Future jet engines with shorter and thinner intakes present a greater risk of intake separation. This leads to a complex tip-low total pressure distortion pattern of varying circumferential extent. In this paper, an experimental study has been completed to determine the impact of such distortion patterns on the operating range and stalling behaviour of a low-speed fan rig. Unsteady casing static pressure measurements have been made during stall events in 11 circumferential extents of tip-low distortion. The performance has been measured and detailed area traverses have been performed at rotor inlet and outlet in 3 of these cases — clean, axisymmetric tip-low and half-annulus tip-low distortion. Axisymmetric tip-low distortion is found to reduce stall margin by 8%. It does not change the stalling mechanism compared to clean inflow. In both cases, high incidence at the tip combined with growth of the casing boundary layer drive instability. In contrast, half-annulus tip-low distortion is found to reduce stall margin by only 4% through a different mechanism. The distortion causes disturbances in the measured casing pressure signals to grow circumferentially in regions of high incidence. Stall occurs when these disturbances do not decay fully in the undistorted region. As the extent of the distorted sector is increased, the stability margin is found to reduce continuously. However, the maximum disturbance size before stall inception is found to occur at intermediate values of distorted sector extent. This corresponds to distortion patterns that provide sufficient circumferential length of undistorted region for disturbances to decay fully before they return to the distorted sector. It is found that as the extent of the tip-low distortion sector is increased, the circumferential size of the stall cell that develops is reduced. However, its speed is found to remain approximately constant at 50% of the rotor blade speed.
{"title":"An Experimental Investigation Into the Impacts of Varying the Circumferential Extent of Tip-Low Total Pressure Distortion on Fan Stability","authors":"Oliver Allen, A. C. Pardo, C. Hall","doi":"10.1115/gt2021-59851","DOIUrl":"https://doi.org/10.1115/gt2021-59851","url":null,"abstract":"Future jet engines with shorter and thinner intakes present a greater risk of intake separation. This leads to a complex tip-low total pressure distortion pattern of varying circumferential extent. In this paper, an experimental study has been completed to determine the impact of such distortion patterns on the operating range and stalling behaviour of a low-speed fan rig. Unsteady casing static pressure measurements have been made during stall events in 11 circumferential extents of tip-low distortion. The performance has been measured and detailed area traverses have been performed at rotor inlet and outlet in 3 of these cases — clean, axisymmetric tip-low and half-annulus tip-low distortion. Axisymmetric tip-low distortion is found to reduce stall margin by 8%. It does not change the stalling mechanism compared to clean inflow. In both cases, high incidence at the tip combined with growth of the casing boundary layer drive instability. In contrast, half-annulus tip-low distortion is found to reduce stall margin by only 4% through a different mechanism. The distortion causes disturbances in the measured casing pressure signals to grow circumferentially in regions of high incidence. Stall occurs when these disturbances do not decay fully in the undistorted region. As the extent of the distorted sector is increased, the stability margin is found to reduce continuously. However, the maximum disturbance size before stall inception is found to occur at intermediate values of distorted sector extent. This corresponds to distortion patterns that provide sufficient circumferential length of undistorted region for disturbances to decay fully before they return to the distorted sector. It is found that as the extent of the tip-low distortion sector is increased, the circumferential size of the stall cell that develops is reduced. However, its speed is found to remain approximately constant at 50% of the rotor blade speed.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"14 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":"115417879","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 deviation between the actual processed blade and the designed blade shape inevitably occurs in the process of compressor blade manufacturing. Rotor37 was used as the research object and a three-dimensional steady Reynolds averaged Navier-Stokes simulation method was adopted in order to study the influence mechanism of blade thickness deviation on blade performance. The blade was parameterized and the blade thicknesses were increased or decreased uniformly, with changes of 0.06mm and 0.1mm respectively. Results illustrate that the blade thickness deviation affects the total pressure ratio, isentropic efficiency and stability margin of the single-stage rotor. Increasing the blade thickness will inhibit the transport of low speed airflow from blade root area to blade tip area along the radial direction. In the peak efficiency condition, this inhibit will cause low speed airflow to converge in the middle of the blade and increase the flow separation loss; while in the reference near stall condition, the inhibition of low speed airflow transport will weaken the accumulation of low energy airflow in the tip area, reduce the loss in the corner area, and expand the stable working range of the blade. Further, increasing the blade thickness causes the shock wave position to move backward and the shock wave intensity will decrease.
