H. Kodama, K. Funazaki, Kodai Ito, R. Ikehata, Y. Koda
This paper describes the method for accurately estimating profile loss using measured flow parameters in a low speed cascade test. One of the issues of the control volume analysis for obtaining the mixed-out profile loss is to give realistic flow conditions on the control surface at the trailing edge plane. In this paper, the flow conditions at the trailing edge plane are related to the measured velocity distribution on the suction surface from throat to the trailing edge using the concept of circulation. The validity of the estimation method is verified by using the results of steady Reynolds Averaged Navier-Stokes (RANS) simulation. In this verification, the verification profile loss of the RANS simulation is obtained by using the calculated drag forces of surface pressure and surface shear stress. The verification is conducted for two kinds of blade profiles at three different Reynolds numbers, 57,000, 100,000 and 147,000, for each blade profile. It is found that all the estimations by the current method are in a range between −10% to +6% of the verification profile losses of the RANS simulations. This method is applied to the data analysis of low speed cascade tests for the two kinds of blade profiles used in the verification. It is shown that the measured pressure loss downstream of the cascade includes an additional loss to the profile loss even in the steady flow.
{"title":"A Method for Estimating Profile Loss of Low Pressure Turbine Blades From the Low Speed Cascade Test Data","authors":"H. Kodama, K. Funazaki, Kodai Ito, R. Ikehata, Y. Koda","doi":"10.1115/gt2019-90401","DOIUrl":"https://doi.org/10.1115/gt2019-90401","url":null,"abstract":"\u0000 This paper describes the method for accurately estimating profile loss using measured flow parameters in a low speed cascade test. One of the issues of the control volume analysis for obtaining the mixed-out profile loss is to give realistic flow conditions on the control surface at the trailing edge plane. In this paper, the flow conditions at the trailing edge plane are related to the measured velocity distribution on the suction surface from throat to the trailing edge using the concept of circulation.\u0000 The validity of the estimation method is verified by using the results of steady Reynolds Averaged Navier-Stokes (RANS) simulation. In this verification, the verification profile loss of the RANS simulation is obtained by using the calculated drag forces of surface pressure and surface shear stress. The verification is conducted for two kinds of blade profiles at three different Reynolds numbers, 57,000, 100,000 and 147,000, for each blade profile. It is found that all the estimations by the current method are in a range between −10% to +6% of the verification profile losses of the RANS simulations.\u0000 This method is applied to the data analysis of low speed cascade tests for the two kinds of blade profiles used in the verification. It is shown that the measured pressure loss downstream of the cascade includes an additional loss to the profile loss even in the steady flow.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"6 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129111719","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}
A. Walker, I. Mariah, D. Tsakmakidou, H. Vadhvana, C. Hall
To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet. Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.
{"title":"The Influence of Fan Root Flow on the Aerodynamics of a Low-Pressure Compressor Transition Duct","authors":"A. Walker, I. Mariah, D. Tsakmakidou, H. Vadhvana, C. Hall","doi":"10.1115/gt2019-92048","DOIUrl":"https://doi.org/10.1115/gt2019-92048","url":null,"abstract":"\u0000 To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the higher required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. Additionally, the desire for shorter, lighter nacelles means that the engine intake may be unable provide a uniform inlet flow to the fan when the aircraft is at an angle of attack or subject to cross winds. Any inlet distortion this generates can also further degrade the quality of the flow entering the core of the engine. This paper uses a combination of experiments and CFD to examine the effects of the inlet flow on the aerodynamics of an engine section splitter and transition duct designed to feed the low-pressure spool of a high bypass ratio turbofan. A fully annular test facility incorporating a 1½ stage axial compressor was used to compare the system performance of a rotor that produced a nominally flat profile with one that had a notably hub deficient flow. A RANS CFD model, employing a mixing plane between the rotor and Engine Section Stator (ESS) and a Reynolds Stress turbulence model, was then validated and used to further investigate the effects of increased inlet boundary layer thickness and bulk swirl distortion at rotor inlet.\u0000 Overall, changes to the inlet condition were seen to have a surprisingly small effect on the flow at duct exit — i.e. the flow presented to the downstream compressor. Changes to the inlet did, however, generate increased secondary flows and degrade the performance of the ESS. This resulted in notably increased total pressure loss; in excess of 12% for the hub-low inlet and in excess of 30% at high inlet swirl where the flow in the ESS separated. However, the increased ESS wake structures, and the enhanced mixing, delayed separation in the duct suggesting that, overall the design was reasonably robust, albeit with a significant penalty in system loss.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"68 2","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114124544","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}
