H. Seehausen, Philipp Gilge, A. Kellersmann, J. Friedrichs, Florian Herbst
The objective of this study is to quantify the sensitivity of blade roughness on the overall performance of a 10-stage high-pressure compressor of the jet engine type V2500-A1. The Reynolds-Aver-aged Navier-Stokes flow solver TRACE is used to study the multistage compressor. The three-dimensional numerical setup contains all geometric and aerodynamic features such as bleed ports and the variable stator vanes system. In order to estimate the effect of stage roughness on overall compressor performance, compressor maps of the CFD-model are created by modeling the surface rough- ness separately for a single stage and combinations of stages. The surface roughness values are applied to the blade’s suction side of the first, center and last stage in the CFD-model by setting an equivalent sand-grain value. This equivalent sand-grain roughness is determined from non-intrusive measurements of blade surfaces from an equivalent real aircraft engine for the first, center and last stage. In addition, further simulations are conducted to analyze the performance drop of a fully rough HPC due to surface roughness. The studies are performed at the operating conditions ‘cruise’ and ‘take-off’ to cover two different Reynolds number regimes. The results show that the models with roughness in a single stage already lead to significantly lower mass flow rates because of higher blockage compared to the smooth compressor. In fact, roughness at the first stage has the biggest effect on the overall performance with a drop in performance of about 0.1% while the effect of the last stage is the smallest. This behavior is mainly caused by enhanced instabilities through the compressor changing the stage-by-stage match-ing of the stages downstream. In addition to the displacement of the compressor maps to a lower mass flow, a reduction of stall and choke margins is noticeable.
本研究的目的是量化叶片粗糙度对V2500-A1型喷气发动机10级高压压气机整体性能的敏感性。采用reynolds - average -age Navier-Stokes流动求解器TRACE对多级压气机进行了研究。三维数值设置包含所有几何和空气动力学特征，如排气口和可变定子叶片系统。为了估计一级粗糙度对压缩机整体性能的影响，通过对单级和多级组合的表面粗糙度分别建模，建立了cfd模型的压缩机图。通过设置等效砂粒值，将表面粗糙度值应用于cfd模型中第一、中心和最后一级叶片吸力侧。该等效砂粒粗糙度是通过对等效真实飞机发动机叶片表面的第一、中、末级进行非侵入式测量确定的。此外，还通过仿真分析了表面粗糙度对全粗糙HPC性能的影响。研究是在“巡航”和“起飞”操作条件下进行的，以涵盖两种不同的雷诺数制度。结果表明，与光滑压气机相比，单级粗糙压气机由于堵塞较大，导致质量流量明显降低。事实上，第一阶段的粗糙度对整体性能的影响最大，性能下降约0.1%，而最后阶段的影响最小。这种行为主要是由于压气机改变了下游各级的逐级匹配，从而增强了不稳定性。除了压缩机的位移映射到更低的质量流量之外，失速和节流余量的减少也是显而易见的。
{"title":"Numerical Study of Stage Roughness Variations in a High Pressure Compressor","authors":"H. Seehausen, Philipp Gilge, A. Kellersmann, J. Friedrichs, Florian Herbst","doi":"10.38036/jgpp.11.3_16","DOIUrl":"https://doi.org/10.38036/jgpp.11.3_16","url":null,"abstract":"The objective of this study is to quantify the sensitivity of blade roughness on the overall performance of a 10-stage high-pressure compressor of the jet engine type V2500-A1. The Reynolds-Aver-aged Navier-Stokes flow solver TRACE is used to study the multistage compressor. The three-dimensional numerical setup contains all geometric and aerodynamic features such as bleed ports and the variable stator vanes system. In order to estimate the effect of stage roughness on overall compressor performance, compressor maps of the CFD-model are created by modeling the surface rough- ness separately for a single stage and combinations of stages. The surface roughness values are applied to the blade’s suction side of the first, center and last stage in the CFD-model by setting an equivalent sand-grain value. This equivalent sand-grain roughness is determined from non-intrusive measurements of blade surfaces from an equivalent real aircraft engine for the first, center and last stage. In addition, further simulations are conducted to analyze the performance drop of a fully rough HPC due to surface roughness. The studies are performed at the operating conditions ‘cruise’ and ‘take-off’ to cover two different Reynolds number regimes. The results show that the models with roughness in a single stage already lead to significantly lower mass flow rates because of higher blockage compared to the smooth compressor. In fact, roughness at the first stage has the biggest effect on the overall performance with a drop in performance of about 0.1% while the effect of the last stage is the smallest. This behavior is mainly caused by enhanced instabilities through the compressor changing the stage-by-stage match-ing of the stages downstream. In addition to the displacement of the compressor maps to a lower mass flow, a reduction of stall and choke margins is noticeable.","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2020-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70080063","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}