{"title":"Mechanism Analysis of the Influence of Blade Thickness Deviation on The Performance of Axial Flow Compressor","authors":"Tian Ji, W. Chu, Zhengtao Guo, Jibo Yang","doi":"10.1115/gt2021-58823","DOIUrl":"https://doi.org/10.1115/gt2021-58823","url":null,"abstract":"\u0000 The deviation between the actual processed blade and the designed blade shape inevitably occurs in the process of compressor blade manufacturing. Rotor37 was used as the research object and a three-dimensional steady Reynolds averaged Navier-Stokes simulation method was adopted in order to study the influence mechanism of blade thickness deviation on blade performance. The blade was parameterized and the blade thicknesses were increased or decreased uniformly, with changes of 0.06mm and 0.1mm respectively. Results illustrate that the blade thickness deviation affects the total pressure ratio, isentropic efficiency and stability margin of the single-stage rotor. Increasing the blade thickness will inhibit the transport of low speed airflow from blade root area to blade tip area along the radial direction. In the peak efficiency condition, this inhibit will cause low speed airflow to converge in the middle of the blade and increase the flow separation loss; while in the reference near stall condition, the inhibition of low speed airflow transport will weaken the accumulation of low energy airflow in the tip area, reduce the loss in the corner area, and expand the stable working range of the blade. Further, increasing the blade thickness causes the shock wave position to move backward and the shock wave intensity will decrease.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"112 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":"124249864","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 flow in shrouded stator cavities can be quite complex with axial, radial, and circumferential variations. As the leakage flow recirculates and is re-injected into the main flow path upstream of the stator, it deteriorates the near-hub flow field and, thus, degrades the overall aerodynamic performance of the compressor. In addition, the windage heating in the cavity can raise thermal-mechanical concerns. Fully understanding the details of the shrouded-hub cavity flow in a multi-stage environment can enable better hub cavity designs. Since the majority of the open literature presents limited details about the structure of compressor cavity flows in the stator wells and how the cavity wells affect the leakage flow, there is a lack of wholistic knowledge of how these flow parameters are interdependent. To shed light on this topic, a coupled CFD model with inclusion of the stator cavity wells for the Purdue 3-Stage (P3S) Axial Compressor Research Facility using the PAX100 configuration was developed and validated against experimental data. Such a model not only quantifies the impact of cavity leakage flow on compressor performance, but it also provides the capability to investigate the flow structure details including the path of the fluid into and out of the cavity. With the model in place, in this part 1 paper, the influence of the hub leakage flow on compressor performance and its interactions with the primary flow were investigated by varying the clearance ratio of a single stator. The understanding of the primary-hub-leakage flow interactions can offer insights leading to better designs of hub cavities.
