Bjoern F. Klose, Christian Morsbach, Michael Bergmann, Alexander Hergt, Joachim Klinner, Sebastian Grund, Edmund Kuegeler
Abstract In the second paper of this three-part series, we focus on the simulation of transonic test cases for turbomachinery applications using a high-order discontinuous Galerkin spectral element method (DGSEM). High-fidelity simulations of transonic compressors and turbines are particularly challenging, as they typically occur at high Reynolds numbers and require additional treatment to reliably capture the shock waves characterizing such flows. A recently developed finite-volume subcell shock capturing scheme tailored for the DGSEM is applied and evaluated with regard to the shock sensor. To this end, we conduct implicit large eddy simulations of a high-pressure turbine cascade from the public literature and a transonic compressor cascade measured at the German Aerospace Center, both at a high Reynolds number above 106. Based on the results, we examine modal-energy and flow-feature based shock indicator functions, compare the simulation data to experimental and numerical studies, and present an analysis of the unsteady features of the flows.
{"title":"A Numerical Test Rig for Turbomachinery Flows Based on Large Eddy Simulations With a High-Order Discontinuous Galerkin Scheme - Part 2: Shock-Capturing and Transonic Flows","authors":"Bjoern F. Klose, Christian Morsbach, Michael Bergmann, Alexander Hergt, Joachim Klinner, Sebastian Grund, Edmund Kuegeler","doi":"10.1115/1.4063827","DOIUrl":"https://doi.org/10.1115/1.4063827","url":null,"abstract":"Abstract In the second paper of this three-part series, we focus on the simulation of transonic test cases for turbomachinery applications using a high-order discontinuous Galerkin spectral element method (DGSEM). High-fidelity simulations of transonic compressors and turbines are particularly challenging, as they typically occur at high Reynolds numbers and require additional treatment to reliably capture the shock waves characterizing such flows. A recently developed finite-volume subcell shock capturing scheme tailored for the DGSEM is applied and evaluated with regard to the shock sensor. To this end, we conduct implicit large eddy simulations of a high-pressure turbine cascade from the public literature and a transonic compressor cascade measured at the German Aerospace Center, both at a high Reynolds number above 106. Based on the results, we examine modal-energy and flow-feature based shock indicator functions, compare the simulation data to experimental and numerical studies, and present an analysis of the unsteady features of the flows.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"4 6","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134993764","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}
Christian Morsbach, Michael Bergmann, Adem Tosun, Bjoern F. Klose, Patrick Bechlars, Edmund Kuegeler
Abstract In this final paper of a three-part series, we apply the numerical test rig based on a high-order discontinuous Galerkin scheme to the MTU T161 low-pressure turbine with diverging end walls at off-design Reynolds number of 90,000, Mach number of 0.6, and inflow angle of 41 deg. The inflow end wall boundary layers are prescribed in accordance with the experiment. Validation of the setup is shown against recent numerical references and the corresponding experimental data. Additionally, we propose and conduct a purely numerical experiment with upstream bar wake generators at a Strouhal number of 1.25, which is well above what was possible in the experiment. We discuss the flow physics at midspan and in the end wall region and highlight the influence of the wakes from the upstream row on the complex secondary flow system using instantaneous flow visualization, phase averages, and modal decomposition techniques.
