The flow physics in a large rotor tip gap in a 1.5-stage axial compressor is investigated in the current study. The flow structure in the rotor tip region is complex with several dominant vortical structures of opposite rotation, resulting in inhomogeneous and highly anisotropic turbulence. Earlier measurements show that eddy viscosity is negative over large parts of the tip region and eddy viscosity varies among stress/strain components. The present study aims to understand how the complex nature of rotor tip leakage flow affects compressor performance when the tip gap size is greater than 4–5% of the rotor span, which is typical of advanced small core engines. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques are applied to study flow physics in a large rotor tip gap (5.5% of rotor span) in a 1.5-stage axial compressor. Calculated flow fields from the two different approaches are compared with available measurements and examined in detail. LES calculates the pressure rise in the present compressor fairly well, while URANS with a standard two-equation turbulence closure underpredicts the pressure rise by 15–20% of the measured values. The current study shows that URANS with the current turbulence closure produces much higher all-positive eddy viscosity in the tip-gap region compared to measurements and LES. The distribution of eddy viscosity in the URANS simulation is also wrong. Consequently, the flow in the tip region is highly damped with significantly larger blockage generation, which results in the tip leakage vortex (TLV) staying closer to the blade suction side compared to the measurement. When the TLV stays closer to the blade, both flow turning and the pressure rise across the compressor are reduced compared to the measurements. It appears that this effect is amplified by a large rotor tip gap.
{"title":"Flow Physics in a Large Rotor Tip Gap in a Multi-Stage Axial Compressor","authors":"C. Hah","doi":"10.1115/gt2021-58745","DOIUrl":"https://doi.org/10.1115/gt2021-58745","url":null,"abstract":"\u0000 The flow physics in a large rotor tip gap in a 1.5-stage axial compressor is investigated in the current study. The flow structure in the rotor tip region is complex with several dominant vortical structures of opposite rotation, resulting in inhomogeneous and highly anisotropic turbulence. Earlier measurements show that eddy viscosity is negative over large parts of the tip region and eddy viscosity varies among stress/strain components. The present study aims to understand how the complex nature of rotor tip leakage flow affects compressor performance when the tip gap size is greater than 4–5% of the rotor span, which is typical of advanced small core engines. Unsteady Reynolds-averaged Navier-Stokes (URANS) and Large Eddy Simulation (LES) techniques are applied to study flow physics in a large rotor tip gap (5.5% of rotor span) in a 1.5-stage axial compressor. Calculated flow fields from the two different approaches are compared with available measurements and examined in detail. LES calculates the pressure rise in the present compressor fairly well, while URANS with a standard two-equation turbulence closure underpredicts the pressure rise by 15–20% of the measured values. The current study shows that URANS with the current turbulence closure produces much higher all-positive eddy viscosity in the tip-gap region compared to measurements and LES. The distribution of eddy viscosity in the URANS simulation is also wrong. Consequently, the flow in the tip region is highly damped with significantly larger blockage generation, which results in the tip leakage vortex (TLV) staying closer to the blade suction side compared to the measurement. When the TLV stays closer to the blade, both flow turning and the pressure rise across the compressor are reduced compared to the measurements. It appears that this effect is amplified by a large rotor tip gap.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"19 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125342702","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
As a compressor is throttled three-dimensional separations develop in the corners between the blades and annulus endwall. Surprisingly, little is understood about the unsteady topology of these separations. One of the problems with studying corner separations is that it is often difficult to understand whether a particular flow structure in the separation is inherent to the separation itself, or due to the response of the separation to changes in the inlet flow. In this paper a novel experimental approach is taken with the aim of isolating the corner separation from external influences. A cascade is designed with the specific aim of precisely controlling the inlet flow. Contrary to previous work, it is shown that the key saddle and focus pair, which describes the time-mean topology of the corner separation on the endwall, moves smoothly and continuously as the incidence of the flow is raised. This behavior is shown to be the result of the time-resolved topology of the flow field, which comprises numerous saddle and focus pairs which are produced stochastically in regions of high shear strain rate. Most importantly, the separation is shown to exhibit an extremely low frequency behavior, changing in topology over timescales which are approximately 80 times the convection time through the blade passage. The behavior is shown to be intrinsic to the separation and causes the separation, for periods, to completely disappear from the endwall. This underlying unsteady structure of the separation is shown to have implications for the ability of RANS-based design codes to be able to accurately predict corner separations.
