive incident angles.

Behaviors of Surge Frequencies in Multi-stage Axial Flow Compressors over a Wide Range of Speeds

The fundamental behaviors of deep surge frequencies in axial-flow compressors have come to be described as a rather simple framework for low pressure-ratio compressors, in terms of two essential reduced frequencies concerning resonance frequencies and surge frequencies, on the basis of numerical-experimental results. For high-pressure-ratio multi-stage compressors, however, the frequency behaviors over a wide range of compressor speeds tend to give drastic deviations from the simple framework, including abrupt changes and apparent irregularities. The cause was investigated here by numerical simulations on a nine-stage axial flow compressor. Generally, deep surge frequencies of high-pressure ratio compressors tend to lower gradually, from low speeds toward higher ones. At somewhat below the design speed, the frequencies tend to decrease precipitously, settling to some very low levels of values. The discontinuous drops in the frequencies are shown to result from occurrences of short-term incomplete surge recoveries in the stalling phase of surge. The phenomenon tends to elongate the duration of the stalled phase and the whole surge duration, resulting in the lowered surge frequency. The phenomenon is caused by the flow reversal of high-temperature air in the stalling phase of surge and the re-suction of the hot air in the unstalling phase. The degree of the peak temperature and the intensity of the interferences between the hot-air re-suction phenomenon and the intrinsic surge appear to affect significantly the additional short-term premature stalling of the compressor. It could be strongly related with the high average temperature due to the hot air re-suction, and the duration time of the suction. The complicated behaviors thus become qualitatively explainable, though only by examinations on numerical-simulation results.

Aerodynamic development of a high-pressure ratio compressor for an advanced microturbine powerplant

Microturbine engines are increasing in popularity both for small poly-generation power plants and unmanned aerial vehicle; however, their overall efficiency is severely limited by their thermodynamic cycle and component efficiency. This compressor allows the real cycle of the engine to operate with lower specific fuel consumption, extending the rage of the vehicle and lowering pollutant emissions. Optimization methods were implemented in the compressor design, in order to increase the aerodynamic efficiency. A 4.5:1 pressure ratio was selected in order to keep the exhaust jet subsonic. Linear models were used to optimize the speed and gross geometry while the TEIS and later fully viscous 3D models were used successively to optimize the end walls and blade shapes. For all numerical simulations, the Menter SST turbulence model was used. In order to minimize modelling errors, the rotor-stator interface was placed as far as possible downstream of the impeller.

Effects of Transient Heat Transfer on Compressor Stability

The effects of heat transfer between the compressor structure and the primary gas path flow on compressor stability are investigated during hot engine re-acceleration transients. A mean line analysis of an advanced, high-pressure ratio compressor is extended to include the effects of heat transfer on both stage matching and blade row flow angle deviation. A lumped capacitance model is used to compute the heat transfer of the compressor blades, hub, and casing to the primary gas path. The inputs to the compressor model with heat transfer are based on a combination of full engine data, compressor test rig measurements, and detailed heat transfer computations. Nonadiabatic transient calculations show a 8.0 point reduction in stall margin from the adiabatic case, with heat transfer predominantly altering the transient stall line. 3.4 points of the total stall margin reduction are attributed to the effect of heat transfer on blade row deviation, with the remainder attributed to stage rematching. Heat transfer increases loading in the front stages and destabilizes the front block. Sensitivity studies show a strong dependence of stall margin to heat transfer magnitude and flow angle deviation at low speed, due to the effects of compressibility. Computations for the same transient using current cycle models with bulk heat transfer effects only capture 1.2 points of the 8.0 point stall margin reduction. Based on this new capability, opportunities exist early in the design process to address potential stability issues due to transient heat transfer.

An Investigation of the Stability Enhancement of a Centrifugal Compressor Stage Using a Porous Throat Diffuser

The stable operating range of a centrifugal compressor stage of an engine turbocharger is limited at low mass flow rates by aerodynamic instabilities which can lead to the onset of rotating stall or surge. There have been many techniques employed to increase the stable operating range of centrifugal compressor stages. The literature demonstrates that there are various possibilities for adding special treatments to the nominal diffuser vane geometry, or including injection or bleed flows to modify the diffuser flow field in order to influence diffuser stability. One such treatment is the porous throat diffuser (PTD). Although the benefits of this technique have been proven in the existing literature, a comprehensive understanding of how this technique operates is not yet available. This paper uses experimental measurements from a high pressure ratio (PR) compressor stage to acquire a sound understanding of the flow features within the vaned diffuser which affect the stability of the overall compression system and investigate the stabilizing mechanism of the porous throat diffuser. The nonuniform circumferential pressure imposed by the asymmetric volute is experimentally and numerically examined to understand if this provides a preferential location for stall inception in the diffuser. The following hypothesis is confirmed: linking of the diffuser throats via the side cavity equalizes the diffuser throat pressure, thus creating a more homogeneous circumferential pressure distribution, which delays stall inception to lower flow rates. The results of the porous throat diffuser configuration are compared to a standard vaned diffuser compressor stage in terms of overall compressor performance parameters, circumferential pressure nonuniformity at various locations through the compressor stage and diffuser subcomponent analysis. The diffuser inlet region was found to be the element most influenced by the porous throat diffuser, and the stability limit is mainly governed by this element.

