Stochastic particle tracking models coupled to RANS fluid simulations are frequently used to simulate particulate transport and hence predict component damage in gas turbines. In simple flows the Continuous Random Walk (CRW) model has been shown to model particulate motion in the diffusion-impaction regime significantly more accurately than Discrete Random Walk implementations. To date, the CRW model has used turbulent flow statistics determined from DNS in channels and experiments in pipes. Robust extension of the CRW model to accelerating flows modelled using RANS is important to enable its use in design studies of rotating engine-realistic geometries of complex curvature. This paper builds on previous work by the authors to use turbulent statistics in the CRW model directly from Reynolds Stress Models (RSM) in RANS simulations. Further improvements are made to this technique to account for strong gradients in Reynolds Stresses in all directions; improve the robustness of the model to the chosen time-step; and to eliminate the need for DNS/experimentally derived statistical flow properties. The effect of these changes were studied using a commercial CFD solver for a simple pipe flow, for which integral deposition prediction accuracy equal to that using the original CRW was achieved. These changes enable the CRW to be applied to more complex flow cases. To demonstrate why this development is important, in a more complex flow case with acceleration, deposition in a turbulent 90° bend was investigated. Critical differences in the predicted deposition are apparent when the results are compared to the alternative tracking models suitable for RANS solutions. The modified CRW model was the only model which captured the more complex deposition distribution, as predicted by published LES studies. Particle tracking models need to be accurate in the spatial distribution of deposition they predict in order to enable more sophisticated engineering design studies.
{"title":"Development and Validation of a Continuous Random Walk Model for Particle Tracking in Accelerating Flows","authors":"Nikul Vadgama, Marios Kapsis, Peter Forsyth, M. McGilvray, D. Gillespie","doi":"10.1115/GT2020-16026","DOIUrl":"https://doi.org/10.1115/GT2020-16026","url":null,"abstract":"\u0000 Stochastic particle tracking models coupled to RANS fluid simulations are frequently used to simulate particulate transport and hence predict component damage in gas turbines. In simple flows the Continuous Random Walk (CRW) model has been shown to model particulate motion in the diffusion-impaction regime significantly more accurately than Discrete Random Walk implementations. To date, the CRW model has used turbulent flow statistics determined from DNS in channels and experiments in pipes. Robust extension of the CRW model to accelerating flows modelled using RANS is important to enable its use in design studies of rotating engine-realistic geometries of complex curvature.\u0000 This paper builds on previous work by the authors to use turbulent statistics in the CRW model directly from Reynolds Stress Models (RSM) in RANS simulations. Further improvements are made to this technique to account for strong gradients in Reynolds Stresses in all directions; improve the robustness of the model to the chosen time-step; and to eliminate the need for DNS/experimentally derived statistical flow properties. The effect of these changes were studied using a commercial CFD solver for a simple pipe flow, for which integral deposition prediction accuracy equal to that using the original CRW was achieved. These changes enable the CRW to be applied to more complex flow cases.\u0000 To demonstrate why this development is important, in a more complex flow case with acceleration, deposition in a turbulent 90° bend was investigated. Critical differences in the predicted deposition are apparent when the results are compared to the alternative tracking models suitable for RANS solutions. The modified CRW model was the only model which captured the more complex deposition distribution, as predicted by published LES studies. Particle tracking models need to be accurate in the spatial distribution of deposition they predict in order to enable more sophisticated engineering design studies.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"55 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129289520","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}
Performance degradation of gas turbine is a common phenomenon during operation. Maintenance of the degraded gas turbines and improving their performance at a low cost are important in engineering. In this paper, the maintenance method based on reassembling degraded components of existing gas turbines was studied. This research was based on a type of 2MW gas turbine engine. Blue ray scanning was carried out to rebuild the 3D flow-path geometries of the compressor and turbine of a degraded engine. Then CFD simulations were carried out to compare the characteristic maps of new and degraded components. Secondly, performance tests of six engines were carried out. A correction method was developed to get the specific component characteristics using test data, which can also analyze and quantify the degradations. Also, a gas turbine performance prediction program was used to find the promising component-exchange plan within 5 given gas turbines to improve total thermal efficiency. Finally, additional test was carried out to verify the performance of the reassembled gas turbine. Through the developed method including 3D scanning, CFD simulation, and correction of component characteristics with engine test data, the component performance degradation of a specific gas turbine can be obtained in quantity. The gas turbine performance predictions based on the acquired characteristic maps showed good agreement with test data. With the help of the method developed in this work, a new gas turbine engine was obtained through exchanging the components of degraded engines, which is at a very low cost and in a short time. The improvement in total thermal efficiency was about 0.3 percentage, which was verified by engine tests.