{"title":"A Numerical Test Rig for Turbomachinery Flows Based on Large Eddy Simulations With a High-Order Discontinuous Galerkin Scheme - Part 3: Secondary Flow Effects","authors":"Christian Morsbach, Michael Bergmann, Adem Tosun, Bjoern F. Klose, Patrick Bechlars, Edmund Kuegeler","doi":"10.1115/1.4063511","DOIUrl":"https://doi.org/10.1115/1.4063511","url":null,"abstract":"Abstract In this final paper of a three-part series, we apply the numerical test rig based on a high-order discontinuous Galerkin scheme to the MTU T161 low-pressure turbine with diverging end walls at off-design Reynolds number of 90,000, Mach number of 0.6, and inflow angle of 41 deg. The inflow end wall boundary layers are prescribed in accordance with the experiment. Validation of the setup is shown against recent numerical references and the corresponding experimental data. Additionally, we propose and conduct a purely numerical experiment with upstream bar wake generators at a Strouhal number of 1.25, which is well above what was possible in the experiment. We discuss the flow physics at midspan and in the end wall region and highlight the influence of the wakes from the upstream row on the complex secondary flow system using instantaneous flow visualization, phase averages, and modal decomposition techniques.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"2 11","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134992696","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 time-averaged aerodynamics of a high-speed low-pressure turbine (LPT) cascade were investigated under the impact of unsteady wakes and purge flows. The tests were performed at an outlet Mach (M) and Reynolds numbers (Re) of 0.90 and 70 k, respectively. Unsteady wakes were simulated by means of a spoked-wheel type wake generator (WG), and a reduced frequency of 0.95 was achieved. The development of a purge flow system to operate at low-pressure levels is presented alongside guidelines for the operation of the circuit. The new purge system was commissioned in terms of its long-term stability and flow uniformity at the cascade inlet. The impact of varying purge flows from a cavity versus a flat endwall was assessed by means of static pressure measurements on the blade pressure side (PS) and suction side (SS) and traverses performed downstream with a miniaturized multi-hole probe. Differences in the secondary flow structures resultant from injecting purge flow into the flowfield are observed, namely, the intensification of the trailing shed vortex (TSV) and passage vortex (PV) that is also displaced away from the endwall. An increase of the endwall losses occurs as the flat endwall was replaced with a cavity geometry and the cavity purge flowrate increases.
{"title":"Aerodynamics of a High-Speed Low-Pressure Turbine Cascade With Cavity Purge and Unsteady Wakes","authors":"Gustavo Lopes, Loris Simonasis, Sergio Lavagnoli","doi":"10.1115/1.4063878","DOIUrl":"https://doi.org/10.1115/1.4063878","url":null,"abstract":"Abstract The time-averaged aerodynamics of a high-speed low-pressure turbine (LPT) cascade were investigated under the impact of unsteady wakes and purge flows. The tests were performed at an outlet Mach (M) and Reynolds numbers (Re) of 0.90 and 70 k, respectively. Unsteady wakes were simulated by means of a spoked-wheel type wake generator (WG), and a reduced frequency of 0.95 was achieved. The development of a purge flow system to operate at low-pressure levels is presented alongside guidelines for the operation of the circuit. The new purge system was commissioned in terms of its long-term stability and flow uniformity at the cascade inlet. The impact of varying purge flows from a cavity versus a flat endwall was assessed by means of static pressure measurements on the blade pressure side (PS) and suction side (SS) and traverses performed downstream with a miniaturized multi-hole probe. Differences in the secondary flow structures resultant from injecting purge flow into the flowfield are observed, namely, the intensification of the trailing shed vortex (TSV) and passage vortex (PV) that is also displaced away from the endwall. An increase of the endwall losses occurs as the flat endwall was replaced with a cavity geometry and the cavity purge flowrate increases.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"3 3","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134993612","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}
Deepanshu Singh, Paul F. Beard, David Cardwell, Kam S. Chana
Abstract The Oxford Turbine Research Facility (OTRF) is a high-speed rotating transient test facility, which allows unsteady aerodynamic and heat transfer measurements at engine representative conditions. In addition, a variety of inlet temperature profiles can be simulated in the rig including radial distortion, circumferential distortion, and swirl. However, the engine representative flows cause complications in the processing of heat transfer data. The unsteadiness in temperature data was found to significantly rise as temperature distortions were introduced in the nozzle guide vane (NGV) inlet profile, to model a lean-burn combustor exit. Using the NGV inlet temperature profile survey data, the thermal unsteadiness has been quantified and compared with a uniform inlet. The experiments with a radially varying NGV inlet temperature profile showed up to nine times higher thermal unsteadiness, compared to the uniform inlet. The second part of the paper is a continuation of the work presented in a previous paper by Singh et al. and describes improved methodologies for derived heat transfer calculations using thin-film gauges. In addition, the uncertainty associated with the derived heat transfer parameters, such as the heat transfer coefficient and adiabatic wall temperature has been quantified. The refined processing techniques have been demonstrated on casing heat transfer measurements, acquired in the OTRF with two inlet temperature profiles.