{"title":"Details of Shrouded Stator Hub Cavity Flow in a Multi-Stage Axial Compressor Part 1: Interactions With the Primary Flow","authors":"Nitya Kamdar, Fangyuan Lou, N. Key","doi":"10.1115/gt2021-60103","DOIUrl":"https://doi.org/10.1115/gt2021-60103","url":null,"abstract":"\u0000 The flow in shrouded stator cavities can be quite complex with axial, radial, and circumferential variations. As the leakage flow recirculates and is re-injected into the main flow path upstream of the stator, it deteriorates the near-hub flow field and, thus, degrades the overall aerodynamic performance of the compressor. In addition, the windage heating in the cavity can raise thermal-mechanical concerns. Fully understanding the details of the shrouded-hub cavity flow in a multi-stage environment can enable better hub cavity designs. Since the majority of the open literature presents limited details about the structure of compressor cavity flows in the stator wells and how the cavity wells affect the leakage flow, there is a lack of wholistic knowledge of how these flow parameters are interdependent. To shed light on this topic, a coupled CFD model with inclusion of the stator cavity wells for the Purdue 3-Stage (P3S) Axial Compressor Research Facility using the PAX100 configuration was developed and validated against experimental data. Such a model not only quantifies the impact of cavity leakage flow on compressor performance, but it also provides the capability to investigate the flow structure details including the path of the fluid into and out of the cavity. With the model in place, in this part 1 paper, the influence of the hub leakage flow on compressor performance and its interactions with the primary flow were investigated by varying the clearance ratio of a single stator. The understanding of the primary-hub-leakage flow interactions can offer insights leading to better designs of hub cavities.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"29 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":"129306310","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}
Ongoing experiments conducted in a one-and-half stages axial compressor installed in the JHU refractive index-matched facility investigate the evolution of flow structure across blade rows. After previously focusing only on the rotor tip region, the present stereo-PIV (SPIV) measurements are performed in a series of axial planes covering an entire passage across the machine, including upstream of the IGV, IGV-rotor gap, rotor-stator gap, and downstream of the stator. The measurements are performed at flow rates corresponding to pre-stall condition and best efficiency point (BEP). Data are acquired for various rotor-blade orientations relative to the IGV and stator blades. The results show that at BEP, the wakes of IGV and rotor are much more distinct and the wake signatures of one row persists downstream of the next, e.g., the flow downstream of the stator is strongly affected by the rotor orientation. In contrast, under pre-stall conditions, the rotor orientation has minimal effect on the flow structure downstream of the stator. However, the wakes of the stator blades, where the axial momentum is low, are now wider. For both conditions, the flow downstream of the rotor is characterized by two regions of axial momentum deficit in addition to the rotor wake. A deficit on the pressure side of the rotor wake tip is associated with the tip leakage vortex (TLV) of the previous rotor blade, and is much broader at pre-stall condition. A deficit on the suction side of the rotor wake near the hub appears to be associated with the hub vortex generated by the neighboring blade, and is broader at BEP. At pre-stall, while the axial momentum upstream of the rotor decreases over the entire tip region, it is particularly evident near the rotor blade tip, where the instantaneous axial velocity becomes intermittently negative. Downstream of the rotor, there is a substantial reduction in mean axial momentum in the upper half of the passage, concurrently with an increase in the circumferential velocity. Consequently, the incidence angle upstream of the stator increases in certain regions by up to 30 degrees. These observations suggest that while the onset of the stall originates from the rotor tip flow, one must examine its impact on the flow structure in the stator passage as well.
{"title":"Experimental Characterization of the Evolution of Global Flow Structure in the Passage of an Axial Compressor","authors":"Ayush Saraswat, S. Koley, J. Katz","doi":"10.1115/gt2021-60325","DOIUrl":"https://doi.org/10.1115/gt2021-60325","url":null,"abstract":"\u0000 Ongoing experiments conducted in a one-and-half stages axial compressor installed in the JHU refractive index-matched facility investigate the evolution of flow structure across blade rows. After previously focusing only on the rotor tip region, the present stereo-PIV (SPIV) measurements are performed in a series of axial planes covering an entire passage across the machine, including upstream of the IGV, IGV-rotor gap, rotor-stator gap, and downstream of the stator. The measurements are performed at flow rates corresponding to pre-stall condition and best efficiency point (BEP). Data are acquired for various rotor-blade orientations relative to the IGV and stator blades. The results show that at BEP, the wakes of IGV and rotor are much more distinct and the wake signatures of one row persists downstream of the next, e.g., the flow downstream of the stator is strongly affected by the rotor orientation. In contrast, under pre-stall conditions, the rotor orientation has minimal effect on the flow structure downstream of the stator. However, the wakes of the stator blades, where the axial momentum is low, are now wider. For both conditions, the flow downstream of the rotor is characterized by two regions of axial momentum deficit in addition to the rotor wake. A deficit on the pressure side of the rotor wake tip is associated with the tip leakage vortex (TLV) of the previous rotor blade, and is much broader at pre-stall condition. A deficit on the suction side of the rotor wake near the hub appears to be associated with the hub vortex generated by the neighboring blade, and is broader at BEP. At pre-stall, while the axial momentum upstream of the rotor decreases over the entire tip region, it is particularly evident near the rotor blade tip, where the instantaneous axial velocity becomes intermittently negative. Downstream of the rotor, there is a substantial reduction in mean axial momentum in the upper half of the passage, concurrently with an increase in the circumferential velocity. Consequently, the incidence angle upstream of the stator increases in certain regions by up to 30 degrees. These observations suggest that while the onset of the stall originates from the rotor tip flow, one must examine its impact on the flow structure in the stator passage as well.