{"title":"Improving Gas Turbine Performance Through Reassembling Degraded Components: An Experimental and Computational Study","authors":"Shuocheng Xia, Zhongran Chi, S. Zang, Hui Wang","doi":"10.1115/GT2020-14627","DOIUrl":"https://doi.org/10.1115/GT2020-14627","url":null,"abstract":"\u0000 Performance degradation of gas turbine is a common phenomenon during operation. Maintenance of the degraded gas turbines and improving their performance at a low cost are important in engineering. In this paper, the maintenance method based on reassembling degraded components of existing gas turbines was studied.\u0000 This research was based on a type of 2MW gas turbine engine. Blue ray scanning was carried out to rebuild the 3D flow-path geometries of the compressor and turbine of a degraded engine. Then CFD simulations were carried out to compare the characteristic maps of new and degraded components. Secondly, performance tests of six engines were carried out. A correction method was developed to get the specific component characteristics using test data, which can also analyze and quantify the degradations. Also, a gas turbine performance prediction program was used to find the promising component-exchange plan within 5 given gas turbines to improve total thermal efficiency. Finally, additional test was carried out to verify the performance of the reassembled gas turbine.\u0000 Through the developed method including 3D scanning, CFD simulation, and correction of component characteristics with engine test data, the component performance degradation of a specific gas turbine can be obtained in quantity. The gas turbine performance predictions based on the acquired characteristic maps showed good agreement with test data. With the help of the method developed in this work, a new gas turbine engine was obtained through exchanging the components of degraded engines, which is at a very low cost and in a short time. The improvement in total thermal efficiency was about 0.3 percentage, which was verified by engine tests.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"120 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116366639","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}
Due to the transition of the energy system to more decentralized sector-coupled technologies, the demand on small, highly efficient and compact turbines is steadily growing. Therefore, supersonic impulse turbines have been subject of academic research for many years because of their compact and low-cost conditions. However, specific loss models for this type of turbine are still missing. In this paper, a CFD-simulation-based surrogate model for the velocity coefficient, unique incidence as well as outflow deviation of the blade, is introduced. This surrogate model forms the basis for an exemplary efficiency optimization of the “Colclough cascade”. In a first step, an automatic and robust blade design methodology for constant-channel blades based on the supersonic turbine blade design of Stratford and Sansome is shown. The blade flow is fully described by seven geometrical and three aerodynamic design parameters. After that, an automated numerical flow simulation (CFD) workflow for supersonic turbine blades is developed. The validation of the CFD setup with a published supersonic axial turbine blade (Colclough design) shows a high consistency in the shock waves, separation zones and boundary layers as well as velocity coefficients. A design of experiments (DOE) with latin hypercube sampling and 1300 sample points is calculated. This CFD data forms the basis for a highly accurate surrogate model of supersonic turbine blade flow suitable for Mach numbers between 1.1 and 1.6. The throat-based Reynolds number is varied between 1*104 and 4*105. Additionally, an optimization is introduced, based on the surrogate model for the Reynolds number and Mach number of Colclough and no degree of reaction (equal inlet and outlet static pressure). The velocity coefficient is improved by up to 3 %.
