The paper presents the results of aeromechanical design of a large-scale model stage for a high-efficient low-noise fan designed for an advanced civil geared turbofan engines with ultra-low rotational speeds of rotor blades (313.4 m/s), high flow specific capacity (up to 202 kg/m2/s) and high bypass ratio (13.5). Total pressure ratio in the bypass duct of the fan model stage is 1.38. To ensure the experimental studies, characteristics are calculated from choking to a surge line within a wide range of rotational speeds. For the studies of the experimental fan model (EFM), a design project is developed and used in manufacturing a fan stage with 0.7-m rotor diameter for tests at the C-3A acoustic test facility. The manufacturing technology for blades made of polymer composite materials (PCM) is of particular importance. Rotor blades of the geared fan model are made of PCM. The analysis of experimental data and their comparison with the computation results within the range of corrected rotational speeds from 0.325 to 1.0 are presented. At first, only gas-dynamic and strength characteristics of the stage are studied. The analysis shows a good agreement of calculated integral parameters with the experimental data. Acoustic performances of the EFM will be studied later on.
{"title":"Computational and Experimental Studies of Model Fans for Advanced Turbofan Engines","authors":"S. Pankov, V. Korzhnev, V. Mileshin, V. Fateev","doi":"10.1115/gt2021-59524","DOIUrl":"https://doi.org/10.1115/gt2021-59524","url":null,"abstract":"\u0000 The paper presents the results of aeromechanical design of a large-scale model stage for a high-efficient low-noise fan designed for an advanced civil geared turbofan engines with ultra-low rotational speeds of rotor blades (313.4 m/s), high flow specific capacity (up to 202 kg/m2/s) and high bypass ratio (13.5). Total pressure ratio in the bypass duct of the fan model stage is 1.38.\u0000 To ensure the experimental studies, characteristics are calculated from choking to a surge line within a wide range of rotational speeds.\u0000 For the studies of the experimental fan model (EFM), a design project is developed and used in manufacturing a fan stage with 0.7-m rotor diameter for tests at the C-3A acoustic test facility.\u0000 The manufacturing technology for blades made of polymer composite materials (PCM) is of particular importance. Rotor blades of the geared fan model are made of PCM.\u0000 The analysis of experimental data and their comparison with the computation results within the range of corrected rotational speeds from 0.325 to 1.0 are presented. At first, only gas-dynamic and strength characteristics of the stage are studied. The analysis shows a good agreement of calculated integral parameters with the experimental data. Acoustic performances of the EFM will be studied later on.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"16 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"124326907","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}
In the present paper, the performance of compressor blades modified with leading edge tubercles was evaluated and compared with that of a baseline profile at a high subsonic Mach number in a 2-D cascade. Specific tubercle geometries were selected based on an extensive literature survey and a Self-Organizing Map analysis. The compressor blade geometry of a popular aero-engine was reverse-engineered using laser-scanning. Baseline and tubercled compressor blades were 3-D printed and tested. Two sinusoidal tubercle shapes based on different amplitudes and wavelengths and one with a power law profile were selected. A 2-D compressor cascade was designed and commissioned to test these blades at high subsonic Mach number in the transonic wind tunnel at Royal Military College of Canada. Surface flow visualizations were performed with oil for observing and locating compressor blade stall for different sets of blades. Flow direction and the total pressure at the cascade exit were measured using a 5-hole, fast-response, traversing probe. Compressor blade performance was measured and compared with various tubercled blades at various angles of incidence, while maintaining periodicity at the inlet and exit planes. Total pressure loss coefficients were calculated for all 4 blades and compared for 6 positive angle of incidence. Power series tubercled profile resulted in slight improvements in the loss coefficient at 0° incidence and none of tubercled geometry compromized performance at the design point. The baseline blade stalled at 8° and tubercles were capable of delaying stall at this condition. Power series profile outperformed the baseline at all angle of incidence (AOI) with significant improvements at 8° AOI. Power series tubercled profiles performed better than other tubercled geometries at almost all AOI except 10° where sinusoidal tubercled profiles performed better. The presence of smaller valley and broader peaks is attributed with the performance improvement, supported by the flow visualization results.