F. Rubechini, M. Giovannini, A. Arnone, D. Simoni, F. Bertini
This paper deals with the design of passive control devices for reducing the impact of secondary flows on the aerodynamics of low-pressure turbine (LPT) stages. A novel kind of device is introduced which consists of shelf-like fences to be added to the blade surface. Such a device is intended to hinder the development of secondary flows, thus reducing losses and flow turning deviation with respect to the straight blade. The first part of this work is devoted to the design of the blade fences, whereas the second part addresses the experimental validation of the device. The blade fences are designed on a LPT stator vane, in an engine-like environment. As secondary flows generated by one blade row produce their major effects on the downstream one, and hence on the stage performance, the assessment is performed on a stator-rotor configuration. Steady calculations are considered for the design, then the optimal geometry is verified via unsteady calculations to include the effects of the actual interaction. The geometry and layout of the blade fences are effectively handled by means of a parametric approach, which enables the fast generation of several configurations. An optimization procedure, based on Artificial Neural Networks (ANNs) is exploited to drive the fences design. The analysis of the relative merit of each solution is carried out using a state-of-the-art CFD approach. Finally, a detailed comparison between the original blade and the one equipped with fences is presented, and the physical mechanisms responsible for the mitigation of secondary flow losses are discussed in detail.
{"title":"Reducing Secondary Flow Losses in Low-Pressure Turbines With Blade Fences: Part I — Design in an Engine-Like Environment","authors":"F. Rubechini, M. Giovannini, A. Arnone, D. Simoni, F. Bertini","doi":"10.1115/gt2019-91280","DOIUrl":"https://doi.org/10.1115/gt2019-91280","url":null,"abstract":"\u0000 This paper deals with the design of passive control devices for reducing the impact of secondary flows on the aerodynamics of low-pressure turbine (LPT) stages. A novel kind of device is introduced which consists of shelf-like fences to be added to the blade surface. Such a device is intended to hinder the development of secondary flows, thus reducing losses and flow turning deviation with respect to the straight blade.\u0000 The first part of this work is devoted to the design of the blade fences, whereas the second part addresses the experimental validation of the device. The blade fences are designed on a LPT stator vane, in an engine-like environment. As secondary flows generated by one blade row produce their major effects on the downstream one, and hence on the stage performance, the assessment is performed on a stator-rotor configuration. Steady calculations are considered for the design, then the optimal geometry is verified via unsteady calculations to include the effects of the actual interaction. The geometry and layout of the blade fences are effectively handled by means of a parametric approach, which enables the fast generation of several configurations. An optimization procedure, based on Artificial Neural Networks (ANNs) is exploited to drive the fences design. The analysis of the relative merit of each solution is carried out using a state-of-the-art CFD approach. Finally, a detailed comparison between the original blade and the one equipped with fences is presented, and the physical mechanisms responsible for the mitigation of secondary flow losses are discussed in detail.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"47 4","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114126527","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 the load of the turbine components of aircraft engines continuously increases, shock loss becomes the dominant factor of turbine stage loss and has become a hot topic. Schlieren technique is one of the few effective experimental methods to observe and study shock wave and, thus, has been widely used. Nevertheless, limited by camera accuracy and computer image processing technology, quantitative schlieren analysis methods were difficult to achieve in engineering applications. Fortunately, several quantitative schlieren methods have been developed with the help of new digital technology. Applying schlieren technique to the highly-loaded turbine cascade test is of great significance to the study of shock wave in highly-loaded turbine cascades. In this paper, the results of quantitative density field and shock intensity and loss in the cascade are obtained by using a double reflection type monochrome schlieren device. The boundary condition of density field is obtained by pressure test, and MATLAB software is used as image processing calculation tool. The quantitative results of this paper prove the feasibility of applying quantitative schlieren method to highly-loaded turbine cascade tests. Also, the implemented image processing method and density boundary condition acquisition method are suitable and convenient for cascade flow and shock measurement tests.
