Previous investigations performed at the von Karman Institute for Fluid Dynamics (VKI) have shown an influence of the gas-to-wall temperature ratio on the bypass transition development along the VKI LS89 blade suction side. In the present work, the influence of this quantity on the flow field downstream of this highly-loaded nozzle guide vane is studied through the evaluation of the aerodynamic losses. The investigation is organized in three sections with different combinations of exit Mach numbers and freestream turbulence intensity (FSTI) while Tgas/Twall is varied between 1.1 and 1.3 for all the tests. The Isentropic Compression Tube facility (CT-2) at VKI allowed the determination of the total pressure loss across the cascade by means of a Pitot tube in the upstream region and a downstream three-hole needle probe. The latter is traversed in the pitch-wise direction by a pneumatic traversing system. Finally, the cascade aerodynamic efficiency is quantified by means of the kinetic energy loss coefficient ζ and the total pressure drop profile distortions in the wake region.
{"title":"Influence of the Gas-to-Wall Temperature Ratio on the Boundary Layer Transition: Investigation of the Wake Behind a Turbine Nozzle Guide Vane","authors":"P. Formisano, T. S. C. Ferreira, T. Arts","doi":"10.1115/gt2019-91834","DOIUrl":"https://doi.org/10.1115/gt2019-91834","url":null,"abstract":"\u0000 Previous investigations performed at the von Karman Institute for Fluid Dynamics (VKI) have shown an influence of the gas-to-wall temperature ratio on the bypass transition development along the VKI LS89 blade suction side. In the present work, the influence of this quantity on the flow field downstream of this highly-loaded nozzle guide vane is studied through the evaluation of the aerodynamic losses. The investigation is organized in three sections with different combinations of exit Mach numbers and freestream turbulence intensity (FSTI) while Tgas/Twall is varied between 1.1 and 1.3 for all the tests. The Isentropic Compression Tube facility (CT-2) at VKI allowed the determination of the total pressure loss across the cascade by means of a Pitot tube in the upstream region and a downstream three-hole needle probe. The latter is traversed in the pitch-wise direction by a pneumatic traversing system. Finally, the cascade aerodynamic efficiency is quantified by means of the kinetic energy loss coefficient ζ and the total pressure drop profile distortions in the wake region.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"33 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127649977","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 study carries out parametric investigations on aerodynamic loss of various types of LP turbine airfoils characterized with different flow deceleration rates (DR) on their suction surfaces under the realistic flow conditions such as wake inflow and freestream turbulence. The Reynolds number examined in this study ranges from 57,000 to 170,000. As for the freestream turbulence, two levels of the turbulence are used, i.e., about 1.2% and 3.5%. Stagnation pressure distributions downstream of each of the airfoil cascades are measured by use of a Pitot tube, while steady-state and unsteady boundary-layers are measured over the rear part of suction surface and pressure side near the trailing edge using a single hot-wire probe. The measured boundary-layer data are used to estimate the cascade loss along with RANS (Reynolds-Averaged Navier-Stokes) simulations by taking advantage of the momentum-theory based Denton’s method. First, relationships between the cascade loss for each flow condition and DR are examined. The estimated loss values are then compared with the measured cascade loss to check the validity of the loss estimation method, which is a derivative of Denton’s method, under the realistic flow conditions.