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"187 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":"115728254","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 investigates the surface boundary layer and wake development of a compressor blade at a range of low Reynolds number from 45000 to 120000. Experiments in a miniature linear compressor cascade facility have been performed with detailed surface pressure measurements and flow visualization to track variations in the separation bubble size. These have been combined with high resolution pneumatic pressure and hot wire probe traverses in the downstream wake. High fidelity DNS simulations have been completed on the same compressor blade section across the same range of operating conditions. The results show that large laminar separation bubbles exist on both blade surfaces. As Reynolds number increases, these separation bubbles shorten in length and reduce in thickness. Correspondingly, the downstream wake narrows, although the peak wake loss coefficient remains approximately constant. As the Reynolds number is increased from 45000 to 120000 the bubble length on the suction side reduced from 48% to 28% chord and on the pressure side reduced from 35% to 20% chord, while the loss coefficient reduced from 9% to 5%. The flow features are examined further within the high-fidelity computations, which reveal the dependence of the wake turbulence on the laminar separation bubbles. The separation bubbles are found to generate turbulent kinetic energy, which convects downstream to form the outer part of wake. As Re increases, a shorter bubble produces less turbulence in the outer part of the boundary layer leading to a narrower wake. However, the trailing edge separation is largely independent of Reynolds number, leading to the constant peak loss coefficient observed. The overall loss is shown to vary linearly with the total turbulence production, and this depends on the size of the separation bubbles. Overall, this research provides new insight into the connection between the blade surface flow field and the wake characteristics at low Reynolds number. The findings suggest that changes that minimize the extent of the blade separation bubbles could provide significant improvements to both the steady and unsteady properties of the wake.
{"title":"Low Reynolds Number Effects on the Separation and Wake of a Compressor Blade","authors":"Qiang Liu, W. Ager, C. Hall, Andrew P. S. Wheeler","doi":"10.1115/gt2021-59284","DOIUrl":"https://doi.org/10.1115/gt2021-59284","url":null,"abstract":"\u0000 This paper investigates the surface boundary layer and wake development of a compressor blade at a range of low Reynolds number from 45000 to 120000. Experiments in a miniature linear compressor cascade facility have been performed with detailed surface pressure measurements and flow visualization to track variations in the separation bubble size. These have been combined with high resolution pneumatic pressure and hot wire probe traverses in the downstream wake. High fidelity DNS simulations have been completed on the same compressor blade section across the same range of operating conditions.\u0000 The results show that large laminar separation bubbles exist on both blade surfaces. As Reynolds number increases, these separation bubbles shorten in length and reduce in thickness. Correspondingly, the downstream wake narrows, although the peak wake loss coefficient remains approximately constant. As the Reynolds number is increased from 45000 to 120000 the bubble length on the suction side reduced from 48% to 28% chord and on the pressure side reduced from 35% to 20% chord, while the loss coefficient reduced from 9% to 5%. The flow features are examined further within the high-fidelity computations, which reveal the dependence of the wake turbulence on the laminar separation bubbles. The separation bubbles are found to generate turbulent kinetic energy, which convects downstream to form the outer part of wake. As Re increases, a shorter bubble produces less turbulence in the outer part of the boundary layer leading to a narrower wake. However, the trailing edge separation is largely independent of Reynolds number, leading to the constant peak loss coefficient observed. The overall loss is shown to vary linearly with the total turbulence production, and this depends on the size of the separation bubbles.\u0000 Overall, this research provides new insight into the connection between the blade surface flow field and the wake characteristics at low Reynolds number. The findings suggest that changes that minimize the extent of the blade separation bubbles could provide significant improvements to both the steady and unsteady properties of the wake.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"16 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":"131928709","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}
Surge margin is critical for the safety of aeronautical compressors, hence predicting it early in the design process using CFD is mandatory. However, close to surge, steady-state Reynolds Averaged Navier-Stokes (RANS) simulations are proven inadequate. Unsteady techniques such as Unsteady RANS (URANS) and Large Eddy Simulation (LES) can provide more reliable predictions. Nevertheless, the accuracy of such methods are dependent on the method used to handle the rotor/stator interfaces. The most precise method, the sliding mesh, requires simulating the full annulus or a periodic sector, which can be very costly. Other techniques to reduce the domain exist, such as the phase-lagged approach or geometric blade scaling, but introduce restrictive assumptions on the flow at near-stall conditions. The objective of this paper is to investigate the near-stall flow of a low-pressure compressor using unsteady methods of varying fidelity: URANS with the phase lag assumption, URANS on a periodic sector and a high-fidelity LES on a smaller periodic sector achieved using geometric blade scaling. Results are compared to experimental measurements. An overall good agreement is found. Results show that the tip leakage vortex is not the origin of the stall on the studied configuration and a hub corner separation is initiated. LES further validates the (U)RANS flow predictions and brings additional insight on unsteady flow separations.
