微型发动机部件改进与整机性能测试
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摘要
微型涡轮发动机具有高能量密度和高推重比优势,是满足先进低成本微小型空中武器系统推进动力需求的先进动力装置。深入开展微型涡轮发动机技术研究,对加速推进微型涡轮发动机的应用进程、提高我军的快速侦察打击能力具有重要意义。
本文基于直径16厘米MTE-C原理样机的部件数值模拟及试验结果,采用Gasturb软件建立了整机性能仿真模型,完成了MTE-C发动机设计点及非设计点的性能估算,获得了压气机和涡轮在地面及不同海拔高度条件下的共同工作线,计算了发动机推力、单位推力、耗油率、空气流量、压气机压比和燃烧室出口温度随转速的变化规律,获得了该发动机的速度高度特性,为该发动机的进一步改进和试验研究提供了理论依据和指导。
论文对保形通道式扩压器三维设计的参数分布进行了对比研究,得出了上凸式分布设计效果较好的结果。采用样条插值控制通道面积分布,改进了保形通道式扩压器设计技术,并总结了保形通道式扩压器的3个设计要点。完成了MTE-C发动机原型扩压器的改进设计,通过优选获得的保形通道式扩压器在设计点总压恢复系数达到0.875,比原型扩压器提高了5.5个百分点,同时扩压度明显提高,静压比从原型的1.14提高到1.23。
探讨了整体叶片式导向器的设计方法,研究了子午流道型线、叶根叶片角分布和叶尖叶片角分布对导向器性能的影响规律,获得了整体叶片式导向器的基本设计准则。完成了MTE-C发动机原型导向器的数值模拟分析,针对其不足之处,采用整体叶片式导向器设计准则进行了改进设计,导向器总压恢复系数从0.909提高到0.918,且出口流场均匀度明显改善。并完成了原型及改进型涡轮级的数值模拟,涡轮级效率从70.9%大幅提高到84.2%,这些数值结果显示了导向器改进具有明显的效果。
针对扩压器及导向器改进效果,采用整机性能仿真模型,预估了发动机总体性能改善效果,通过调节发动机尾喷管出口面积,使发动机部件达到最优匹配效果,指导发动机地面台架试车试验。完成了原型发动机的地面台架试车,运转至85000rpm,此时发动机空气流量约为0.618kg/s,压气机级压比约3.45,发动机推力295N,耗油率0.0488g/(N·s)(即1.76kg/(DaN·h))。完成了换装保形通道式扩压器的发动机地面台架试车,在转速70000rpm时,发动机空气流量增加约6.5%,推力提高11%,耗油率降低7%。完成了换装整体叶片式导向器的发动机地面台架试车,在转速69000rpm时,发动机流量增大4.0%,推力提高36.8%,耗油率高了5.1%。台架试车结果既验证了数值模拟的正确,又证明了本文研究所发展的两项改进措施是有效的。
Micro turbine engine, with its merits in high power density and thrust-weight ratio, can be used as propulsion system and satisfy the requirements of advanced micro vehicles at low payment. Development of micro turbine engine technique can contribute to the using process of micro turbine engine and enhance our army ability of spy-beating quickly.
Based on the simulation and experiment results of MTE-C (16 centimeters), a new integer engine model is built using the software Gasturb. At design and off-design points, the performance of MTE-C is obtained, as well as working line of compressor and turbine at different height. Changing rules of rotate speed to thrust、unit thrust、specific fuel consumption、air flux、pressure ratio of compressor and temperature at combustor’s exit are also gained. These all contribute to the next stage of development and experiments of micro turbine engine.
Comparison of prototype diffuser performance at different design parameters is also finished in this paper, and design-rules of distribution on the convex are obtained. In order to control the area of flow path, spline interpolation is used in this paper, which improves the design technic of prototype diffuser. Three main design features are summarized as well. With improvement, the total pressure recovery of this new diffuser used in MTE-C can be 0.875, 5.5% higher than the former. At the same time, diffuser degree is increased greatly, pressure ratio raised from 1.14 to 1.23.
The influences of meridian-passage and vane angle distribution are researched, and the design cretirion of the vane turbine nozzle is obtained. The numerical simulation of original turbine nozzle is completed, and a vane turbine nozzle is designed to promote the turbine performance. The total pressure recovery of turbine nozzle increases from 0.909 to 0.918 and the quality of nozzle exit is improved obviously. The numerical simulation results of turbine stages with original and modified turbine nozzles show that the performance of turbine with improved vane nozzle promote greatly. The efficiency of turbine stage at design point increases from 70.9% to 84.2%.