在尾流和自由流湍流等实际流动条件下，对具有不同减速率(DR)的各类低压涡轮翼型吸力面气动损失进行了参数化研究。本研究检测的雷诺数范围为57,000至170,000。对于自由流湍流，我们使用了两个级别的湍流，即约1.2%和3.5%。每个翼型叶栅下游的滞止压力分布由皮托管测量，而稳态和非定常边界层在吸力面后部和靠近尾缘的压力侧使用单个热线探头测量。利用基于动量理论的Denton方法，利用实测边界层数据和RANS (reynolds - average Navier-Stokes)模拟来估计叶栅损失。首先，研究了每个流动条件下的叶栅损失与DR之间的关系。然后将估计的损失值与实测的叶栅损失进行比较，以验证损失估计方法(Denton方法的导数)在实际流动条件下的有效性。
{"title":"Parametric Studies on Aerodynamic Performance of Various Types of LP Turbine Airfoils for Aero-Engines Under the Influence Periodic Wakes and Freestream Turbulence","authors":"K. Funazaki, Daichi Murakami, Y. Okamura","doi":"10.1115/gt2019-90408","DOIUrl":"https://doi.org/10.1115/gt2019-90408","url":null,"abstract":"\u0000 This study carries out parametric investigations on aerodynamic loss of various types of LP turbine airfoils characterized with different flow deceleration rates (DR) on their suction surfaces under the realistic flow conditions such as wake inflow and freestream turbulence. The Reynolds number examined in this study ranges from 57,000 to 170,000. As for the freestream turbulence, two levels of the turbulence are used, i.e., about 1.2% and 3.5%. Stagnation pressure distributions downstream of each of the airfoil cascades are measured by use of a Pitot tube, while steady-state and unsteady boundary-layers are measured over the rear part of suction surface and pressure side near the trailing edge using a single hot-wire probe. The measured boundary-layer data are used to estimate the cascade loss along with RANS (Reynolds-Averaged Navier-Stokes) simulations by taking advantage of the momentum-theory based Denton’s method. First, relationships between the cascade loss for each flow condition and DR are examined. The estimated loss values are then compared with the measured cascade loss to check the validity of the loss estimation method, which is a derivative of Denton’s method, under the realistic flow conditions.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"215 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121720723","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The large superconducting magnet are widely used and play an important role in nuclear fusion device, high-energy accelerator, space target infrared and other forefront fields. In this paper, we design several cryogenic centrifugal compressor stages for helium refrigerator with 3.3K∼11K inlet temperature, used in exhausting the low-pressure and low-temperature saturated helium steam, further producing superfluid helium and cooling the large superconducting magnet. Due to the cryogenic compressor running environment and the internal structure, about 30W heat leakage from impeller hubs and the end of shafts into the flow passages are unavoidable, and it is necessary to consider this heat leakage in the centrifugal compressor impeller design. Firstly, adding the heat leakage into thermodynamic process and centrifugal compressor effective formula, then adjusting the inlet and outlet design parameters according to the heat transfer rate and make the preliminarily designs of each stage of the compressor. Four different methods for varying p design condition parameters are compared by CFD simulation. The 3-D impellers are designed by the Streamline Curvature Method, and the vaneless diffusers and volutes are designed by the inlet and outlet section aerodynamic parameters. When the best design method is determined, the blade profiles are further adjusted in order to improve the stage performance. Finally, the characteristic curves under variable working conditions and the generalized stage performance curves of each stage are calculated. In this paper, we delve into a high efficiency centrifugal compressor stage design method with the consideration of heat transfer effect. In the next step, more running data will be obtained after the further experiments and the long-term operation.
{"title":"Design of Cryogenic Centrifugal Compressor Stages With Heat Transfer Considerations","authors":"Jin Xiong, Yingkun Zhang, Penghua Guo, Jingyin Li","doi":"10.1115/gt2019-90614","DOIUrl":"https://doi.org/10.1115/gt2019-90614","url":null,"abstract":"\u0000 The large superconducting magnet are widely used and play an important role in nuclear fusion device, high-energy accelerator, space target infrared and other forefront fields. In this paper, we design several cryogenic centrifugal compressor stages for helium refrigerator with 3.3K∼11K inlet temperature, used in exhausting the low-pressure and low-temperature saturated helium steam, further producing superfluid helium and cooling the large superconducting magnet.\u0000 Due to the cryogenic compressor running environment and the internal structure, about 30W heat leakage from impeller hubs and the end of shafts into the flow passages are unavoidable, and it is necessary to consider this heat leakage in the centrifugal compressor impeller design. Firstly, adding the heat leakage into thermodynamic process and centrifugal compressor effective formula, then adjusting the inlet and outlet design parameters according to the heat transfer rate and make the preliminarily designs of each stage of the compressor. Four different methods for varying p design condition parameters are compared by CFD simulation. The 3-D impellers are designed by the Streamline Curvature Method, and the vaneless diffusers and volutes are designed by the inlet and outlet section aerodynamic parameters. When the best design method is determined, the blade profiles are further adjusted in order to improve the stage performance. Finally, the characteristic curves under variable working conditions and the generalized stage performance curves of each stage are calculated. In this paper, we delve into a high efficiency centrifugal compressor stage design method with the consideration of heat transfer effect. In the next step, more running data will be obtained after the further experiments and the long-term operation.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"9 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126352329","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}
