Naren Shankar Radha Krishnan, Sathish Kumar Sarath Kumar, Vijayaraja Kengaiah, S. Chidambaram, E. Rathakrishnan
Abstract The characteristics of a sonic under-expanded coaxial jet with lip thickness 1.5D p (where D p is the exit diameter of primary jet equals 10 mm) with the primary jet operating at nozzle pressure ratio (NPR) of 3, 4 and 5. For NPR 3 operating primary jet, the secondary jet operates at NPR 2.5, 1.4 and 1.27. For primary jet NPR 4, the secondary jet operating NPR is 3.2, 1.6 and 1.4. For NPR 5 primary jet, the secondary jet NPR is 3.8, 1.89 and 1.52. The study is performed using a co-flow nozzle of bypass ratio (BR) 6.4, 1.4 and 0.7. The core length of the primary jet is used as a measure to quantify the mixing of the primary jet in the presence of coaxial jet. The shock structure present in the near field was viewed using shadowgraph technique. Centreline pitot pressure distribution, radial spread and waves present in the jet core were analyzed. The results show that the mixing associated with the high bypass coaxial jet is superior to the low bypass coaxial jet. This mixing superiority associated with high bypass coaxial jet prevails all levels of expansion.
{"title":"Effect of bypass ratio on sonic underexpanded co-flow jets with finite lip thickness","authors":"Naren Shankar Radha Krishnan, Sathish Kumar Sarath Kumar, Vijayaraja Kengaiah, S. Chidambaram, E. Rathakrishnan","doi":"10.1515/tjj-2022-0068","DOIUrl":"https://doi.org/10.1515/tjj-2022-0068","url":null,"abstract":"Abstract The characteristics of a sonic under-expanded coaxial jet with lip thickness 1.5D p (where D p is the exit diameter of primary jet equals 10 mm) with the primary jet operating at nozzle pressure ratio (NPR) of 3, 4 and 5. For NPR 3 operating primary jet, the secondary jet operates at NPR 2.5, 1.4 and 1.27. For primary jet NPR 4, the secondary jet operating NPR is 3.2, 1.6 and 1.4. For NPR 5 primary jet, the secondary jet NPR is 3.8, 1.89 and 1.52. The study is performed using a co-flow nozzle of bypass ratio (BR) 6.4, 1.4 and 0.7. The core length of the primary jet is used as a measure to quantify the mixing of the primary jet in the presence of coaxial jet. The shock structure present in the near field was viewed using shadowgraph technique. Centreline pitot pressure distribution, radial spread and waves present in the jet core were analyzed. The results show that the mixing associated with the high bypass coaxial jet is superior to the low bypass coaxial jet. This mixing superiority associated with high bypass coaxial jet prevails all levels of expansion.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-11-16","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"41946926","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Kuanxin Hou, H. Xiang, Jie Gao, Binbin Lin, Qiulin Dai, Yao Fu, Yong Yang, Lei Huang
Abstract Aiming at the problem of aerodynamic coupling interference between disturbance of the built-in measurement probe and internal flow of high load compressor, the experiment of axial position and structure dimension of cylindrical probe support effect on the performance of the compressor cascade was carried out on the transonic plane cascade test facility. The effect characteristics of the distribution of probe support on aerodynamic performance of downstream cascade under different inlet Mach numbers were analyzed through the experimental data. The experimental results show that the axial position of probe support has little impact on the total pressure loss coefficient of cascade under the low-speed inlet flow condition. However, the variety of the axial position of probe support has a great impact on the total pressure loss coefficient of cascade under the high-speed inlet flow condition. With the decrease of the axial distance between probe support and cascade, the total pressure loss coefficient of cascade increases. Under different inlet flow conditions, the variety of the structure dimension of probe support has a significant impact on the total pressure loss coefficient of cascade. With the increase of the structure dimension of probe support, the total pressure loss coefficient of cascade increases.