As part of the design of a new particle accelerator at CERN, a research is conducted to study the challenges and opportunities of multi-stage turbocompressor machines operating with light gases and more specifically with a mixture of helium and neon. First, a 1D stage performance prediction model is implemented and coupled with a genetic algorithm in order to generate an impeller database. Then, a stacking method is developed considering design philoso-phies and technological limitations observed in the industry. This model is coupled with a second loop of the same genetic algorithm, which provides multi-stage architectures optimised for either com-pactness, i.e. number of stages, or efficiency. For both objectives, an ideal number of stages can be determined which increases signif-icantly as the operating gas becomes lighter. The impellers diversity within the database also plays an important role on the overall machine architecture. Finally, in alignment with potential technological improvements, the motor maximum rotational speed is varied to study the achievable reduction in the required number of stages.
{"title":"Impeller design and multi-stage architecture optimisation for turbocompressors operating with a helium-neon gas mixture","authors":"M. Podeur, D. Vogt, S. Mauri, P. Jenny","doi":"10.38036/jgpp.11.4_1","DOIUrl":"https://doi.org/10.38036/jgpp.11.4_1","url":null,"abstract":"As part of the design of a new particle accelerator at CERN, a research is conducted to study the challenges and opportunities of multi-stage turbocompressor machines operating with light gases and more specifically with a mixture of helium and neon. First, a 1D stage performance prediction model is implemented and coupled with a genetic algorithm in order to generate an impeller database. Then, a stacking method is developed considering design philoso-phies and technological limitations observed in the industry. This model is coupled with a second loop of the same genetic algorithm, which provides multi-stage architectures optimised for either com-pactness, i.e. number of stages, or efficiency. For both objectives, an ideal number of stages can be determined which increases signif-icantly as the operating gas becomes lighter. The impellers diversity within the database also plays an important role on the overall machine architecture. Finally, in alignment with potential technological improvements, the motor maximum rotational speed is varied to study the achievable reduction in the required number of stages.","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2020-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70080117","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}
Robert Krewinkel, Anders Such, A. D. Torre, A. Wiedermann, Daniel Castillo, Silvia Araguas Rodriguez, J. Schleifenbaum, M. Blaswich
The use of additive manufacturing (AM), for example Selective Laser Melting (SLM), is poised to spark a revolution in the way high-temperature components for gas turbines are designed, but a number of grave uncertainties remain. These lie mainly with the materials sciences, but some questions with regard to manufacturing and operating SLM-parts as hot gas path components and the demands on the tolerances of the cooling features associated therewith remain as well. In order to quantify the impact of these uncertainties, Nozzle Guide Vanes (NGVs) with a geometry that would normally be investment-cast were produced with SLM. A back-to-back comparison of vanes from the two manufacturing processes was performed. The design of the SLM-vanes will be described and the SLM-manufacturing process of the NGVs will be touched upon, especially the use of MAR M-509, which is seldom used for SLM. In addition, characterization of the NGVs with 3D-scans of the outer geometry and the pin-fin matrix shall be discussed. The NGVs were operated for approximately 70 hours at relevant load conditions in a highly-instrumented test engine on a test bed at the Oberhausen plant of MAN. The temperatures of the AM and investment-cast vanes were measured using Thermal History Paints (THPs); a comparison between these different kinds of parts will be drawn.