{"title":"Performance Evaluation of Leading Edge Tubercles Applied to the Blades in a 2-D Compressor Cascade","authors":"S. Sidhu, Asad Asghar, W. Allan","doi":"10.1115/gt2021-58798","DOIUrl":"https://doi.org/10.1115/gt2021-58798","url":null,"abstract":"\u0000 In the present paper, the performance of compressor blades modified with leading edge tubercles was evaluated and compared with that of a baseline profile at a high subsonic Mach number in a 2-D cascade. Specific tubercle geometries were selected based on an extensive literature survey and a Self-Organizing Map analysis. The compressor blade geometry of a popular aero-engine was reverse-engineered using laser-scanning. Baseline and tubercled compressor blades were 3-D printed and tested. Two sinusoidal tubercle shapes based on different amplitudes and wavelengths and one with a power law profile were selected. A 2-D compressor cascade was designed and commissioned to test these blades at high subsonic Mach number in the transonic wind tunnel at Royal Military College of Canada. Surface flow visualizations were performed with oil for observing and locating compressor blade stall for different sets of blades. Flow direction and the total pressure at the cascade exit were measured using a 5-hole, fast-response, traversing probe. Compressor blade performance was measured and compared with various tubercled blades at various angles of incidence, while maintaining periodicity at the inlet and exit planes. Total pressure loss coefficients were calculated for all 4 blades and compared for 6 positive angle of incidence. Power series tubercled profile resulted in slight improvements in the loss coefficient at 0° incidence and none of tubercled geometry compromized performance at the design point. The baseline blade stalled at 8° and tubercles were capable of delaying stall at this condition. Power series profile outperformed the baseline at all angle of incidence (AOI) with significant improvements at 8° AOI. Power series tubercled profiles performed better than other tubercled geometries at almost all AOI except 10° where sinusoidal tubercled profiles performed better. The presence of smaller valley and broader peaks is attributed with the performance improvement, supported by the flow visualization results.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"57 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129618472","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}
Jaehyoun Lee, S. Lim, Hyoun-Woo Shin, Sungryong Lee, S. Song
Periodic unsteady flow kinematics in a shrouded multistage low-speed axial compressor has been measured for the first time. Data have been acquired at the inlet and exit of a shrouded 3rd-stage stator with a particular focus on the hub flows. The newly found features of the hub flow in a shrouded multistage compressor are different from those at the midspan or in unshrouded (i.e., cantilevered) compressors. First, the merging of the 2nd-stage stator and 3rd-stage rotor wakes causes positive radial migration near the rotor wake pressure surface at the hub of the 3rd-stage stator inlet. Second, the low-momentum labyrinth seal leakage flow of the 3rd-stage stator merges with the 3rd-stage rotor wake to create streamwise vorticity at the 3rd-stage stator inlet hub. Third, contrary to unshrouded stators, suction side hub corner separation in the shrouded 3rd-stage stator reduces rotor wake stretching. Thus, velocity disturbances are attenuated less, and amplitudes of periodic fluctuations in flow angles are larger at the 3rd-stage stator exit hub than at midspan. The positive radial migration of the rotor wake hub flow and wake stretching reduction are expected to decrease efficiency, whereas streamwise vorticity generation is expected to increase efficiency.