{"title":"Effect of the distribution of the upstream probe support on performance of compressor cascade","authors":"Kuanxin Hou, H. Xiang, Jie Gao, Binbin Lin, Qiulin Dai, Yao Fu, Yong Yang, Lei Huang","doi":"10.1515/tjeng-2022-0027","DOIUrl":"https://doi.org/10.1515/tjeng-2022-0027","url":null,"abstract":"Abstract Aiming at the problem of aerodynamic coupling interference between disturbance of the built-in measurement probe and internal flow of high load compressor, the experiment of axial position and structure dimension of cylindrical probe support effect on the performance of the compressor cascade was carried out on the transonic plane cascade test facility. The effect characteristics of the distribution of probe support on aerodynamic performance of downstream cascade under different inlet Mach numbers were analyzed through the experimental data. The experimental results show that the axial position of probe support has little impact on the total pressure loss coefficient of cascade under the low-speed inlet flow condition. However, the variety of the axial position of probe support has a great impact on the total pressure loss coefficient of cascade under the high-speed inlet flow condition. With the decrease of the axial distance between probe support and cascade, the total pressure loss coefficient of cascade increases. Under different inlet flow conditions, the variety of the structure dimension of probe support has a significant impact on the total pressure loss coefficient of cascade. With the increase of the structure dimension of probe support, the total pressure loss coefficient of cascade increases.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-11-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"41446578","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Q. Wang, Zhourong Zhang, Shuning Xiao, Qingsong Hong
Abstract The variations of inlet environment parameters can make significant effects on the compressor performance. This paper aims to investigate the effects of inlet total pressure and total temperature changes on the rated condition performance of a nine-stage HPC. Different cases of total pressure and total temperature boundary conditions at this compressor inlet are studied by 3-D numerical simulations with experimental validations. The numerical results confirm that the variations of inlet total pressure and total temperature make different effects on the rated condition performance of compressor. The overall performance parameters, such as the corrected mass flow and isentropic efficiency, will increase with inlet total pressure increasing and decrease with inlet total temperature increasing by different change rules. The flow similarity is also investigated by comparing the calculated results of critical quantities in different cases. The results indicate that the rising inlet total pressure can increase the Reynolds number and it is beneficial to reduce the viscous influence so that it is available to improve the performance; the rising inlet total temperature can decrease both the specific heat ratio and Reynolds number so that it will lead to the compressor performance decline inevitably.
{"title":"Impact investigation of inlet environmental changes on the rated condition performance of a high-pressure compressor","authors":"Q. Wang, Zhourong Zhang, Shuning Xiao, Qingsong Hong","doi":"10.1515/tjj-2022-0022","DOIUrl":"https://doi.org/10.1515/tjj-2022-0022","url":null,"abstract":"Abstract The variations of inlet environment parameters can make significant effects on the compressor performance. This paper aims to investigate the effects of inlet total pressure and total temperature changes on the rated condition performance of a nine-stage HPC. Different cases of total pressure and total temperature boundary conditions at this compressor inlet are studied by 3-D numerical simulations with experimental validations. The numerical results confirm that the variations of inlet total pressure and total temperature make different effects on the rated condition performance of compressor. The overall performance parameters, such as the corrected mass flow and isentropic efficiency, will increase with inlet total pressure increasing and decrease with inlet total temperature increasing by different change rules. The flow similarity is also investigated by comparing the calculated results of critical quantities in different cases. The results indicate that the rising inlet total pressure can increase the Reynolds number and it is beneficial to reduce the viscous influence so that it is available to improve the performance; the rising inlet total temperature can decrease both the specific heat ratio and Reynolds number so that it will lead to the compressor performance decline inevitably.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-10-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44795898","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract The passive control of jets using vanes as vortex generators is studied by numerical simulation in this paper. The vanes are positioned inside the nozzle near the exit, inclined to the flow with the longitudinal direction of the jet. Two configurations namely, straight vanes (k = 0 mm−1) and curved vanes (k = 0.05 mm−1) are considered. Curvature k is defined as the reciprocal of the radius of the vanes. The blockage due to the presence of the vanes is 0.5%. The total pressure variation along the jet centreline and along the radial distance is determined from nozzle exit at a Mach number of 0.4, 0.6 and 0.8. It is found that the vanes cause faster decay of the jet, both in the near field and far field compared to the base nozzle (plain circular nozzle) and the curved vanes perform better than the straight vanes in promoting the jet mixing. A maximum of 54% reduction in jet potential core length is achieved by the curved vanes and the jet becomes asymmetrical due to the presence of the vanes inside the nozzle, as observed in the radial pressure decay plots and Mach number contours.