{"title":"Design and Characterization of Additively Manufactured NGVs Operated in a Small Industrial Gas Turbine","authors":"Robert Krewinkel, Anders Such, A. D. Torre, A. Wiedermann, Daniel Castillo, Silvia Araguas Rodriguez, J. Schleifenbaum, M. Blaswich","doi":"10.38036/jgpp.11.4_36","DOIUrl":"https://doi.org/10.38036/jgpp.11.4_36","url":null,"abstract":"The use of additive manufacturing (AM), for example Selective Laser Melting (SLM), is poised to spark a revolution in the way high-temperature components for gas turbines are designed, but a number of grave uncertainties remain. These lie mainly with the materials sciences, but some questions with regard to manufacturing and operating SLM-parts as hot gas path components and the demands on the tolerances of the cooling features associated therewith remain as well. In order to quantify the impact of these uncertainties, Nozzle Guide Vanes (NGVs) with a geometry that would normally be investment-cast were produced with SLM. A back-to-back comparison of vanes from the two manufacturing processes was performed. The design of the SLM-vanes will be described and the SLM-manufacturing process of the NGVs will be touched upon, especially the use of MAR M-509, which is seldom used for SLM. In addition, characterization of the NGVs with 3D-scans of the outer geometry and the pin-fin matrix shall be discussed. The NGVs were operated for approximately 70 hours at relevant load conditions in a highly-instrumented test engine on a test bed at the Oberhausen plant of MAN. The temperatures of the AM and investment-cast vanes were measured using Thermal History Paints (THPs); a comparison between these different kinds of parts will be drawn.","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2020-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70080225","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}
Ayaka Kamiyama, A. Murata, Shohei Yamamoto, K. Iwamoto, Y. Okita
Electrification of aircraft has been realizing improvement in efficiency, reliability, and safety of the aircraft by substituting hydraulic and mechanical system with electric system. On the other hand, in future electric aircraft partially replacing the fan driving engines with motors, the thermal management of heat generation from the electric system will become crucial problem to be solved. In this study, for the future electric aircraft, thermal management in oil-cooling motors and air-cooling motor controllers was considered. Three-dimensional steady thermal network analysis (TNA) was performed for analyzing temperature field in an oil cooler for motor cooling and a heat sink for motor controller air-cooling. The present numerical procedure was verified by comparing the results of TNA with those of three-dimensional fluid-solid conjugate heat transfer analysis (3D-CFD). After the verification, TNA was performed for several aircraft flight scenarios, and the optimum geometry of the oil cooler and the heat sink was investigated under the constraints of allowable pressure loss of air flow and outlet oil temperature for oil cooler or maximum local wall temperature for heat sink using the weight (mass) as the object function to be minimized. Furthermore, the water-mist injection to the air flow was considered for lowering the air temperature and the weight of the heat sink.
{"title":"Three-dimensional Thermal Network Analysis of Multi-stage Heat Sink with Water-mist Injection Applied to Thermal Management of Electric Aircraft","authors":"Ayaka Kamiyama, A. Murata, Shohei Yamamoto, K. Iwamoto, Y. Okita","doi":"10.38036/jgpp.11.4_45","DOIUrl":"https://doi.org/10.38036/jgpp.11.4_45","url":null,"abstract":"Electrification of aircraft has been realizing improvement in efficiency, reliability, and safety of the aircraft by substituting hydraulic and mechanical system with electric system. On the other hand, in future electric aircraft partially replacing the fan driving engines with motors, the thermal management of heat generation from the electric system will become crucial problem to be solved. In this study, for the future electric aircraft, thermal management in oil-cooling motors and air-cooling motor controllers was considered. Three-dimensional steady thermal network analysis (TNA) was performed for analyzing temperature field in an oil cooler for motor cooling and a heat sink for motor controller air-cooling. The present numerical procedure was verified by comparing the results of TNA with those of three-dimensional fluid-solid conjugate heat transfer analysis (3D-CFD). After the verification, TNA was performed for several aircraft flight scenarios, and the optimum geometry of the oil cooler and the heat sink was investigated under the constraints of allowable pressure loss of air flow and outlet oil temperature for oil cooler or maximum local wall temperature for heat sink using the weight (mass) as the object function to be minimized. Furthermore, the water-mist injection to the air flow was considered for lowering the air temperature and the weight of the heat sink.","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2020-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70080342","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper presents an analysis of the vibration-induced effects on the aerofoil aerodynamics and boundary-layer development of a low-pressure–turbine blade. Large-eddy simulations of an MTU-T161 low-pressure–turbine blade with imposed sinusoidal rigid-body oscillations were conducted for frequencies of 50 and 100 Hz as well as for a fixed reference blade. The oscillations are shown to impact both the time-averaged flow field and unsteady velocity fluctuations. These changes appear most markedly as a reduction in the stagnation-point pressure and a par-tial suppression of the separation bubble on the suction side of the aerofoil. The results suggest that the deterministic velocity fluctuations introduced by the oscillating blade promote transition on the suction side and expedite the generation of turbulence. is presented for the investigation of vibration-induced effects on the aerofoil aerodynamics and boundary-layer development of low-pressure–turbine blades. To achieve this goal, large-eddy simulations of an MTU-T161 LPT profile with imposed sinusoidal rigid-body oscillations are analysed. The oscillations are shown to impact both the time-averaged flow field and the unsteady fluctuations. use of periodic boundary conditions normal to the blade motion and allows the study of oscillation-induced effects on the profile boundary layer.