{"title":"Optimization of Supersonic Axial Turbine Blades Based on Surrogate Models","authors":"Markus Waesker, B. Buelten, N. Kienzle, C. Doetsch","doi":"10.1115/GT2020-14465","DOIUrl":"https://doi.org/10.1115/GT2020-14465","url":null,"abstract":"\u0000 Due to the transition of the energy system to more decentralized sector-coupled technologies, the demand on small, highly efficient and compact turbines is steadily growing. Therefore, supersonic impulse turbines have been subject of academic research for many years because of their compact and low-cost conditions. However, specific loss models for this type of turbine are still missing. In this paper, a CFD-simulation-based surrogate model for the velocity coefficient, unique incidence as well as outflow deviation of the blade, is introduced. This surrogate model forms the basis for an exemplary efficiency optimization of the “Colclough cascade”.\u0000 In a first step, an automatic and robust blade design methodology for constant-channel blades based on the supersonic turbine blade design of Stratford and Sansome is shown. The blade flow is fully described by seven geometrical and three aerodynamic design parameters. After that, an automated numerical flow simulation (CFD) workflow for supersonic turbine blades is developed. The validation of the CFD setup with a published supersonic axial turbine blade (Colclough design) shows a high consistency in the shock waves, separation zones and boundary layers as well as velocity coefficients.\u0000 A design of experiments (DOE) with latin hypercube sampling and 1300 sample points is calculated. This CFD data forms the basis for a highly accurate surrogate model of supersonic turbine blade flow suitable for Mach numbers between 1.1 and 1.6. The throat-based Reynolds number is varied between 1*104 and 4*105.\u0000 Additionally, an optimization is introduced, based on the surrogate model for the Reynolds number and Mach number of Colclough and no degree of reaction (equal inlet and outlet static pressure). The velocity coefficient is improved by up to 3 %.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"112 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128058315","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 numerical simulation model of electrothermal de-icing process on carbon fiber reinforced polymer (CFRP) composite is conducted to study the effect of thermal properties of the substrate on the ice melting process. A novel melting model which is based on the enthalpy-porosity method is applied to study the transient ice melting process and heat transfer of the de-icing sys-tem. Multi-layered electrothermal de-icing systems including composites with different fiber orientation are used to analyze the effects of orthotropic heat conductivity of the CFRP composite on the ice melting process and heat transfer. Movement of the ice-water interface, the melted zone thickness and the melted zone area on CFRP composite are investigated on the three-dimensional electrothermal de-icing unit. The effects of thermal properties of substrate on the temperature distribution of the ice-airfoil interface are analyzed. The computational results show that the thermal properties of substrates affect the temperature on the ice-airfoil interface, the temperature distribution in the substrate, ice melting area, ice melting rate and ice melting volume significantly. The time that ice starts to melt on the CFRP composite substrate is earlier than that on the metal substrate. However, it takes more time for the ice to melt completely on the ice-CFRP interface than that on the ice-metal inter-face. The orthotropic heat conductivity of CFRP composite results in strong directivity of the melting area on the ice-CFRP in-terface. A ratio parameter is defined to represent the matching degree of substrate materials and geometry model of de-icing system. The simulation model can be applied to study electrothermal de-icing system of nacelle inlet and airfoil made of composite. The results in present work is also helpful to predict the change of temperature during de-icing process and provide guidelines for the optimizing the electrothermal de-icing system to reduce power consumption according to the fiber structure of composite.