{"title":"The Unsteady Topology of Corner Separations","authors":"I. Dawkins, James V. Taylor, X. Ottavy, R. Miller","doi":"10.1115/GT2021-58730","DOIUrl":"https://doi.org/10.1115/GT2021-58730","url":null,"abstract":"\u0000 As a compressor is throttled three-dimensional separations develop in the corners between the blades and annulus endwall. Surprisingly, little is understood about the unsteady topology of these separations. One of the problems with studying corner separations is that it is often difficult to understand whether a particular flow structure in the separation is inherent to the separation itself, or due to the response of the separation to changes in the inlet flow. In this paper a novel experimental approach is taken with the aim of isolating the corner separation from external influences. A cascade is designed with the specific aim of precisely controlling the inlet flow.\u0000 Contrary to previous work, it is shown that the key saddle and focus pair, which describes the time-mean topology of the corner separation on the endwall, moves smoothly and continuously as the incidence of the flow is raised. This behavior is shown to be the result of the time-resolved topology of the flow field, which comprises numerous saddle and focus pairs which are produced stochastically in regions of high shear strain rate. Most importantly, the separation is shown to exhibit an extremely low frequency behavior, changing in topology over timescales which are approximately 80 times the convection time through the blade passage. The behavior is shown to be intrinsic to the separation and causes the separation, for periods, to completely disappear from the endwall. This underlying unsteady structure of the separation is shown to have implications for the ability of RANS-based design codes to be able to accurately predict corner separations.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"35 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127714100","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
For over forty years, coatings have been used as a means to prevent the build-up of foulants on the flow passages of centrifugal compressors. In general, this has been a “One Size Fits All” solution, with most OEMs and applicators offering nearly identical coating systems, despite the fact that there are distinct differences in the types of foulant that occurs in different services. To develop new coatings, there are two options, test “on the job” or test in a laboratory setting. This paper looks at various test methods to evaluate coatings by isolating the important aspects that a good fouling resistant coating should possess, namely fouling, erosion, and corrosion resistance. Each of these will be investigated separately in a battery of easy to perform tests. To the extent possible, the tests for each aspect attempt to replicate the complex conditions that occur in a centrifugal compressor that impact that specific aspect. The fouling and corrosion tests were performed with considerations for two different types of service. To illustrate the ability of the test to evaluate various coatings, the results from three different coatings will be discussed. These test methods will also be compared to other test methods used in literature.
{"title":"Considerations for the Evaluation of Fouling Resistant Coatings for Centrifugal Compressors","authors":"S. Laney","doi":"10.1115/gt2021-59099","DOIUrl":"https://doi.org/10.1115/gt2021-59099","url":null,"abstract":"\u0000 For over forty years, coatings have been used as a means to prevent the build-up of foulants on the flow passages of centrifugal compressors. In general, this has been a “One Size Fits All” solution, with most OEMs and applicators offering nearly identical coating systems, despite the fact that there are distinct differences in the types of foulant that occurs in different services. To develop new coatings, there are two options, test “on the job” or test in a laboratory setting. This paper looks at various test methods to evaluate coatings by isolating the important aspects that a good fouling resistant coating should possess, namely fouling, erosion, and corrosion resistance. Each of these will be investigated separately in a battery of easy to perform tests. To the extent possible, the tests for each aspect attempt to replicate the complex conditions that occur in a centrifugal compressor that impact that specific aspect. The fouling and corrosion tests were performed with considerations for two different types of service. To illustrate the ability of the test to evaluate various coatings, the results from three different coatings will be discussed. These test methods will also be compared to other test methods used in literature.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"76 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"132556796","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Jaewoo Choi, D. Šimurda, Jae-Wook Song, M. Luxa, Sungryong Lee, J. Hála, J. Lepicovsky, T. Radnic, Jun-Sik Seo