{"title":"Profile Aerodynamics of an Oscillating Low-Pressure–Turbine Blade","authors":"Felix Schwarzbach, Dajan Mimic, Florian Herbst","doi":"10.38036/jgpp.11.4_68","DOIUrl":"https://doi.org/10.38036/jgpp.11.4_68","url":null,"abstract":"This paper presents an analysis of the vibration-induced effects on the aerofoil aerodynamics and boundary-layer development of a low-pressure–turbine blade. Large-eddy simulations of an MTU-T161 low-pressure–turbine blade with imposed sinusoidal rigid-body oscillations were conducted for frequencies of 50 and 100 Hz as well as for a fixed reference blade. The oscillations are shown to impact both the time-averaged flow field and unsteady velocity fluctuations. These changes appear most markedly as a reduction in the stagnation-point pressure and a par-tial suppression of the separation bubble on the suction side of the aerofoil. The results suggest that the deterministic velocity fluctuations introduced by the oscillating blade promote transition on the suction side and expedite the generation of turbulence. is presented for the investigation of vibration-induced effects on the aerofoil aerodynamics and boundary-layer development of low-pressure–turbine blades. To achieve this goal, large-eddy simulations of an MTU-T161 LPT profile with imposed sinusoidal rigid-body oscillations are analysed. The oscillations are shown to impact both the time-averaged flow field and the unsteady fluctuations. use of periodic boundary conditions normal to the blade motion and allows the study of oscillation-induced effects on the profile boundary layer.","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2020-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70080358","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 this paper, an alternative micro-gas turbine is proposed, where the traditional compressor-turbine arrangement is replaced by an axial, throughflow wave rotor. The investigated wave rotor features symmetrically cambered wall profiles and angled port arrangement for shaft power extraction and uses shock and rarefaction waves for pressure exchange and to achieve gas compression and expansion within a single device. A validated quasi-one-dimensional model that solves the laminar Navier-Stokes equations using a two-step Richtmyer scheme with minmod flux limiter is employed to characterise and examine microgas turbine behaviour. The model accounts for wall heat transfer, flow leakage, wall friction and inviscid blade forces. In addition, modified boundary conditions consider finite passage opening effects and a simple steady-flow combustor model is defined that links the high pressure inand outlet ports. The model is used to conduct a parametric study to investigate the effects of leakage gap, heat release rate, exhaust backpressure, as well as profile camber on gas turbine performance with a focus on generated combustor compression and expansion efficiency, shaft power and system efficiency. The implications of combustor pressure loss as well as effects of a potential recuperator are discussed as well. The results identify axial leakage and combustor pressure loss as primary drivers for enhanced performance. Finally, the results reinforce the capacity of wave rotors to compress and expand gas efficiently, while thermal efficiency remains below 10 percent.