{"title":"Numerical Simulation Model of Electrothermal De-Icing Process on Composite Substrate","authors":"Xiaofeng Guo, Zhiqiang Guo, Qiang Yang, W. Dong","doi":"10.1115/GT2020-16116","DOIUrl":"https://doi.org/10.1115/GT2020-16116","url":null,"abstract":"\u0000 A numerical simulation model of electrothermal de-icing process on carbon fiber reinforced polymer (CFRP) composite is conducted to study the effect of thermal properties of the substrate on the ice melting process. A novel melting model which is based on the enthalpy-porosity method is applied to study the transient ice melting process and heat transfer of the de-icing sys-tem. Multi-layered electrothermal de-icing systems including composites with different fiber orientation are used to analyze the effects of orthotropic heat conductivity of the CFRP composite on the ice melting process and heat transfer. Movement of the ice-water interface, the melted zone thickness and the melted zone area on CFRP composite are investigated on the three-dimensional electrothermal de-icing unit. The effects of thermal properties of substrate on the temperature distribution of the ice-airfoil interface are analyzed. The computational results show that the thermal properties of substrates affect the temperature on the ice-airfoil interface, the temperature distribution in the substrate, ice melting area, ice melting rate and ice melting volume significantly. The time that ice starts to melt on the CFRP composite substrate is earlier than that on the metal substrate. However, it takes more time for the ice to melt completely on the ice-CFRP interface than that on the ice-metal inter-face. The orthotropic heat conductivity of CFRP composite results in strong directivity of the melting area on the ice-CFRP in-terface. A ratio parameter is defined to represent the matching degree of substrate materials and geometry model of de-icing system. The simulation model can be applied to study electrothermal de-icing system of nacelle inlet and airfoil made of composite. The results in present work is also helpful to predict the change of temperature during de-icing process and provide guidelines for the optimizing the electrothermal de-icing system to reduce power consumption according to the fiber structure of composite.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"23 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121927565","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}
Typically, Gas turbine clearance between stator and rotor parts are governed by radial clearance which determines performance, cooling flow requirements, part performance etc. These clearances must be optimized to meet product requirements. Clearance which are kept too tight at assembly condition causes excessive rubbing during starting or shutdown of gas turbine which results in excessive heat generation and damage to rotating and statoric parts. In some cases, rubbing could cause tip liberations and damages to flow path which results in aero dynamic losses. Similarly, if rotor tip clearance is large at assembly condition results in aerodynamic losses. While designing component design at interface locations, tip clearance should be kept as minimum as possible to have proper balance between tip rubs at low speed and aero dynamic loss. In this paper describes the experience of Baker Hughes, wherein different design options which are considered to arrive at optimized clearance in low pressure turbine section of Gas turbine have been discussed. Typically, in low pressure turbine wherein the turbine airfoils were provided with tip rails and shrouds are brazed with honeycombs. These were allowed to rub during startup of Gas turbine engine and provide a tighter clearance at steady state operating condition. In this paper, to have tighter clearance in operating condition few methods of incorporating axial clearance in addition to radial tip clearance are discussed. By incorporating tighter axial clearance, it was found to be evident that performance estimated has improved compared to having only radial tip clearance. This paper also describes design considerations for honeycomb pertaining to axial clearance control and the methods for rubbing tip shroud with honeycomb. This paper highlights worm chart developed for a rotor blade and stator shroud configuration and highlight benefit of axial clearance control mechanism. This paper also covers different designs of heat dissipation methods in tip rails of a bucket while it is rubbing with honeycomb. In this paper, based analytical case study proves that, when axial clearance was made tighter, it reduces leakage flows and improves Low pressure turbine stage efficiency significantly.
