Abstract Three-dimensional unsteady numerical simulations were conducted to investigate the detailed film cooling mechanism of the high-pressure turbine shroud with the first-stage turbine blade and guide vane for an aero-engine under the high-speed rotation of blades and rotor-stator interaction. The slip mesh was used to realize the relative motion between the rotating blade and the stationary turbine shroud. It is found that the coolant jet is alternately influenced by the hot mainstream, tip clearance leakage flow and leakage vortex due to the high rotational speed of blades. The film cooling characteristics of the turbine shroud significantly present an unsteady and periodic flow and heat transfer phenomenon. The insufficient cooling margin for film holes at the upstream of the blade leading edge can occur not only under high blowing ratios due to the coolant jet liftoff, but also at low blowing ratios due to the insufficient coolant flow rate as a result of the high exit pressure. A novel shroud cooling structure with coolant supply by the added throttle chamber is put forward, and expected to provide better thermal protection for the high-pressure turbine shroud near the leading edge of blades with no extra increase in the total mass flow rate of coolant.
{"title":"Unsteady film cooling characteristics of the high-pressure turbine shroud with blade rotation in an aero-engine","authors":"Z. Kou, Zihao Bao, Guang-chao Li, Xunyan Yin","doi":"10.1515/tjj-2021-0032","DOIUrl":"https://doi.org/10.1515/tjj-2021-0032","url":null,"abstract":"Abstract Three-dimensional unsteady numerical simulations were conducted to investigate the detailed film cooling mechanism of the high-pressure turbine shroud with the first-stage turbine blade and guide vane for an aero-engine under the high-speed rotation of blades and rotor-stator interaction. The slip mesh was used to realize the relative motion between the rotating blade and the stationary turbine shroud. It is found that the coolant jet is alternately influenced by the hot mainstream, tip clearance leakage flow and leakage vortex due to the high rotational speed of blades. The film cooling characteristics of the turbine shroud significantly present an unsteady and periodic flow and heat transfer phenomenon. The insufficient cooling margin for film holes at the upstream of the blade leading edge can occur not only under high blowing ratios due to the coolant jet liftoff, but also at low blowing ratios due to the insufficient coolant flow rate as a result of the high exit pressure. A novel shroud cooling structure with coolant supply by the added throttle chamber is put forward, and expected to provide better thermal protection for the high-pressure turbine shroud near the leading edge of blades with no extra increase in the total mass flow rate of coolant.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-07-11","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"49344481","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 Leading edge (LE) plays a prominent role in compressor flow. The asymmetric leading edge (ASYLE) has shown superiorities to symmetric LE in blade aerodynamic performance. However, the influencing rules of ASYLE design parameters are still ambiguous. In this work, numerical calculations were conducted to investigate the influencing effects of LE point curvature and position. The results show that the operating range of ASYLE blades expand with the decrease of LE point curvature, which helps to moderate LE flow acceleration, while the LE point position mainly affects the operating range. It is also revealed that the SSLE curvature peak is supposed to be close to LE point, and the maximum value of PSLE curvature should be restricted.
{"title":"Influence of leading edge point on aerodynamic performance of asymmetric leading edge compressor airfoils","authors":"Guanhua Yang, Limin Gao, Haohao Wang, Longrui Chang","doi":"10.1515/tjj-2021-0054","DOIUrl":"https://doi.org/10.1515/tjj-2021-0054","url":null,"abstract":"Abstract Leading edge (LE) plays a prominent role in compressor flow. The asymmetric leading edge (ASYLE) has shown superiorities to symmetric LE in blade aerodynamic performance. However, the influencing rules of ASYLE design parameters are still ambiguous. In this work, numerical calculations were conducted to investigate the influencing effects of LE point curvature and position. The results show that the operating range of ASYLE blades expand with the decrease of LE point curvature, which helps to moderate LE flow acceleration, while the LE point position mainly affects the operating range. It is also revealed that the SSLE curvature peak is supposed to be close to LE point, and the maximum value of PSLE curvature should be restricted.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-05-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46314035","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 paper outlines a finite volumes refined mathematical model of the working gas flow in the flow path of the three stage modern single shaft gas turbine engine that can be used in floating power plants. Such mathematical model based on the finite volumes of hexagonal-type was constructed using the three-dimensional Navier–Stokes equations for the case of viscous working fluid flow. For the problem solution such boundary conditions as “inlet’, “outlet” and “wall” have been used. The calculation is carried out in a non-stationary setting with a time step of 1.5974 × 10−6 s, which corresponds to the angle of rotation of the rotor, relative to the stator, of 0.09°. The total number of time iterations is 350. Also, it was shown that the variation field of pressure on the blades feather surfaces and the gas flow velocity due to rotation are the critical factors, causing the blades vibration. The result was confirmed with the experiment. The obtained results would be used as a base for further investigations of gas flow pressure field on the blades surface, because the gas flow pressure are key factors, causing the rotor forced vibration, and as initial data for their fatigue strength and crack study.