{"title":"Parametric Numerical Study on the Performance Characteristics of a Micro-Wave Rotor Gas Turbine","authors":"Stefan Tuechler, C. Copeland","doi":"10.38036/JGPP.12.1_10","DOIUrl":"https://doi.org/10.38036/JGPP.12.1_10","url":null,"abstract":"In this paper, an alternative micro-gas turbine is proposed, where the traditional compressor-turbine arrangement is replaced by an axial, throughflow wave rotor. The investigated wave rotor features symmetrically cambered wall profiles and angled port arrangement for shaft power extraction and uses shock and rarefaction waves for pressure exchange and to achieve gas compression and expansion within a single device. A validated quasi-one-dimensional model that solves the laminar Navier-Stokes equations using a two-step Richtmyer scheme with minmod flux limiter is employed to characterise and examine microgas turbine behaviour. The model accounts for wall heat transfer, flow leakage, wall friction and inviscid blade forces. In addition, modified boundary conditions consider finite passage opening effects and a simple steady-flow combustor model is defined that links the high pressure inand outlet ports. The model is used to conduct a parametric study to investigate the effects of leakage gap, heat release rate, exhaust backpressure, as well as profile camber on gas turbine performance with a focus on generated combustor compression and expansion efficiency, shaft power and system efficiency. The implications of combustor pressure loss as well as effects of a potential recuperator are discussed as well. The results identify axial leakage and combustor pressure loss as primary drivers for enhanced performance. Finally, the results reinforce the capacity of wave rotors to compress and expand gas efficiently, while thermal efficiency remains below 10 percent.","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2019-05-12","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44361598","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}
{"title":"FLOW AND HEAT TRANSFER IN A RIB-ROUGHENED TRAILING-EDGE COOLING CHANNEL WITH CROSSOVER IMPINGEMENT","authors":"F. Xue, M. Taslim","doi":"10.38036/jgpp.10.1_1","DOIUrl":"https://doi.org/10.38036/jgpp.10.1_1","url":null,"abstract":"","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2019-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70079724","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}
Soufiane Ramdani, N. Yamasaki, Y. Inokuchi, T. Ishii
Experimental and computational studies are performed on slit resonators, i.e., one aperture with a high aspect ratio that spans through the center of the plate. An impedance tube experiment is conducted to investigate the macroscopic response of slit apertures. The test specimens include straight and tapered apertures. Subsequently, the effect of introducing a bias flow is investigated. The absorption performance increases when the sound pressure level increases to a level that causes a production of shed vortices. When a bias flow is introduced, the absorption coefficient reaches its maximum absorption of the incident sound wave in the region near the resonant frequency for a Mach number close to 9.71 . 2D numerical simulations are performed and validated with the experimental results. Good agreement is obtained for the majority of the simulated cases. Vortex shedding and its effect on the absorption coefficients is also investigated. INTRODUCTION Acoustic liners are widely used devices in aero-engines, and they are installed on the inner side of the nacelle to reduce fan noise in turbofan engines of commercial airplanes (conventional acoustic liners without bias flow). They are also used in the combustion chamber to reduce the acoustic instabilities caused by the combustion (acoustic liners with bias flow where the bias flow is caused by secondary air flow). An acoustic liner is typically made of a perforated metal sheet backed by a cavity. Each aperture of the perforated sheet and the cavity form a Helmholtz resonator. The resonator effectively absorbs the sound near the resonant frequency, however, its absorbing performance decreases at off-resonant frequencies. Howe [1] theoretically proposed that a low frequency (low Strouhal number) sound wave can be significantly attenuated by a jet flow by converting the acoustical energy into energy of fluctuating vorticity, which is shed from the nozzle edge. Bechert [2] proposed another theory to explain this phenomenon, and this was supported via experimental data. Bechert [2] also proposed a simple theory to predict the optimum Mach number of bias flow to obtain the perfect attenuation. On the other hand, Howe's theory to predict the sound absorption coefficient including the effects of a bias flow is well supported by an experiment by Hughes and Dowling [3]. Hence, this led to the idea that the off-resonant performance of a resonator can be improved if a jet (or a bias flow) is introduced from an aperture of an acoustic liner. Lahiri et al. [4] collected this type of experimental data and showed that the application of a bias flow through the aperture widens the frequency range of dissipation, with the penalty of reduced peak performance near the resonant frequency. Zhao and Li [5] wrote a summary on tunable acoustic liners including a liner with bias flow. In the field of numerical simulation, Mendez and Eldredge [6] performed a 3D large eddy simulation (LES), and Ji and Zhao [7] performed a 2D la