L. Mare, M. Carnevale, M. Rife, D. Kulkarni, Richard Northall
The growth in diameter of turbofan engines exacerbates problems related to the interaction of the Outlet Guide Vanes (OGV), pylon and intake because it reduces the ratio between components gaps and disturbance wavelength. The main components of this interaction are the potential fields generated by the intake and by structural components in the bypass, the pylon and the Radial Drive Fairing (RDF). The OGV bladerow and the fan are immersed in these potential fields and suffer performance degradation as well as integrity issues as a result. Simple actuator-disc analysis shows that a uniform OGV cascade amplifies the effect of the pylon potential flow. Therefore, a number of methods have been proposed over the years to compute OGV exit flow angle patterns that result in an approximately circumferentially uniform static pressure field at fan exit. Within actuator disc approximations, the determination of the optimal exit flow angle pattern can be accomplished analytically but little information is obtained on how the geometry of the vanes ought to be modified. Consequently, it is not difficult to generate by this method OGV cascades that stall or choke locally. More recent contributions use CFD computations coupled to optimization methods to determine OGV patterns that reduce the distortion at the fan exit, while minimising some measure of OGV loss. Whilst in principle more rational, these methods encounter practical difficulties due the computational power needed to obtain reliable loss estimates while exploring large design spaces. In this paper the view is taken that the performance of the OGV bladerow can be preserved during the optimization process if the loading distribution of each vane is made to match the loading distribution of the nominal vane (i.e. the aerodynamic design intent with axisymmetric inlet and exit flow). As loading distributions are readily available from inviscid-type analysis, the generation of optimal OGV patterns can be accomplished with very reasonable computational expense using a method based on the model described in part I of this paper.
{"title":"Modelling and Design of LPC Components With Semi-Analytical Models: Part II — Design of Optimal OGV Cascades","authors":"L. Mare, M. Carnevale, M. Rife, D. Kulkarni, Richard Northall","doi":"10.1115/GT2019-90481","DOIUrl":"https://doi.org/10.1115/GT2019-90481","url":null,"abstract":"\u0000 The growth in diameter of turbofan engines exacerbates problems related to the interaction of the Outlet Guide Vanes (OGV), pylon and intake because it reduces the ratio between components gaps and disturbance wavelength. The main components of this interaction are the potential fields generated by the intake and by structural components in the bypass, the pylon and the Radial Drive Fairing (RDF). The OGV bladerow and the fan are immersed in these potential fields and suffer performance degradation as well as integrity issues as a result. Simple actuator-disc analysis shows that a uniform OGV cascade amplifies the effect of the pylon potential flow. Therefore, a number of methods have been proposed over the years to compute OGV exit flow angle patterns that result in an approximately circumferentially uniform static pressure field at fan exit. Within actuator disc approximations, the determination of the optimal exit flow angle pattern can be accomplished analytically but little information is obtained on how the geometry of the vanes ought to be modified. Consequently, it is not difficult to generate by this method OGV cascades that stall or choke locally. More recent contributions use CFD computations coupled to optimization methods to determine OGV patterns that reduce the distortion at the fan exit, while minimising some measure of OGV loss. Whilst in principle more rational, these methods encounter practical difficulties due the computational power needed to obtain reliable loss estimates while exploring large design spaces. In this paper the view is taken that the performance of the OGV bladerow can be preserved during the optimization process if the loading distribution of each vane is made to match the loading distribution of the nominal vane (i.e. the aerodynamic design intent with axisymmetric inlet and exit flow). As loading distributions are readily available from inviscid-type analysis, the generation of optimal OGV patterns can be accomplished with very reasonable computational expense using a method based on the model described in part I of this paper.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"63 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"131463189","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
M. Asli, Cleopatra Cuciumita, P. Stathopoulos, C. Paschereit