{"title":"Application of Quantitative Monochrome Calibrated Schlieren Technique in Highly-Loaded Turbine Cascade Test","authors":"Zhihong Zhou, Haotian Wang, Huo-xing Liu","doi":"10.1115/gt2019-90342","DOIUrl":"https://doi.org/10.1115/gt2019-90342","url":null,"abstract":"\u0000 As the load of the turbine components of aircraft engines continuously increases, shock loss becomes the dominant factor of turbine stage loss and has become a hot topic. Schlieren technique is one of the few effective experimental methods to observe and study shock wave and, thus, has been widely used. Nevertheless, limited by camera accuracy and computer image processing technology, quantitative schlieren analysis methods were difficult to achieve in engineering applications. Fortunately, several quantitative schlieren methods have been developed with the help of new digital technology. Applying schlieren technique to the highly-loaded turbine cascade test is of great significance to the study of shock wave in highly-loaded turbine cascades. In this paper, the results of quantitative density field and shock intensity and loss in the cascade are obtained by using a double reflection type monochrome schlieren device. The boundary condition of density field is obtained by pressure test, and MATLAB software is used as image processing calculation tool. The quantitative results of this paper prove the feasibility of applying quantitative schlieren method to highly-loaded turbine cascade tests. Also, the implemented image processing method and density boundary condition acquisition method are suitable and convenient for cascade flow and shock measurement tests.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"299 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"132719425","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}
During the lifetime of a turbine stage, some of the blade tips may undergo changes due to mechanical rubbing with casing surface and also due to thermal oxidation. Understanding the effect these damaged blades have over the undamaged blades is essential to estimate the performance of the turbine stage in the operable tip clearance range. In this paper, the passage to passage aerodynamic interaction in a turbine stage is studied by modifying the tip gap of selected turbine blades and analyzing their effect on the neighboring blade passage flows. The experiments in this study are carried out in a single-stage low-speed axial turbine facility. All measurements are taken in the stationary frame of reference using a time-accurate differential dynamic pressure transducer mounted in a Kiel probe head. The experimental results from this study show that even with a significant increase on a selected blade’s tip clearance, its effect on the AFTRF turbine flow is only confined to its neighboring blade passage. The disturbances due to the altered tip clearance of one passage are not measurably propagated to its neighboring turbine passages. The changes made in one of the blades in a turbine stage do not significantly alter the aerodynamic performance of other blades. This result is particularly important for large-scale turbine research rigs such as AFTRF where the unsteady total pressure field is mapped in a time-efficient and phase-locked manner.