{"title":"Methods to Have Tighter Clearance in Gas Turbine: Turbine Section","authors":"S. Babu, K. Ramesh, L. Tognarelli","doi":"10.1115/GT2020-15973","DOIUrl":"https://doi.org/10.1115/GT2020-15973","url":null,"abstract":"Typically, Gas turbine clearance between stator and rotor parts are governed by radial clearance which determines performance, cooling flow requirements, part performance etc. These clearances must be optimized to meet product requirements. Clearance which are kept too tight at assembly condition causes excessive rubbing during starting or shutdown of gas turbine which results in excessive heat generation and damage to rotating and statoric parts. In some cases, rubbing could cause tip liberations and damages to flow path which results in aero dynamic losses. Similarly, if rotor tip clearance is large at assembly condition results in aerodynamic losses. While designing component design at interface locations, tip clearance should be kept as minimum as possible to have proper balance between tip rubs at low speed and aero dynamic loss. In this paper describes the experience of Baker Hughes, wherein different design options which are considered to arrive at optimized clearance in low pressure turbine section of Gas turbine have been discussed. Typically, in low pressure turbine wherein the turbine airfoils were provided with tip rails and shrouds are brazed with honeycombs. These were allowed to rub during startup of Gas turbine engine and provide a tighter clearance at steady state operating condition. In this paper, to have tighter clearance in operating condition few methods of incorporating axial clearance in addition to radial tip clearance are discussed. By incorporating tighter axial clearance, it was found to be evident that performance estimated has improved compared to having only radial tip clearance. This paper also describes design considerations for honeycomb pertaining to axial clearance control and the methods for rubbing tip shroud with honeycomb. This paper highlights worm chart developed for a rotor blade and stator shroud configuration and highlight benefit of axial clearance control mechanism. This paper also covers different designs of heat dissipation methods in tip rails of a bucket while it is rubbing with honeycomb. In this paper, based analytical case study proves that, when axial clearance was made tighter, it reduces leakage flows and improves Low pressure turbine stage efficiency significantly.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"17 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128151383","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}
Particles ingested by aero gas turbines are capable of melting in the combustor and depositing on high pressure turbine vane surfaces, where they degrade aerodynamic and thermodynamic performance. The extent of the damage caused is a complex physical process dependent on the thermal and inertial properties of the particles, the operating state of the engine and importantly, engine architecture. The dominant architecture considerations are the position of the burner flames relative to the nozzle guide vane leading edges and the temperature difference across the burner flames. In this work, we investigate the influence of this on particle deposition by approximating the temperature variation of the hot streak as a sinusoidal profile. A parametric analysis is carried out using numerical simulations and an elastic-plastic particle deposition model, to evaluate the effect of mean temperature, temperature difference across the hot streak, and hot streak position on the deposition rate of a generic particle size distribution. Results show that the dominant effect driving particulate deposition is a combination of the gas temperature, hot streak position relative to the vane leading edge and the particulate type. The rate of deposition on a vane for sub-bituminous ash particles may be reduced by up to 56% if the combination of mean temperature, temperature difference across the hot streak, and hot streak position are chosen carefully.
{"title":"Influence of Turboshaft Engine Architecture on Ash Particle Deposition: Reduced Order Model Application","authors":"Matthew Ellis, N. Bojdo, A. Filippone","doi":"10.1115/GT2020-15203","DOIUrl":"https://doi.org/10.1115/GT2020-15203","url":null,"abstract":"\u0000 Particles ingested by aero gas turbines are capable of melting in the combustor and depositing on high pressure turbine vane surfaces, where they degrade aerodynamic and thermodynamic performance. The extent of the damage caused is a complex physical process dependent on the thermal and inertial properties of the particles, the operating state of the engine and importantly, engine architecture. The dominant architecture considerations are the position of the burner flames relative to the nozzle guide vane leading edges and the temperature difference across the burner flames. In this work, we investigate the influence of this on particle deposition by approximating the temperature variation of the hot streak as a sinusoidal profile. A parametric analysis is carried out using numerical simulations and an elastic-plastic particle deposition model, to evaluate the effect of mean temperature, temperature difference across the hot streak, and hot streak position on the deposition rate of a generic particle size distribution. Results show that the dominant effect driving particulate deposition is a combination of the gas temperature, hot streak position relative to the vane leading edge and the particulate type. The rate of deposition on a vane for sub-bituminous ash particles may be reduced by up to 56% if the combination of mean temperature, temperature difference across the hot streak, and hot streak position are chosen carefully.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"59 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126071353","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}
Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency. In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.