{"title":"Periodic Unsteady Kinematics of Hub Flows in a Shrouded Multistage Compressor","authors":"Jaehyoun Lee, S. Lim, Hyoun-Woo Shin, Sungryong Lee, S. Song","doi":"10.1115/gt2021-58551","DOIUrl":"https://doi.org/10.1115/gt2021-58551","url":null,"abstract":"Periodic unsteady flow kinematics in a shrouded multistage low-speed axial compressor has been measured for the first time. Data have been acquired at the inlet and exit of a shrouded 3rd-stage stator with a particular focus on the hub flows. The newly found features of the hub flow in a shrouded multistage compressor are different from those at the midspan or in unshrouded (i.e., cantilevered) compressors. First, the merging of the 2nd-stage stator and 3rd-stage rotor wakes causes positive radial migration near the rotor wake pressure surface at the hub of the 3rd-stage stator inlet. Second, the low-momentum labyrinth seal leakage flow of the 3rd-stage stator merges with the 3rd-stage rotor wake to create streamwise vorticity at the 3rd-stage stator inlet hub. Third, contrary to unshrouded stators, suction side hub corner separation in the shrouded 3rd-stage stator reduces rotor wake stretching. Thus, velocity disturbances are attenuated less, and amplitudes of periodic fluctuations in flow angles are larger at the 3rd-stage stator exit hub than at midspan. The positive radial migration of the rotor wake hub flow and wake stretching reduction are expected to decrease efficiency, whereas streamwise vorticity generation is expected to increase efficiency.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128549057","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 increased need to design higher performing aerodynamic shapes has led to design optimisation cycles requiring high-fidelity CFD models and high-dimensional parametrisation schemes. The computational cost of employing global search algorithms on such scenarios has typically been prohibitive for most academic and industrial environments. In this paper, a novel strategy is presented that leverages the capabilities of Artificial Neural Networks for regressing complex unstructured data, while coupling them with dimensionality reduction algorithms. This approach enables employing global-based optimisation methods on high-dimensional applications through a reduced computational cost. This methodology is demonstrated on the efficiency optimisation of a modern jet engine fan blade with constrained pressure ratio. The outcome is compared against a state-of-the-art adjoint-based approach. Results indicate the strategy proposed achieves comparable improvements to its adjoint counterpart with a reduced computational cost, and can scale better to multi-objective optimisation applications.
{"title":"Global Optimisation of a Transonic Fan Blade Through AI-Enabled Active Subspaces","authors":"Diego I. Lopez, T. Ghisu, S. Shahpar","doi":"10.1115/GT2021-59166","DOIUrl":"https://doi.org/10.1115/GT2021-59166","url":null,"abstract":"\u0000 The increased need to design higher performing aerodynamic shapes has led to design optimisation cycles requiring high-fidelity CFD models and high-dimensional parametrisation schemes. The computational cost of employing global search algorithms on such scenarios has typically been prohibitive for most academic and industrial environments. In this paper, a novel strategy is presented that leverages the capabilities of Artificial Neural Networks for regressing complex unstructured data, while coupling them with dimensionality reduction algorithms. This approach enables employing global-based optimisation methods on high-dimensional applications through a reduced computational cost. This methodology is demonstrated on the efficiency optimisation of a modern jet engine fan blade with constrained pressure ratio. The outcome is compared against a state-of-the-art adjoint-based approach. Results indicate the strategy proposed achieves comparable improvements to its adjoint counterpart with a reduced computational cost, and can scale better to multi-objective optimisation applications.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"3 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123869752","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
A. Budo, V. Terrapon, M. Arnst, K. Hillewaert, S. Mouriaux, B. Rodriguez, Jules Bartholet
This paper describes the evaluation of a newly developed viscous time-marching through-flow solver to two test cases to assess the applicability of the method using correlations from the literature to modern blade designs. The test cases are the classic axial compressor stage CME2 and a modern highly loaded multi-stage axial low-pressure compressor developed by Safran Aero Boosters. The through-flow solver is based on the Navier-Stokes equations and uses a pseudo-time marching method. The closure models currently include terms of major importance: the blade forces and the Reynolds stress. The results are compared to higher-fidelity results including three-dimensional RANS simulations to assess their reliability for design and off-design conditions. The main originality of this work is the evaluation of the CFD-based method in the context of a compressor with highly three-dimensional blades, as such an analysis is not commonly found in the literature. The solver gives realistic predictions of loss and deviation for the compressor stage CME2 at both design and off-design operating conditions. Regarding the second test case, the through-flow simulations based on theoretically non-adapted correlations for such a compressor are still in good agreement with RANS simulations, although the results for the 2nd test case are probably not as good as for the first. These results are a promising first step towards the use of this through-flow model for industrial design. Regarding the ongoing closure models development, suggestions to extend the loss models to a larger range of designs are discussed.