{"title":"An Experimental Study of Passage-to-Passage Flow Interactions in a Single Stage Axial Flow Research Turbine Rotor","authors":"Veerandra C. Andichamy, C. Camcı, Y. Kim","doi":"10.1115/gt2019-91629","DOIUrl":"https://doi.org/10.1115/gt2019-91629","url":null,"abstract":"\u0000 During the lifetime of a turbine stage, some of the blade tips may undergo changes due to mechanical rubbing with casing surface and also due to thermal oxidation. Understanding the effect these damaged blades have over the undamaged blades is essential to estimate the performance of the turbine stage in the operable tip clearance range. In this paper, the passage to passage aerodynamic interaction in a turbine stage is studied by modifying the tip gap of selected turbine blades and analyzing their effect on the neighboring blade passage flows. The experiments in this study are carried out in a single-stage low-speed axial turbine facility. All measurements are taken in the stationary frame of reference using a time-accurate differential dynamic pressure transducer mounted in a Kiel probe head. The experimental results from this study show that even with a significant increase on a selected blade’s tip clearance, its effect on the AFTRF turbine flow is only confined to its neighboring blade passage. The disturbances due to the altered tip clearance of one passage are not measurably propagated to its neighboring turbine passages. The changes made in one of the blades in a turbine stage do not significantly alter the aerodynamic performance of other blades. This result is particularly important for large-scale turbine research rigs such as AFTRF where the unsteady total pressure field is mapped in a time-efficient and phase-locked manner.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"81 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"133233139","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}
M. Zenz, Asim Hafizovic, L. Simonassi, P. Leitl, F. Heitmeir, A. Marn
One of the main goals for modern aircrafts is to lower the fuel consumption and noise emissions without worsening the aerodynamic performance. One possibility to lower the fuel consumption is to reduce the skin-friction losses of vanes and blades inside the engine. Therefore, this paper is about the aeroacoustical as well as the aerodynamical effects of a riblet foil applied on the suction side surface of turbine exit guide vanes (TEGVs) of a 1½ stage low pressure turbine (LPT). There have been numerous studies concerning riblets but none using them in a LPT. In general, if riblets are applied on the suction side of vanes or blades, they lower the drag and increase the lift. Test runs were performed under two different operating points in a subsonic test turbine facility for aerodynamic, aeroacoustic, and aeroelastic investigations (STTF-AAAI) located at the Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology. One operating point was the design point of the riblets and the second one an off-design point. During the test campaign, two different set-ups have been investigated. One configuration with riblets applied on the suction side of the TEGVs, and one configuration with a smooth foil on the vanes to achieve the same thickness as the first set-up. This smooth configuration serves as a reference case. The tested riblet structure was of trapezoid type with 45 μm tip distance and a height to tip distance ratio of 0.45. The acoustical data has been obtained by using flush mounted condenser microphones, rotated over 360 deg around the flow channel. The aerodynamical data was obtained by using an aerodynamic five-hole-probe as well as a trailing edge probe. Measuring in planes up- and downstream of each TEGV allowed the comparison of a rough pressure loss estimation between the two studied set-ups. The present work gives a closer insight into the change of the acoustical and aerodynamical behaviour by applying riblets to LPT vanes.
{"title":"Aeroacoustical and Aerodynamical Investigations of Riblets Applied on Low Pressure Turbine Exit Guide Vanes for Two Different Operating Points","authors":"M. Zenz, Asim Hafizovic, L. Simonassi, P. Leitl, F. Heitmeir, A. Marn","doi":"10.1115/gt2019-90283","DOIUrl":"https://doi.org/10.1115/gt2019-90283","url":null,"abstract":"\u0000 One of the main goals for modern aircrafts is to lower the fuel consumption and noise emissions without worsening the aerodynamic performance. One possibility to lower the fuel consumption is to reduce the skin-friction losses of vanes and blades inside the engine. Therefore, this paper is about the aeroacoustical as well as the aerodynamical effects of a riblet foil applied on the suction side surface of turbine exit guide vanes (TEGVs) of a 1½ stage low pressure turbine (LPT). There have been numerous studies concerning riblets but none using them in a LPT. In general, if riblets are applied on the suction side of vanes or blades, they lower the drag and increase the lift.\u0000 Test runs were performed under two different operating points in a subsonic test turbine facility for aerodynamic, aeroacoustic, and aeroelastic investigations (STTF-AAAI) located at the Institute for Thermal Turbomachinery and Machine Dynamics at Graz University of Technology. One operating point was the design point of the riblets and the second one an off-design point. During the test campaign, two different set-ups have been investigated. One configuration with riblets applied on the suction side of the TEGVs, and one configuration with a smooth foil on the vanes to achieve the same thickness as the first set-up. This smooth configuration serves as a reference case. The tested riblet structure was of trapezoid type with 45 μm tip distance and a height to tip distance ratio of 0.45.\u0000 The acoustical data has been obtained by using flush mounted condenser microphones, rotated over 360 deg around the flow channel. The aerodynamical data was obtained by using an aerodynamic five-hole-probe as well as a trailing edge probe. Measuring in planes up- and downstream of each TEGV allowed the comparison of a rough pressure loss estimation between the two studied set-ups.\u0000 The present work gives a closer insight into the change of the acoustical and aerodynamical behaviour by applying riblets to LPT vanes.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"489 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116162821","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}