{"title":"Control of subsonic jets using vanes as vortex generators","authors":"Balamani Gandhinathan, Thanigaiarasu Subramanian","doi":"10.1515/tjj-2022-0062","DOIUrl":"https://doi.org/10.1515/tjj-2022-0062","url":null,"abstract":"Abstract The passive control of jets using vanes as vortex generators is studied by numerical simulation in this paper. The vanes are positioned inside the nozzle near the exit, inclined to the flow with the longitudinal direction of the jet. Two configurations namely, straight vanes (k = 0 mm−1) and curved vanes (k = 0.05 mm−1) are considered. Curvature k is defined as the reciprocal of the radius of the vanes. The blockage due to the presence of the vanes is 0.5%. The total pressure variation along the jet centreline and along the radial distance is determined from nozzle exit at a Mach number of 0.4, 0.6 and 0.8. It is found that the vanes cause faster decay of the jet, both in the near field and far field compared to the base nozzle (plain circular nozzle) and the curved vanes perform better than the straight vanes in promoting the jet mixing. A maximum of 54% reduction in jet potential core length is achieved by the curved vanes and the jet becomes asymmetrical due to the presence of the vanes inside the nozzle, as observed in the radial pressure decay plots and Mach number contours.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-10-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44626541","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract In view of the problem that the vibration of the centrifugal impeller of the compressor exceeds the standard when a micro turbojet engine crosses the first-order critical speed, this research proposes a vibration control technology for the engine rotor system based on the new integral squeeze film damper (NISFD) according to the unique structure and assembly characteristics of the rotor bearing seat of the micro turbojet engine. Through numerical analysis and experimental research, the dynamic characteristics of NISFD are analyzed, and the vibration reduction mechanism of NISFD is revealed. The engine test verification based on NISFD is carried out. The results show that NISFD has good damping characteristics, which can improve the bending stress and strain energy distribution of the rotor. The engine test proves that NISFD can well solve the problem of excessive vibration of the engine when passing through the first-order critical speed, greatly reduce the vibration level within the working speed range, and significantly improve the reliability of the engine during operation. It provides a solid theoretical basis for the design of aeroengine, and provides a reference for the practical engineering application of NISFD in aeroengine.
{"title":"Research on new integral squeeze film damper vibration control of micro turbojet engine at high speed","authors":"Wenhao Wang, Lidong He, Yangliu Sun, Zijian Wang, Hua-Liang D. Zhang, Qiaoqiao Bao, Xingyun Jia, Hua-jun Dong","doi":"10.1515/tjj-2022-0032","DOIUrl":"https://doi.org/10.1515/tjj-2022-0032","url":null,"abstract":"Abstract In view of the problem that the vibration of the centrifugal impeller of the compressor exceeds the standard when a micro turbojet engine crosses the first-order critical speed, this research proposes a vibration control technology for the engine rotor system based on the new integral squeeze film damper (NISFD) according to the unique structure and assembly characteristics of the rotor bearing seat of the micro turbojet engine. Through numerical analysis and experimental research, the dynamic characteristics of NISFD are analyzed, and the vibration reduction mechanism of NISFD is revealed. The engine test verification based on NISFD is carried out. The results show that NISFD has good damping characteristics, which can improve the bending stress and strain energy distribution of the rotor. The engine test proves that NISFD can well solve the problem of excessive vibration of the engine when passing through the first-order critical speed, greatly reduce the vibration level within the working speed range, and significantly improve the reliability of the engine during operation. It provides a solid theoretical basis for the design of aeroengine, and provides a reference for the practical engineering application of NISFD in aeroengine.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-10-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47541151","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Le Zhang, Lidong He, Xiaodong Wang, Xinwei Huang, Xingyun Jia, Chunrui Liu
Abstract Aiming at the problems of large leakage and high friction heat generation of the piston ring under high pressure and high speed, based on the previous research on the V-shaped groove piston ring, the research on the design method of the V-shaped groove piston ring is carried out in depth. The numerical analysis is carried out focusing on the number of V-shaped grooves and the depth of the V-shaped grooves of the piston ring. The influence law is verified and analyzed by using a high-pressure and high-speed rotating test bench. The simulation and experimental results show that the leakage of 4 grooves and 8 grooves piston ring is reduced by 14.9 and 28.5%, and the temperature is reduced by 10.9 and 4.7%; When there are 8 V-shaped grooves on the piston ring, the leakage is reduced by 28.5 and 21.7%; when the V-shaped groove depth is 0.04 and 0.08 mm, the temperature is reduced by 4.1%. However, when the V-shaped groove depth is 0.12 mm, the leakage is increased by 2.7% compared with the traditional piston ring. Appropriate groove number and groove depth can realize the coordinated design of low leakage and low friction temperature rise.