Issues in performance measurement of CO2 compressor near the critical point

Interest in the supercritical carbon dioxide (S-CO2) Brayton cycle technology is gradually increasing due to high efficiency at moderate turbine inlet temperature and potential to achieve high compactness of the whole power conversion system. As a part of research activities on the S-CO2 Brayton cycle technology development, Supercritical CO2 Pressurization Experiment (SCO2PE) facility was constructed for lab scale CO2 compression experiment near the critical point. SCO2PE facility operation is focused on resolving issues in a compressor design and operation near the critical point of CO2. From the SCO2PE operation experience, performance measurement uncertainty issues in a low pressure ratio compressor were identified and these issues become generic if a compressor operates near the critical point. Very high uncertainty on performance measurement was observed due to the peak region in thermodynamic property variation of CO2 near the critical point and low pressure ratio operation of the compressor. High pressure ratio compressor design and operation can mitigate high uncertainty issues in performance measurement. However, low pressure ratio operation cannot be ignored since low pressure ratio operation is always involved during a part load operation or transient operation of a power conversion system. Thus, an alternative approach was suggested to secure reliability of performance estimation from the basic measurement.

Miniature Piezoelectric Compressor for Joule-Thomson Cryocoolers

Joule-Thomson (JT) cryocoolers operate with a continuous flow of the working fluid that enters the cooler at a high pressure and leaves it at a lower pressure. Ideally, the temperature of the outgoing fluid equals the temperature of the entering fluid. JT cryocoolers that operate with pure refrigerants require high pressure of a few tens of MPa where the low pressure is usually around 0.1MPa. Circulation of the working fluid in such cases requires high pressure ratio compressors that evidently have large dimensions. JT cryocoolers can operate with much lower pressure ratios by using mixed-refrigerants. Cooling from 300K to about 80K in a single stage cryocooler normally requires a pressure ratio of about 1:25.In the present research a miniature compressor driven by piezoelectric elements is developed in collaboration between Rafael and the Technion. This type of compressor has the advantage of improved long life compared to other mechanical compressors, very low vibrations, and silent operation. In the current case, the design goal of the intake and discharge pressures has been 0.1 and 2.5MPa, respectively, with a flow rate of 0.06g/s. The compressor has two compression stages; 1:5 and 5:25.Several configurations have been considered, fabricated, and tested. The performance of the last configuration approaches the desired specification and is presented in the current paper together with the design concept.

Study on the performance deterioration of mixed flow impeller due to change in tip clearance

Performance of mixed flow compressor with un-shrouded impeller strongly depends upon unsteady, asymmetrical flow fields in the axial directions. The flow through the mixed flow impeller is complex due to three-dimensional nature of geometry. In mixed flow impeller, there are clearances between the rotating impeller blades and the casing as the high pressure ratio compressors are usually open shrouded impellers. As a result, certain amount of reduction in the performance is unavoidable due to clearance flows. In the present investigations, numerical analysis is performed using a commercial code to investigate tip clearance effects on through flow. The performance of mixed flow impeller with four different clearances between impeller and stationary shroud are evaluated and compared with experimental results. The impeller performance map was obtained for different operating speeds and mass flow rates with different tip clearances. The result shows that the tip leakage flow strongly interacts with mainstream and contributes to total pressure loss and performance reduction. The pressure and performance decrement are approximately linearly proportional to the gap between impeller and stationary shroud. The analysis showed scope for improvement in design of the compressor for better performance in terms of efficiency and operating range.

Experimental Investigation of High Pressure Ratio Centrifugal Compressor With Axisymmetric and Nonaxisymmetric Recirculation Device