Detonative Pressure Gain Combustion has the potential to increase the propulsion efficiency of aero-engines and the thermal efficiency of stationary gas turbines. Important advances were made in this field, especially in the case of Rotating Detonation Combustion (RDC). Although experimental and numerical studies reported in the literature have significantly increased in number, the major open problem is a lack of efficient turbomachinery to transform the fluctuating potential energy from an RDC into power output. For this problem to be properly addressed, time resolved data at the outlet of an RDC needs to be collected. As a first step, numerical data can be used to generate a geometry for the turbine, which must be validated experimentally. To determine the performance of a turbine vane row, total pressure losses need to be measured. There are several challenges in measuring the total pressure between the outlet of an RDC and the inlet of a turbine vane row. The high temperature values, the distance of the pressure transducer from the outlet of the combustor lead to a lower time resolution of the pressure signal. The confined space is also an issue, allowing for very few options in measuring the total pressure. Another major problem is the shock wave that may form as a detached shock wave with respect to the body of the pressure probe at certain moments in the flow cycle, which leads to measuring a different value rather than the actual value of the flow field. To address these issues, the current study presents a numerical investigation of a guide vane row that was experimentally tested at the outlet of an RDC working on hydrogen and air under stoichiometric conditions. One of the vane rows was 3D printed with a geometry allowing the measurement of total pressure. Static pressure at the outlet of the RDC was also measured. It was observed that the measured pressures are average values in time. Based on these averages, the total inlet pressure and velocity variations in time were reconstructed in an exponential trend, according to the ones reported in the literature and the aforementioned experiments. These variations were set as inlet conditions for transient numerical simulations. Results show that the total pressure amplitude decreases significantly when the flow passes the annulus and the vanes as well. By looking in to the flow field detail, the presence of shock wave in front of the blade is investigated. Additionally, it is calculated that the average total pressure decreases 7.9% by the vane row.
{"title":"Numerical Investigation of a Turbine Guide Vane Exposed to Rotating Detonation Exhaust Flow","authors":"M. Asli, Cleopatra Cuciumita, P. Stathopoulos, C. Paschereit","doi":"10.1115/gt2019-91263","DOIUrl":"https://doi.org/10.1115/gt2019-91263","url":null,"abstract":"\u0000 Detonative Pressure Gain Combustion has the potential to increase the propulsion efficiency of aero-engines and the thermal efficiency of stationary gas turbines. Important advances were made in this field, especially in the case of Rotating Detonation Combustion (RDC). Although experimental and numerical studies reported in the literature have significantly increased in number, the major open problem is a lack of efficient turbomachinery to transform the fluctuating potential energy from an RDC into power output. For this problem to be properly addressed, time resolved data at the outlet of an RDC needs to be collected. As a first step, numerical data can be used to generate a geometry for the turbine, which must be validated experimentally. To determine the performance of a turbine vane row, total pressure losses need to be measured. There are several challenges in measuring the total pressure between the outlet of an RDC and the inlet of a turbine vane row. The high temperature values, the distance of the pressure transducer from the outlet of the combustor lead to a lower time resolution of the pressure signal. The confined space is also an issue, allowing for very few options in measuring the total pressure. Another major problem is the shock wave that may form as a detached shock wave with respect to the body of the pressure probe at certain moments in the flow cycle, which leads to measuring a different value rather than the actual value of the flow field. To address these issues, the current study presents a numerical investigation of a guide vane row that was experimentally tested at the outlet of an RDC working on hydrogen and air under stoichiometric conditions. One of the vane rows was 3D printed with a geometry allowing the measurement of total pressure. Static pressure at the outlet of the RDC was also measured. It was observed that the measured pressures are average values in time. Based on these averages, the total inlet pressure and velocity variations in time were reconstructed in an exponential trend, according to the ones reported in the literature and the aforementioned experiments. These variations were set as inlet conditions for transient numerical simulations. Results show that the total pressure amplitude decreases significantly when the flow passes the annulus and the vanes as well. By looking in to the flow field detail, the presence of shock wave in front of the blade is investigated. Additionally, it is calculated that the average total pressure decreases 7.9% by the vane row.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"59 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116947990","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
M. S. Chiong, S. Rajoo, R. Martinez-Botas, Torsten Palenschat, P. Weitzman, M. Anderson, T. Ebel