Overall efficiency of an axial compressor is largely affected by its front stage when it is operating under transonic flow conditions. For this reason, many manufacturers and researchers are advancing research and development of transonic airfoils in these days. Doosan, in frame of a development of high efficiency gas turbine, developed high efficiency airfoil for a transonic rotor and conducted cascade tests. Therefore, this study deals with a test of two compressor transonic blade cascades at inlet Mach number over 1.1. To improve the efficiency and operating range, two kinds of thickness distribution type based on Enhanced Doosan Airfoil (EDA), which applied unique rule, were applied and assessed. The first airfoil consists of polynomial thickness distribution and the second airfoil consists of new thickness distribution with specially tailored leading edge. In order to ensure accurate geometry of a model, a detailed checkout process upon production of model blades used in the test was performed. This is because, in the case of transonic airfoil, if the inlet leading edge shape differs by more than 0.2% than designed airfoil of leading edge, the result will be completely different. Therefore, not only the tolerance within 0.1% was confirmed but also the shape produced through simulation and 3D CMM scan data. The main parameters for the comparison are an inlet Mach number, an axial velocity density ratio (AVDR) and the kind of thickness distribution. Results of tests and CFD blade to blade analysis using MISES 2.70 are compared. The flow field was visualized using schlieren technique and parameters of the suction side boundary layer were evaluated at several locations based on Pitot probe traverses. The results confirm that a suction peak at the round leading edge disappears in the case of the new thickness type distribution with tailored leading edge. This confirms that the profile shaping without jump in curvature in the leading edge region leads to smooth acceleration without peaks. Nevertheless, results show that the new thickness distribution type is not absolutely good in comparison with the polynomial thickness distribution type with respect to the total pressure loss coefficient. Moreover, bucket range (operating range) is also almost the same. Results of the suction side boundary layer traversing suggest that the transition of the boundary layer takes place beyond the location x/cax > 0.088. The MISES results show that a shock location and the boundary layer parameters are similar to test results. However, values of the loss coefficient show some difference. Therefore, a new correlation in particular transonic flow condition was developed.
{"title":"Development of Loss Correlation and Tool Validation at Transonic Condition Based on Cascade Test","authors":"Jaewoo Choi, D. Šimurda, Jae-Wook Song, M. Luxa, Sungryong Lee, J. Hála, J. Lepicovsky, T. Radnic, Jun-Sik Seo","doi":"10.1115/gt2021-01771","DOIUrl":"https://doi.org/10.1115/gt2021-01771","url":null,"abstract":"\u0000 Overall efficiency of an axial compressor is largely affected by its front stage when it is operating under transonic flow conditions. For this reason, many manufacturers and researchers are advancing research and development of transonic airfoils in these days. Doosan, in frame of a development of high efficiency gas turbine, developed high efficiency airfoil for a transonic rotor and conducted cascade tests. Therefore, this study deals with a test of two compressor transonic blade cascades at inlet Mach number over 1.1. To improve the efficiency and operating range, two kinds of thickness distribution type based on Enhanced Doosan Airfoil (EDA), which applied unique rule, were applied and assessed. The first airfoil consists of polynomial thickness distribution and the second airfoil consists of new thickness distribution with specially tailored leading edge. In order to ensure accurate geometry of a model, a detailed checkout process upon production of model blades used in the test was performed. This is because, in the case of transonic airfoil, if the inlet leading edge shape differs by more than 0.2% than designed airfoil of leading edge, the result will be completely different. Therefore, not only the tolerance within 0.1% was confirmed but also the shape produced through simulation and 3D CMM scan data. The main parameters for the comparison are an inlet Mach number, an axial velocity density ratio (AVDR) and the kind of thickness distribution. Results of tests and CFD blade to blade analysis using MISES 2.70 are compared. The flow field was visualized using schlieren technique and parameters of the suction side boundary layer were evaluated at several locations based on Pitot probe traverses. The results confirm that a suction peak at the round leading edge disappears in the case of the new thickness type distribution with tailored leading edge. This confirms that the profile shaping without jump in curvature in the leading edge region leads to smooth acceleration without peaks. Nevertheless, results show that the new thickness distribution type is not absolutely good in comparison with the polynomial thickness distribution type with respect to the total pressure loss coefficient. Moreover, bucket range (operating range) is also almost the same. Results of the suction side boundary layer traversing suggest that the transition of the boundary layer takes place beyond the location x/cax > 0.088. The MISES results show that a shock location and the boundary layer parameters are similar to test results. However, values of the loss coefficient show some difference. Therefore, a new correlation in particular transonic flow condition was developed.