Till M. Biedermann, Youssef Moutamassik, Frank Kameier
Abstract Heavy-duty centrifugal fans require high reliability and first-class performance. Besides, extreme conditions and harsh environments are often encountered, such as in the papermaking process, in steel or cement plants or the chemical and petrochemical industry. Therefore, the design of high-performance heavy-duty industrial fans requires robust yet efficient solutions. The previous work indicates a high aerodynamic and aeroacoustic sensitivity concerning the specific position of the volute cutoff (tongue). This effect will be further investigated, not by directly changing the orientation of the cutoff, but by varying the position of the impeller relative to a fixed volute casing. The initial evaluation is done through a numerical study of three influencing parameters, which allow the aerodynamic dependencies to be modeled using low-layer artificial networks. Subsequently, extensive experimental studies were carried out to validate the aerodynamic dependencies and also to incorporate information on the aeroacoustic performance. The obtained results show that the operating point represents the key factor in determining the optimal positioning, with qualitatively comparable dependencies found for both tested fans. From an aeroacoustic point of view, the determined optimal configuration does not necessarily coincide with the observed aerodynamic desires, so careful analysis and a reasonable compromise are required, motivating for a multi-objective optimization process.
{"title":"Assessment of the Impeller/Volute Relationship of Centrifugal Fans From an Aerodynamic and Aeroacoustic Perspective","authors":"Till M. Biedermann, Youssef Moutamassik, Frank Kameier","doi":"10.1115/1.4062243","DOIUrl":"https://doi.org/10.1115/1.4062243","url":null,"abstract":"Abstract Heavy-duty centrifugal fans require high reliability and first-class performance. Besides, extreme conditions and harsh environments are often encountered, such as in the papermaking process, in steel or cement plants or the chemical and petrochemical industry. Therefore, the design of high-performance heavy-duty industrial fans requires robust yet efficient solutions. The previous work indicates a high aerodynamic and aeroacoustic sensitivity concerning the specific position of the volute cutoff (tongue). This effect will be further investigated, not by directly changing the orientation of the cutoff, but by varying the position of the impeller relative to a fixed volute casing. The initial evaluation is done through a numerical study of three influencing parameters, which allow the aerodynamic dependencies to be modeled using low-layer artificial networks. Subsequently, extensive experimental studies were carried out to validate the aerodynamic dependencies and also to incorporate information on the aeroacoustic performance. The obtained results show that the operating point represents the key factor in determining the optimal positioning, with qualitatively comparable dependencies found for both tested fans. From an aeroacoustic point of view, the determined optimal configuration does not necessarily coincide with the observed aerodynamic desires, so careful analysis and a reasonable compromise are required, motivating for a multi-objective optimization process.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-05-22","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135335627","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract The relative casing motion can significantly influence the turbine blade tip aerothermal performance. In this study, experimental investigation was conducted in a newly developed high-speed disk rotor rig which can mimic engine realistic high-speed casing relative motion while enabling full optical access to a transonic turbine blade tip surface. Spatially-resolved tip heat transfer data, including heat transfer coefficient and film cooling effectiveness, were obtained for a cooled transonic squealer tip by infrared transient thermal measurement. Combined with closely coupled Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) analysis, this paper reveals an interesting interaction mechanism between the cooling injections from the pressure side and the cavity floor with and without the effect of relative casing motion. Both experimental data and CFD results show a consistent trend in both heat transfer and cooling performance. With cavity cooling only, the cooling performance reduces with the effect of relative casing motion. However, with additional cooling injection from the pressure side, a significant improvement in the combined cooling performance with the relative casing motion can be observed. Such opposite trend highlights the importance of relative casing motion when ranking different tip cooling designs. With the consideration of relative casing motion, extra tip cooling benefit can be obtained by combining cooling injections from two different locations.