The meanline modeling of asymmetrical twin-scroll turbocharger turbine performance has been in continuous development since its first introduction in 2008 by Daimler AG. Asymmetric turbine features scrolls of different sizes to enable high-pressure exhaust gas recirculation (EGR), which has been proven to be effective. However, the asymmetric feature of the scrolls has been identified as the major difficulty in the meanline modeling. This is particularly true at unequal-admission operating conditions, where the two volute scrolls are subjected to different magnitude of pressure and mass flow rate. Tuning a mean-line model for one particular turbine unit is usually possible, depending on the amount of flexibility in its loss sub-models. However the transferability of loss coefficients for the use of different turbine sizes has not been studied in the case of asymmetrical twin-scroll turbine. The feasibility in doing so, and the consequences, difficulties and outcomes, will be presented in this paper. The modeling work in this study will be carried out using commercial meanline tool — RITAL™ from Concepts NREC. The loss model is first tuned for the largest asymmetrical twin-scroll turbine in Daimler AG’s product line, at different admission conditions. After the model prediction has been validated with coldflow experimental data, the same loss coefficient settings are then used for performance prediction of a smaller asymmetrical twin-scroll turbine. The finding of this paper suggests that the fraction of the volute flow exit area varies with turbine admission state. The variation trend exhibits a unique function of the turbine geometry, and found to be consistent across the turbine sizes. On the other hand, the rotor loss coefficient may remain constant regardless of the turbine operation. The goal of this study is to establish the confidence level for blind performance prediction in the future.
{"title":"Meanline Modeling of Asymmetrical Twin-Scroll Turbine: Loss Coefficient Transferability Between Turbines of Different Sizes","authors":"M. S. Chiong, S. Rajoo, R. Martinez-Botas, Torsten Palenschat, P. Weitzman, M. Anderson, T. Ebel","doi":"10.1115/gt2019-91982","DOIUrl":"https://doi.org/10.1115/gt2019-91982","url":null,"abstract":"\u0000 The meanline modeling of asymmetrical twin-scroll turbocharger turbine performance has been in continuous development since its first introduction in 2008 by Daimler AG. Asymmetric turbine features scrolls of different sizes to enable high-pressure exhaust gas recirculation (EGR), which has been proven to be effective. However, the asymmetric feature of the scrolls has been identified as the major difficulty in the meanline modeling. This is particularly true at unequal-admission operating conditions, where the two volute scrolls are subjected to different magnitude of pressure and mass flow rate. Tuning a mean-line model for one particular turbine unit is usually possible, depending on the amount of flexibility in its loss sub-models. However the transferability of loss coefficients for the use of different turbine sizes has not been studied in the case of asymmetrical twin-scroll turbine. The feasibility in doing so, and the consequences, difficulties and outcomes, will be presented in this paper. The modeling work in this study will be carried out using commercial meanline tool — RITAL™ from Concepts NREC. The loss model is first tuned for the largest asymmetrical twin-scroll turbine in Daimler AG’s product line, at different admission conditions. After the model prediction has been validated with coldflow experimental data, the same loss coefficient settings are then used for performance prediction of a smaller asymmetrical twin-scroll turbine. The finding of this paper suggests that the fraction of the volute flow exit area varies with turbine admission state. The variation trend exhibits a unique function of the turbine geometry, and found to be consistent across the turbine sizes. On the other hand, the rotor loss coefficient may remain constant regardless of the turbine operation. The goal of this study is to establish the confidence level for blind performance prediction in the future.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"87 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"131833077","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}
Modern turbomachinery faces increased performance demands in terms of efficiency, compactness, and pressure-rise. Advancements in computational technology have allowed numerical methods to become the backbone of design development efforts. However, the unique complexities of centrifugal compressor flow-fields pose difficult computational problems. As such, advanced experimental methods must be used to obtain high-quality datasets to further inform, improve, and validate computational methods in complex flow regimes. Recent experimental work on a high-speed centrifugal compressor has provided detailed, unsteady, three-component velocity data using Laser Doppler Velocimetry. A passage vortex is present and its nascent tied to the increased incidence at mid-span associated with impeller wake flow. This vortex begins in the hub-pressure side corner and grows to fill the passage and become temporally stable. The vortex development is unsteady in nature and the unsteady effects persist 40% downstream of the throat. Distinct jet and wake flow patterns from the impeller also do not agglomerate until 40% downstream of the throat. Additionally, the critical impact of the unsteady flow development on the time-averaged flow-field is explained.