{"title":"The design method of the V-shaped groove piston ring","authors":"Le Zhang, Lidong He, Xiaodong Wang, Xinwei Huang, Xingyun Jia, Chunrui Liu","doi":"10.1515/tjj-2022-0033","DOIUrl":"https://doi.org/10.1515/tjj-2022-0033","url":null,"abstract":"Abstract Aiming at the problems of large leakage and high friction heat generation of the piston ring under high pressure and high speed, based on the previous research on the V-shaped groove piston ring, the research on the design method of the V-shaped groove piston ring is carried out in depth. The numerical analysis is carried out focusing on the number of V-shaped grooves and the depth of the V-shaped grooves of the piston ring. The influence law is verified and analyzed by using a high-pressure and high-speed rotating test bench. The simulation and experimental results show that the leakage of 4 grooves and 8 grooves piston ring is reduced by 14.9 and 28.5%, and the temperature is reduced by 10.9 and 4.7%; When there are 8 V-shaped grooves on the piston ring, the leakage is reduced by 28.5 and 21.7%; when the V-shaped groove depth is 0.04 and 0.08 mm, the temperature is reduced by 4.1%. However, when the V-shaped groove depth is 0.12 mm, the leakage is increased by 2.7% compared with the traditional piston ring. Appropriate groove number and groove depth can realize the coordinated design of low leakage and low friction temperature rise.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-10-18","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47995925","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Mohan Agrawal, Jai Bhan Verma, G. Joshi, S. Chandel, Ved Prakash, R. K. Mishra
Abstract This paper reports the numerical investigations to analyze the effect of the application of Gurney Flaps on various configurations in a Low-Pressure Turbine (LPT) cascade in mitigating the laminar flow separation during low Reynolds number operations. T106 LPT blade of chord 60 mm has been selected for the present study. Gurney Flaps of flat type and quarter round type of a particular height of the chord length are selected. Flaps are provided near the blades’ trailing edges for numerical analysis. The numerical computations are performed using STARCCM+ software, and the K-ω SST turbulence Model is used for turbulence closure. The studies are performed at various Reynolds numbers ranging from 37,500 to 138,750 to understand the application of various configurations of Gurney Flap vis a vis turbine blade without Gurney Flap. Performance parameters such as lift coefficient, drag coefficient, aerodynamic efficiency, and static pressure distributions over the blade surface are used to analyse the alterations in the performance of aerodynamic characteristics of the LPT Blade. All the Gurney Flap configurations improved the overall blade aerodynamics. Among these configurations, the flat Gurney Flap configuration is found to be superior as it improves the aerodynamic efficiency of the blade.