N. Jain, L. Bravo, Do Kyun Kim, M. Murugan, A. Ghoshal, F. Ham, A. Flatau
In this work, massively parallel wall-modeled Large Eddy Simulations (LES) are conducted to simulate flow through a single stage power turbine sector of a gas-turbine engine under realistic operating conditions. The numerical framework in the current work uses finite volume based compressible CharLES solver that utilizes a moving Voronoi diagram based grid generation. To test grid sensitivity and evaluate the capability of the solver in predicting turbomachinery flows, three grids of varying resolution are used to simulate flow through the baseline gas-turbine under design operating conditions. After assessing the flow solution quality and establishing simulation parameters, LES simulations are conducted to investigate the performance of gas-turbine at off-design conditions. The conditions include the rotor design point at 100% speed, and off-design points at 75%, and 50% speeds subject to high temperatures from the combustor exit flow. The results showed that the internal flow becomes highly unsteady as the rotational speed of rotor deviates from the design point leading to reduced aerodynamic performance. This study demonstrates that the current framework is able to robustly simulate the unsteady flow in a three-dimensional moving rotor environment towards the design of variable speed gas-turbine engines for US Army Future Vertical Lift program.
{"title":"Towards Large Eddy Simulation of Rotating Turbomachinery for Variable Speed Gas Turbine Engine Operation","authors":"N. Jain, L. Bravo, Do Kyun Kim, M. Murugan, A. Ghoshal, F. Ham, A. Flatau","doi":"10.1115/gt2019-91592","DOIUrl":"https://doi.org/10.1115/gt2019-91592","url":null,"abstract":"\u0000 In this work, massively parallel wall-modeled Large Eddy Simulations (LES) are conducted to simulate flow through a single stage power turbine sector of a gas-turbine engine under realistic operating conditions. The numerical framework in the current work uses finite volume based compressible CharLES solver that utilizes a moving Voronoi diagram based grid generation. To test grid sensitivity and evaluate the capability of the solver in predicting turbomachinery flows, three grids of varying resolution are used to simulate flow through the baseline gas-turbine under design operating conditions. After assessing the flow solution quality and establishing simulation parameters, LES simulations are conducted to investigate the performance of gas-turbine at off-design conditions. The conditions include the rotor design point at 100% speed, and off-design points at 75%, and 50% speeds subject to high temperatures from the combustor exit flow. The results showed that the internal flow becomes highly unsteady as the rotational speed of rotor deviates from the design point leading to reduced aerodynamic performance. This study demonstrates that the current framework is able to robustly simulate the unsteady flow in a three-dimensional moving rotor environment towards the design of variable speed gas-turbine engines for US Army Future Vertical Lift program.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"10 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"125046835","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}
M. Giovannini, F. Rubechini, G. Amato, A. Arnone, D. Simoni, V. Yepmo, F. Satta, F. Bertini
This paper deals with the design of passive control devices for reducing the impact of secondary flows on the aerodynamics of low-pressure turbine (LPT) stages. A novel kind of device is introduced which consists of shelf-like fences to be added to the blade surface. Such a device is intended to contrast the development of secondary flows, thus reducing losses and flow turning deviation with respect to the straight blade. In this second part, an experimental campaign on a linear cascade is presented which is aimed at proving the beneficial impact of the blade fences. Experiments were carried out on a low-speed test-rig, equipped with a large scale blade representative of the stators of the engine-like environment considered in part I. Measurements are mainly focused on the stator losses and on the flow field at the stator exit. The performance of the blade fences was evaluated by comparing the straight cascade and the fenced ones. The measurements highlighted the impact of the blade fences on the development of the secondary flows, affecting both the stator losses and the non-uniformity of the flow field over the exit plane, which, in the actual stage environment, impacts the operation of the downstream blade row. Moreover, the comparison between CFD and experiments proved the accuracy of the CFD setup, thus suggesting its reliability in predicting the stage performance in the engine-like configuration.