{"title":"Low-Pressure Compressor Near-Stall Predictions Using Unsteady CFD Methods","authors":"D. Vanpouille, D. Papadogiannis, S. Hiernaux","doi":"10.1115/gt2021-59186","DOIUrl":"https://doi.org/10.1115/gt2021-59186","url":null,"abstract":"\u0000 Surge margin is critical for the safety of aeronautical compressors, hence predicting it early in the design process using CFD is mandatory. However, close to surge, steady-state Reynolds Averaged Navier-Stokes (RANS) simulations are proven inadequate. Unsteady techniques such as Unsteady RANS (URANS) and Large Eddy Simulation (LES) can provide more reliable predictions. Nevertheless, the accuracy of such methods are dependent on the method used to handle the rotor/stator interfaces. The most precise method, the sliding mesh, requires simulating the full annulus or a periodic sector, which can be very costly. Other techniques to reduce the domain exist, such as the phase-lagged approach or geometric blade scaling, but introduce restrictive assumptions on the flow at near-stall conditions. The objective of this paper is to investigate the near-stall flow of a low-pressure compressor using unsteady methods of varying fidelity: URANS with the phase lag assumption, URANS on a periodic sector and a high-fidelity LES on a smaller periodic sector achieved using geometric blade scaling. Results are compared to experimental measurements. An overall good agreement is found. Results show that the tip leakage vortex is not the origin of the stall on the studied configuration and a hub corner separation is initiated. LES further validates the (U)RANS flow predictions and brings additional insight on unsteady flow separations.","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":"130554568","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 development of viscous endwall flow is of major importance when considering highly-loaded compressor stages. Essentially, all losses occurring in a subsonic compressor are caused by viscous shear stresses building up boundary layers on individual aerofoils and endwall surfaces. These boundary layers cause significant aerodynamic blockage and cause a reduction in effective flow area, depending on the specifics of the stage design. The presented work describes the numerical investigation of blockage development in a 3.5-stage low-speed compressor with tandem stator vanes. The research is aimed at understanding the mechanism of blockage generation and growth in tandem vane rows and across the entire compressor. Therefore, the blockage generation is investigated as a function of the operating point, the rotational speed and the inlet boundary layer thickness.
{"title":"Endwall Boundary Layer Development in a Multistage Low-Speed Compressor With Tandem Stator Vanes","authors":"Michael Hopfinger, V. Gümmer","doi":"10.1115/gt2021-58742","DOIUrl":"https://doi.org/10.1115/gt2021-58742","url":null,"abstract":"\u0000 The development of viscous endwall flow is of major importance when considering highly-loaded compressor stages. Essentially, all losses occurring in a subsonic compressor are caused by viscous shear stresses building up boundary layers on individual aerofoils and endwall surfaces. These boundary layers cause significant aerodynamic blockage and cause a reduction in effective flow area, depending on the specifics of the stage design. The presented work describes the numerical investigation of blockage development in a 3.5-stage low-speed compressor with tandem stator vanes. The research is aimed at understanding the mechanism of blockage generation and growth in tandem vane rows and across the entire compressor. Therefore, the blockage generation is investigated as a function of the operating point, the rotational speed and the inlet boundary layer thickness.","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":"129641229","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 most common compressor map framework, referred to here as the β-framework, will be shown to suffer from limitations that grow more troublesome in the multiple-map environment. When maps are coupled in series in the β-framework, it is very common to find operating points that are physically unrealizable, but these cannot generally be avoided without first generating them. A feasible situation is described in which the β-framework leads to an apparent physical paradox. In the proposed S-framework, the map itself is recast in terms of independent variables (corrected speed and exit corrected flow) and dependent variables (inlet corrected flow and temperature ratio). The propagation of information in map coupling is split into an upstream-marching corrected flow ‘flux’ and a downstream-marching temperature ‘flux’. Finding the equilibrium operating point requires only finding a simple intersection between curves. The S-framework is then developed further into a more compact S’-framework that exhibits a natural set of qualitative symmetries. The S- and S’-frameworks are shown to simplify compressor map expression, resolve the problems shown with the β-framework, and aid intuition with regard to off-design phenomena. The resolution of the paradox using the S’-framework is a new description of multistage compressor performance hysteresis.