{"title":"Application of a Viscous Through-Flow Model to a Modern Axial Low-Pressure Compressor","authors":"A. Budo, V. Terrapon, M. Arnst, K. Hillewaert, S. Mouriaux, B. Rodriguez, Jules Bartholet","doi":"10.1115/gt2021-59926","DOIUrl":"https://doi.org/10.1115/gt2021-59926","url":null,"abstract":"\u0000 This paper describes the evaluation of a newly developed viscous time-marching through-flow solver to two test cases to assess the applicability of the method using correlations from the literature to modern blade designs. The test cases are the classic axial compressor stage CME2 and a modern highly loaded multi-stage axial low-pressure compressor developed by Safran Aero Boosters. The through-flow solver is based on the Navier-Stokes equations and uses a pseudo-time marching method. The closure models currently include terms of major importance: the blade forces and the Reynolds stress. The results are compared to higher-fidelity results including three-dimensional RANS simulations to assess their reliability for design and off-design conditions. The main originality of this work is the evaluation of the CFD-based method in the context of a compressor with highly three-dimensional blades, as such an analysis is not commonly found in the literature.\u0000 The solver gives realistic predictions of loss and deviation for the compressor stage CME2 at both design and off-design operating conditions. Regarding the second test case, the through-flow simulations based on theoretically non-adapted correlations for such a compressor are still in good agreement with RANS simulations, although the results for the 2nd test case are probably not as good as for the first. These results are a promising first step towards the use of this through-flow model for industrial design. Regarding the ongoing closure models development, suggestions to extend the loss models to a larger range of designs are discussed.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"14 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121276093","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}
Bhanu Pratap Singh Tanwar, Ajey Singh, Chetan S. Mistry
Adoption of a tandem bladed rotor configuration brings special flow features at the exit compared to the conventional rotor. For tandem bladed rotor, there is the presence of strong dual-tip leakage flow, atypical exit flow angle distributions, corner blade separations leading to thicker dual wakes at the exit of the rotor to name a few. This makes the aerodynamic design of downstream stator more challenging in terms of overall performance as well as operational stability. The modern compressor requisite of being lighter and cost-efficient needs to be taken care of both aerodynamic and mechanical requirements. To overcome all these challenges, the cantilever type stator (without hub rotation) has been chosen and been analyzed for the present study. The effects of different hub gap sizes of the cantilever stator in combination with the tandem bladed axial compressor stage are investigated in order to explore passive flow control mechanism near the hub. The goal of the work is to get further insights into the aerodynamic aspects of flow using a detailed flow field analysis. The numerical study was performed using ANSYS TurboGrid® for mesh generation and the commercial package ANSYS CFX® 18.0 was used as solver for steady-state simulation. Stationary hub boundary conditions have been employed for the stator in all 3 cases [baseline, 1% and 2% (of span) part clearance]. For no clearance case, the regions of momentum deficit were observed in the vicinity of the hub endwall and suction surface of the stator. The region keeps growing along both streamwise and spanwise direction as a low momentum bubble is formed near trailing edge. This low momentum bubble seems to be transported along the span and moved more towards the suction surface. The solution strategy explored to mitigate the effect of hub corner separation by adapting hub clearance. The role played by secondary flow in feeding the low momentum flow along the span is seen to be moderated by the high momentum leakage flow from the pressure side. The hub leakage flow from the blade pressure side reenergized the low momentum fluid on the suction side refraining it to travel along the span and mitigate its effect by suppressing the separation tendency near end wall region. The formation of large size bubble gets reduced in overall size both in the circumferential and span-wise direction. This phenomenon compels the low momentum flow to pass along the low span region. Numerically obtained results provide an insightful mechanism of the interaction of secondary flow structures and the influence of hub clearance flow. Hub corner stall, which is the consequence of low momentum fluid sweeping across the blade passage near the end wall got wiped out in the presence of hub clearance. This phenomenon diminishes the extent and overall effect of the hub corner stall. The interaction of hub leakage vortex and passage vortex leads to mitigation of overall secondary flow adverse effects. As a result, perform