The performance improvements of MTE-C are predicted with the engine performance simulation model, for the prototype diffuser and modified turbine nozzle respectively. The area of engine nozzle is adjusted for optimal matching of engine parts. It directs the ground test of the micro turbine engine.
The ground test of MTE-C is completed and the ration speed reached 85krpm. The air mass is 0.618kg/s, the pressure ratio is 3.45, the engine thrust is 295N, and the specific fuel consumption is 0.0488g/(N·s), at this speed. The ground test of MTE-C with prototype diffuser is completed. At 70000rpm, the air mass and engine thrust increase by 6.5% and 11% respectively, and the specific fuel consumption decreases by 7%. The ground test of MTE-C with modified turbine nozzle is completed. At 69000rpm, the air mass, engine thrust and the specific fuel consumption increase by 4%, 36.8% and 5.1% respectively.
本文基于直径16厘米MTE-C原理样机的部件数值模拟及试验结果,采用Gasturb软件建立了整机性能仿真模型,完成了MTE-C发动机设计点及非设计点的性能估算,获得了压气机和涡轮在地面及不同海拔高度条件下的共同工作线,计算了发动机推力、单位推力、耗油率、空气流量、压气机压比和燃烧室出口温度随转速的变化规律,获得了该发动机的速度高度特性,为该发动机的进一步改进和试验研究提供了理论依据和指导。
论文对保形通道式扩压器三维设计的参数分布进行了对比研究,得出了上凸式分布设计效果较好的结果。采用样条插值控制通道面积分布,改进了保形通道式扩压器设计技术,并总结了保形通道式扩压器的3个设计要点。完成了MTE-C发动机原型扩压器的改进设计,通过优选获得的保形通道式扩压器在设计点总压恢复系数达到0.875,比原型扩压器提高了5.5个百分点,同时扩压度明显提高,静压比从原型的1.14提高到1.23。
探讨了整体叶片式导向器的设计方法,研究了子午流道型线、叶根叶片角分布和叶尖叶片角分布对导向器性能的影响规律,获得了整体叶片式导向器的基本设计准则。完成了MTE-C发动机原型导向器的数值模拟分析,针对其不足之处,采用整体叶片式导向器设计准则进行了改进设计,导向器总压恢复系数从0.909提高到0.918,且出口流场均匀度明显改善。并完成了原型及改进型涡轮级的数值模拟,涡轮级效率从70.9%大幅提高到84.2%,这些数值结果显示了导向器改进具有明显的效果。
针对扩压器及导向器改进效果,采用整机性能仿真模型,预估了发动机总体性能改善效果,通过调节发动机尾喷管出口面积,使发动机部件达到最优匹配效果,指导发动机地面台架试车试验。完成了原型发动机的地面台架试车,运转至85000rpm,此时发动机空气流量约为0.618kg/s,压气机级压比约3.45,发动机推力295N,耗油率0.0488g/(N·s)(即1.76kg/(DaN·h))。完成了换装保形通道式扩压器的发动机地面台架试车,在转速70000rpm时,发动机空气流量增加约6.5%,推力提高11%,耗油率降低7%。完成了换装整体叶片式导向器的发动机地面台架试车,在转速69000rpm时,发动机流量增大4.0%,推力提高36.8%,耗油率高了5.1%。台架试车结果既验证了数值模拟的正确,又证明了本文研究所发展的两项改进措施是有效的。
Micro turbine engine, with its merits in high power density and thrust-weight ratio, can be used as propulsion system and satisfy the requirements of advanced micro vehicles at low payment. Development of micro turbine engine technique can contribute to the using process of micro turbine engine and enhance our army ability of spy-beating quickly.
Based on the simulation and experiment results of MTE-C (16 centimeters), a new integer engine model is built using the software Gasturb. At design and off-design points, the performance of MTE-C is obtained, as well as working line of compressor and turbine at different height. Changing rules of rotate speed to thrust、unit thrust、specific fuel consumption、air flux、pressure ratio of compressor and temperature at combustor’s exit are also gained. These all contribute to the next stage of development and experiments of micro turbine engine.