{"title":"Interaction Mechanism of Transonic Squealer Tip Cooling With the Effect of High-Speed Relative Casing Motion","authors":"Wenbo Xie, Shaopeng Lu, Hongmei Jiang, Xu Peng, Qiang Zhang","doi":"10.1115/1.4062278","DOIUrl":"https://doi.org/10.1115/1.4062278","url":null,"abstract":"Abstract The relative casing motion can significantly influence the turbine blade tip aerothermal performance. In this study, experimental investigation was conducted in a newly developed high-speed disk rotor rig which can mimic engine realistic high-speed casing relative motion while enabling full optical access to a transonic turbine blade tip surface. Spatially-resolved tip heat transfer data, including heat transfer coefficient and film cooling effectiveness, were obtained for a cooled transonic squealer tip by infrared transient thermal measurement. Combined with closely coupled Reynolds-averaged Navier–Stokes computational fluid dynamics (CFD) analysis, this paper reveals an interesting interaction mechanism between the cooling injections from the pressure side and the cavity floor with and without the effect of relative casing motion. Both experimental data and CFD results show a consistent trend in both heat transfer and cooling performance. With cavity cooling only, the cooling performance reduces with the effect of relative casing motion. However, with additional cooling injection from the pressure side, a significant improvement in the combined cooling performance with the relative casing motion can be observed. Such opposite trend highlights the importance of relative casing motion when ranking different tip cooling designs. With the consideration of relative casing motion, extra tip cooling benefit can be obtained by combining cooling injections from two different locations.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-05-22","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135335628","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract Rotating detonation engines (RDEs) are characterized by a thermodynamic cycle with an efficiency gain up to 15% at medium pressure ratios with respect to systems based on the conventional Joule–Bryton cycle. Multiple turbine designs can be considered, and this article deals with the supersonic inlet configuration. After having reviewed the main design steps of an exemplary RDE supersonic turbine, the article focuses on the considerable effects that endwall losses have on the performance of supersonic-inlet turbines and on the reasons why endwall contouring is strongly recommended for an efficient design. Parametric analyses, carried out by a novel in-house mean-line code validated against computational fluid dynamics (CFD), reveal that endwall friction losses contribute significantly to the overall stage loss. Endwall boundary layers also reduce the effective area, which can be critical for the self-starting capability of the supersonic channel. Therefore, a variable blade height geometry is necessary to extend the design space and guarantee a higher efficiency with respect to a constant-span configuration. The in-house CFD-based evolutionary shape optimization code was adapted to search for the optimal endwall shape for these unconventional machines. The optimal shape reduces shock losses and deviation angles and provides a significant gain in efficiency and work extraction. Finally, a novel technique is proposed to design the three-dimensional shape of the rotor based on the method of characteristics and tailored on the flow delivered by the stator.
{"title":"The Role of Endwall Shape Optimization in the Design of Supersonic Turbines for Rotating Detonation Engines","authors":"Noraiz Mushtaq, Giacomo Persico, Paolo Gaetani","doi":"10.1115/1.4062277","DOIUrl":"https://doi.org/10.1115/1.4062277","url":null,"abstract":"Abstract Rotating detonation engines (RDEs) are characterized by a thermodynamic cycle with an efficiency gain up to 15% at medium pressure ratios with respect to systems based on the conventional Joule–Bryton cycle. Multiple turbine designs can be considered, and this article deals with the supersonic inlet configuration. After having reviewed the main design steps of an exemplary RDE supersonic turbine, the article focuses on the considerable effects that endwall losses have on the performance of supersonic-inlet turbines and on the reasons why endwall contouring is strongly recommended for an efficient design. Parametric analyses, carried out by a novel in-house mean-line code validated against computational fluid dynamics (CFD), reveal that endwall friction losses contribute significantly to the overall stage loss. Endwall boundary layers also reduce the effective area, which can be critical for the self-starting capability of the supersonic channel. Therefore, a variable blade height geometry is necessary to extend the design space and guarantee a higher efficiency with respect to a constant-span configuration. The in-house CFD-based evolutionary shape optimization code was adapted to search for the optimal endwall shape for these unconventional machines. The optimal shape reduces shock losses and deviation angles and provides a significant gain in efficiency and work extraction. Finally, a novel technique is proposed to design the three-dimensional shape of the rotor based on the method of characteristics and tailored on the flow delivered by the stator.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-05-22","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135287983","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The flow through the predominantly two-dimensional geometries of cascades of blades is intrinsically three-dimensional and unsteady. Direct Numerical Simulation, Large Eddy Simulations, and time-resolved Particle Image Velocimetry resolve the full flow physics, relevant to aerodynamic loss and heat management. Such studies build upon earlier insight drawn from quasi two-dimensional investigations that identified the key areas where progress in understanding was most needed. These areas stretch across the full passage, from the leading edge of the blade to the passage outflow. Streamwise surface vorticity, transition, the calmed region, shock-boundary layer interaction, and vortex shedding are considered in detail, specifically (i) on what gaps in their physical understanding the works of Jonathan Paul Gostelow exposed and (ii) what gaps were present in the two-dimensional computational approaches used to represent these flows in these works. These useful insights are obtained from the geometrically simpler settings of circular cylinders in cross-flow and from flat plate experiments, as well as from cascades of blades. This paper presents an overview of the physical understanding of the flow features that underpins the more recent time-resolved three-dimensional investigations, led by the late Emeritus Professor Jonathan Paul Gostelow. This work celebrates some of Paul Gostelow's 50+ years of turbomachinery research achievements and develops awareness about their significance towards reaching a more complete knowledge of the flow physics in turbomachinery, using the more recent time-resolved three-dimensional modelling capability of Computational Fluid Dynamics software.