{"title":"Numerical Investigations on Application of Cantilever Stator on Aerodynamic Performance of Tandem Bladed Axial-Flow Compressor","authors":"Bhanu Pratap Singh Tanwar, Ajey Singh, Chetan S. Mistry","doi":"10.1115/gt2021-60233","DOIUrl":"https://doi.org/10.1115/gt2021-60233","url":null,"abstract":"\u0000 Adoption of a tandem bladed rotor configuration brings special flow features at the exit compared to the conventional rotor. For tandem bladed rotor, there is the presence of strong dual-tip leakage flow, atypical exit flow angle distributions, corner blade separations leading to thicker dual wakes at the exit of the rotor to name a few. This makes the aerodynamic design of downstream stator more challenging in terms of overall performance as well as operational stability. The modern compressor requisite of being lighter and cost-efficient needs to be taken care of both aerodynamic and mechanical requirements. To overcome all these challenges, the cantilever type stator (without hub rotation) has been chosen and been analyzed for the present study. The effects of different hub gap sizes of the cantilever stator in combination with the tandem bladed axial compressor stage are investigated in order to explore passive flow control mechanism near the hub. The goal of the work is to get further insights into the aerodynamic aspects of flow using a detailed flow field analysis.\u0000 The numerical study was performed using ANSYS TurboGrid® for mesh generation and the commercial package ANSYS CFX® 18.0 was used as solver for steady-state simulation. Stationary hub boundary conditions have been employed for the stator in all 3 cases [baseline, 1% and 2% (of span) part clearance].\u0000 For no clearance case, the regions of momentum deficit were observed in the vicinity of the hub endwall and suction surface of the stator. The region keeps growing along both streamwise and spanwise direction as a low momentum bubble is formed near trailing edge. This low momentum bubble seems to be transported along the span and moved more towards the suction surface. The solution strategy explored to mitigate the effect of hub corner separation by adapting hub clearance. The role played by secondary flow in feeding the low momentum flow along the span is seen to be moderated by the high momentum leakage flow from the pressure side. The hub leakage flow from the blade pressure side reenergized the low momentum fluid on the suction side refraining it to travel along the span and mitigate its effect by suppressing the separation tendency near end wall region. The formation of large size bubble gets reduced in overall size both in the circumferential and span-wise direction. This phenomenon compels the low momentum flow to pass along the low span region.\u0000 Numerically obtained results provide an insightful mechanism of the interaction of secondary flow structures and the influence of hub clearance flow. Hub corner stall, which is the consequence of low momentum fluid sweeping across the blade passage near the end wall got wiped out in the presence of hub clearance. This phenomenon diminishes the extent and overall effect of the hub corner stall. The interaction of hub leakage vortex and passage vortex leads to mitigation of overall secondary flow adverse effects. As a result, perform","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"22 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122347469","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}
Xiao-chen Wang, Xiao-dong Ren, Xue-song Li, Hong Wu, C. Gu
In this paper, a novel airfoil design method is proposed by introducing multiple physical characteristics to the process of parameterization. The new design method is to operate the camber and thickness distributions in a way that the chordwise mapping information is adjusted without using local geometrical parameters. A proper design space is expected to be constructed, in which the laws of optimization can be simply transmitted from a single airfoil to the whole blade and even stages. The capability of the method is firstly validated in the Pareto optimization of a 2D compressor airfoil, and then specialized for the optimization of a rotor in a 3-stage transonic compressor. The 3-stage transonic compressor studied in this paper is representative of the inlet stages of an industrial gas turbine, and the shock structure of the third rotor (R3) presents the peak strength near the mid span. The optimized design reduces the peak Mach number on the suction surface, and increases the efficiency of the third stage and the whole compressor by nearly 1% and 0.3% respectively. The mechanism of the loss optimization is analyzed in detail.