{"title":"UNDERSTANDING THERMAL UNSTEADINESS IN ENGINE REPRESENTATIVE FLOWS AND IMPROVED METHODOLOGIES FOR DERIVED HEAT TRANSFER CALCULATIONS USING THIN-FILM GAUGES","authors":"Deepanshu Singh, Paul F. Beard, David Cardwell, Kam S. Chana","doi":"10.1115/1.4063735","DOIUrl":"https://doi.org/10.1115/1.4063735","url":null,"abstract":"Abstract The Oxford Turbine Research Facility (OTRF) is a high-speed rotating transient test facility, which allows unsteady aerodynamic and heat transfer measurements at engine representative conditions. In addition, a variety of inlet temperature profiles can be simulated in the rig including radial distortion, circumferential distortion, and swirl. However, the engine representative flows cause complications in the processing of heat transfer data. The unsteadiness in temperature data was found to significantly rise as temperature distortions were introduced in the nozzle guide vane (NGV) inlet profile, to model a lean-burn combustor exit. Using the NGV inlet temperature profile survey data, the thermal unsteadiness has been quantified and compared with a uniform inlet. The experiments with a radially varying NGV inlet temperature profile showed up to nine times higher thermal unsteadiness, compared to the uniform inlet. The second part of the paper is a continuation of the work presented in a previous paper by Singh et al. and describes improved methodologies for derived heat transfer calculations using thin-film gauges. In addition, the uncertainty associated with the derived heat transfer parameters, such as the heat transfer coefficient and adiabatic wall temperature has been quantified. The refined processing techniques have been demonstrated on casing heat transfer measurements, acquired in the OTRF with two inlet temperature profiles.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"8 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134992702","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}
Michael Bergmann, Christian Morsbach, Bjoern F. Klose, Graham Ashcroft, Edmund Küegeler
Abstract In this first paper of a three-part series, we present the extension and validation of the high-order discontinuous Galerkin scheme in DLR’s CFD-solver trace for scale-resolving simulations of unsteady row interactions. The translational movement of rows in linear cascade experiments is represented in the numerical model by solving the equations in the relative frame of reference. To couple rows in different frames of reference, a sliding interface approach based on the mortar technique for non-conforming meshes has been developed. The verification of the approach is exemplified by three canonical test cases. First, the experimental order of convergence is verified for the isentropic vortex convection. Subsequently, the suitability of the sliding interface approach for scale-resolving simulations is tested on the Taylor–Green vortex flow and a turbulent cylinder flow. Finally, the LES solver is applied to the T106D cascade with upstream moving bars at an exit Reynolds number of 200,000 and exit Mach number of 0.4. The flow physics with and without bars is discussed in terms of the instantaneous flow field, and time- and phase-averaged quantities. The comparison with experimental data shows overall a good agreement, especially for the total pressure losses in the wake, but also reveals uncertainties related to the reproduction of an experiment in the numerical model.
{"title":"A Numerical Test Rig for Turbomachinery Flows Based on Large Eddy Simulations With a High-Order Discontinuous Galerkin Scheme - Part 1: Sliding Interfaces and Unsteady Row Interactions","authors":"Michael Bergmann, Christian Morsbach, Bjoern F. Klose, Graham Ashcroft, Edmund Küegeler","doi":"10.1115/1.4063734","DOIUrl":"https://doi.org/10.1115/1.4063734","url":null,"abstract":"Abstract In this first paper of a three-part series, we present the extension and validation of the high-order discontinuous Galerkin scheme in DLR’s CFD-solver trace for scale-resolving simulations of unsteady row interactions. The translational movement of rows in linear cascade experiments is represented in the numerical model by solving the equations in the relative frame of reference. To couple rows in different frames of reference, a sliding interface approach based on the mortar technique for non-conforming meshes has been developed. The verification of the approach is exemplified by three canonical test cases. First, the experimental order of convergence is verified for the isentropic vortex convection. Subsequently, the suitability of the sliding interface approach for scale-resolving simulations is tested on the Taylor–Green vortex flow and a turbulent cylinder flow. Finally, the LES solver is applied to the T106D cascade with upstream moving bars at an exit Reynolds number of 200,000 and exit Mach number of 0.4. The flow physics with and without bars is discussed in terms of the instantaneous flow field, and time- and phase-averaged quantities. The comparison with experimental data shows overall a good agreement, especially for the total pressure losses in the wake, but also reveals uncertainties related to the reproduction of an experiment in the numerical model.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"7 5","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134992706","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}
Filip Wasilczuk, Pawel Flaszynski, Piotr Doerffer, Krzysztof Marugi, Tomasz Borzecki
Abstract This paper presents a study conducted on a new gas turbine, designed to limit leakage in the labyrinth seal. The slots in the fin are used to generate a bypass flow, which obstructs the flow in the gap above the fin. The method was tested numerically and experimentally beforehand using a simplified model without rotation or blade passages. In this paper, the validation of the method using a model of a turbine stage is shown. RANS simulations using two turbulence models – Spalart-Allmaras (SA) and k-ω EARSM were conducted. Comparisons of leakage flow and stage efficiency for reference and fluidic sealing configurations are presented. Fluidic sealing configuration is effective and reduces the leakage flow by 13-18.5% (depending on the turbulence model). The analysis of the flow structure in the seal region revealed, that the use of fluidic sealing resulted in significant circumferential flow anisotropy.