{"title":"An insight into in quasi two-dimensional flow features over turbine blading from the works of Jonathan Paul Gostelow","authors":"A. Rona, D. Adebayo, Jonathan P Gostelow","doi":"10.1115/1.4062552","DOIUrl":"https://doi.org/10.1115/1.4062552","url":null,"abstract":"\u0000 The flow through the predominantly two-dimensional geometries of cascades of blades is intrinsically three-dimensional and unsteady. Direct Numerical Simulation, Large Eddy Simulations, and time-resolved Particle Image Velocimetry resolve the full flow physics, relevant to aerodynamic loss and heat management. Such studies build upon earlier insight drawn from quasi two-dimensional investigations that identified the key areas where progress in understanding was most needed. These areas stretch across the full passage, from the leading edge of the blade to the passage outflow. Streamwise surface vorticity, transition, the calmed region, shock-boundary layer interaction, and vortex shedding are considered in detail, specifically (i) on what gaps in their physical understanding the works of Jonathan Paul Gostelow exposed and (ii) what gaps were present in the two-dimensional computational approaches used to represent these flows in these works. These useful insights are obtained from the geometrically simpler settings of circular cylinders in cross-flow and from flat plate experiments, as well as from cascades of blades. This paper presents an overview of the physical understanding of the flow features that underpins the more recent time-resolved three-dimensional investigations, led by the late Emeritus Professor Jonathan Paul Gostelow. This work celebrates some of Paul Gostelow's 50+ years of turbomachinery research achievements and develops awareness about their significance towards reaching a more complete knowledge of the flow physics in turbomachinery, using the more recent time-resolved three-dimensional modelling capability of Computational Fluid Dynamics software.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-05-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"48361137","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Jingtian Duan, Kecheng Zhang, Jin Xu, J. Lei, Junmei Wu
Flow statistic in the mid-plane of a rectangular channel with diamond-s pin fins was obtained by means of particle imaging velocimetry (PIV) at Re=10000. Large scale and small scale fluctuation was separated using proper orthogonal decomposition (POD). The flow characteristics were compared to Nusselt number distribution on the endwall acquired by thermochromic liquid crystal (TLC) to reveal flow mechanism driving heat transfer enhancement. Results indicate that local vorticity plays an important role on strengthening Nu on both sides of leading point (Zone 1). Downstream of the two sharp edges on both sides (Zone 2), small size disturbances from shear layer eddies drive local heat transfer. The flow characteristics and heat transfer distribution downstream of the 1st row (Zone 3) presents alternated feature along Y direction due to the interaction between shear layers of neighboring pin fins. Lateral velocity fluctuation induced by large vortex shedding drives the heat transfer augmentation in Zone 3 where there is violent large vortex shedding. While small size disturbances of the shear layer drive local heat transfer enhancement in Zone 3 downstream of pin fins where large vortex shedding is suppressed. For the 2nd and 3rd row, there is no difference in the flow characteristics downstream of neighboring pin fins. Small size fluctuations distributed uniformly downstream of large vortex shedding (Zone 4) resulting in a uniformly distributed Nu.