{"title":"A Novel Airfoil Design Method and its Application to the Optimization Design of a 3-Stage Transonic Compressor","authors":"Xiao-chen Wang, Xiao-dong Ren, Xue-song Li, Hong Wu, C. Gu","doi":"10.1115/gt2021-59231","DOIUrl":"https://doi.org/10.1115/gt2021-59231","url":null,"abstract":"\u0000 In this paper, a novel airfoil design method is proposed by introducing multiple physical characteristics to the process of parameterization. The new design method is to operate the camber and thickness distributions in a way that the chordwise mapping information is adjusted without using local geometrical parameters. A proper design space is expected to be constructed, in which the laws of optimization can be simply transmitted from a single airfoil to the whole blade and even stages. The capability of the method is firstly validated in the Pareto optimization of a 2D compressor airfoil, and then specialized for the optimization of a rotor in a 3-stage transonic compressor. The 3-stage transonic compressor studied in this paper is representative of the inlet stages of an industrial gas turbine, and the shock structure of the third rotor (R3) presents the peak strength near the mid span. The optimized design reduces the peak Mach number on the suction surface, and increases the efficiency of the third stage and the whole compressor by nearly 1% and 0.3% respectively. The mechanism of the loss optimization is analyzed in detail.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"23 2 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"123437120","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}
Gonçalo Cruz, Cedric Babin, X. Ottavy, F. Fontaneto
As the next generation of turbomachinery components becomes more sensitive to instrumentation intrusiveness, a reduction of the number of measurement devices required for the evaluation of performance is a possible and cost-effective way to mitigate the arising of non-mastered experimental errors. A first approach to a data assimilation methodology based on Bayesian inference is developed with the aim of reducing the instrumentation effort. A numerical model is employed to provide an initial belief of the flow, that is then updated based on experimental observations, using an ensemble Kalman filter algorithm for inverse problems. Validation of the algorithm is achieved with the usage of experimental measurements not used in the data assimilation process. The methodology is tested for a low aspect ratio axial compressor stage, showing a good prediction of the corrected compressor map, as well as a promising prediction of the inter-row radial pressure distribution and 2D flow field.
{"title":"Bayesian Inference of Experimental Data for Axial Compressor Performance Assessment","authors":"Gonçalo Cruz, Cedric Babin, X. Ottavy, F. Fontaneto","doi":"10.1115/gt2021-59200","DOIUrl":"https://doi.org/10.1115/gt2021-59200","url":null,"abstract":"\u0000 As the next generation of turbomachinery components becomes more sensitive to instrumentation intrusiveness, a reduction of the number of measurement devices required for the evaluation of performance is a possible and cost-effective way to mitigate the arising of non-mastered experimental errors. A first approach to a data assimilation methodology based on Bayesian inference is developed with the aim of reducing the instrumentation effort. A numerical model is employed to provide an initial belief of the flow, that is then updated based on experimental observations, using an ensemble Kalman filter algorithm for inverse problems. Validation of the algorithm is achieved with the usage of experimental measurements not used in the data assimilation process. The methodology is tested for a low aspect ratio axial compressor stage, showing a good prediction of the corrected compressor map, as well as a promising prediction of the inter-row radial pressure distribution and 2D flow field.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"90 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122655096","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 Reynolds number effects on performance of a Controlled Diffusion Airfoil (CDA) geometry in a linear compressor cascade setup are presented. A set of highly-resolved Large Eddy Simulations at Reynolds numbers from 300,000 to 1,000,000 at a Mach number of 0.7 are performed, with one of the simulations including incident wakes at engine relevant reduced frequency of 1.56. The focus of this work is to investigate loss variation of a modern blade geometry by applying both the novel entropy analysis introduced by (Zhao & Sandberg, GT2019-90126) and the mechanical work potential analysis by Miller (GT2013-95488) in the form presented by Leggett et al. (GT2019-91253). The loss analysis, applied both to the whole blade, and to carefully selected subvolumes, including suction and pressure surface boundary layers, highlights regions of increased loss production and the effects of Reynolds number on the location and strength of different flow processes, such as transition and separation. The results are further broken down to show the contribution to loss of individual terms of the analyses and the changes that occur due to changes in Reynolds number. The effects of incident wakes are also captured using this analysis technique and detailed in the paper.