{"title":"LDV Characterization of Unsteady Vaned Diffuser Flow in a Centrifugal Compressor","authors":"W. Gooding, John C. Fabian, N. Key","doi":"10.1115/gt2019-90476","DOIUrl":"https://doi.org/10.1115/gt2019-90476","url":null,"abstract":"\u0000 Modern turbomachinery faces increased performance demands in terms of efficiency, compactness, and pressure-rise. Advancements in computational technology have allowed numerical methods to become the backbone of design development efforts. However, the unique complexities of centrifugal compressor flow-fields pose difficult computational problems. As such, advanced experimental methods must be used to obtain high-quality datasets to further inform, improve, and validate computational methods in complex flow regimes.\u0000 Recent experimental work on a high-speed centrifugal compressor has provided detailed, unsteady, three-component velocity data using Laser Doppler Velocimetry. A passage vortex is present and its nascent tied to the increased incidence at mid-span associated with impeller wake flow. This vortex begins in the hub-pressure side corner and grows to fill the passage and become temporally stable. The vortex development is unsteady in nature and the unsteady effects persist 40% downstream of the throat. Distinct jet and wake flow patterns from the impeller also do not agglomerate until 40% downstream of the throat. Additionally, the critical impact of the unsteady flow development on the time-averaged flow-field is explained.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"351 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134364582","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}
S. Gövert, F. Ferraro, A. Krumme, Clemens Buske, M. Tegeler, F. Kocian, F. Mare
Reducing the uncertainties in the prediction of turbine inlet conditions is a crucial aspect to improve aero engine designs and further increase engine efficiencies. To meet constantly stricter emission regulations, lean burn combustion could play a key role for future engine designs. However, these combustion systems are characterized by significant swirl for flame stabilization and reduced cooling air mass flows. As a result, substantial spatial and transient variations of the turbine inlet conditions are encountered. To investigate the effect of the combustor on the high pressure turbine, a rotating cooled transonic high-pressure configuration has been designed and investigated experimentally at the DLR turbine test facility ‘NG-Turb’ in Göttingen, Germany. It is a rotating full annular 1.5 stage turbine configuration which is coupled to a combustor simulator. The combustor simulator is designed to create turbine inlet conditions which are hydrodynamically representative for a lean-burn aero engine. A detailed description of the test rig and its instrumentation as well as a discussion of the measurement results is presented in part I of this paper. Part II focuses on numerical modeling of the test rig to further extend the understanding of the measurement results. Integrated simulations of the configuration including combustor simulator and nozzle guide vanes are performed for leading edge and passage clocking position and the effect on the hot streak migration is discussed. The simulation and experimental results at the combustor-turbine interface are compared showing a good overall agreement. The relevant flow features are correctly predicted in the simulations, proving the suitability of the numerical model for application to integrated combustor-turbine interaction analysis.