{"title":"Investigation on Flow Characteristics and Its Effect on Heat Transfer Enhancement in a Wide Channel with Staggered Diamond-S Pin Fins","authors":"Jingtian Duan, Kecheng Zhang, Jin Xu, J. Lei, Junmei Wu","doi":"10.1115/1.4062502","DOIUrl":"https://doi.org/10.1115/1.4062502","url":null,"abstract":"\u0000 Flow statistic in the mid-plane of a rectangular channel with diamond-s pin fins was obtained by means of particle imaging velocimetry (PIV) at Re=10000. Large scale and small scale fluctuation was separated using proper orthogonal decomposition (POD). The flow characteristics were compared to Nusselt number distribution on the endwall acquired by thermochromic liquid crystal (TLC) to reveal flow mechanism driving heat transfer enhancement. Results indicate that local vorticity plays an important role on strengthening Nu on both sides of leading point (Zone 1). Downstream of the two sharp edges on both sides (Zone 2), small size disturbances from shear layer eddies drive local heat transfer. The flow characteristics and heat transfer distribution downstream of the 1st row (Zone 3) presents alternated feature along Y direction due to the interaction between shear layers of neighboring pin fins. Lateral velocity fluctuation induced by large vortex shedding drives the heat transfer augmentation in Zone 3 where there is violent large vortex shedding. While small size disturbances of the shear layer drive local heat transfer enhancement in Zone 3 downstream of pin fins where large vortex shedding is suppressed. For the 2nd and 3rd row, there is no difference in the flow characteristics downstream of neighboring pin fins. Small size fluctuations distributed uniformly downstream of large vortex shedding (Zone 4) resulting in a uniformly distributed Nu.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-05-05","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"49314331","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Izhar Ullah, L. Wright, Chao-Cheng Shiau, Je-Chin Han, Zhihong Gao, Andrea Stanton
This work is an experimental study of film cooling effectiveness on two real scale turbine vanes in a 3-vane annular cascade. The cascade is connected to a high flow steady compressor to provide the mainstream flow. The inlet velocity is maintained at 35 m/s at the center line of the annular cascade. Two heavily cooled, real scale turbine vanes are tested with cooling holes on the pressure and suction surfaces. Vane 1 has 645 cooling holes distributed around the vane. Vane 2 has an additional two row of holes at the near leading edge SS of the vane. The MFR is varied from 3.12% to 4.82%. Increasing the MFR increases the film effectiveness. The introduction of additional rows of cooling holes resulted in re-distribution of coolant from the pressure surface to the suction surface. This study provides designers with more insight into how to place rows of cooling holes to have improved effectiveness.
{"title":"FILM COOLING COMPARISON OF FULL-SCALE TURBINE VANES USING THE PRESSURE SENSITIVE PAINT TECHNIQUE","authors":"Izhar Ullah, L. Wright, Chao-Cheng Shiau, Je-Chin Han, Zhihong Gao, Andrea Stanton","doi":"10.1115/1.4062436","DOIUrl":"https://doi.org/10.1115/1.4062436","url":null,"abstract":"\u0000 This work is an experimental study of film cooling effectiveness on two real scale turbine vanes in a 3-vane annular cascade. The cascade is connected to a high flow steady compressor to provide the mainstream flow. The inlet velocity is maintained at 35 m/s at the center line of the annular cascade. Two heavily cooled, real scale turbine vanes are tested with cooling holes on the pressure and suction surfaces. Vane 1 has 645 cooling holes distributed around the vane. Vane 2 has an additional two row of holes at the near leading edge SS of the vane. The MFR is varied from 3.12% to 4.82%. Increasing the MFR increases the film effectiveness. The introduction of additional rows of cooling holes resulted in re-distribution of coolant from the pressure surface to the suction surface. This study provides designers with more insight into how to place rows of cooling holes to have improved effectiveness.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-04-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"48365755","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Emanuel Chirayath, Haosen Xu, Xiang I. A. Yang, R. Kunz