{"title":"Highly Resolved Simulations of a CDA Compressor Cascade: Effect of Reynolds Number on Losses","authors":"J. Leggett, R. Sandberg","doi":"10.1115/gt2021-58665","DOIUrl":"https://doi.org/10.1115/gt2021-58665","url":null,"abstract":"\u0000 The Reynolds number effects on performance of a Controlled Diffusion Airfoil (CDA) geometry in a linear compressor cascade setup are presented. A set of highly-resolved Large Eddy Simulations at Reynolds numbers from 300,000 to 1,000,000 at a Mach number of 0.7 are performed, with one of the simulations including incident wakes at engine relevant reduced frequency of 1.56. The focus of this work is to investigate loss variation of a modern blade geometry by applying both the novel entropy analysis introduced by (Zhao & Sandberg, GT2019-90126) and the mechanical work potential analysis by Miller (GT2013-95488) in the form presented by Leggett et al. (GT2019-91253). The loss analysis, applied both to the whole blade, and to carefully selected subvolumes, including suction and pressure surface boundary layers, highlights regions of increased loss production and the effects of Reynolds number on the location and strength of different flow processes, such as transition and separation. The results are further broken down to show the contribution to loss of individual terms of the analyses and the changes that occur due to changes in Reynolds number. The effects of incident wakes are also captured using this analysis technique and detailed in the paper.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"45 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"131295264","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. Popov, M. Miheev, A. Vorobyev, O. Baturin, V. Zubanov
The paper describes the process of gas-dynamic modernization of a 16-stage axial compressor of an industrial gas turbine unit. Tests of the baseline variant of the compressor revealed a significant shortfall of efficiency, pressure ratio, and stability margins. In addition, the ongoing work on the modernization of the entire engine sets the task to the authors of not just achieving design parameters but significantly exceeding them (air mass flow rate by 6%, pressure ratio by 2%, adiabatic efficiency by 1% relative to the design values). To achieve these goals, a numerical model of the compressor was developed and validated. The characteristics obtained with its help were carefully analyzed. It was found that the front stage group has low efficiency, and the rear stage group is significantly oversized in terms of mass flow rate. Modernization works were significantly hampered by the presence of many stages and many independent variables. For this reason, the problem was solved in several stages. A separate modernization of the first and rear groups of stages was performed. Moreover, methods of mathematical optimization were used when developing the rear block of 10 stages. Then the working processes of the compressor parts were matched. As a result of the research, a variant was found to modernize the existing 16-stage axial compressor, providing an increase in the air mass flow rate by 18%, adiabatic efficiency by 3.5%, and margins of gas-dynamic stability up to 16%.
{"title":"Comprehensive Performance Improvement for a 16-Stage Axial Compressor","authors":"G. Popov, M. Miheev, A. Vorobyev, O. Baturin, V. Zubanov","doi":"10.1115/gt2021-59589","DOIUrl":"https://doi.org/10.1115/gt2021-59589","url":null,"abstract":"\u0000 The paper describes the process of gas-dynamic modernization of a 16-stage axial compressor of an industrial gas turbine unit. Tests of the baseline variant of the compressor revealed a significant shortfall of efficiency, pressure ratio, and stability margins. In addition, the ongoing work on the modernization of the entire engine sets the task to the authors of not just achieving design parameters but significantly exceeding them (air mass flow rate by 6%, pressure ratio by 2%, adiabatic efficiency by 1% relative to the design values).\u0000 To achieve these goals, a numerical model of the compressor was developed and validated. The characteristics obtained with its help were carefully analyzed. It was found that the front stage group has low efficiency, and the rear stage group is significantly oversized in terms of mass flow rate.\u0000 Modernization works were significantly hampered by the presence of many stages and many independent variables. For this reason, the problem was solved in several stages. A separate modernization of the first and rear groups of stages was performed. Moreover, methods of mathematical optimization were used when developing the rear block of 10 stages. Then the working processes of the compressor parts were matched.\u0000 As a result of the research, a variant was found to modernize the existing 16-stage axial compressor, providing an increase in the air mass flow rate by 18%, adiabatic efficiency by 3.5%, and margins of gas-dynamic stability up to 16%.","PeriodicalId":257596,"journal":{"name":"Volume 2A: Turbomachinery — Axial Flow Fan and Compressor Aerodynamics","volume":"42 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126875335","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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