多级轴流压气机甚高压部件设计探究
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摘要
以大涵道比涡扇发动机总增压比由50∶1提升至70∶1为目标,设计了串接在某10级23∶1高压压气机之后的5级2.2∶1轴流压气机甚高压部件,探究以全轴流方式提高总增压比的方案的可行性。通过部件总体与一维设计,S2通流反问题与叶片造型,计算流体力学验证,在采用了各级正预旋、转子尖部大落后角、静子正弯等措施后实现了该设计。研究表明:在达到设计指标的情况下,该多级轴流甚高压部件的叶尖间隙可选择为0.2mm,若取较为常规的0.3mm叶尖间隙,则其大轮毂比、相对大叶尖间隙等几何特征,将导致失速裕度下降明显,稳定工作范围变窄。另外,在结构方面,全轴流甚高压部件方案还需要解决叶片数量巨大,级成本提高等问题。
For increasing the overall pressure ratio from 50∶1 to 70∶1 of the high bypass ratio turbofan engine,a 5-stage ultra-high-pressure-ratio axial-flow component with pressure ratio of 2.2∶1 was designed,and connected to the exit of a 10-stage high-pressure compressor with pressure ratio of 23∶1.The feasibility to improve the overall pressure ratio with a complete axial-flow configuration was explored.The design was implemented through the process including component 1-D overall aerodynamic design,S2 through-flow inverse problem and blade modeling and 3-D computational fluid dynamics verification.The specific methods such as positive pre-whirl at each stage,large-deviation-angle rotor tip and positive bowed stator were used in the process.Under the condition of achieving the prescribed aerodynamic design parameters,it's recommended to set the rotor tip clearances to 0.2 mm rather than 0.3 mm,otherwise the compressor stall margin would drop significantly and its stable working range would be narrowed due to the geometric characteristics of large hub-tip ratio and the relatively larger rotor tip clearances.In the terms of structure,the design may lead to some issues that must be solved before the compressors enter into service,such as a huge number of blades and the higher machining costs of each stage.
For increasing the overall pressure ratio from 50∶1 to 70∶1 of the high bypass ratio turbofan engine,a 5-stage ultra-high-pressure-ratio axial-flow component with pressure ratio of 2.2∶1 was designed,and connected to the exit of a 10-stage high-pressure compressor with pressure ratio of 23∶1.The feasibility to improve the overall pressure ratio with a complete axial-flow configuration was explored.The design was implemented through the process including component 1-D overall aerodynamic design,S2 through-flow inverse problem and blade modeling and 3-D computational fluid dynamics verification.The specific methods such as positive pre-whirl at each stage,large-deviation-angle rotor tip and positive bowed stator were used in the process.Under the condition of achieving the prescribed aerodynamic design parameters,it's recommended to set the rotor tip clearances to 0.2 mm rather than 0.3 mm,otherwise the compressor stall margin would drop significantly and its stable working range would be narrowed due to the geometric characteristics of large hub-tip ratio and the relatively larger rotor tip clearances.In the terms of structure,the design may lead to some issues that must be solved before the compressors enter into service,such as a huge number of blades and the higher machining costs of each stage.
引文
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[23]DENTON J D.The calculation of three-dimensional viscous flow through multistage turbomachines[J].Journal of Turbomachinery,1992,114(1):18-26.
[2]LIANG Chunhua,SUN Mingxia,LIU Hongxia.Feature analysis on US military aircraft engine advanced technology programs[J].Journal of Aerospace Power,2014,29(7):1701-1709.(in English)
[3]孙明霞,梁春华.VAATE计划下的革新性发动机[J].航空科学技术,2011(5):18-21.SUN Mingxia,LIANG Chunhua.Innovative engine under VAATE program[J].Aeronautical Science and Technology,2011(5):18-21.(in Chinese)
[4]郑天慧,孟令勇.美国下一代军用发动机项目HEETE持续推进[J].国际航空,2009(6):76-77.ZHENG Tianhui,MENG Lingyong.The United States continues advancement of next generation military engine project HEETE[J].International Aviation,2009(6):76-77.(in Chinese)
[5]AIAA Air Breathing Propulsion Technical Committee.The versatile affordable advanced turbine engines(VAATE)initiative[R].AIAA Position Paper,2006.
[6]林左鸣.世界航空发动机手册[M].北京:航空工业出版社,2012.
[7]WARWICK G.R-Rs US unit aims for combat engine business[EB/OL].[2008-06-23].http:∥aviationweek.com/awin/r-r-s-us-unit-aims-combat-engine-business.
[8]LUNDBLADH A,BANK R V D,AVELLN R,et al.Turbines and structures for ultra-high pressure ratio aero-engines[R].ASME Paper GT2016-57307,2016.
[9]BANK R V D,DONNERHACK S,RAE A,et al.LEM-COTEC:improving the core-engine thermal efficiency[R].ASME Paper GT2014-25040,2014.
[10]BANK R V D,DONNERHACK S,RAE A,et al.Compressors for ultra-high-pressure-ratio aero-engines[J].CEAS Aeronautical Journal,2016,7(3):455-470.
[11]NORRIS G.Rolls-Royce details advance and ultrafan test plan[EB/OL].[2014-08-25].http:∥aviationweek.com/commercial-aviation/rolls-royce-details-advance-and-ultrafan-test-plan.
[12]WENNERSTROM A J.Aircraft turbofan engine with multiple high-pressure core modules not concentric with the engine centerline:US20150260127[P].2015-09-17.
[13]高效节能发动机文集编委会.高效节能发动机文集:第3分册风扇、压气机设计与试验[M].北京:航空工业出版社,1991.
[14]李杰,陈光,吕跃进.世界著名商用航空发动机要览[M].北京:航空工业出版社,2016.
[15]陈光.航空发动机结构设计分析[M].2版.北京:北京航空航天大学出版社,2014.
[16]HEARSEY R M.A revised computer program for axial compressor design[M].[S.l.]:Aerospace Research Laboratories,1975.
[17]FROST G R,WENNERSTROM A J.The design of axial compressor airfoils using arbitrary camber lines[R].[S.l.]:Aerospace Reserch Laboratories 73-0107,AD 765165,1973.
[18]《航空发动机设计手册》总编委会.航空发动机设计手册:第8册压气机[M].北京:航空工业出版社,2000.
[19]桂幸民,腾金芳,刘宝杰,等.航空压气机气动热力学理论与应用[M].上海:上海交通大学出版社,2014.
[20]章诚.多级轴流压气机甚高压部件设计探究[D].北京:北京航空航天大学,2018.ZHANG Cheng.Design inquiry of an ultra-high-pressureratio component of multi-stage axial flow compressors[D].Beijing:Beijing University of Aeronautics and Astronautics,2018.(in Chinese)
[21]郭然,贾力平,樊小莉,等.NUMECA系列教程[M].北京:机械工业出版社,2013.
[22]刘宝杰,张帅,于贤君,等.单级低速模拟轴流压气机实验台改进设计[J].航空动力学报,2018,33(7):1665-1675.LIU Baojie,ZHANG Shuai,YU Xianjun,et al.Optimization design of a single stage low speed simulation axial compressor test facility[J].Journal of Aerospace Power,2018,33(7):1665-1675.(in Chinese)
[23]DENTON J D.The calculation of three-dimensional viscous flow through multistage turbomachines[J].Journal of Turbomachinery,1992,114(1):18-26.