The damage due to particulate matter ingestion by propulsion gas turbine engines can be significant, impacting the operability and performance of plant components. Here, we focus on the axial compressor whose blades become damaged when operated in dusty/sandy environments, resulting in significant performance degradation. In this work, CFD studies are performed to model the effects of airfoil damage on the first stage rotor blading of a GE T700-401C compressor. We use thermoplastic additive manufacturing to construct representative physical models of three damage morphologies – ballistically bent/curved leading edges, cragged erosion of leading edges, and eroded leading/tailing edges at outer span locations. The resultant damaged plastic geometries, and a baseline undamaged configuration are then optically scanned and incorporated into sublayer resolved full stage, unsteady RANS analyses. Boundary conditions are imposed that conform to damaged compressor operation protocols, and this iterative process for accommodating corrected mass flow and off-design powering is presented. The results for the three damaged and one undamaged configuration are studied in terms of compressible wave field and secondary/tip flows, spanwise performance parameter distributions and efficiency. A method to estimate the effect of rotor damage on engine SFC is presented. The code, modeling and meshing strategies pursued here are consistent with a validation study carried out for NASA Rotor 37 – these results are briefly included, and provide confidence in the predictions of the T700 geometry studied. The results provide quantitative comparisons of, and insight into, the physical mechanisms associated with damaged compressor performance degradation.
{"title":"Full Stage Axial Compressor Performance Modeling Incorporating the Effects of Blade Damage due to Particle Ingestion","authors":"Emanuel Chirayath, Haosen Xu, Xiang I. A. Yang, R. Kunz","doi":"10.1115/1.4062397","DOIUrl":"https://doi.org/10.1115/1.4062397","url":null,"abstract":"\u0000 The damage due to particulate matter ingestion by propulsion gas turbine engines can be significant, impacting the operability and performance of plant components. Here, we focus on the axial compressor whose blades become damaged when operated in dusty/sandy environments, resulting in significant performance degradation. In this work, CFD studies are performed to model the effects of airfoil damage on the first stage rotor blading of a GE T700-401C compressor. We use thermoplastic additive manufacturing to construct representative physical models of three damage morphologies – ballistically bent/curved leading edges, cragged erosion of leading edges, and eroded leading/tailing edges at outer span locations. The resultant damaged plastic geometries, and a baseline undamaged configuration are then optically scanned and incorporated into sublayer resolved full stage, unsteady RANS analyses. Boundary conditions are imposed that conform to damaged compressor operation protocols, and this iterative process for accommodating corrected mass flow and off-design powering is presented. The results for the three damaged and one undamaged configuration are studied in terms of compressible wave field and secondary/tip flows, spanwise performance parameter distributions and efficiency. A method to estimate the effect of rotor damage on engine SFC is presented. The code, modeling and meshing strategies pursued here are consistent with a validation study carried out for NASA Rotor 37 – these results are briefly included, and provide confidence in the predictions of the T700 geometry studied. The results provide quantitative comparisons of, and insight into, the physical mechanisms associated with damaged compressor performance degradation.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-04-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47088969","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract Non-axisymmetric endwall profiling offers features to simultaneously mitigate aerodynamic losses and hot gas ingestion in axial turbines. This paper presents an optimization methodology to generate a contoured surface integrated with real geometrical effects such as blade fillets and a rim seal channel with the aim of achieving higher efficiencies while reducing hot gas ingestion. The contoured rotor platform is constructed using a B-spline surface clamped in the axial direction. In the azimuthal direction, the surface is unclamped to allow geometrical continuity across the periodic boundaries. The endwall parameterization is used to optimize a rotor hub platform of a high-pressure turbine stage. A differential evolution optimizer is used to rank individuals in terms of efficiency. The single-objective optimization is set to maximize the aerodynamic efficiency and it is defined such that it accounts for the flow non-uniformity through a mixed-out averaging procedure. Engine representative conditions typical of a two-stage high-pressure turbine are used as boundary conditions. Geometrical and aerodynamic constraints are set to guarantee a fair comparison among individuals and to meet engine requirements. Two surface parameterizations, which use a different number of design variables but share the same construction strategy, are presented to show the trade-off between the number of degrees-of-freedom and the aerodynamic improvement. Different purge flow conditions are considered to assess the robustness of the optimization results at off-design conditions for relevant geometries. The aim of this paper is to show the advanced shape flexibility of the implemented parameterization for contoured platforms featuring technological effects such as blade fillet and rim seal channel. The work provides design guidelines to setup engine-realistic constraints for endwall contour optimization of turbine stages.
