Abstract Film and jet impingement cooling are widely used techniques in gas turbine vane and blade cooling. The present work investigates and compares the flow structure of a film-cooled flat plate with a hybrid cooling scheme. The hybrid cooling scheme combines both impingement hole and film holes and is named combined impingement-film (IFC) cooling. Experimental validation and computational analyses are carried out on a flat plate with film holes. Different flow parameters, such as velocity pattern, Turbulent kinetic energy, and streamline flow structure, are compared for the two cases in different regions of the flat plate. It is observed that the hybrid scheme shows advantages over film cooling. The jet-to-jet interaction, jet crossflow interaction, and vortex formation are the main factors affecting film cooling performance. There is a 52 % drop in turbulent kinetic energy for the hybrid cooling compared to the film cooling at the film hole exit. More mixing in the coolant and mainstream interaction is observed for the FC case than in the IFC.
{"title":"Flow structure comparison of film cooling versus hybrid cooling: a CFD study","authors":"R. Panda, A. K. Pujari, B. Gudla","doi":"10.1515/tjj-2022-0058","DOIUrl":"https://doi.org/10.1515/tjj-2022-0058","url":null,"abstract":"Abstract Film and jet impingement cooling are widely used techniques in gas turbine vane and blade cooling. The present work investigates and compares the flow structure of a film-cooled flat plate with a hybrid cooling scheme. The hybrid cooling scheme combines both impingement hole and film holes and is named combined impingement-film (IFC) cooling. Experimental validation and computational analyses are carried out on a flat plate with film holes. Different flow parameters, such as velocity pattern, Turbulent kinetic energy, and streamline flow structure, are compared for the two cases in different regions of the flat plate. It is observed that the hybrid scheme shows advantages over film cooling. The jet-to-jet interaction, jet crossflow interaction, and vortex formation are the main factors affecting film cooling performance. There is a 52 % drop in turbulent kinetic energy for the hybrid cooling compared to the film cooling at the film hole exit. More mixing in the coolant and mainstream interaction is observed for the FC case than in the IFC.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-05-10","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47246041","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}
Lei Shi, Ruilin Jiang, Chen-lei Huang, Shuhan Guo, Guoshun Chen
Abstract The entire compression system of DGEN380 was selected to study the degradation of aerodynamic characteristics caused by erosion of the leading edge. Numerical simulations under different degrees of leading edge erosion were performed. It was found that the leading edge erosion causes a significant decrease in operability from 70 % of the blade height to the tip, and mainly affects the subsequent flow field in the bypass. At 90 % of blade height, the isentropic efficiency and total pressure ratio of the blade with leading edge erosion are reduced compared to the origin blade. Afterwards, the thrust equation was introduced to analyze the effect of leading edge erosion on the thrust loss under different operating conditions of the aircraft. Among the three operating conditions, the thrust loss rate of the cruise condition was the largest, which provided a reference for the subsequent evaluation of the blade repair condition.
{"title":"The aerodynamic performance degradation analysis of a small high bypass turbofan engine compression system with fan rotor blade leading edge erosion","authors":"Lei Shi, Ruilin Jiang, Chen-lei Huang, Shuhan Guo, Guoshun Chen","doi":"10.1515/tjj-2023-0031","DOIUrl":"https://doi.org/10.1515/tjj-2023-0031","url":null,"abstract":"Abstract The entire compression system of DGEN380 was selected to study the degradation of aerodynamic characteristics caused by erosion of the leading edge. Numerical simulations under different degrees of leading edge erosion were performed. It was found that the leading edge erosion causes a significant decrease in operability from 70 % of the blade height to the tip, and mainly affects the subsequent flow field in the bypass. At 90 % of blade height, the isentropic efficiency and total pressure ratio of the blade with leading edge erosion are reduced compared to the origin blade. Afterwards, the thrust equation was introduced to analyze the effect of leading edge erosion on the thrust loss under different operating conditions of the aircraft. Among the three operating conditions, the thrust loss rate of the cruise condition was the largest, which provided a reference for the subsequent evaluation of the blade repair condition.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44986717","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 With the advancement of the supersonic aero propulsion system, optimizing the combined performance of inlet/engine integration has become increasingly crucial. To solve the coupling inlet/engine problem, a quasi-one-dimensional inlet modeling and drag calculation method are proposed, integrated performance seeking control (PSC) based on the neighborhood-based speciation differential evolution-grey wolf optimizer (NSDE-GWO) is presented and quantitatively analyses the influence of variable geometry inlet regulation on performance. The results reveal that the optimization effect of the ramp angle adjustment is generally better than that of the bleed adjustment, and the NSDE-GWO hybrid algorithm achieves remarkable optimization solutions in all three different modes. The PSC with variable geometry inlet adjustment provides more additional potential for optimization compared with fixed geometry inlet, and the performance can be maximized by adjusting both the bleed adjustment and the ramp angle. This study maximizes the exploitation of potential and has theoretical guidance and practical engineering significance.