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"130462838","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The drive for ever higher aircraft efficiency inevitably leads to a reduction in core compressor size; over the past 30 years the height of the rear stages of a typical aeroengine has reduced by 40%. This trend will continue; a further reduction of almost 50% is expected by 2050. It is shown in this paper that the two greatest sources of increased profile loss as core size is reduced comes from the mechanical and manufacturing constraints placed on the maximum and trailing edge thickness. Furthermore, mitigation strategies are presented which successfully minimized or eliminated these loss sources. An experimental study showed that local modification to the trailing edge shape from a semi-circle to a more elliptical shape almost entirely eliminated the increase in trailing edge loss associated with a reduction in compressor size. Elliptical trailing edges improved loss by both narrowing the wake and increasing the base pressure. Despite the local nature of the modification, the elliptical trailing edges had a powerful effect on flow turning. This was accounted for by redesigning each design to achieve the specified exit flow angle. A set of “as manufactured” small core blades was created by combining measured data from existing manufacture methods with airfoils scaled to ensure minimum thicknesses necessary for mechanical integrity along the blade were not breached. Core sizes down to 40% of current designs were run in CFD. It is often assumed that the loss in efficiency as core size is reduced is unavoidable. However, it was shown that approximately half of this is directly due to the increase in thickness to satisfy tolerancing. Part to part deviation has negligible effect at all but the smallest compressor size tested (at 40% of current values).
{"title":"Aerodynamic Mitigation of Mechanical Constraints in Small Compressor Blade Profiles","authors":"T. Dickens, James V. Taylor, C. Hall, R. Miller","doi":"10.1115/gt2021-59452","DOIUrl":"https://doi.org/10.1115/gt2021-59452","url":null,"abstract":"\u0000 The drive for ever higher aircraft efficiency inevitably leads to a reduction in core compressor size; over the past 30 years the height of the rear stages of a typical aeroengine has reduced by 40%. This trend will continue; a further reduction of almost 50% is expected by 2050.\u0000 It is shown in this paper that the two greatest sources of increased profile loss as core size is reduced comes from the mechanical and manufacturing constraints placed on the maximum and trailing edge thickness. Furthermore, mitigation strategies are presented which successfully minimized or eliminated these loss sources.\u0000 An experimental study showed that local modification to the trailing edge shape from a semi-circle to a more elliptical shape almost entirely eliminated the increase in trailing edge loss associated with a reduction in compressor size. Elliptical trailing edges improved loss by both narrowing the wake and increasing the base pressure. Despite the local nature of the modification, the elliptical trailing edges had a powerful effect on flow turning. This was accounted for by redesigning each design to achieve the specified exit flow angle.\u0000 A set of “as manufactured” small core blades was created by combining measured data from existing manufacture methods with airfoils scaled to ensure minimum thicknesses necessary for mechanical integrity along the blade were not breached. Core sizes down to 40% of current designs were run in CFD. It is often assumed that the loss in efficiency as core size is reduced is unavoidable. However, it was shown that approximately half of this is directly due to the increase in thickness to satisfy tolerancing. Part to part deviation has negligible effect at all but the smallest compressor size tested (at 40% of current values).","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"13 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"115034375","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Jinwook Lee, Vaishnavi Ramaswamy, Z. Spakovszky, E. Greitzer, M. Drela, Jérôme Talbotec
Part II describes the experimental assessment and the application of the ideas in Part I concerning the mechanisms that determine the role of blade surface waviness on laminar-turbulent transition and their consequent effect on civil aircraft fan performance. A natural transition wind tunnel was designed and constructed to characterize the impact of surface waviness on transition, using both hotwire anemometry and infrared thermography. The experimental results support the new hypothesis presented in Part I, concerning the way in which blade surface waviness affects fan performance through motion of the transition onset location due to interaction between surface waviness and Tollmien-Schlichting (TS) boundary layer instability. In particular, the theoretical amplification of the TS waves, and the corresponding transition onset location movement due to surface waviness, was borne out over a range of variations in Reynolds number, non-dimensional surface wavelength, non-dimensional surface wave height, and location of surface wave initiation, relevant to composite fan blade parameters. Further, the increase of receptivity coefficient, and thus the initial amplitude of disturbances due to geometric resonance between surface wavelength and TS wavelength, was also confirmed by the experiments. Surface waviness was estimated, in some cases, to result in a nearly 1% decrease in fan efficiency compared to a non-wavy blade. Suggestions are given for mitigation of the effects of waviness, including the idea of blade curvature rescheduling as a method to delay transition and thus decrease loss.