Centrifugal compressors used for turbochargers are required to have a wide operating range. A recirculation device, which consists of a bleed slot, an upstream slot, and an annular cavity connecting both slots, is often used with them. It improves the incidence angle of the impeller leading edge, i.e., the blade loading of the inducer, at low flow rates due to the recirculation flow supplied to the compressor inlet. However, the compressor efficiency drops when there is a recirculation flow from the bleed slot to the upstream slot. A one dimensional analysis in the first section of this paper showed that the reduction in the compressor efficiency can be lowered by decreasing the pressure drop or reducing the recirculation flow rate within the recirculation device. This study examined the possibility of improvement in the compressor efficiency by the use of a recirculation device with an asymmetric bleed slot. An impeller of a turbocharger compressor is normally contained in a volute. Since the geometry of the volute is not axisymmetric, the impeller is surrounded by an asymmetric flow field. Hence each impeller passage, which is formed by two adjacent full blades, is operated at a different operating point. This means that some of the passages need the improvement in the blade loading by the recirculation device but others do not. There is a possibility that this is realized by a recirculation device with an asymmetrically distributed bleed slot, called a nonaxisymmetric recirculation device in this paper. If the asymmetric bleed slot shortens the average distance between the bleed slot and upstream slot or reduces the area of the bleed slot, it can reduce the pressure drop or recirculation flow rate within the recirculation deviceand, hence, can improve the compressor efficiency. This study discusses the characteristics of high pressure ratio compressors for turbochargers without the recirculation device and those with the recirculation device with an axisymmetric bleed slot. Furthermore, the effects of nonaxisymmetric recirculation devices on the compressor characteristics are experimentally investigated. Two types of nonaxisymmetric recirculation devices were tested. One had the bleed slot of a sine wave pattern. The other had the bleed slot partially channeled in the circumferential direction. There were appropriate positions relative to the volute for both nonaxisymmetric recirculation devices. The compressor efficiency with nonaxisymmetric recirculation devices was higher than that with axisymmetric recirculation devices and the surge lines of the compressor with nonaxisymmetric recirculation devices were located at a flow rate lower than or equal to those with the axisymmetric recirculation devices.

3209 高増速比遊星歯車を用いた高圧力比ターボコンプレッサの開発(OS-09 ハイポイドギヤ・ウォームギヤ・遊星歯車)

3209 高増速比遊星歯車を用いた高圧力比ターボコンプレッサの開発(OS-09 ハイポイドギヤ・ウォームギヤ・遊星歯車)

슈라우드 임펠러에서 누설 유동이 압축기 성능에 미치는 영향에 대한 수치해석적 연구

Generally the Shrouded type impeller is considered to free from the loss of tip leakage flow, but is is actually not possible to complete sealing between the rotating impeller and the sealing which is stay still. As a result, there is the possibility of flow leaking between impeller exit to entrance, especially with high pressure ratio compressor machine. The Cavity leakage flow is expected to influence negative effect on a machine performance and also inner flow structure.<BR> In this study, Impeller with shroud-casing gap leakage flow is simulated by numerical method (Using CFX 12.1). The influence of leakage flow on compressor performance and efficiency will be analysed. also detail flowfield change will posted.<BR><BR>

Spring Stiffness Selection Criteria for Nozzle Check Valves Employed in Compressor Stations

Nozzle type check valves are often employed in compressor stations in three locations: compressor outlet, station discharge, and station bypass. The fundamental design concept of these valves is based on creating a converging diverging flow through the valve internal geometry such that a minimum area is achieved at a location corresponding to the back of the check valve disk at the fully open position. This will ensure maximum hydrodynamic force coefficient which allows the valve to be fully open with minimum flow. Spring forces and stiffness determine the performance of this type of check valves and impact the overall operation and integrity of the compressor station. This paper examines the effects of various spring characteristics and stiffness in relation to the compressor and station flow characteristics. The results show that when the spring forces are higher than the maximum hydrodynamic force at minimum flow, the disk will not be at the fully open position, which will give rise to disk fluttering and potential for cyclic high velocity impact between components of the internal valve assembly. This could lead to self destruction of the check valve and subsequent risk of damage to the compressor unit itself. The paper also points to the fact that the spring selection criteria for a unit check valve are different than that for station and bypass check valves. An example of a case study with actual field data from a high pressure ratio compressor station employing this type of check valves is presented to illustrate the associated dynamic phenomena and fluid-structure interaction within the internal assembly of the check valve.

Study of the Flow in Centrifugal Compressor

Reducing the losses of the tip clearance flow is one of the keys in an unshrouded centrifugal compressor design and development because tip clearances are large in relation to the span of the blades and also centrifugal compressors produce a sufficiently large pressure rise in single stage. This problem is more acute for a low flow high-pressure ratio impeller design. The large tip clearance would cause flow separations, and as a result it would drop both the efficiency and surge margin. Thus a design of a high efficiency and wide operation range low flow coefficient centrifugal compressor is a great challenge. This paper describes a recent development of high efficiency and wide surge margin low flow coefficient centrifugal compressor. A viscous turbomachinery optimal design method developed by the authors for axial flow machine was further extended and used in the centrifugal compressor design. The compressor has three main parts: impeller, a low solidity diffuser and volute. The tip clearance is under a special consideration in this design to allow impeller insensitiveness to the clearance. A patented three-dimensional low solidity diffuser design method is used and applied to this design. The compressor test results demonstrated to be successful to extend the low solidity diffusers to high-pressure ratio compressor. The compressor stage performance showed the total to static efficiency of the compressor being about 85% and stability range over 35%. The test results are in good agreement with the design.