{"title":"TURBINE ENDWALL CONTOURING THROUGH ADVANCED OPTIMIZATION TECHNIQUES","authors":"Matteo Burigana, Tom Verstraete, Sergio Lavagnoli","doi":"10.1115/1.4062211","DOIUrl":"https://doi.org/10.1115/1.4062211","url":null,"abstract":"Abstract Non-axisymmetric endwall profiling offers features to simultaneously mitigate aerodynamic losses and hot gas ingestion in axial turbines. This paper presents an optimization methodology to generate a contoured surface integrated with real geometrical effects such as blade fillets and a rim seal channel with the aim of achieving higher efficiencies while reducing hot gas ingestion. The contoured rotor platform is constructed using a B-spline surface clamped in the axial direction. In the azimuthal direction, the surface is unclamped to allow geometrical continuity across the periodic boundaries. The endwall parameterization is used to optimize a rotor hub platform of a high-pressure turbine stage. A differential evolution optimizer is used to rank individuals in terms of efficiency. The single-objective optimization is set to maximize the aerodynamic efficiency and it is defined such that it accounts for the flow non-uniformity through a mixed-out averaging procedure. Engine representative conditions typical of a two-stage high-pressure turbine are used as boundary conditions. Geometrical and aerodynamic constraints are set to guarantee a fair comparison among individuals and to meet engine requirements. Two surface parameterizations, which use a different number of design variables but share the same construction strategy, are presented to show the trade-off between the number of degrees-of-freedom and the aerodynamic improvement. Different purge flow conditions are considered to assess the robustness of the optimization results at off-design conditions for relevant geometries. The aim of this paper is to show the advanced shape flexibility of the implemented parameterization for contoured platforms featuring technological effects such as blade fillet and rim seal channel. The work provides design guidelines to setup engine-realistic constraints for endwall contour optimization of turbine stages.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-12","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135289103","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract Aspect ratio and solidity play complementary roles in the aerodynamic design of axial fan rotor blades. Few studies have experimented the effect of the aspect ratio of rotor blades on the performance of low-speed axial fans or its interaction with blade solidity in terms of fan aerodynamic performance. This study examined the selection of the solidity and blade aspect ratio in the preliminary design of low-pressure industrial fans with minimized hub-to-tip ratios. The aim of this study was to make available to the fan community experimental data that allow the determination of the optimal aspect ratio for practical applications as a function of the blade solidity. Various aspects of the performance of 16 prototypes of a 315-mm-diameter propeller fan were compared. The industrial fan prototypes all had a hub-to-tip ratio of 0.2 and were derived from four baseline designs conceived to ideally achieve the same best efficiency operation with different values of the aspect ratio. In addition, the prototypes’ assemblies were conceived to allow operation with different blade counts, i.e., with different rotor solidities at a fixed blade aspect ratio. The aerodynamic performance of the fans, measured in accordance with the ISO-5801 standard, was evaluated to assess the sensitivity and trends of the fan pressure and efficiency with respect to the blade aspect ratio and solidity at fixed tip clearance. The measured effects of the aspect ratio and solidity are discussed on the basis of data available in the literature. The results of the experimental analysis were used to formulate general guidelines for the preliminary design of propeller fans with minimized hub-to-tip ratios.