{"title":"Compressor Maps and Coupling: Symmetry, Paradox, and Clarity","authors":"Benjamin Iwrey","doi":"10.1115/gt2021-60012","DOIUrl":"https://doi.org/10.1115/gt2021-60012","url":null,"abstract":"\u0000 The most common compressor map framework, referred to here as the β-framework, will be shown to suffer from limitations that grow more troublesome in the multiple-map environment.\u0000 When maps are coupled in series in the β-framework, it is very common to find operating points that are physically unrealizable, but these cannot generally be avoided without first generating them. A feasible situation is described in which the β-framework leads to an apparent physical paradox.\u0000 In the proposed S-framework, the map itself is recast in terms of independent variables (corrected speed and exit corrected flow) and dependent variables (inlet corrected flow and temperature ratio). The propagation of information in map coupling is split into an upstream-marching corrected flow ‘flux’ and a downstream-marching temperature ‘flux’. Finding the equilibrium operating point requires only finding a simple intersection between curves.\u0000 The S-framework is then developed further into a more compact S’-framework that exhibits a natural set of qualitative symmetries. The S- and S’-frameworks are shown to simplify compressor map expression, resolve the problems shown with the β-framework, and aid intuition with regard to off-design phenomena. The resolution of the paradox using the S’-framework is a new description of multistage compressor performance hysteresis.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"132 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":"127085830","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}
Isak Jonsson, C. Xisto, M. Lejon, Anders Dahl, T. Grönstedt
The use of hydrogen as aviation fuel is again resurfacing with unprecedented vigor. It is well known that hydrogen is a formidable heat sink and the use of heat sinks in the compression system of an aero engine may enable not only preheating of the fuel but also improve the gas turbine cycle itself. One such opportunity arises from extracting heat to the fuel as part of the compression process. This work presents the design process and pre-test evaluation of a low-speed compressor test facility dedicated to aerothermal measurements. The design has been derived from a high-speed transonic compressor developed for a large sized geared turbofan engine. The proposed pre-test evaluation methodology provides a comprehensive and affordable way to estimate facility accuracy by virtually addressing all the experimental procedures, from data acquisition to a final performance map. The evaluation of gathering compressor performance parameters via a gas-path investigation process was achieved while relying on results from numerical simulations. The pre-test evaluation details uncertainties introduced throughout this process with transducers, flow and probe specific errors, traverse discretization, and data normalization. A suitable instrumentation configuration is presented which shows that the performance parameters pressure ratio (Π) and isentropic efficiency (ηc) can be determined with uncertainties below 1% for most operating conditions and below 0.5% at design conditions.