{"title":"Gas dynamic analysis of the modern single shaft gas turbine engine flow path","authors":"S. Morhun, S. Vilkul","doi":"10.1515/tjeng-2022-0019","DOIUrl":"https://doi.org/10.1515/tjeng-2022-0019","url":null,"abstract":"Abstract The paper outlines a finite volumes refined mathematical model of the working gas flow in the flow path of the three stage modern single shaft gas turbine engine that can be used in floating power plants. Such mathematical model based on the finite volumes of hexagonal-type was constructed using the three-dimensional Navier–Stokes equations for the case of viscous working fluid flow. For the problem solution such boundary conditions as “inlet’, “outlet” and “wall” have been used. The calculation is carried out in a non-stationary setting with a time step of 1.5974 × 10−6 s, which corresponds to the angle of rotation of the rotor, relative to the stator, of 0.09°. The total number of time iterations is 350. Also, it was shown that the variation field of pressure on the blades feather surfaces and the gas flow velocity due to rotation are the critical factors, causing the blades vibration. The result was confirmed with the experiment. The obtained results would be used as a base for further investigations of gas flow pressure field on the blades surface, because the gas flow pressure are key factors, causing the rotor forced vibration, and as initial data for their fatigue strength and crack study.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-05-26","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"42158372","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 Aero engine performance deterioration highly influences its reliability, availability and life cycle. Predictive maintenance is therefore a key figure within Industry 4.0, which guarantees high availability and reduced downtime thus reduced operational costs for both military and civil engines. This leads to maintenance on demand and needs an effective engine health monitoring system. This paper overviews the work carried out on aero engine health monitoring, diagnostic and prognostic techniques based on gas path performance parameters. The inception of performance monitoring and its evolution over time, techniques used to establish a high-quality data base using engine model performance adaptation, and effects of computationally intelligent techniques on promoting the implementation of engine fault diagnosis are reviewed. Generating dependable information about the health condition of the engine is therefore a requisite for a successful implementation of condition-based maintenance. Based on this study, further research can be attempted to predict residual life of critical components using degradation pattern from aero engine performance data bank which will be an invaluable asset for engine designers as well as for operators.
{"title":"Aero engine health monitoring, diagnostics and prognostics for condition-based maintenance: an overview","authors":"Narahari Rath, R. Mishra, A. Kushari","doi":"10.1515/tjeng-2022-0020","DOIUrl":"https://doi.org/10.1515/tjeng-2022-0020","url":null,"abstract":"Abstract Aero engine performance deterioration highly influences its reliability, availability and life cycle. Predictive maintenance is therefore a key figure within Industry 4.0, which guarantees high availability and reduced downtime thus reduced operational costs for both military and civil engines. This leads to maintenance on demand and needs an effective engine health monitoring system. This paper overviews the work carried out on aero engine health monitoring, diagnostic and prognostic techniques based on gas path performance parameters. The inception of performance monitoring and its evolution over time, techniques used to establish a high-quality data base using engine model performance adaptation, and effects of computationally intelligent techniques on promoting the implementation of engine fault diagnosis are reviewed. Generating dependable information about the health condition of the engine is therefore a requisite for a successful implementation of condition-based maintenance. Based on this study, further research can be attempted to predict residual life of critical components using degradation pattern from aero engine performance data bank which will be an invaluable asset for engine designers as well as for operators.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-05-12","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47590035","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}
Xi Gao, Zhiyuan Cao, Xiang Zhang, Fei Zhang, Jing Yang, Bo Liu
Abstract With the purpose of investigating the effect mechanism of injection on flow separation of the S-duct, different single-hole schemes were investigated and compared with double-hole schemes. Results show that, the performance of S-duct can be improved by using injection. The optimal scheme in this study is a double-hole injection scheme with two holes located at the same axial position. Flow separation reduction and a 16.9% reduction of loss coefficient were achieved by injection with an injection coefficient of 0.46% in each hole. The flow mechanisms are that, firstly, high momentum fluid is injected to separated flow by air injection; secondary, high momentum flow is transported to flow near downside wall by injection vortex. The position effect, injection flow rate effect and hole shape effect were also discussed. For double-hole scheme, the scheme with two holes located different axial positions generates a stronger vortex by mixing two injection vortexes and enhances secondary flow. Though the flow separation is reduced, a severe nonuniform flow field at outlet is formed. Due to the less swirling flow achieved at outlet of S-duct by Double-y scheme, it can offer a more uniform flow field for downstream compressor and has a better control effectiveness.