Development of a Low Flow Coefficient Single Stage Centrifugal Compressor

A low flow coefficient unshrouded centrifugal compressor would give up clearance in relation to the span of the blades, because centrifugal compressors produce a sufficiently large pressure rise in fewer stages. This problem is more acute for a low flow high-pressure ratio impeller. The large tip clearance would cause flow separations, and as a result it would drop both the efficiency and surge margin. Thus a design of a high efficiency and wide operation range for a low flow coefficient centrifugal compressor is a great challenge. This paper describes a new development of high efficiency and large surge margin low flow coefficient (0.145) centrifugal compressor. A viscous turbomachinery optimal design method developed by the authors for axial flow machine was further extended and used in this centrifugal compressor design. The new compressor has three main parts: impeller, a low solidity diffuser, and volute. The tip clearance is under special consideration in this design to allow impeller insensitiveness to the clearance. A three-dimensional low solidity diffuser design method is proposed and applied to this design. This design is successful in extending the low solidarity diffusers to high-pressure ratio compressor. It is demonstrated that the design is a great success. The design performance range of the total to static efficiency of the compressor is about 85% and stability range is over 35%. The experimental results show that the test results are in good agreement with the design.

Parametric Study of Injection Angle Effects on Stability Enhancement of Transonic Axial Compressors

An effective approach in suppressing stall is mass injection upstream of the tip of an axial flow compressor. Developing reliable injection systems for high pressure ratio compressors requires an understanding of the effects of injection parameters, for example, injection angle, on the compressor stability and performance. This paper reports on a numerical investigation of injection angle on the operability of a high-speed compressor rotor. The injection port is set to be choked for all of the injection cases. Simulations indicate that injection angle has a significant impact on stability improvement. Maximum range extension is obtained when an injection yaw angle equal to -30 deg is applied. At this injection angle, the stalling mass flow coefficient was reduced by 17.4% by using an injected mass flow equivalent to 1.55% of the baseline annulus flow. Results also indicate that the best injection case creates an incidence angle of about 0 deg over the tip of the blade and adds the highest momentum in the relative frame of reference.

Detailed Flow Study of Mach Number 1.6 High Transonic Flow With a Shock Wave in a Pressure Ratio 11 Centrifugal Compressor Impeller

Requirements for aeronautical gas turbine engines for helicopters include small size, low weight, high output, and low fuel consumption. In order to achieve these requirements, development work has been carried out on high efficiency and high pressure ratio compressors. As a result, we have developed a single stage centrifugal compressor with a pressure ratio of 11 for a 1000 shp class gas turbine. The centrifugal compressor is a high transonic compressor with an inlet Mach number of about 1.6. In high inlet Mach number compressors, the flow distortion due to the shock wave and the shock boundary layer interaction must have a large effect on the flow in the inducer. In order to ensure the reliability of aerodynamic design technology, the actual supersonic flow phenomena with a shock wave must be ascertained using measurement and Computational Fluid Dynamics (CFD). This report presents the measured results of the high transonic flow at the impeller inlet using Laser Doppler Velocimeter (LDV) and verification of CFD, with respect to the high transonic flow velocity distribution, pressure distribution, and shock boundary layer interaction at the inducer. The impeller inlet tangential velocity is about 460 m/s and the relative Mach number reaches about 1.6. Using a LDV, about 500 m/s relative velocity was measured preceding a steep deceleration of velocity. The following steep deceleration of velocity at the middle of blade pitch clarified the cause as being the pressure rise of a shock wave, through comparison with CFD as well as comparison with the pressure distribution measured using a high frequency pressure transducer. Furthermore, a reverse flow is measured in the vicinity of casing surface. It was clarified by comparison with CFD that the reverse flow is caused by the shock-boundary layer interaction. Generally CFD shows good agreement with the measured velocity distribution at the inducer and splitter inlet, except in the vicinity of the casing surface.

Process Applications and Design of Single and Two Shaft Hot Gas Expander Compression Systems

Single and two shaft hot gas expander compression systems will soon replace conventional multitrain, low speed equipment due to significant economic and process advantages. The single train concept is a direct result of the recent advances in high pressure ratio compressor, high performance radial inflow turbine and high performance steam turbine technology. The application and custom engineering of single train compression systems is discussed, whereby emphasis is placed on performance and reliability.

Potential Technology Improvements for the CF6 Turbofan

THE CF6 engine is a dual rotor, high bypass ratio turbofan engine designed for airline operation. It incorporates a variable stator, high pressure ratio compressor, an annular combustor, an air‐cooled core engine turbine and a co‐axial front fan driven by a five‐stage low pressure turbine (Fig. 1). The core engine has an exhaust thrust spoiler which operates in conjunction with the fan cascade reverser during landing roll.