{"title":"Installed performance seeking control based on supersonic variable inlet/engine coupling model","authors":"Chen Wang, Ximing Sun, Xian Du","doi":"10.1515/tjj-2023-0030","DOIUrl":"https://doi.org/10.1515/tjj-2023-0030","url":null,"abstract":"Abstract With the advancement of the supersonic aero propulsion system, optimizing the combined performance of inlet/engine integration has become increasingly crucial. To solve the coupling inlet/engine problem, a quasi-one-dimensional inlet modeling and drag calculation method are proposed, integrated performance seeking control (PSC) based on the neighborhood-based speciation differential evolution-grey wolf optimizer (NSDE-GWO) is presented and quantitatively analyses the influence of variable geometry inlet regulation on performance. The results reveal that the optimization effect of the ramp angle adjustment is generally better than that of the bleed adjustment, and the NSDE-GWO hybrid algorithm achieves remarkable optimization solutions in all three different modes. The PSC with variable geometry inlet adjustment provides more additional potential for optimization compared with fixed geometry inlet, and the performance can be maximized by adjusting both the bleed adjustment and the ramp angle. This study maximizes the exploitation of potential and has theoretical guidance and practical engineering significance.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"42531928","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}
S. P. Sanaka, Ramanaiah Kandula, Khyathi Sree Chalamalasetty, Durga Rao Kappala
Abstract The objective of the study is realizing the effect of fuel mass flow rate and flight speed on combustion in scramjet engine. DLR conical strut based scramjet combustor configuration was chosen and simulated the chemical reaction between the air and hydrogen fuel. A slot of size 40 mm × 0.295 mm provided at the center of the strut to injected hydrogen fuel from the rare side in to the downstream flow. ICEM CFD software is used for the generation of structured elements in computational domain for three dimensional flow analyses. Standard k-epsilon turbulence model and species transport equation is used in ANSYS fluent solver. The predicted temperature, velocity distribution along the axial length was compared with the experimental results and validated. The temperature distribution at different Mach numbers and mass flow rate reveals that the peak temperature increased with the flight speed and inlet fuel mass flow rate. The peak temperature noticed at the center of the combustor is around 3500 K at a flight speed of Mach 4. The predicted variation of temperature, pressure, velocity in the combustor and the flow structure for reacting flow facilitate good understanding of the combustion process in scramjet combustor.