{"title":"Bias Flow Acoustic Liner with Straight and Tapered Apertures","authors":"Soufiane Ramdani, N. Yamasaki, Y. Inokuchi, T. Ishii","doi":"10.38036/jgpp.10.3_1","DOIUrl":"https://doi.org/10.38036/jgpp.10.3_1","url":null,"abstract":"Experimental and computational studies are performed on slit resonators, i.e., one aperture with a high aspect ratio that spans through the center of the plate. An impedance tube experiment is conducted to investigate the macroscopic response of slit apertures. The test specimens include straight and tapered apertures. Subsequently, the effect of introducing a bias flow is investigated. The absorption performance increases when the sound pressure level increases to a level that causes a production of shed vortices. When a bias flow is introduced, the absorption coefficient reaches its maximum absorption of the incident sound wave in the region near the resonant frequency for a Mach number close to 9.71 . 2D numerical simulations are performed and validated with the experimental results. Good agreement is obtained for the majority of the simulated cases. Vortex shedding and its effect on the absorption coefficients is also investigated. INTRODUCTION Acoustic liners are widely used devices in aero-engines, and they are installed on the inner side of the nacelle to reduce fan noise in turbofan engines of commercial airplanes (conventional acoustic liners without bias flow). They are also used in the combustion chamber to reduce the acoustic instabilities caused by the combustion (acoustic liners with bias flow where the bias flow is caused by secondary air flow). An acoustic liner is typically made of a perforated metal sheet backed by a cavity. Each aperture of the perforated sheet and the cavity form a Helmholtz resonator. The resonator effectively absorbs the sound near the resonant frequency, however, its absorbing performance decreases at off-resonant frequencies. Howe [1] theoretically proposed that a low frequency (low Strouhal number) sound wave can be significantly attenuated by a jet flow by converting the acoustical energy into energy of fluctuating vorticity, which is shed from the nozzle edge. Bechert [2] proposed another theory to explain this phenomenon, and this was supported via experimental data. Bechert [2] also proposed a simple theory to predict the optimum Mach number of bias flow to obtain the perfect attenuation. On the other hand, Howe's theory to predict the sound absorption coefficient including the effects of a bias flow is well supported by an experiment by Hughes and Dowling [3]. Hence, this led to the idea that the off-resonant performance of a resonator can be improved if a jet (or a bias flow) is introduced from an aperture of an acoustic liner. Lahiri et al. [4] collected this type of experimental data and showed that the application of a bias flow through the aperture widens the frequency range of dissipation, with the penalty of reduced peak performance near the resonant frequency. Zhao and Li [5] wrote a summary on tunable acoustic liners including a liner with bias flow. In the field of numerical simulation, Mendez and Eldredge [6] performed a 3D large eddy simulation (LES), and Ji and Zhao [7] performed a 2D la","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2019-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70080407","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}
Power output and efficiency of gas turbines depend strongly upon the achievable pressure rise in the subsequent diffuser. In combination with the requirement to keep diffuser length to a minimum, ever steeper opening angles are sought, while avoiding diffuser stall. In terms of diffuser pressure rise, the boundaries of what is achievable can be pushed further if the tip leakage vortices from the last stage are used to re-accelerate the diffuser boundary layer, thus delaying separation onset. Such measures have been shown to decrease total pressure losses as well. In this paper, we show that the benefit of total pressure loss reduction in vortex-stabilised diffusers becomes more pronounced for steeper opening angles by means of a numerically and experimentally validated approach. In extension, we provide evidence that the loss production in highly loaded vortex-stabilised diffusers, which would stall otherwise, can be brought down to the level of non-stalling diffusers. Furthermore, we present a detailed analysis of the different loss mechanisms and their response to vortex-stabilisation of the diffuser. NOMENCLATURE Symbols � cross-sectional area of the diffuser AR area ratio of the diffuser �, � flow velocity �� pressure recovery coefficient �� specific isobaric heat capacity �r reduced frequency h enthalpy (default: static) l chord length meridional coordinate number of blades � rotational speed in