{"title":"Design and Pre-Test Evaluation of a Low-Pressure Compressor Test Facility for Cryogenic Hydrogen Fuel Integration","authors":"Isak Jonsson, C. Xisto, M. Lejon, Anders Dahl, T. Grönstedt","doi":"10.1115/gt2021-58946","DOIUrl":"https://doi.org/10.1115/gt2021-58946","url":null,"abstract":"\u0000 The use of hydrogen as aviation fuel is again resurfacing with unprecedented vigor. It is well known that hydrogen is a formidable heat sink and the use of heat sinks in the compression system of an aero engine may enable not only preheating of the fuel but also improve the gas turbine cycle itself. One such opportunity arises from extracting heat to the fuel as part of the compression process. This work presents the design process and pre-test evaluation of a low-speed compressor test facility dedicated to aerothermal measurements. The design has been derived from a high-speed transonic compressor developed for a large sized geared turbofan engine. The proposed pre-test evaluation methodology provides a comprehensive and affordable way to estimate facility accuracy by virtually addressing all the experimental procedures, from data acquisition to a final performance map. The evaluation of gathering compressor performance parameters via a gas-path investigation process was achieved while relying on results from numerical simulations. The pre-test evaluation details uncertainties introduced throughout this process with transducers, flow and probe specific errors, traverse discretization, and data normalization. A suitable instrumentation configuration is presented which shows that the performance parameters pressure ratio (Π) and isentropic efficiency (ηc) can be determined with uncertainties below 1% for most operating conditions and below 0.5% at design conditions.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"28 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":"127393381","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}
Sweep in a transonic fan is conventionally used to reduce design point losses by inclining the passage shock relative to the incoming flow. However, future low pressure ratio fans operate to lower Mach numbers meaning the role of sweep at cruise is diminished. Instead, sweep might be repurposed to improve the performance of critical high Mach number off-design conditions such as high angle of attack (AOA). In this paper, we use unsteady computational fluid dynamics to compare two transonic low pressure ratio fans, one radially stacked and one highly swept, coupled to a short intake design, at the high AOA flight condition. The AOA considered is 35°, which is sufficient to separate the intake bottom lip. The midspan of the swept fan was shifted upstream to add positive sweep to the outer span. Based on previous design experience, it was hypothesised the swept fan would reduce transonic losses when operating at high AOA. However, it was found the swept fan increased the rotor loss by 24% relative to the radial fan. Loss was increased through two key mechanisms. i) Rotor choking: flow is redistributed around the intake separation and enters the rotor midspan with high Mach numbers. Sweeping the fan upstream reduced the effective intake length, which increased the inlet relative Mach number and amplified choking losses. ii): Rotor-separation interaction (RSI): the rotor tip experiences low mass flow inside the separation, which increases the pressure rise across the casing to a point where the boundary layer separates. The swept fan diffused the casing streamtube, causing the casing separation to increase in size and persist in the passage for longer. High RSI loss indicated the swept fan was operating closer to the rotating stall point.
{"title":"Sweep Effects on Fan-Intake Aerodynamics at High Angle of Attack","authors":"B. Mohankumar, C. Hall, M. Wilson","doi":"10.1115/gt2021-58569","DOIUrl":"https://doi.org/10.1115/gt2021-58569","url":null,"abstract":"\u0000 Sweep in a transonic fan is conventionally used to reduce design point losses by inclining the passage shock relative to the incoming flow. However, future low pressure ratio fans operate to lower Mach numbers meaning the role of sweep at cruise is diminished. Instead, sweep might be repurposed to improve the performance of critical high Mach number off-design conditions such as high angle of attack (AOA). In this paper, we use unsteady computational fluid dynamics to compare two transonic low pressure ratio fans, one radially stacked and one highly swept, coupled to a short intake design, at the high AOA flight condition. The AOA considered is 35°, which is sufficient to separate the intake bottom lip.\u0000 The midspan of the swept fan was shifted upstream to add positive sweep to the outer span. Based on previous design experience, it was hypothesised the swept fan would reduce transonic losses when operating at high AOA. However, it was found the swept fan increased the rotor loss by 24% relative to the radial fan. Loss was increased through two key mechanisms. i) Rotor choking: flow is redistributed around the intake separation and enters the rotor midspan with high Mach numbers. Sweeping the fan upstream reduced the effective intake length, which increased the inlet relative Mach number and amplified choking losses. ii): Rotor-separation interaction (RSI): the rotor tip experiences low mass flow inside the separation, which increases the pressure rise across the casing to a point where the boundary layer separates. The swept fan diffused the casing streamtube, causing the casing separation to increase in size and persist in the passage for longer. High RSI loss indicated the swept fan was operating closer to the rotating stall point.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"61 Suppl 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":"129123268","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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