{"title":"Investigation of Combustor-Turbine-Interaction in a Rotating Cooled Transonic High-Pressure Turbine Test Rig: Part 2 — Numerical Modelling and Simulation","authors":"S. Gövert, F. Ferraro, A. Krumme, Clemens Buske, M. Tegeler, F. Kocian, F. Mare","doi":"10.1115/gt2019-90736","DOIUrl":"https://doi.org/10.1115/gt2019-90736","url":null,"abstract":"\u0000 Reducing the uncertainties in the prediction of turbine inlet conditions is a crucial aspect to improve aero engine designs and further increase engine efficiencies. To meet constantly stricter emission regulations, lean burn combustion could play a key role for future engine designs. However, these combustion systems are characterized by significant swirl for flame stabilization and reduced cooling air mass flows. As a result, substantial spatial and transient variations of the turbine inlet conditions are encountered. To investigate the effect of the combustor on the high pressure turbine, a rotating cooled transonic high-pressure configuration has been designed and investigated experimentally at the DLR turbine test facility ‘NG-Turb’ in Göttingen, Germany. It is a rotating full annular 1.5 stage turbine configuration which is coupled to a combustor simulator. The combustor simulator is designed to create turbine inlet conditions which are hydrodynamically representative for a lean-burn aero engine. A detailed description of the test rig and its instrumentation as well as a discussion of the measurement results is presented in part I of this paper. Part II focuses on numerical modeling of the test rig to further extend the understanding of the measurement results. Integrated simulations of the configuration including combustor simulator and nozzle guide vanes are performed for leading edge and passage clocking position and the effect on the hot streak migration is discussed. The simulation and experimental results at the combustor-turbine interface are compared showing a good overall agreement. The relevant flow features are correctly predicted in the simulations, proving the suitability of the numerical model for application to integrated combustor-turbine interaction analysis.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"30 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134634311","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}
Launch systems development today is heading in two, seemingly divergent, directions. A first direction is towards bigger launchers, designed to carry more than 50 tons into low Earth orbit (LEO). On the other hand, there’s growing interest from government agencies and start-up companies alike in very small vehicles for dedicated launches of small satellites, vehicles that can place just a few hundred, or even a few dozen, kilograms into LEO. At the same time, space programs are starting to look into the possibility of using the existing turbines of rocket engines turbopumps working on classical fuels to work on alternative fuels. The desired characteristics of being a simple, lightweight, high specific work output, low mass flow rate turbine easily translate into supersonic turbines and outweigh the disadvantage of having low efficiency compared to subsonic turbines. Such a turbine needs proof of concept because, at small scales, the flow changes dramatically due to end wall losses which may cause the turbine to choke prematurely and combined with the effect of shock wave losses, characteristic to a supersonic flow, this effect can turn out to be critical. This paper presents a methodology to design such a turbine, taking into account the requirements derived from an application represented by a micro launcher with a maximum payload of 100 kg. As compared to the design of a classical turbine, this methodology is focused on geometrical limitations, to ensure the manufacturability of the turbine, as well as aerodynamic efficiency. The methodology was applied for obtaining the geometry of a turbine for the aforementioned application using a classical fuel as design point. 3D numerical simulations were computed for this geometry, and the efficiency of the turbine was obtained within 8% of the analytical data. To facilitate the use of different fuels, a simple and fast method was also developed for predicting the performance of a turbine of known geometry and performance for an initial working fluid when changing the nature of this working fluid. Benefiting from having the performance already estimated for another working fluid (the design fluid), the Mach numbers similarity criterion can be used to estimate the performance of the same geometry when changing the working fluid, as a known practice in the gas-turbine field. The Mach number appears as a scaling parameter in many of the equations for compressible flows, shock waves, and expansions. Not only is it suited for similarity of a classical turbine, it is also appropriate for a supersonic turbine, due to the fact the conditions behind shock waves are only dependent on the Mach number and the fluid’s properties. Using this method, the performance of the designed turbine was computed for a completely different working fluid than the design one. The largest difference in power output generated by changing the working fluid is of 170%. A new set of numerical simulations was done, and the results co