{"title":"The Application of Fluidic Sealing in Shrouded Gas Turbine Blades","authors":"Filip Wasilczuk, Pawel Flaszynski, Piotr Doerffer, Krzysztof Marugi, Tomasz Borzecki","doi":"10.1115/1.4064045","DOIUrl":"https://doi.org/10.1115/1.4064045","url":null,"abstract":"Abstract This paper presents a study conducted on a new gas turbine, designed to limit leakage in the labyrinth seal. The slots in the fin are used to generate a bypass flow, which obstructs the flow in the gap above the fin. The method was tested numerically and experimentally beforehand using a simplified model without rotation or blade passages. In this paper, the validation of the method using a model of a turbine stage is shown. RANS simulations using two turbulence models – Spalart-Allmaras (SA) and k-ω EARSM were conducted. Comparisons of leakage flow and stage efficiency for reference and fluidic sealing configurations are presented. Fluidic sealing configuration is effective and reduces the leakage flow by 13-18.5% (depending on the turbulence model). The analysis of the flow structure in the seal region revealed, that the use of fluidic sealing resulted in significant circumferential flow anisotropy.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"20 5","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-10","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135092537","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}
Zhibo Chen, Marshall Galbraith, Zoltan Spakovszky, Edward Greitzer, Jayant Sabnis
Abstract This article presents conceptual design guidelines and results for a tail-integrated propulsion system for a turbo-electric civil transport aircraft with boundary layer ingestion (BLI). The aerodynamic performance goal is separation-free and shock-free operation at cruise with fuel burn reduction, compared with a baseline conventional aircraft for the same mission. The assessment of BLI benefits is based on calculations using CFD and TASOPT software, both to characterize the design challenges and to establish the physical mechanisms for resolving these challenges. The guidelines include a “horseshoe” inlet to accept the non-uniform flow without incurring separation, a nacelle profile similar to supercritical airfoils to reduce shock strength, and an annular nozzle to eliminate flow separation between tail-BLI propulsors. The conceptual design has nine BLI propulsors with electric fans on an axisymmetric tail of a single-aisle aircraft. The fans are powered by twin underwing turbofans. The estimated benefit of the tail-BLI, twin underwing turbofan aircraft is 10.4% in Payload-Range Fuel Consumption (PRFC) at a cruise Mach number of 0.8, compared to a baseline twin underwing turbofan configuration. Sensitivity studies further show that a 1% increase in installed (i.e., with BLI) fan isentropic efficiency translates to 0.8% rise in PRFC benefit.