研究的目的是实现燃料质量、流量和飞行速度对超燃冲压发动机燃烧的影响。选择了基于DLR锥形支杆的超燃冲压发动机燃烧室构型，并对空气与氢燃料的化学反应进行了模拟。在支板中心有一个40 mm × 0.295 mm的槽，用于从稀有侧向下游流动注入氢燃料。采用ICEM CFD软件在计算域生成结构单元，进行三维流动分析。ANSYS fluent求解器采用标准k-epsilon湍流模型和种输运方程。将预测的温度、速度沿轴向长度的分布与实验结果进行了比较，并得到了验证。不同马赫数和质量流量下的温度分布表明，峰值温度随飞行速度和进口燃油质量流量的增加而增加。在4马赫的飞行速度下，燃烧室中心的峰值温度约为3500k。预测了燃烧室内温度、压力、速度的变化以及反应流的流动结构，有助于更好地理解超燃冲压发动机燃烧过程。
{"title":"Reacting flow analysis in scramjet engine: effect of mass flow rate of fuel and flight velocity","authors":"S. P. Sanaka, Ramanaiah Kandula, Khyathi Sree Chalamalasetty, Durga Rao Kappala","doi":"10.1515/tjeng-2023-0029","DOIUrl":"https://doi.org/10.1515/tjeng-2023-0029","url":null,"abstract":"Abstract The objective of the study is realizing the effect of fuel mass flow rate and flight speed on combustion in scramjet engine. DLR conical strut based scramjet combustor configuration was chosen and simulated the chemical reaction between the air and hydrogen fuel. A slot of size 40 mm × 0.295 mm provided at the center of the strut to injected hydrogen fuel from the rare side in to the downstream flow. ICEM CFD software is used for the generation of structured elements in computational domain for three dimensional flow analyses. Standard k-epsilon turbulence model and species transport equation is used in ANSYS fluent solver. The predicted temperature, velocity distribution along the axial length was compared with the experimental results and validated. The temperature distribution at different Mach numbers and mass flow rate reveals that the peak temperature increased with the flight speed and inlet fuel mass flow rate. The peak temperature noticed at the center of the combustor is around 3500 K at a flight speed of Mach 4. The predicted variation of temperature, pressure, velocity in the combustor and the flow structure for reacting flow facilitate good understanding of the combustion process in scramjet combustor.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43841774","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}
Kaleeswaran Periyasamy, Kadiresh P. Natarajan, Bogadi Surendra, Khandai Suresh Chandra
Abstract An experimental study is conducted to determine the effect of chevrons with zero penetration angles at the CD nozzle exit on an emitted noise field. The implication of passive control is to reduce the blockage of the nozzle exit area with minimal engine thrust penalty. The cold air jets issued at design Mach numbers 1.5 and 1.75 from the De Laval nozzles of the circular section were investigated. This passive control eliminates screech tones at the over and ideally expanded conditions at 60° and 90° in the azimuth plane. The acoustic data measurements have also been observed for the chosen jet Mach numbers. The schlieren images reveal the shock cell pattern to eliminate the effect of shock-associated noise levels at supersonic jets. The results show that 10 chevrons with no penetration act as an effective eliminator of screech tone and noise suppression average ∆OASPL value up to 3 dB at Mach number 1.75.
{"title":"Effect of zero penetration angle chevrons in supersonic jet noise and screech tone mitigation","authors":"Kaleeswaran Periyasamy, Kadiresh P. Natarajan, Bogadi Surendra, Khandai Suresh Chandra","doi":"10.1515/tjj-2022-0073","DOIUrl":"https://doi.org/10.1515/tjj-2022-0073","url":null,"abstract":"Abstract An experimental study is conducted to determine the effect of chevrons with zero penetration angles at the CD nozzle exit on an emitted noise field. The implication of passive control is to reduce the blockage of the nozzle exit area with minimal engine thrust penalty. The cold air jets issued at design Mach numbers 1.5 and 1.75 from the De Laval nozzles of the circular section were investigated. This passive control eliminates screech tones at the over and ideally expanded conditions at 60° and 90° in the azimuth plane. The acoustic data measurements have also been observed for the chosen jet Mach numbers. The schlieren images reveal the shock cell pattern to eliminate the effect of shock-associated noise levels at supersonic jets. The results show that 10 chevrons with no penetration act as an effective eliminator of screech tone and noise suppression average ∆OASPL value up to 3 dB at Mach number 1.75.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-17","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"42725090","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 Circulation control is a kind of efficient flow control technology which can improve aircraft aerodynamic performance and reduce fuel consumption. However, improving the aerodynamic efficiency of a circulation device to enhance flight endurance and achieve environmentally flying is a challenging problem for the application of circulation control. This paper presents an efficient flow control technique that combines co-flow jet and multi-stage slot circulation control. The combinational flow control technique is applied to a supersonic airfoil to test its energy consumption and aerodynamic benefit achievement. Results show that both the single and double slot circulation control can improve the maximum lift-drag ratio of the baseline airfoil, with an increment of 11.3% and 19.1%, respectively. Compared with the single application of co-flow jet control which can increase the lift-drag ratio of the baseline airfoil by 16.3% and extend the stall angle of attack from 6° to 8°, the combinational flow control can obtain a more significant lift-drag ratio increment by about 27.3% and eliminate flow separations at high angle of attack. The stall angle of attack can even be increased to about 10°. Additionally, the blowing efficiency of the circulation control airfoil has been comprehensively analyzed. The results show that the maximum effective lift-drag ratio and highest blowing efficiency can be achieved at a blowing coefficient of 0.00235.