Comparison of prototype diffuser performance at different design parameters is also finished in this paper, and design-rules of distribution on the convex are obtained. In order to control the area of flow path, spline interpolation is used in this paper, which improves the design technic of prototype diffuser. Three main design features are summarized as well. With improvement, the total pressure recovery of this new diffuser used in MTE-C can be 0.875, 5.5% higher than the former. At the same time, diffuser degree is increased greatly, pressure ratio raised from 1.14 to 1.23.
The influences of meridian-passage and vane angle distribution are researched, and the design cretirion of the vane turbine nozzle is obtained. The numerical simulation of original turbine nozzle is completed, and a vane turbine nozzle is designed to promote the turbine performance. The total pressure recovery of turbine nozzle increases from 0.909 to 0.918 and the quality of nozzle exit is improved obviously. The numerical simulation results of turbine stages with original and modified turbine nozzles show that the performance of turbine with improved vane nozzle promote greatly. The efficiency of turbine stage at design point increases from 70.9% to 84.2%.
The performance improvements of MTE-C are predicted with the engine performance simulation model, for the prototype diffuser and modified turbine nozzle respectively. The area of engine nozzle is adjusted for optimal matching of engine parts. It directs the ground test of the micro turbine engine.
The ground test of MTE-C is completed and the ration speed reached 85krpm. The air mass is 0.618kg/s, the pressure ratio is 3.45, the engine thrust is 295N, and the specific fuel consumption is 0.0488g/(N·s), at this speed. The ground test of MTE-C with prototype diffuser is completed. At 70000rpm, the air mass and engine thrust increase by 6.5% and 11% respectively, and the specific fuel consumption decreases by 7%. The ground test of MTE-C with modified turbine nozzle is completed. At 69000rpm, the air mass, engine thrust and the specific fuel consumption increase by 4%, 36.8% and 5.1% respectively.
引文
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6.黄国平,梁德旺,何志强,大型飞机辅助动力装置与微型涡轮发动机技术特点对比,航空动力学报,Vol.23(2):383-388。
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3.周俊虎,刘茂省,杨卫娟,等,微燃烧发动机应用于飞行器的性能分析,燃烧科学与技术,2006,Vol.12 (3): 193-196。
4.肖鹏,“捕食者”的新利爪-LOCAAS系统,科学(中文版),2007,5:30-31。
5.赵豫,于尔铿,新型发散式发电装置,电网技术,2004,Vol.28 (4):47-50。
6.黄国平,梁德旺,何志强,大型飞机辅助动力装置与微型涡轮发动机技术特点对比,航空动力学报,Vol.23(2):383-388。
7.张宝诚,航空发动机实验和测试技术,北京,北京航空航天大学出版社,2005。
8. Alan H Epstein.Micro Turbine Engines For Soldier Power. The Defense Science and Technology Seminar, 2000.
9. Epstein AH, etc. Micro Gas Turbine Generators. Mas-sachusetts Inst. Of Tech., Gas Turbine Lab, Final Rept.1995-2000.
10. Han Sojrg Schlip, Fabrication of Turbine-Compressor-Shaft Assembly For Micro Gas Turbine Engine. Mater'sThe-sis, Stanford Univ., December2000.
11. Han Sojrg Schlip, KangS, StampflJ, CooperAG, PrinzFB. Application of the Mold SDM Process to the Fabrication of Ceramic Parts for a Micro Gas Turbine engine. HeinrichJG(ed.), Proceedings Ceramics Materials and Components for Engines, Germany, 2000.
12. Kousuke Isomura, etc: Development of Micro-Turbo Charger and Micro-Combustor as Feasibility Studies of Three-Dimensional Gas Turbine at Micro-Scale[R]. Atlanta: Proceedings of ASME Turbo Expo 2003, No.GT-2003-38151.
13. Jones, M. M. A. C. E. J. A., MINIATURE TURBOJET DEVELOPMENT AT HAMILTON SUNDSTRAND THE TJ-50,TJ-120 AND TJ-30[C], In 2nd AIAA "Unmanned Unlimited" Systems, Technologies, and Operations—Aerospac, : San Diego, California, (2003) .
14. H. S. Yang,M. S. Rho,etc. Permanent Magnet High Speed Stater/Generator System Development Directly Coupled to Gas Turbine Engine for Mobile Auxiliary Power Unit[R]. Vienna, Austria: Proceedings of ASME Turbo Expo 2004, No.GT2004-53165.
15. Yungmo Kang, Robert Mckeirnan. Annular Recuperator Development and Performance Test for 200kW Microturbine[R]. Atlanta, Georgia: Proceedings of ASME Turbo Expo 2003, No.GT2003-38522.
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