{"title":"Effects of Surface Waviness on Fan Blade Boundary Layer Transition and Profile Loss — Part II: Experimental Assessments and Applications","authors":"Jinwook Lee, Vaishnavi Ramaswamy, Z. Spakovszky, E. Greitzer, M. Drela, Jérôme Talbotec","doi":"10.1115/GT2021-58678","DOIUrl":"https://doi.org/10.1115/GT2021-58678","url":null,"abstract":"\u0000 Part II describes the experimental assessment and the application of the ideas in Part I concerning the mechanisms that determine the role of blade surface waviness on laminar-turbulent transition and their consequent effect on civil aircraft fan performance. A natural transition wind tunnel was designed and constructed to characterize the impact of surface waviness on transition, using both hotwire anemometry and infrared thermography. The experimental results support the new hypothesis presented in Part I, concerning the way in which blade surface waviness affects fan performance through motion of the transition onset location due to interaction between surface waviness and Tollmien-Schlichting (TS) boundary layer instability. In particular, the theoretical amplification of the TS waves, and the corresponding transition onset location movement due to surface waviness, was borne out over a range of variations in Reynolds number, non-dimensional surface wavelength, non-dimensional surface wave height, and location of surface wave initiation, relevant to composite fan blade parameters. Further, the increase of receptivity coefficient, and thus the initial amplitude of disturbances due to geometric resonance between surface wavelength and TS wavelength, was also confirmed by the experiments. Surface waviness was estimated, in some cases, to result in a nearly 1% decrease in fan efficiency compared to a non-wavy blade. Suggestions are given for mitigation of the effects of waviness, including the idea of blade curvature rescheduling as a method to delay transition and thus decrease loss.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"80 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"115047130","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper considers the impact of a damaged leading edge on the stall margin and stall inception mechanisms of a transonic, low pressure ratio fan. The damage takes the form of a squared-off leading edge over the upper half of the blade. Full-annulus, unsteady CFD simulations are used to explain the stall inception mechanisms for the fan at low- and high-speed operating points. A combination of steady and unsteady simulations show that the fan is predicted to be sensitive to leading edge damage at low speed, but insensitive at high speed. This blind prediction aligns well with rig test data. The difference in response is shown to be caused by the change between subsonic and supersonic flow regimes at the leading edge. Where the inlet relative flow is subsonic, rotating stall is initiated by growth and propagation of a subsonic leading edge flow separation. This separation is shown to be triggered at higher mass flow rates when the leading edge is damaged, reducing the stable flow range. Where the inlet relative flow is supersonic, the flow undergoes a supersonic expansion around the leading edge, creating a supersonic flow patch terminated by a shock on the suction surface. Rotating stall is triggered by growth of this separation, which is insensitive to leading edge shape. This creates a marked difference in sensitivity to damage at low- and high-speed operating points.