{"title":"Design of a Working Fluid Versatile Supersonic Turbine for Micro Launchers","authors":"Cleopatra Cuciumita, V. Vilag","doi":"10.1115/gt2019-91255","DOIUrl":"https://doi.org/10.1115/gt2019-91255","url":null,"abstract":"\u0000 Launch systems development today is heading in two, seemingly divergent, directions. A first direction is towards bigger launchers, designed to carry more than 50 tons into low Earth orbit (LEO). On the other hand, there’s growing interest from government agencies and start-up companies alike in very small vehicles for dedicated launches of small satellites, vehicles that can place just a few hundred, or even a few dozen, kilograms into LEO. At the same time, space programs are starting to look into the possibility of using the existing turbines of rocket engines turbopumps working on classical fuels to work on alternative fuels. The desired characteristics of being a simple, lightweight, high specific work output, low mass flow rate turbine easily translate into supersonic turbines and outweigh the disadvantage of having low efficiency compared to subsonic turbines. Such a turbine needs proof of concept because, at small scales, the flow changes dramatically due to end wall losses which may cause the turbine to choke prematurely and combined with the effect of shock wave losses, characteristic to a supersonic flow, this effect can turn out to be critical. This paper presents a methodology to design such a turbine, taking into account the requirements derived from an application represented by a micro launcher with a maximum payload of 100 kg. As compared to the design of a classical turbine, this methodology is focused on geometrical limitations, to ensure the manufacturability of the turbine, as well as aerodynamic efficiency. The methodology was applied for obtaining the geometry of a turbine for the aforementioned application using a classical fuel as design point. 3D numerical simulations were computed for this geometry, and the efficiency of the turbine was obtained within 8% of the analytical data. To facilitate the use of different fuels, a simple and fast method was also developed for predicting the performance of a turbine of known geometry and performance for an initial working fluid when changing the nature of this working fluid. Benefiting from having the performance already estimated for another working fluid (the design fluid), the Mach numbers similarity criterion can be used to estimate the performance of the same geometry when changing the working fluid, as a known practice in the gas-turbine field. The Mach number appears as a scaling parameter in many of the equations for compressible flows, shock waves, and expansions. Not only is it suited for similarity of a classical turbine, it is also appropriate for a supersonic turbine, due to the fact the conditions behind shock waves are only dependent on the Mach number and the fluid’s properties. Using this method, the performance of the designed turbine was computed for a completely different working fluid than the design one. The largest difference in power output generated by changing the working fluid is of 170%. A new set of numerical simulations was done, and the results co","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"38 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116478319","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}
G. Xi, Yubao Tian, Y. Guo, Yonghong Tang, Zhiheng Wang
The flow losses in a U-bend and return channel system of an inter stage has an important influence on the stage performance. To study the flow characteristics in such system, a pseudo-stage which consists of pseudo guide vanes, U-bend and return channel, combined with a normal stage, was designed and its performance curves were measured at different machine Mach numbers. Simultaneously, the numerical simulations, in which different data processing methods of CFD post process were adopted, were performed and the prediction accuracy was verified by comparison with the experimental results. The results indicated that the design of pseudo guide vanes provides an approximately constant flow direction at the inlet of U-bend with the variation of inlet Reynolds number, which matches the blade metal angle of the return channel well. The further analyses on the measured and numerical results showed that the loss characteristics vary with the inlet Reynolds number. When the Reynolds number is greater than a critical Reynolds number, the loss coefficient in the U-bend and return channel keeps nearly constant. The non-uniformity of flow angle and total pressure at the outlet of return channel were also illustrated and discussed.
{"title":"Experimental and Numerical Investigations on Flow Losses of a U-Bend and Return Channel System for Centrifugal Compressor","authors":"G. Xi, Yubao Tian, Y. Guo, Yonghong Tang, Zhiheng Wang","doi":"10.1115/gt2019-90559","DOIUrl":"https://doi.org/10.1115/gt2019-90559","url":null,"abstract":"\u0000 The flow losses in a U-bend and return channel system of an inter stage has an important influence on the stage performance. To study the flow characteristics in such system, a pseudo-stage which consists of pseudo guide vanes, U-bend and return channel, combined with a normal stage, was designed and its performance curves were measured at different machine Mach numbers. Simultaneously, the numerical simulations, in which different data processing methods of CFD post process were adopted, were performed and the prediction accuracy was verified by comparison with the experimental results. The results indicated that the design of pseudo guide vanes provides an approximately constant flow direction at the inlet of U-bend with the variation of inlet Reynolds number, which matches the blade metal angle of the return channel well. The further analyses on the measured and numerical results showed that the loss characteristics vary with the inlet Reynolds number. When the Reynolds number is greater than a critical Reynolds number, the loss coefficient in the U-bend and return channel keeps nearly constant. The non-uniformity of flow angle and total pressure at the outlet of return channel were also illustrated and discussed.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"111 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2019-06-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124882667","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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