{"title":"Numerical investigation of flow control in low-pressure turbine cascade using Gurney Flaps","authors":"Mohan Agrawal, Jai Bhan Verma, G. Joshi, S. Chandel, Ved Prakash, R. K. Mishra","doi":"10.1515/tjeng-2022-0045","DOIUrl":"https://doi.org/10.1515/tjeng-2022-0045","url":null,"abstract":"Abstract This paper reports the numerical investigations to analyze the effect of the application of Gurney Flaps on various configurations in a Low-Pressure Turbine (LPT) cascade in mitigating the laminar flow separation during low Reynolds number operations. T106 LPT blade of chord 60 mm has been selected for the present study. Gurney Flaps of flat type and quarter round type of a particular height of the chord length are selected. Flaps are provided near the blades’ trailing edges for numerical analysis. The numerical computations are performed using STARCCM+ software, and the K-ω SST turbulence Model is used for turbulence closure. The studies are performed at various Reynolds numbers ranging from 37,500 to 138,750 to understand the application of various configurations of Gurney Flap vis a vis turbine blade without Gurney Flap. Performance parameters such as lift coefficient, drag coefficient, aerodynamic efficiency, and static pressure distributions over the blade surface are used to analyse the alterations in the performance of aerodynamic characteristics of the LPT Blade. All the Gurney Flap configurations improved the overall blade aerodynamics. Among these configurations, the flat Gurney Flap configuration is found to be superior as it improves the aerodynamic efficiency of the blade.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-10-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46345779","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Keran Song, Linyuan Jia, Yu-chun Chen, Tian Tan, Peiyang Fan
Abstract Aiming to optimize the mode transition control schedule of the adaptive cycle engine (ACE), an optimization method based on a gradient algorithm was proposed. During the mode transition, the compressor surge margin, the total turbine inlet temperature and the combustion chamber fuel-air ratio are not exceeded. The integration of thrust by time was selected as the optimization objective. The geometric parameters were updated via the optimization target derivative. The speed limit of the geometric adjustment mechanism was also considered in the optimization process. The control schedules of two mode transition processes, from triple to double bypass mode and from double to triple-bypass mode were optimized. The optimization method presented in this paper can be applied to various mode transitions and the geometric adjustment speed meets all the constraints.
{"title":"Optimization of ACE mode transition control schedule considering geometric adjustment speed","authors":"Keran Song, Linyuan Jia, Yu-chun Chen, Tian Tan, Peiyang Fan","doi":"10.1515/tjeng-2022-0010","DOIUrl":"https://doi.org/10.1515/tjeng-2022-0010","url":null,"abstract":"Abstract Aiming to optimize the mode transition control schedule of the adaptive cycle engine (ACE), an optimization method based on a gradient algorithm was proposed. During the mode transition, the compressor surge margin, the total turbine inlet temperature and the combustion chamber fuel-air ratio are not exceeded. The integration of thrust by time was selected as the optimization objective. The geometric parameters were updated via the optimization target derivative. The speed limit of the geometric adjustment mechanism was also considered in the optimization process. The control schedules of two mode transition processes, from triple to double bypass mode and from double to triple-bypass mode were optimized. The optimization method presented in this paper can be applied to various mode transitions and the geometric adjustment speed meets all the constraints.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-10-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"48309878","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract To improve the prediction accuracy of profile loss at low Reynolds number, a typical low-pressure turbine cascade T106D-EIZ was selected to numerically investigate the effect of Reynolds number on turbine cascade flow. A detailed analysis of profile loss was performed and a profile loss model considering the low-Re effect was developed. Results showed that the incidence angle has a great effect on the inlet and outlet Mach number at low Reynolds number, and the variation of inlet and outlet Mach number further affects the blade profile loss. A correction factor was introduced to consider the effect of incidence angle and Mach number on the profile loss. The profile loss coefficient and stalling incidence angle were both extended to lower Reynolds number based on the numerical results. A Smart Through Flow Analysis Program (STFAP) was developed using the finite volume method to solve the circumferentially averaged Euler equations of S2 surface. Aerodynamic performance of E3 5-stage low-pressure turbine was predicted by STFAP coupled with low-Re profile loss model. Compared with K-O model, the prediction accuracy of efficiency of low-pressure turbine last stage is improved by nearly 1.1 percentage points when the 5-stage low-pressure turbine is in a low Reynolds number state.