{"title":"An efficient flow control technique based on co-flow jet and multi-stage slot circulation control applied to a supercritical airfoil","authors":"Lei Wang, Hanan Lu, Yue Xu, Q. Li","doi":"10.1515/tjj-2023-0027","DOIUrl":"https://doi.org/10.1515/tjj-2023-0027","url":null,"abstract":"Abstract Circulation control is a kind of efficient flow control technology which can improve aircraft aerodynamic performance and reduce fuel consumption. However, improving the aerodynamic efficiency of a circulation device to enhance flight endurance and achieve environmentally flying is a challenging problem for the application of circulation control. This paper presents an efficient flow control technique that combines co-flow jet and multi-stage slot circulation control. The combinational flow control technique is applied to a supersonic airfoil to test its energy consumption and aerodynamic benefit achievement. Results show that both the single and double slot circulation control can improve the maximum lift-drag ratio of the baseline airfoil, with an increment of 11.3% and 19.1%, respectively. Compared with the single application of co-flow jet control which can increase the lift-drag ratio of the baseline airfoil by 16.3% and extend the stall angle of attack from 6° to 8°, the combinational flow control can obtain a more significant lift-drag ratio increment by about 27.3% and eliminate flow separations at high angle of attack. The stall angle of attack can even be increased to about 10°. Additionally, the blowing efficiency of the circulation control airfoil has been comprehensively analyzed. The results show that the maximum effective lift-drag ratio and highest blowing efficiency can be achieved at a blowing coefficient of 0.00235.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-10","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"42465910","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}
Goveraiahgari Venkatesh, R. Meenakshi Reddy, Pabbisetty Mallikarjuna Rao
Abstract Turbine inlet air temperatures are extremely high, which can result in blade material damage. As a result, cooling the turbine blades is required, and a variety of cooling techniques have been introduced. The majority of the previous research on pin fins has focused on circular fins using a wedge duct to apply a constant temperature and uniform heat flux to the end wall and pin fin surfaces. The present study compares seven oblong pin-fins to seven circular pin fins in a wedge duct with a Reynolds number range of 10,000–50,000 and a constant heat flux (surface) of 3280 W/m2 applied to the endwall and surfaces of the oblong pin fin. The results indicate that the friction factor for oblong fins is 14% lower than for circular pin fins. The thermal performance factor is increased by 11.4%. The thermal performance factor can be improved by using oblong pin fins with higher Reynolds numbers.
{"title":"Numerical investigations of heat transfer characteristics using oblong fins and circular fins in a wedge channel","authors":"Goveraiahgari Venkatesh, R. Meenakshi Reddy, Pabbisetty Mallikarjuna Rao","doi":"10.1515/tjj-2022-0055","DOIUrl":"https://doi.org/10.1515/tjj-2022-0055","url":null,"abstract":"Abstract Turbine inlet air temperatures are extremely high, which can result in blade material damage. As a result, cooling the turbine blades is required, and a variety of cooling techniques have been introduced. The majority of the previous research on pin fins has focused on circular fins using a wedge duct to apply a constant temperature and uniform heat flux to the end wall and pin fin surfaces. The present study compares seven oblong pin-fins to seven circular pin fins in a wedge duct with a Reynolds number range of 10,000–50,000 and a constant heat flux (surface) of 3280 W/m2 applied to the endwall and surfaces of the oblong pin fin. The results indicate that the friction factor for oblong fins is 14% lower than for circular pin fins. The thermal performance factor is increased by 11.4%. The thermal performance factor can be improved by using oblong pin fins with higher Reynolds numbers.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-04","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43607522","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}
Lingling Chen, Y. Shi, Haoqi Yang, Yalin Shi, Qingzhen Yang
Abstract This paper numerically studied supersonic film cooling performance in a convergent divergent nozzle, for the purpose of exploring and extending the knowledge of high-efficient cooling techniques for exhaust nozzles in real working conditions. The work was conducted with a steady state RANS approach with the SST turbulence model. The boundary conditions were chosen to mimic a real engine condition. The slot height and the inlet pressure ratio were varied to investigate the effect of the geometrical condition and flow condition. The flow field was analyzed in detail to study the phenomena of the supersonic secondary flow injected into the transonic flow. Besides the cooling effectiveness, the thrust coefficient and the discharge coefficient for the cooling cases were discussed together with the baseline case, to evaluate the influence of the supersonic coolant injection on the nozzle performance. The work can be a basis for the design of cooling schemes in an aero-engine exhaust nozzle.