{"title":"Fan Stability With Leading Edge Damage: Blind Prediction and Validation","authors":"E. Gunn, T. Brandvik, M. Wilson, R. Maxwell","doi":"10.1115/gt2021-59495","DOIUrl":"https://doi.org/10.1115/gt2021-59495","url":null,"abstract":"\u0000 This paper considers the impact of a damaged leading edge on the stall margin and stall inception mechanisms of a transonic, low pressure ratio fan. The damage takes the form of a squared-off leading edge over the upper half of the blade.\u0000 Full-annulus, unsteady CFD simulations are used to explain the stall inception mechanisms for the fan at low- and high-speed operating points. A combination of steady and unsteady simulations show that the fan is predicted to be sensitive to leading edge damage at low speed, but insensitive at high speed. This blind prediction aligns well with rig test data.\u0000 The difference in response is shown to be caused by the change between subsonic and supersonic flow regimes at the leading edge. Where the inlet relative flow is subsonic, rotating stall is initiated by growth and propagation of a subsonic leading edge flow separation. This separation is shown to be triggered at higher mass flow rates when the leading edge is damaged, reducing the stable flow range. Where the inlet relative flow is supersonic, the flow undergoes a supersonic expansion around the leading edge, creating a supersonic flow patch terminated by a shock on the suction surface. Rotating stall is triggered by growth of this separation, which is insensitive to leading edge shape. This creates a marked difference in sensitivity to damage at low- and high-speed operating points.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"121 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125513414","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
T. Eggers, J. Friedrichs, J. Goessling, J. Seume, N. Natale, J. Flüh, N. Paletta
In the CA3ViAR (Composite fan Aerodynamic, Aeroelastic, and Aeroacoustic Validation Rig) project, a composite low-transonic fan is designed and tested. The aim is a scaled ultra-high bypass ratio (UHBR) fan with state-of-the-art aerodynamic performance and composite rotor blades, which features aeroelastic phenomena, e.g. forced response by inlet distortions and flutter, under certain operating points within the wind tunnel. In this paper, the aerodynamic and aeroelastic design process starting from the overall performance specifications to a threedimensional numerical model is described. A target of eigen-frequency and twist-to-plunge ratio is specified such that flutter occurs at desired operating conditions with a sufficient margin with respect to the working line. Different materials and layups of the composite blade are analyzed to reach the structural target. The fan should serve as an open test case to advance the future research on aerodynamic, aeroelastic, and aeroacoustic performance investigations in a wide range of operating conditions. A preliminary fan stage design is presented in this paper.
{"title":"Composite UHBR Fan for Forced Response and Flutter Investigations","authors":"T. Eggers, J. Friedrichs, J. Goessling, J. Seume, N. Natale, J. Flüh, N. Paletta","doi":"10.1115/gt2021-58941","DOIUrl":"https://doi.org/10.1115/gt2021-58941","url":null,"abstract":"\u0000 In the CA3ViAR (Composite fan Aerodynamic, Aeroelastic, and Aeroacoustic Validation Rig) project, a composite low-transonic fan is designed and tested. The aim is a scaled ultra-high bypass ratio (UHBR) fan with state-of-the-art aerodynamic performance and composite rotor blades, which features aeroelastic phenomena, e.g. forced response by inlet distortions and flutter, under certain operating points within the wind tunnel. In this paper, the aerodynamic and aeroelastic design process starting from the overall performance specifications to a threedimensional numerical model is described. A target of eigen-frequency and twist-to-plunge ratio is specified such that flutter occurs at desired operating conditions with a sufficient margin with respect to the working line. Different materials and layups of the composite blade are analyzed to reach the structural target. The fan should serve as an open test case to advance the future research on aerodynamic, aeroelastic, and aeroacoustic performance investigations in a wide range of operating conditions. A preliminary fan stage design is presented in this paper.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"120 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123391314","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Previous research into axial compressor stall has mainly focused on stall inception and methods to extend the stable operating range. This paper considers the performance of an axial compressor beyond stall and investigates how the characteristics of stall cells depend on Reynolds number. An experimental study has been conducted using a single-stage axial compressor capable of operating across the Reynolds number range of 10,000–100,000. Detailed unsteady measurements have been used to measure the behaviour across a range of install flow coefficients. These measurements have been used to extract the stall hysteresis and to determine the size, speed, number, and spanwise extent of the stall cells. The results show that for the stalled compressor, as Reynolds number increases, the size of the minimum stable stall cell decreases. This means that a larger change in throttle area is needed to reduce the stall cell down to a size where the compressor can recover from stall. At Re = 100,000, the stall hysteresis is six times greater than at Re = 20,000. At the design Reynolds number, the number of stall cells that form transitions from one, to two, and then to four stall cells as the flow coefficient is reduced. At lower Reynolds numbers, the two stall cell state becomes unstable; instead, a single stall cell transitions directly into five cells. In all cases, as the number of stall cells increases, so do the speed of the stall cells and the total size. Further reductions in the flow coefficient cause an increase in the total size of the stall cells and a decrease in the stall cell speed.