{"title":"Performance enhancement and flow separation control in an S-duct by air injection","authors":"Xi Gao, Zhiyuan Cao, Xiang Zhang, Fei Zhang, Jing Yang, Bo Liu","doi":"10.1515/tjj-2021-0063","DOIUrl":"https://doi.org/10.1515/tjj-2021-0063","url":null,"abstract":"Abstract With the purpose of investigating the effect mechanism of injection on flow separation of the S-duct, different single-hole schemes were investigated and compared with double-hole schemes. Results show that, the performance of S-duct can be improved by using injection. The optimal scheme in this study is a double-hole injection scheme with two holes located at the same axial position. Flow separation reduction and a 16.9% reduction of loss coefficient were achieved by injection with an injection coefficient of 0.46% in each hole. The flow mechanisms are that, firstly, high momentum fluid is injected to separated flow by air injection; secondary, high momentum flow is transported to flow near downside wall by injection vortex. The position effect, injection flow rate effect and hole shape effect were also discussed. For double-hole scheme, the scheme with two holes located different axial positions generates a stronger vortex by mixing two injection vortexes and enhances secondary flow. Though the flow separation is reduced, a severe nonuniform flow field at outlet is formed. Due to the less swirling flow achieved at outlet of S-duct by Double-y scheme, it can offer a more uniform flow field for downstream compressor and has a better control effectiveness.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-05-12","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"49652196","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 further reveal the interaction between crossflow and jet, single oblique jet impingement in crossflow is studied. The influences of Reynolds number (Re = 3750–15,000), oblique angle (θ = 50°–90°), jet intake mode (jet normal to inlet plane and jet along hole axis), crossflow ratio (CR = 0.1–0.3) and temperature ratio (TR = 1.0–1.2) are considered. For both jet intake modes, with the decrease of θ, the peak value of Nusselt number increases and its location shifts upstream. The Nusselt number for case of jet normal to inlet plane is higher than that of jet along hole axis with wider lateral coverage range. The CR affects the convection heat transfer between crossflow and target surface, and it also affects the interaction between crossflow and jet. Synthetically, the best heat transfer appears at CR = 0.2. The Nusselt number decreases sharply when TR goes from 1.0 to 1.1, while the decline trend slows down when TR goes from 1.1 to 1.2.
{"title":"Heat transfer distribution of single oblique jet impingement in crossflow under different inlet conditions","authors":"Juan He, Qinghua Deng, K. Xiao, Z. Feng","doi":"10.1515/tjeng-2021-0029","DOIUrl":"https://doi.org/10.1515/tjeng-2021-0029","url":null,"abstract":"Abstract To further reveal the interaction between crossflow and jet, single oblique jet impingement in crossflow is studied. The influences of Reynolds number (Re = 3750–15,000), oblique angle (θ = 50°–90°), jet intake mode (jet normal to inlet plane and jet along hole axis), crossflow ratio (CR = 0.1–0.3) and temperature ratio (TR = 1.0–1.2) are considered. For both jet intake modes, with the decrease of θ, the peak value of Nusselt number increases and its location shifts upstream. The Nusselt number for case of jet normal to inlet plane is higher than that of jet along hole axis with wider lateral coverage range. The CR affects the convection heat transfer between crossflow and target surface, and it also affects the interaction between crossflow and jet. Synthetically, the best heat transfer appears at CR = 0.2. The Nusselt number decreases sharply when TR goes from 1.0 to 1.1, while the decline trend slows down when TR goes from 1.1 to 1.2.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-04-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"48216610","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}
C. Vinay, Kumar Gottegere Narayanappa, Y. Giridhara Babu
Abstract Turboprop engines require an exhaust nozzle or stub to duct the engine exhaust flue gas outboard of the aircraft. The design of these exhaust stubs are dictated primarily by the aircraft’s configuration. In pusher aircraft, the exhaust stubs are designed to minimize the exposure of the flue gases from the engine exhaust on the propeller blades and fuselage. A fluid-thermal-structure coupling analysis is performed to understand the thermal effects of the engine exhaust jet flow on the thermo-mechanical behavior of pusher configured light transport aircraft propeller and structure. The steady thermal flow field of the aircraft with forward and reverse thrust, in which propeller blade angle variations were analyzed for different aircraft speed. The present work investigates a three-dimensional analysis of flow around the nacelle-airframe and the effect of exhaust flue gas impingement on the propeller blade surface. Based on the insights from the numerical results, the designed exhaust duct was integrated on the aircraft and carried out ground static and flight testing for various flight operating conditions in which propeller blade and fuselage surface temperature were measured. Numerical and experimental results are compared and validated for certain flight conditions and found satisfactory.