{"title":"Effects of Shock Wave Development on Secondary Flow Behavior in Linear Turbine Cascade at Transonic Condition","authors":"H. Tsujita, Masanao Kaneko","doi":"10.1115/GT2020-15772","DOIUrl":"https://doi.org/10.1115/GT2020-15772","url":null,"abstract":"\u0000 Gas turbines widely applied to power generation and aerospace propulsion systems are continuously enhanced in efficiency for the reduction of environmental load. The energy recovery efficiency from working fluid in a turbine component constituting gas turbines can be enhanced by the increase of turbine blade loading. However, the increase of turbine blade loading inevitably intensifies the secondary flows, and consequently increases the associated loss generation. The development of the passage vortex is strongly influenced by the pitchwise pressure gradient on the endwall in the cascade passage. In addition, a practical high pressure turbine stage is generally driven under transonic flow conditions where the shock wave strongly influences the pressure distribution on the endwall. Therefore, it becomes very important to clarify the effects of the shock wave formation on the secondary flow behavior in order to increase the turbine blade loading without the deterioration of efficiency.\u0000 In this study, the two-dimensional and the three-dimensional transonic flows in the HS1A linear turbine cascade at the design incidence angle were analyzed numerically by using the commercial CFD code with the assumption of steady compressible flow. The isentropic exit Mach number was varied from the subsonic to the supersonic conditions in order to examine the effects of development of shock wave caused by the increase of exit Mach number on the secondary flow behavior. The increase of exit Mach number induced the shock across the passage and increased its obliqueness. The increase of obliqueness reduced the cross flow on the endwall by moving the local minimum point of static pressure along the suction surface toward the trailing edge. As a consequence, the increase of exit Mach number attenuated the passage vortex.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134035792","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 describes a coupled experimental and CFD campaign conducted on a 1.5 intermediate turbine stage in the full range of operating conditions, from start-up to design point under variable expansion ratio and physical speed. The test maintains engine similitude conditions and allows direct comparison with CFD data to assess the predictions accuracy. The choice of variables to describe the speedlines is also addressed by using both measured and predicted data. A discussion on velocity ratio versus corrected speed illustrates the advantages of the former parameter the adoption of which produces constant shape curves in a very wide range of operating conditions. The comparison between measurements and predictions suggests that CFD, in conjunction with performance correlations, is a viable tool to predict speedlines in a fairly wide range of conditions, provided that geometrical and operational details are carefully matched.
{"title":"Analysis of Measured and Predicted Turbine Maps From Start-Up to Design Point","authors":"Alberto Greco, S. Biagiotti, V. Michelassi, Tomasz Jurek, D. D. Benedetto, S. Francini, M. Marconcini","doi":"10.1115/GT2020-15484","DOIUrl":"https://doi.org/10.1115/GT2020-15484","url":null,"abstract":"\u0000 This paper describes a coupled experimental and CFD campaign conducted on a 1.5 intermediate turbine stage in the full range of operating conditions, from start-up to design point under variable expansion ratio and physical speed. The test maintains engine similitude conditions and allows direct comparison with CFD data to assess the predictions accuracy. The choice of variables to describe the speedlines is also addressed by using both measured and predicted data. A discussion on velocity ratio versus corrected speed illustrates the advantages of the former parameter the adoption of which produces constant shape curves in a very wide range of operating conditions. The comparison between measurements and predictions suggests that CFD, in conjunction with performance correlations, is a viable tool to predict speedlines in a fairly wide range of conditions, provided that geometrical and operational details are carefully matched.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128596293","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}
Gas turbine nozzle cooling often uses inserts to intentionally distribute cool air through impingement holes to hot spots, especially at leading edge regions of stage-1 nozzles. Owing to the variations of engine operation environment, inlet air filtration systems, upstream component surface coating layers, and oil/air sealing material choices, solid particles could enter the engine from outside or be generated inside from rotor rubbing, seal debris, peeled off coating layers or rusty surfaces, etc. These particles and their agglomerates can be carried by airflow to enter nozzle inserts and clog impingement holes, which may reduce cooling air significantly, resulting in severe engine failures. To reduce the risk of insert clogging, particle separation devices and filtration mesh screens have been implemented in front of nozzle inlets to prevent larger particles from entering. Therefore, designing appropriate nozzle insert hole sizes becomes very critical to let smaller particles pass through and exhaust from the nozzle exit. This experimental study focuses on finding a correlation between the impingement hole sizes of a nozzle insert and their clogged areas caused by seeded fine and medium size particles of Fe2O3 under specific pressure ratios. A nozzle insert was first chosen, and a single row of cylindrical impingement holes was machined at the leading edge of the insert. Measurements were conducted in a pressurized vessel at Reynolds numbers from 7,000 to 62,000 and pressure ratios from 1.01 to 1.10. Results indicate that insert clogging is a strong function of the hole size, particle size, and pressure ratio. The tested particles showed a bimodal distribution of fine and medium sizes, and the medium size particles played a major role in clogging holes. A clogging diagram with 3 zones (fully clogged zone, partially clogged zone, and no clogging zone) is generated from the test data, which can provide important design criteria for sizing the insert leading edge impingement holes to reduce the risk of clogging.