{"title":"Reducing Secondary Flow Losses in Low-Pressure Turbines With Blade Fences: Part II — Experimental Validation on Linear Cascades","authors":"M. Giovannini, F. Rubechini, G. Amato, A. Arnone, D. Simoni, V. Yepmo, F. Satta, F. Bertini","doi":"10.1115/gt2019-91284","DOIUrl":"https://doi.org/10.1115/gt2019-91284","url":null,"abstract":"\u0000 This paper deals with the design of passive control devices for reducing the impact of secondary flows on the aerodynamics of low-pressure turbine (LPT) stages. A novel kind of device is introduced which consists of shelf-like fences to be added to the blade surface. Such a device is intended to contrast the development of secondary flows, thus reducing losses and flow turning deviation with respect to the straight blade.\u0000 In this second part, an experimental campaign on a linear cascade is presented which is aimed at proving the beneficial impact of the blade fences. Experiments were carried out on a low-speed test-rig, equipped with a large scale blade representative of the stators of the engine-like environment considered in part I. Measurements are mainly focused on the stator losses and on the flow field at the stator exit. The performance of the blade fences was evaluated by comparing the straight cascade and the fenced ones. The measurements highlighted the impact of the blade fences on the development of the secondary flows, affecting both the stator losses and the non-uniformity of the flow field over the exit plane, which, in the actual stage environment, impacts the operation of the downstream blade row. Moreover, the comparison between CFD and experiments proved the accuracy of the CFD setup, thus suggesting its reliability in predicting the stage performance in the engine-like configuration.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"3 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123947488","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}
E. Perez, J. Schmitz, Nicholas A. Jaffa, A. Jemcov, J. Cameron, S. Morris
The aerodynamic characteristics of high–lift airfoil designs is of interest for improved performance and reduced blade count in Low–Pressure Turbine (LPT) design. The present paper presents both experimental measurements as well as numerical simulation results from a single-stage LPT. The airfoils were designed for an embedded stage with a total pressure expansion ratio of 1.75 and a rotor Zweifel coefficient of 1.35. The measurement program was highly unique in that detailed measurements were obtained using a variety of different probe types, including time–resolved total pressure and hot–wires. Agreement between various measurement types was generally good, but differences beyond typically stated uncertainty bounds were noted. The computations were done using RANS and a mixing model via commercially available software. The numerical results were evaluated to determine the efficacy of this type of model for prediction and design of high–lift airfoils. The computations agreed very well with the experimental results in the midspan region, but losses were over–predicted in the lower 40% span near the hub. A basic description and understanding of the flow physics in the LPT stage are presented based on the relative agreement between the experiments and computations.