{"title":"Tail-Integrated Boundary Layer Ingesting Propulsion Systems for Turbo-Electric Aircraft","authors":"Zhibo Chen, Marshall Galbraith, Zoltan Spakovszky, Edward Greitzer, Jayant Sabnis","doi":"10.1115/1.4063930","DOIUrl":"https://doi.org/10.1115/1.4063930","url":null,"abstract":"Abstract This article presents conceptual design guidelines and results for a tail-integrated propulsion system for a turbo-electric civil transport aircraft with boundary layer ingestion (BLI). The aerodynamic performance goal is separation-free and shock-free operation at cruise with fuel burn reduction, compared with a baseline conventional aircraft for the same mission. The assessment of BLI benefits is based on calculations using CFD and TASOPT software, both to characterize the design challenges and to establish the physical mechanisms for resolving these challenges. The guidelines include a “horseshoe” inlet to accept the non-uniform flow without incurring separation, a nacelle profile similar to supercritical airfoils to reduce shock strength, and an annular nozzle to eliminate flow separation between tail-BLI propulsors. The conceptual design has nine BLI propulsors with electric fans on an axisymmetric tail of a single-aisle aircraft. The fans are powered by twin underwing turbofans. The estimated benefit of the tail-BLI, twin underwing turbofan aircraft is 10.4% in Payload-Range Fuel Consumption (PRFC) at a cruise Mach number of 0.8, compared to a baseline twin underwing turbofan configuration. Sensitivity studies further show that a 1% increase in installed (i.e., with BLI) fan isentropic efficiency translates to 0.8% rise in PRFC benefit.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" 60","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135340591","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 In supersonic aerospace applications, aerospike nozzles have been subject of growing interest. This study sheds light on the noise components of a cold jet exhausting an aerospike nozzle. Implicit large eddy simulations (ILES) are deployed to simulate the jet at a nozzle pressure ratio (NPR)=3. For far-field acoustic computation, the Ffowcs Williams–Hawkings (FWH) equation is applied. A mesh sensitivity study is performed and the jet instantaneous and time-averaged flow characteristics are analyzed. The annular shock structure displays short non-attached shock-cells and longer attached shock-cells. Downstream of the aerospike, a circular shock-cell structure is formed with long shock-cells. Two-point cross-correlations of data acquired at monitoring points located along the shear layers allow to identify upstream propagating waves associated to screech. Power spectral density at monitoring points in the annular shock-cell structure allows to identify its radial oscillation modes. Furthermore, a vortex sheet model is adapted to predict the annular shock-cells length and the BBSAN central frequency. High sound pressure levels (SPL) are detected at the determined BBSAN central frequencies. Finally, high SPL are obtained at the radial oscillation frequencies for the annular shock-cell structure.
摘要在超声速航空应用中，气动喷管受到越来越多的关注。本研究揭示了喷气喷嘴冷射流的噪声成分。采用隐式大涡模拟(ILES)对喷嘴压力比(NPR)=3时的射流进行了模拟。对于远场声学计算，采用Ffowcs williams - hawkins (FWH)方程。进行了网格灵敏度研究，分析了射流的瞬时和时均流动特性。环形激波结构具有较短的非附着激波单元和较长的附着激波单元。在气柱的下游，形成了由长激波组成的环形激波结构。沿剪切层监测点采集的两点互相关数据可以识别与尖啸有关的上游传播波。环形冲击单元结构监测点的功率谱密度允许识别其径向振荡模式。此外，采用涡片模型预测了环形激波单元的长度和BBSAN的中心频率。在确定的BBSAN中心频率处检测到高声压级(SPL)。最后，环形激波室结构在径向振荡频率下获得了较高的声压级。
{"title":"Computational Aeroacoustics for a Cold, Non-Ideally Expanded Aerospike Nozzle","authors":"Thomas Golliard, Mihai Mihaescu","doi":"10.1115/1.4063877","DOIUrl":"https://doi.org/10.1115/1.4063877","url":null,"abstract":"Abstract In supersonic aerospace applications, aerospike nozzles have been subject of growing interest. This study sheds light on the noise components of a cold jet exhausting an aerospike nozzle. Implicit large eddy simulations (ILES) are deployed to simulate the jet at a nozzle pressure ratio (NPR)=3. For far-field acoustic computation, the Ffowcs Williams–Hawkings (FWH) equation is applied. A mesh sensitivity study is performed and the jet instantaneous and time-averaged flow characteristics are analyzed. The annular shock structure displays short non-attached shock-cells and longer attached shock-cells. Downstream of the aerospike, a circular shock-cell structure is formed with long shock-cells. Two-point cross-correlations of data acquired at monitoring points located along the shear layers allow to identify upstream propagating waves associated to screech. Power spectral density at monitoring points in the annular shock-cell structure allows to identify its radial oscillation modes. Furthermore, a vortex sheet model is adapted to predict the annular shock-cells length and the BBSAN central frequency. High sound pressure levels (SPL) are detected at the determined BBSAN central frequencies. Finally, high SPL are obtained at the radial oscillation frequencies for the annular shock-cell structure.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135293132","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}