{"title":"Turbine Nozzle Insert Clogging With Seeded Medium Size Particles","authors":"Hongzhou Xu, Kevin Liu, M. Fox","doi":"10.1115/GT2020-14522","DOIUrl":"https://doi.org/10.1115/GT2020-14522","url":null,"abstract":"\u0000 Gas turbine nozzle cooling often uses inserts to intentionally distribute cool air through impingement holes to hot spots, especially at leading edge regions of stage-1 nozzles. Owing to the variations of engine operation environment, inlet air filtration systems, upstream component surface coating layers, and oil/air sealing material choices, solid particles could enter the engine from outside or be generated inside from rotor rubbing, seal debris, peeled off coating layers or rusty surfaces, etc. These particles and their agglomerates can be carried by airflow to enter nozzle inserts and clog impingement holes, which may reduce cooling air significantly, resulting in severe engine failures. To reduce the risk of insert clogging, particle separation devices and filtration mesh screens have been implemented in front of nozzle inlets to prevent larger particles from entering. Therefore, designing appropriate nozzle insert hole sizes becomes very critical to let smaller particles pass through and exhaust from the nozzle exit. This experimental study focuses on finding a correlation between the impingement hole sizes of a nozzle insert and their clogged areas caused by seeded fine and medium size particles of Fe2O3 under specific pressure ratios. A nozzle insert was first chosen, and a single row of cylindrical impingement holes was machined at the leading edge of the insert. Measurements were conducted in a pressurized vessel at Reynolds numbers from 7,000 to 62,000 and pressure ratios from 1.01 to 1.10. Results indicate that insert clogging is a strong function of the hole size, particle size, and pressure ratio. The tested particles showed a bimodal distribution of fine and medium sizes, and the medium size particles played a major role in clogging holes. A clogging diagram with 3 zones (fully clogged zone, partially clogged zone, and no clogging zone) is generated from the test data, which can provide important design criteria for sizing the insert leading edge impingement holes to reduce the risk of clogging.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"26 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134347482","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 numerical analysis of the effect of Mach number on the boundary layer development and aerodynamic performance of a high-lift, after loaded low pressure turbine blade is presented in this paper. The turbine blade is designed for the GTF engine and works in a low Reynolds number, high Mach number environment. Three different isentropic exit Mach numbers (0.14, 0.87 and 1.17) are simulated by large eddy simulation method, while the Reynolds number based on the axial chord length of the blade and the exit flow velocity is kept the same (1 × 105). The condition Mais,2 = 0.14 represents the lowspeeed wind tunnel environment which is usually used in the low pressure turbine investigation. The condition Mais,2 = 0.87 represents the design point of the turbine blade. The condition Mais,2 = 1.17 represents the severe environment when the shock wave shows up. A comparison of the boundary layer development is made and the total pressure loss results from the boundary layer is discussed.
{"title":"The Influence of Mach Number on the Boundary Layer Development of High-Lift Low Pressure Turbine in Low Reynolds Number","authors":"Wenhua Duan, Jian Liu, W. Qiao","doi":"10.1115/GT2020-16278","DOIUrl":"https://doi.org/10.1115/GT2020-16278","url":null,"abstract":"\u0000 A numerical analysis of the effect of Mach number on the boundary layer development and aerodynamic performance of a high-lift, after loaded low pressure turbine blade is presented in this paper. The turbine blade is designed for the GTF engine and works in a low Reynolds number, high Mach number environment. Three different isentropic exit Mach numbers (0.14, 0.87 and 1.17) are simulated by large eddy simulation method, while the Reynolds number based on the axial chord length of the blade and the exit flow velocity is kept the same (1 × 105). The condition Mais,2 = 0.14 represents the lowspeeed wind tunnel environment which is usually used in the low pressure turbine investigation. The condition Mais,2 = 0.87 represents the design point of the turbine blade. The condition Mais,2 = 1.17 represents the severe environment when the shock wave shows up. A comparison of the boundary layer development is made and the total pressure loss results from the boundary layer is discussed.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"71 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124443530","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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