{"title":"Effect of Solidity and Aspect Ratio on the Aerodynamic Performance of Axial-Flow Fans With 0.2 Hub-to-Tip Ratio","authors":"Massimo Masi, Piero Danieli, Andrea Lazzaretto","doi":"10.1115/1.4057063","DOIUrl":"https://doi.org/10.1115/1.4057063","url":null,"abstract":"Abstract Aspect ratio and solidity play complementary roles in the aerodynamic design of axial fan rotor blades. Few studies have experimented the effect of the aspect ratio of rotor blades on the performance of low-speed axial fans or its interaction with blade solidity in terms of fan aerodynamic performance. This study examined the selection of the solidity and blade aspect ratio in the preliminary design of low-pressure industrial fans with minimized hub-to-tip ratios. The aim of this study was to make available to the fan community experimental data that allow the determination of the optimal aspect ratio for practical applications as a function of the blade solidity. Various aspects of the performance of 16 prototypes of a 315-mm-diameter propeller fan were compared. The industrial fan prototypes all had a hub-to-tip ratio of 0.2 and were derived from four baseline designs conceived to ideally achieve the same best efficiency operation with different values of the aspect ratio. In addition, the prototypes’ assemblies were conceived to allow operation with different blade counts, i.e., with different rotor solidities at a fixed blade aspect ratio. The aerodynamic performance of the fans, measured in accordance with the ISO-5801 standard, was evaluated to assess the sensitivity and trends of the fan pressure and efficiency with respect to the blade aspect ratio and solidity at fixed tip clearance. The measured effects of the aspect ratio and solidity are discussed on the basis of data available in the literature. The results of the experimental analysis were used to formulate general guidelines for the preliminary design of propeller fans with minimized hub-to-tip ratios.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-10","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135543258","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper describes low-order-model-based analysis of the design of an experiment to be used for parametric studies of adiabatic film and overall cooling effectiveness for fully cooled systems (internal and film) under wide ranges of mainstream-to-coolant temperature ratio variation, in the range 0.50 < T0m/T0c < 2.30. The purpose is to improve understanding of—and validation of—the scaling process from typical rig conditions to engine conditions. We are primarily interested in the variation in overall effectiveness when the controlling non-dimensional groups change in a natural co-dependent way with changes in temperature ratio: that is, the practical situation of interest to engine designers. We distinguish this from the situation in which individual non-dimensional groups are varied in isolation: a situation that we believe is essentially impossible to meaningfully approximate in practice, despite a body of literature purporting to do the same. Design and commissioning data from a new high temperature (600 K) test facility is presented, with detailed uncertainty analysis. We show that a typical nozzle guide vane which at engine conditions (TR = 2.00) would have overall cooling effectiveness of 0.450, would be expected to have overall effectiveness of 0.418 at typical rig conditions (TR = 1.20). That is, typical scaling from engine-to-rig result is −7.1%, and typical scaling from rig-to-engine is +7.7%, This result is important for first order estimation of overall cooling performance at engine conditions.
{"title":"IMPACT OF TEMPERATURE RATIO ON OVERALL COOLING PERFORMANCE: LOW-ORDER-MODEL-BASED ANALYSIS OF EXPERIMENT DESIGN","authors":"A. D. Naidu, T. Povey","doi":"10.1115/1.4062279","DOIUrl":"https://doi.org/10.1115/1.4062279","url":null,"abstract":"\u0000 This paper describes low-order-model-based analysis of the design of an experiment to be used for parametric studies of adiabatic film and overall cooling effectiveness for fully cooled systems (internal and film) under wide ranges of mainstream-to-coolant temperature ratio variation, in the range 0.50 < T0m/T0c < 2.30. The purpose is to improve understanding of—and validation of—the scaling process from typical rig conditions to engine conditions. We are primarily interested in the variation in overall effectiveness when the controlling non-dimensional groups change in a natural co-dependent way with changes in temperature ratio: that is, the practical situation of interest to engine designers. We distinguish this from the situation in which individual non-dimensional groups are varied in isolation: a situation that we believe is essentially impossible to meaningfully approximate in practice, despite a body of literature purporting to do the same. Design and commissioning data from a new high temperature (600 K) test facility is presented, with detailed uncertainty analysis. We show that a typical nozzle guide vane which at engine conditions (TR = 2.00) would have overall cooling effectiveness of 0.450, would be expected to have overall effectiveness of 0.418 at typical rig conditions (TR = 1.20). That is, typical scaling from engine-to-rig result is −7.1%, and typical scaling from rig-to-engine is +7.7%, This result is important for first order estimation of overall cooling performance at engine conditions.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-04-06","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43651482","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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