{"title":"Development and application of a profile loss model considering the low-Re effect in low-pressure turbine","authors":"W. Jia, Q. Kong, Guanyun Xiao, Handong Mu","doi":"10.1515/tjeng-2022-0052","DOIUrl":"https://doi.org/10.1515/tjeng-2022-0052","url":null,"abstract":"Abstract To improve the prediction accuracy of profile loss at low Reynolds number, a typical low-pressure turbine cascade T106D-EIZ was selected to numerically investigate the effect of Reynolds number on turbine cascade flow. A detailed analysis of profile loss was performed and a profile loss model considering the low-Re effect was developed. Results showed that the incidence angle has a great effect on the inlet and outlet Mach number at low Reynolds number, and the variation of inlet and outlet Mach number further affects the blade profile loss. A correction factor was introduced to consider the effect of incidence angle and Mach number on the profile loss. The profile loss coefficient and stalling incidence angle were both extended to lower Reynolds number based on the numerical results. A Smart Through Flow Analysis Program (STFAP) was developed using the finite volume method to solve the circumferentially averaged Euler equations of S2 surface. Aerodynamic performance of E3 5-stage low-pressure turbine was predicted by STFAP coupled with low-Re profile loss model. Compared with K-O model, the prediction accuracy of efficiency of low-pressure turbine last stage is improved by nearly 1.1 percentage points when the 5-stage low-pressure turbine is in a low Reynolds number state.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-09-23","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"48627356","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Abstract As an essential component, the bleed system plays a critical role in supplying turbine cooling air, guaranteeing stage matching, pressurizing the cabin, and de-icing at the wing and engine inlet. However, the extraction of the bleeding air from the compressor causes the engine efficiency degradation and thrust deficit. Therefore, flow control based on bleed is conducted to compensate the bleed induced disadvantages. The influence of the circumferential bleeding slot location on the tip leakage vortex and passage vortex controlling in a compressor cascade with the tip clearance is numerically studied using large eddy simulation. Three bleed configurations and the smooth casing configuration are investigated. 17.11% loss reduction is obtained through bleeding at 10% c x upstream of the blade leading edge with a bleeding rate of 2.76%. The vortex structures and flow patterns are compared and analyzed to reveal the controlling mechanism. Subsequently, the axial vorticity and loss evolution is discussed, and the interaction between the primary flow and bleeding air is revealed. It’s found that bleeding slot placed within the blade passage is exposed into a highly static pressure gradient, and this causes the bleeding air flows into and spills out the bleeding slot and leads to unnecessary loss. Moreover, the influence of large bleeding rate and inlet boundary layer is assessed.
{"title":"Vortex structure control based bleed in axial compressor cascade with tip clearance using large eddy simulation","authors":"Yun Gong, Shaowen Chen, Cong Zeng","doi":"10.1515/tjj-2022-0047","DOIUrl":"https://doi.org/10.1515/tjj-2022-0047","url":null,"abstract":"Abstract As an essential component, the bleed system plays a critical role in supplying turbine cooling air, guaranteeing stage matching, pressurizing the cabin, and de-icing at the wing and engine inlet. However, the extraction of the bleeding air from the compressor causes the engine efficiency degradation and thrust deficit. Therefore, flow control based on bleed is conducted to compensate the bleed induced disadvantages. The influence of the circumferential bleeding slot location on the tip leakage vortex and passage vortex controlling in a compressor cascade with the tip clearance is numerically studied using large eddy simulation. Three bleed configurations and the smooth casing configuration are investigated. 17.11% loss reduction is obtained through bleeding at 10% c x upstream of the blade leading edge with a bleeding rate of 2.76%. The vortex structures and flow patterns are compared and analyzed to reveal the controlling mechanism. Subsequently, the axial vorticity and loss evolution is discussed, and the interaction between the primary flow and bleeding air is revealed. It’s found that bleeding slot placed within the blade passage is exposed into a highly static pressure gradient, and this causes the bleeding air flows into and spills out the bleeding slot and leads to unnecessary loss. Moreover, the influence of large bleeding rate and inlet boundary layer is assessed.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2022-09-22","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"42372351","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":4,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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