revolutions per minute � pressure (default: static) Euler radius � specific gas constant � temperature (default: static) � rotational velocity � generalised spatial coordinate � axial coordinate � flow angle, whirl angle curve diffuser half-opening angle Δ difference diffuser effectiveness total pressure loss coefficient circumferential coordinate Lamé constant Λ loss rectification number dynamic viscosity � kinetic energy coefficient � rectified total pressure loss coefficient � density � generalised spatial vector Σ stabilisation number , Ψ flow coefficient, loading coefficient Subscripts I, II rotor inlet/outlet plane eff effective corr correlated in, out diffuser inlet/outlet ref reference rel relative t turbulent quantity tot total quantity enthalpy-induced dilatational shearing-induced thermodynamic � vorticity-induced INTRODUCTION Without the use of exhaust diffusers, the expansion of hot gas achievable in turbines is bounded by the ambient pressure. Only the subsequent conversion of kinetic energy into static pressure, realised by an increase in cross-sectional area in the diffuser, allows for considerably higher expansion ratios in the turbine. As a consequence, the power output and—assuming constant heat input—efficiency increase. The resulting main aerodynamic design goal of exhaust gas diffusers is to convert as much kinetic energy as possible into static pressure, i.e., maximise the ratio of the static pressure rise over the diffuser to the kinetic energy at diffuser inlet. Diffuser designers tend to call this ratio pressure recovery and
{"title":"Total Pressure Loss Reduction in Annular Diffusers","authors":"Dajan Mimic, C. Jätz, P. Sauer, Florian Herbst","doi":"10.38036/jgpp.10.2_1","DOIUrl":"https://doi.org/10.38036/jgpp.10.2_1","url":null,"abstract":"Power output and efficiency of gas turbines depend strongly upon the achievable pressure rise in the subsequent diffuser. In combination with the requirement to keep diffuser length to a minimum, ever steeper opening angles are sought, while avoiding diffuser stall. In terms of diffuser pressure rise, the boundaries of what is achievable can be pushed further if the tip leakage vortices from the last stage are used to re-accelerate the diffuser boundary layer, thus delaying separation onset. Such measures have been shown to decrease total pressure losses as well. In this paper, we show that the benefit of total pressure loss reduction in vortex-stabilised diffusers becomes more pronounced for steeper opening angles by means of a numerically and experimentally validated approach. In extension, we provide evidence that the loss production in highly loaded vortex-stabilised diffusers, which would stall otherwise, can be brought down to the level of non-stalling diffusers. Furthermore, we present a detailed analysis of the different loss mechanisms and their response to vortex-stabilisation of the diffuser. NOMENCLATURE Symbols � cross-sectional area of the diffuser AR area ratio of the diffuser �, � flow velocity �� pressure recovery coefficient �� specific isobaric heat capacity �r reduced frequency h enthalpy (default: static) l chord length meridional coordinate number of blades � rotational speed in revolutions per minute � pressure (default: static) Euler radius � specific gas constant � temperature (default: static) � rotational velocity � generalised spatial coordinate � axial coordinate � flow angle, whirl angle curve diffuser half-opening angle Δ difference diffuser effectiveness total pressure loss coefficient circumferential coordinate Lamé constant Λ loss rectification number dynamic viscosity � kinetic energy coefficient � rectified total pressure loss coefficient � density � generalised spatial vector Σ stabilisation number , Ψ flow coefficient, loading coefficient Subscripts I, II rotor inlet/outlet plane eff effective corr correlated in, out diffuser inlet/outlet ref reference rel relative t turbulent quantity tot total quantity enthalpy-induced dilatational shearing-induced thermodynamic � vorticity-induced INTRODUCTION Without the use of exhaust diffusers, the expansion of hot gas achievable in turbines is bounded by the ambient pressure. Only the subsequent conversion of kinetic energy into static pressure, realised by an increase in cross-sectional area in the diffuser, allows for considerably higher expansion ratios in the turbine. As a consequence, the power output and—assuming constant heat input—efficiency increase. The resulting main aerodynamic design goal of exhaust gas diffusers is to convert as much kinetic energy as possible into static pressure, i.e., maximise the ratio of the static pressure rise over the diffuser to the kinetic energy at diffuser inlet. Diffuser designers tend to call this ratio pressure recovery and","PeriodicalId":38948,"journal":{"name":"International Journal of Gas Turbine, Propulsion and Power Systems","volume":"1 1","pages":""},"PeriodicalIF":0.0,"publicationDate":"2019-01-01","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"70079775","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}
京公网安备 11010802042870号
京ICP备2023020795号-1
扫码关注我们
求助内容：
应助结果提醒方式：