{"title":"Detailed investigation of supersonic film cooling performance in a convergent divergent nozzle","authors":"Lingling Chen, Y. Shi, Haoqi Yang, Yalin Shi, Qingzhen Yang","doi":"10.1515/tjj-2022-0080","DOIUrl":"https://doi.org/10.1515/tjj-2022-0080","url":null,"abstract":"Abstract This paper numerically studied supersonic film cooling performance in a convergent divergent nozzle, for the purpose of exploring and extending the knowledge of high-efficient cooling techniques for exhaust nozzles in real working conditions. The work was conducted with a steady state RANS approach with the SST turbulence model. The boundary conditions were chosen to mimic a real engine condition. The slot height and the inlet pressure ratio were varied to investigate the effect of the geometrical condition and flow condition. The flow field was analyzed in detail to study the phenomena of the supersonic secondary flow injected into the transonic flow. Besides the cooling effectiveness, the thrust coefficient and the discharge coefficient for the cooling cases were discussed together with the baseline case, to evaluate the influence of the supersonic coolant injection on the nozzle performance. The work can be a basis for the design of cooling schemes in an aero-engine exhaust nozzle.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-03-27","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46450552","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 A design for a turbine pre-swirl system with impeller cavity is proposed to improve the quality of cooling air supplied to the turbine blades of an aeroengine. Impeller cavity is analyzed in order to increase the system pressure ratio with a low system power consumption at the system outlet. Theoretical and numerical investigation are used to investigate the flow characteristics in an impeller cavity pre-swirl system. The conclusions in this study indicate that the impeller structure can increases the pressure ratio by changing the power consumption and distribution of the absolute velocity in the impeller cavity and system outlet. To obtain high pressure ratio and low power consumption, the impeller should have a structure with a high outlet installation radius and low outlet angle. The highest increase in the pressure ratio compared with the empty cavity pre-swirl system is 6.4% and the corresponding increase in the power consumption is 2620 W.
{"title":"A novel design of impeller cavity pre-swirl system for efficient supercharging and low power consumption","authors":"Wenlei Lian, Yu Zhu, Xiaoming Chen, Zhixiang Zhou, Yong-Yi Huang","doi":"10.1515/tjj-2022-0038","DOIUrl":"https://doi.org/10.1515/tjj-2022-0038","url":null,"abstract":"Abstract A design for a turbine pre-swirl system with impeller cavity is proposed to improve the quality of cooling air supplied to the turbine blades of an aeroengine. Impeller cavity is analyzed in order to increase the system pressure ratio with a low system power consumption at the system outlet. Theoretical and numerical investigation are used to investigate the flow characteristics in an impeller cavity pre-swirl system. The conclusions in this study indicate that the impeller structure can increases the pressure ratio by changing the power consumption and distribution of the absolute velocity in the impeller cavity and system outlet. To obtain high pressure ratio and low power consumption, the impeller should have a structure with a high outlet installation radius and low outlet angle. The highest increase in the pressure ratio compared with the empty cavity pre-swirl system is 6.4% and the corresponding increase in the power consumption is 2620 W.","PeriodicalId":50284,"journal":{"name":"International Journal of Turbo & Jet-Engines","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-03-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"48359430","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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