Xianjun Yu, Shouyang Zhao, Guangfeng An, Yuhang Xu, Xiaobin Xu
Abstract Blade surface roughness could significantly affect the aerodynamics performance of compressors. To explore the influence of roughness magnitude and location on blade performance, experiments were conducted in a low-speed linear compressor cascade with controlled diffusion airfoils (CDA). A part-span roughness method was employed in the experiment to maintain the axial velocity density ratio (AVDR) during the change of blade roughness magnitudes and locations. Five blade surface local roughness schemes, including the leading-edge, the fore- and aft-part of the suction surface, and the pressure surface, which were determined based on geometry sensitivity analysis, were investigated with the variation of the surface roughness magnitude between Ra = 3.1μm to 18.8μm. Cascade inlet and outlet flowfields and the blade surface static pressure distributions were measured, which could help to distinguish the change of blade performance characteristics and even blade surface boundary layer development state. A critical roughness effect was found and significant blade loss increment and available incidence range reduction appear at super-critical roughness states. At the measured maxi-roughness condition, 28.4% loss increase and 41.2% incidence range reduction were reached.
{"title":"Experimental investigations on the effects of surface roughness for compressor cascades with different roughness magnitude and location","authors":"Xianjun Yu, Shouyang Zhao, Guangfeng An, Yuhang Xu, Xiaobin Xu","doi":"10.1115/1.4063973","DOIUrl":"https://doi.org/10.1115/1.4063973","url":null,"abstract":"Abstract Blade surface roughness could significantly affect the aerodynamics performance of compressors. To explore the influence of roughness magnitude and location on blade performance, experiments were conducted in a low-speed linear compressor cascade with controlled diffusion airfoils (CDA). A part-span roughness method was employed in the experiment to maintain the axial velocity density ratio (AVDR) during the change of blade roughness magnitudes and locations. Five blade surface local roughness schemes, including the leading-edge, the fore- and aft-part of the suction surface, and the pressure surface, which were determined based on geometry sensitivity analysis, were investigated with the variation of the surface roughness magnitude between Ra = 3.1μm to 18.8μm. Cascade inlet and outlet flowfields and the blade surface static pressure distributions were measured, which could help to distinguish the change of blade performance characteristics and even blade surface boundary layer development state. A critical roughness effect was found and significant blade loss increment and available incidence range reduction appear at super-critical roughness states. At the measured maxi-roughness condition, 28.4% loss increase and 41.2% incidence range reduction were reached.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"12 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-11-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135935482","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}
Massimiliano Nardini, Thomas Jelly, Melissa Kozul, Richard Sandberg, Paul Vitt, Greg Sluyter
Abstract Turbine blades generally present surface roughness introduced in the manufacturing process or caused by in-service degradation, which can have a significant impact on aero-thermal performance. A better understanding of the fundamental physical mechanisms arising from the interaction between the roughness and the turbine flow at engine-relevant conditions can provide insights for the design of blades with improved efficiency and longer operational life. To this end, a high-fidelity numerical framework combining a well-validated solver for Direct Numerical Simulation and a second-order accurate immersed boundary method is employed to predict roughness-induced aero-thermal effects on an LS89 high-pressure turbine (HPT) blade at engine-relevant conditions. Different amplitudes and distributions of surface roughness are investigated and a reference smooth-blade simulation under the same flow conditions is conducted for comparison. Roughness of increasing amplitude progressively shifts the blade suction side boundary layer transition upstream, producing larger values of the turbulent kinetic energy and higher total wake losses. The on-surface data-capturing capabilities of the numerical framework provide direct measurements of the heat flux and the skin friction coefficient, hence offering quantitative information between the surface topology and engineering-relevant performance parameters. This work may provide a benchmark for future numerical studies of turbomachinery flows with roughness.
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