{"title":"Experimental and numerical investigation on the effect of turboprop engine exhaust gas impingement on pusher aircraft","authors":"C. Vinay, Kumar Gottegere Narayanappa, Y. Giridhara Babu","doi":"10.1515/tjj-2022-0011","DOIUrl":"https://doi.org/10.1515/tjj-2022-0011","url":null,"abstract":"Abstract Turboprop engines require an exhaust nozzle or stub to duct the engine exhaust flue gas outboard of the aircraft. The design of these exhaust stubs are dictated primarily by the aircraft’s configuration. In pusher aircraft, the exhaust stubs are designed to minimize the exposure of the flue gases from the engine exhaust on the propeller blades and fuselage. A fluid-thermal-structure coupling analysis is performed to understand the thermal effects of the engine exhaust jet flow on the thermo-mechanical behavior of pusher configured light transport aircraft propeller and structure. The steady thermal flow field of the aircraft with forward and reverse thrust, in which propeller blade angle variations were analyzed for different aircraft speed. The present work investigates a three-dimensional analysis of flow around the nacelle-airframe and the effect of exhaust flue gas impingement on the propeller blade surface. Based on the insights from the numerical results, the designed exhaust duct was integrated on the aircraft and carried out ground static and flight testing for various flight operating conditions in which propeller blade and fuselage surface temperature were measured. Numerical and experimental results are compared and validated for certain flight conditions and found satisfactory.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-04-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46147717","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}
Chenxi Zhao, Chang-Hyun Yang, Chenqing Zhang, Xiaotian Zhang, G. Xi
Abstract This paper focuses on the investigation of the blade deformation and tip clearance effects of a high-flow-coefficient centrifugal compressor. Blade deformation caused by the pressure load and centrifugal load under operating condition is considered by fluid/solid interaction method to simulate the stage performance accurately. Hence the numerical results are in good agreement with the experimental data. The blade deformation is then analyzed in detail. Also, the effect of the size of tip clearance is studied by evaluating the performance with five various gaps. The result shows that the increment of the tip clearance from 0 to 2.0 mm leads to a 5.14% peak efficiency decline. However, increasing tip gap contributes to enlarge the stable operating range in terms of stall margin and choke margin.
{"title":"Numerical and experimental investigations of tip clearance effects in a high-flow-coefficient centrifugal compressor","authors":"Chenxi Zhao, Chang-Hyun Yang, Chenqing Zhang, Xiaotian Zhang, G. Xi","doi":"10.1515/tjj-2022-0002","DOIUrl":"https://doi.org/10.1515/tjj-2022-0002","url":null,"abstract":"Abstract This paper focuses on the investigation of the blade deformation and tip clearance effects of a high-flow-coefficient centrifugal compressor. Blade deformation caused by the pressure load and centrifugal load under operating condition is considered by fluid/solid interaction method to simulate the stage performance accurately. Hence the numerical results are in good agreement with the experimental data. The blade deformation is then analyzed in detail. Also, the effect of the size of tip clearance is studied by evaluating the performance with five various gaps. The result shows that the increment of the tip clearance from 0 to 2.0 mm leads to a 5.14% peak efficiency decline. However, increasing tip gap contributes to enlarge the stable operating range in terms of stall margin and choke margin.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-03-28","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"41627553","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 solve the problems of the nonlinear damping force in the traditional squeeze film damper (SFD), a novel structure of G-type integral squeeze film damper (GISFD) based on ISFD is proposed for the first time. The finite element model and test rig of the ball bearing-rotor system are established to explore the influence of GISFD and ISFD on the dynamic characteristics of the unbalanced rotor system. The results show that both GISFD and ISFD can change the critical speed of the rotor system, reduce the bending strain energy of the shaft, and reduce the bearing dynamic load of the rotor system. Through comparison, it is found that the effect of GISFD is more obvious. The experimental results show that, compared with the unbalanced rotor system without damper, the peak-peak value of amplitude in the rotor system with GISFD and ISFD at 3000 rpm is reduced by 25.53 and 15.81%. The amplitude in the disk at the first-order critical speed is effectively reduced, and the reduction range reach 52.01 and 35.44%, respectively. GISFD has a more significant effect of suppressing unbalanced vibration, and has superior vibration damping performance when compared with ISFD.