{"title":"Detailed Experimental Measurement and RANS Simulation of a Low Pressure Turbine With High Lift Blading","authors":"E. Perez, J. Schmitz, Nicholas A. Jaffa, A. Jemcov, J. Cameron, S. Morris","doi":"10.1115/gt2019-91820","DOIUrl":"https://doi.org/10.1115/gt2019-91820","url":null,"abstract":"\u0000 The aerodynamic characteristics of high–lift airfoil designs is of interest for improved performance and reduced blade count in Low–Pressure Turbine (LPT) design. The present paper presents both experimental measurements as well as numerical simulation results from a single-stage LPT. The airfoils were designed for an embedded stage with a total pressure expansion ratio of 1.75 and a rotor Zweifel coefficient of 1.35. The measurement program was highly unique in that detailed measurements were obtained using a variety of different probe types, including time–resolved total pressure and hot–wires. Agreement between various measurement types was generally good, but differences beyond typically stated uncertainty bounds were noted. The computations were done using RANS and a mixing model via commercially available software. The numerical results were evaluated to determine the efficacy of this type of model for prediction and design of high–lift airfoils. The computations agreed very well with the experimental results in the midspan region, but losses were over–predicted in the lower 40% span near the hub. A basic description and understanding of the flow physics in the LPT stage are presented based on the relative agreement between the experiments and computations.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"80 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127184204","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}
Two sound power measurement approaches were developped that are easy to install and have the ability to detect the dominant modal content by applying the modern signal processing method, Compressed Sensing. In general Compressed Sensing requires only few measurement positions for an exact reconstruction of sparse acoustic mode fields. For a current study we have chosen two Compressed Sensing algorithms. Each require separate sensor array arrangements and deliver different modal contents, from which the sound power can be derived. Firstly, an Azimuthal Mode Analysis is conducted by applying the Enhanced Orthogonal Matching Pursuit (EOMP) algorithm to a sound pressure measurement vector. The measurements are obtained by using a sensor ring array with optimized positions. In a subsequent step, the sound power is calculated by referring the detected azimuthal mode spectrum to a model describing the energy distribution over the radial mode content. Secondly, using the Block Orthogonal Matching Pursuit (BOMP) algorithm, the radial mode amplitudes are determined directly. This algorithm requires the sensors to be placed at optimized azimuthal and axial positions and reconstructs a set of dominant radial modes that occur in groups. With the objective to verify both methods, the newly designed and optimized arrays in combination with the aforementioned mode reconstruction algorithms are applied to a numerical data set. This data was provided by URANS simulations of a radial compressor set-up, which is an exact replication of an actual test rig located at the RWTH Aachen University. The introduced estimation methods perform well as shown by comparison with an exact and high resolution Radial Mode Analysis Method. In the near future, the presented measurement approaches will be applied in an experimental study performed at the radial compressor test rig.
{"title":"Sound Power Measurements at Radial Compressors Using Compressed Sensing Based Signal Processing Methods","authors":"J. Hurst, Maximilian Behn, U. Tapken, L. Enghardt","doi":"10.1115/gt2019-90782","DOIUrl":"https://doi.org/10.1115/gt2019-90782","url":null,"abstract":"\u0000 Two sound power measurement approaches were developped that are easy to install and have the ability to detect the dominant modal content by applying the modern signal processing method, Compressed Sensing. In general Compressed Sensing requires only few measurement positions for an exact reconstruction of sparse acoustic mode fields. For a current study we have chosen two Compressed Sensing algorithms. Each require separate sensor array arrangements and deliver different modal contents, from which the sound power can be derived. Firstly, an Azimuthal Mode Analysis is conducted by applying the Enhanced Orthogonal Matching Pursuit (EOMP) algorithm to a sound pressure measurement vector. The measurements are obtained by using a sensor ring array with optimized positions. In a subsequent step, the sound power is calculated by referring the detected azimuthal mode spectrum to a model describing the energy distribution over the radial mode content. Secondly, using the Block Orthogonal Matching Pursuit (BOMP) algorithm, the radial mode amplitudes are determined directly. This algorithm requires the sensors to be placed at optimized azimuthal and axial positions and reconstructs a set of dominant radial modes that occur in groups. With the objective to verify both methods, the newly designed and optimized arrays in combination with the aforementioned mode reconstruction algorithms are applied to a numerical data set. This data was provided by URANS simulations of a radial compressor set-up, which is an exact replication of an actual test rig located at the RWTH Aachen University. The introduced estimation methods perform well as shown by comparison with an exact and high resolution Radial Mode Analysis Method. In the near future, the presented measurement approaches will be applied in an experimental study performed at the radial compressor test rig.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"17 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-11-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"133881286","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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