{"title":"In-Stall Compressor Performance and the Effects of Reynolds Number","authors":"J. Hutchings, C. Hall","doi":"10.1115/gt2021-58457","DOIUrl":"https://doi.org/10.1115/gt2021-58457","url":null,"abstract":"\u0000 Previous research into axial compressor stall has mainly focused on stall inception and methods to extend the stable operating range. This paper considers the performance of an axial compressor beyond stall and investigates how the characteristics of stall cells depend on Reynolds number. An experimental study has been conducted using a single-stage axial compressor capable of operating across the Reynolds number range of 10,000–100,000. Detailed unsteady measurements have been used to measure the behaviour across a range of install flow coefficients. These measurements have been used to extract the stall hysteresis and to determine the size, speed, number, and spanwise extent of the stall cells. The results show that for the stalled compressor, as Reynolds number increases, the size of the minimum stable stall cell decreases. This means that a larger change in throttle area is needed to reduce the stall cell down to a size where the compressor can recover from stall. At Re = 100,000, the stall hysteresis is six times greater than at Re = 20,000. At the design Reynolds number, the number of stall cells that form transitions from one, to two, and then to four stall cells as the flow coefficient is reduced. At lower Reynolds numbers, the two stall cell state becomes unstable; instead, a single stall cell transitions directly into five cells. In all cases, as the number of stall cells increases, so do the speed of the stall cells and the total size. Further reductions in the flow coefficient cause an increase in the total size of the stall cells and a decrease in the stall cell speed.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"24 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121742072","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The loss coefficient based only on the stagnation pressure has traditionally been used in the analysis of axial compressors for the comparison between shrouded and cantilevered stator configurations. In recent years, engineers have been able to perform more detailed Computational Fluid Dynamics simulations, allowing them to resolve the flow field in the leakage paths. The two stator hub designs are, however, affected by the rotating surfaces in a different way: in cantilevered stators, the relative rotation between the stator and the hub imparts energy to the hub flow, whereas in shrouded stators, the rotating inner leakage surface imparts energy to the seal cavity leakage flow. The aim of this work is to analyze the performance of a multi-stage axial compressor featuring a change of stator hub configuration, by employing both the conventional loss coefficient based on the stagnation pressure and the loss coefficient based on the entropy change. It is shown, that in the evaluation of the losses of a multi-stage axial machine, it is essential to consider the different 3D distributions of stagnation temperature resulting from the two stator hub configurations, which are transferred to the downstream rows.
{"title":"Numerical Evaluation of Losses in Shrouded and Cantilevered Stators of a Multi-Stage Axial Compressor","authors":"Ilaria De Dominicis, Sebastian Robens, V. Gümmer","doi":"10.1115/gt2021-58388","DOIUrl":"https://doi.org/10.1115/gt2021-58388","url":null,"abstract":"\u0000 The loss coefficient based only on the stagnation pressure has traditionally been used in the analysis of axial compressors for the comparison between shrouded and cantilevered stator configurations. In recent years, engineers have been able to perform more detailed Computational Fluid Dynamics simulations, allowing them to resolve the flow field in the leakage paths. The two stator hub designs are, however, affected by the rotating surfaces in a different way: in cantilevered stators, the relative rotation between the stator and the hub imparts energy to the hub flow, whereas in shrouded stators, the rotating inner leakage surface imparts energy to the seal cavity leakage flow.\u0000 The aim of this work is to analyze the performance of a multi-stage axial compressor featuring a change of stator hub configuration, by employing both the conventional loss coefficient based on the stagnation pressure and the loss coefficient based on the entropy change. It is shown, that in the evaluation of the losses of a multi-stage axial machine, it is essential to consider the different 3D distributions of stagnation temperature resulting from the two stator hub configurations, which are transferred to the downstream rows.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"35 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"115791656","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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