{"title":"Effect of G-type integral squeeze film damper on the dynamic characteristics in rotor system","authors":"Wei Yan, Lidong He, Gang Zhu, Xingyun Jia","doi":"10.1515/tjj-2021-0046","DOIUrl":"https://doi.org/10.1515/tjj-2021-0046","url":null,"abstract":"Abstract To solve the problems of the nonlinear damping force in the traditional squeeze film damper (SFD), a novel structure of G-type integral squeeze film damper (GISFD) based on ISFD is proposed for the first time. The finite element model and test rig of the ball bearing-rotor system are established to explore the influence of GISFD and ISFD on the dynamic characteristics of the unbalanced rotor system. The results show that both GISFD and ISFD can change the critical speed of the rotor system, reduce the bending strain energy of the shaft, and reduce the bearing dynamic load of the rotor system. Through comparison, it is found that the effect of GISFD is more obvious. The experimental results show that, compared with the unbalanced rotor system without damper, the peak-peak value of amplitude in the rotor system with GISFD and ISFD at 3000 rpm is reduced by 25.53 and 15.81%. The amplitude in the disk at the first-order critical speed is effectively reduced, and the reduction range reach 52.01 and 35.44%, respectively. GISFD has a more significant effect of suppressing unbalanced vibration, and has superior vibration damping performance when compared with ISFD.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-03-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44308420","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}
Hano van Eck, S. J. van der Spuy, T. von Backström
Abstract The use of micro gas turbines (MGTs) for once-off and unique applications means that a rapid turnaround design process is required. The development of a MATLAB® application-based program for the initial design and performance analysis of radial and mixed flow compressors is discussed. The program code is based on one dimensional (1D) mean line flow and loss model theory. For verification, 18 test compressors were developed, covering a wide range of design velocities, mass flow rates and meridional exit angles. Predicted performance results were verified using Numeca/FINE™ Turbo software. Initial comparisons between 1D and Computational Fluid Dynamics (CFD) results did not match well. The 1D software over-predicted compressor performance and provided poor choke prediction. Consequently, empirical models to correct these deviations were derived and implemented into the 1D Application code. The updated 1D code provided a mean choke prediction difference of 1.59% compared to an initial 14.98% difference at the design point. Mean total-to-total isentropic efficiency and pressure ratio reduced to differences of 0.74 and 1.24% from values of 11.23 and 9.31% respectively.
{"title":"Development of a one-dimensional code for the initial design of a micro gas turbine mixed flow compressor stage","authors":"Hano van Eck, S. J. van der Spuy, T. von Backström","doi":"10.1515/tjj-2022-0008","DOIUrl":"https://doi.org/10.1515/tjj-2022-0008","url":null,"abstract":"Abstract The use of micro gas turbines (MGTs) for once-off and unique applications means that a rapid turnaround design process is required. The development of a MATLAB® application-based program for the initial design and performance analysis of radial and mixed flow compressors is discussed. The program code is based on one dimensional (1D) mean line flow and loss model theory. For verification, 18 test compressors were developed, covering a wide range of design velocities, mass flow rates and meridional exit angles. Predicted performance results were verified using Numeca/FINE™ Turbo software. Initial comparisons between 1D and Computational Fluid Dynamics (CFD) results did not match well. The 1D software over-predicted compressor performance and provided poor choke prediction. Consequently, empirical models to correct these deviations were derived and implemented into the 1D Application code. The updated 1D code provided a mean choke prediction difference of 1.59% compared to an initial 14.98% difference at the design point. Mean total-to-total isentropic efficiency and pressure ratio reduced to differences of 0.74 and 1.24% from values of 11.23 and 9.31% respectively.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2022-03-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43775797","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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