双模态冲压发动机Ma6性能潜力受燃烧室入口参数影响的灵敏度分析
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摘要
为系统掌握燃烧室入口参数对双模态冲压发动机性能潜力的影响,采用双模态冲压发动机燃烧室工作过程的一维分析方法,在飞行马赫数为6的不同燃烧室工作工况(即不同特征马赫数Mac)条件下,保持进气道捕获流量不变,研究了燃烧室入口马赫数Main和总压恢复系数σin对性能潜力(燃料比冲Isp,f)的影响。获得的数据表明,冲压发动机的性能潜力受燃烧室入口马赫数和总压恢复系数两者的综合影响,燃烧室入口马赫数越低、总压恢复系数越高,发动机性能潜力越大;在常见的燃烧室入口参数范围内(入口马赫数∈[2.4 3.5],入口总压恢复系数∈[0.3 0.7]),入口马赫数每减小0.1,燃料比冲增加约1.1%~1.8%,入口总压恢复系数越小,影响程度越大;入口总压恢复系数每增大0.1,燃料比冲增大约2.4%~4.0%,入口马赫数越大影响程度越大;存在燃烧室入口条件不同,但是发动机的比冲性能潜力相同的情况。对燃烧室分段过程的损失分析表明,在入口为超声速、加热段为亚声速(特征马赫数不大于1)工况条件下,入口马赫数增加导致燃烧区前激波串损失增大,是比冲性能降低的主要原因;燃烧室特征马赫数越大,燃烧过程导致的损失越大;从燃烧室入口到尾喷管出口全过程总压损失越小,获得的比冲性能越大。推导和拟合了冲压发动机冲量差燃料比冲随燃烧室入口马赫数和总压恢复系数的灵敏度关系式,与被拟合数据的差异在3%以内,该关系式可用于双模态冲压发动机部件参数匹配与流道一体化设计工作。
To obtain a systematic understanding about the effects of supersonic combustor incoming parameters on a Dual-mode scramjet performance potential(fuel specific impulse Isp,f),one dimensional analysis method is applied to study the relationship between performance and incoming parameters,including Mach number(Main)and total pressure recovery coefficient(σin),while the flight Mach number is 6,the incoming massflow rate is fixed,and the characteristic Mach number(Mac)in heating sections is varied. Results indicate that the performance potential is influenced synthetically by these incoming parameters. In a common range of the combustor inflow parameters(Main∈[2.4 3.5],σin∈[0.3 0.7]),Isp,f increases by 1.1%~1.8% while Main decrease by 0.1. The higher σin is,the greater influence of Main has. Isp,f increases by 2.4%~4.0% while σin increase by 0.1.The higher Main is,the more influence of σin has. There are cases with the same specific impulse but different incoming parameters. Loss analysis for combustor subsections shows that increasing of Main leads to great loss in its pre-combustion shock train process with supersonic incoming flow and subsonic heating process. This is the main reason to make Isp,f decreased. More loss will be deduced in combustion process when Mac increases in heating section. The formulized sensitivity relationship between the potential performance and combustor incoming parameters is deduced. The difference from the original data is less than 3%. The relationship can be used to match the processes between engine components.
To obtain a systematic understanding about the effects of supersonic combustor incoming parameters on a Dual-mode scramjet performance potential(fuel specific impulse Isp,f),one dimensional analysis method is applied to study the relationship between performance and incoming parameters,including Mach number(Main)and total pressure recovery coefficient(σin),while the flight Mach number is 6,the incoming massflow rate is fixed,and the characteristic Mach number(Mac)in heating sections is varied. Results indicate that the performance potential is influenced synthetically by these incoming parameters. In a common range of the combustor inflow parameters(Main∈[2.4 3.5],σin∈[0.3 0.7]),Isp,f increases by 1.1%~1.8% while Main decrease by 0.1. The higher σin is,the greater influence of Main has. Isp,f increases by 2.4%~4.0% while σin increase by 0.1.The higher Main is,the more influence of σin has. There are cases with the same specific impulse but different incoming parameters. Loss analysis for combustor subsections shows that increasing of Main leads to great loss in its pre-combustion shock train process with supersonic incoming flow and subsonic heating process. This is the main reason to make Isp,f decreased. More loss will be deduced in combustion process when Mac increases in heating section. The formulized sensitivity relationship between the potential performance and combustor incoming parameters is deduced. The difference from the original data is less than 3%. The relationship can be used to match the processes between engine components.
引文
[1]牛志刚,费立群,冯守义,等.冲压发动机模型燃烧室低压燃烧性能试验[J].推进技术,2011,32(4).(NIU Zhi-gang,FEI Li-qun,FENG Shou-yi,et al.Experiment on Combustion Performance of a Ramjet Prototype Combustor in Low Pressure Condition[J].Journal of Propulsion Technology,2011,32(4).)
[2]罗文雷,潘宇,谭建国,等.模型冲压发动机低压条件下燃烧效率试验[J].推进技术,2010,31(3).(LUO Wen-lei,PAN Yu,TAN Jian-guo,et al.Experimental Investigation on Combustion Efficiency of the Ramjet Model at Low Pressure[J].Journal of Propulsion Technology,2010,31(3).)
[3]Heiser W H,Pratt D T.Hypersonic Airbreathing Propulsion[M].Washington DC:American Institute of Aeronautics and Astronautics,Inc,1994.
[4]徐旭,陈兵,徐大军.冲压发动机原理及技术[M].北京:北京航空航天大学出版社,2014.
[5]邓远灏.超燃冲压发动机隔离段流动特性研究[D].西安:西北工业大学,2005.
[6]董贤蒙.超燃冲压发动机建模与仿真研究[D].长沙:国防科学技术大学,2011.
[7]陈军,白菡尘,柳森.燃烧室工作过程对冲压发动机性能潜力的影响研究[J].实验流体力学,2013,(3).
[8]Chen Jun,Bai Hanchen.Study of the Combustion Organization Method’s Effects on Performance of Ramjet[C].Shanghai:Sino-Russia Hypersonic Flow Conference,2011.
[9]陈军,白菡尘,柳森.化学平衡效应对冲压燃烧室性能潜力与工作过程关系的影响分析[J].推进技术,2013,34(10).(CHEN Jun,BAI Han-chen,LIU Sen.Effects of Chemical Equilibrium on Relation Between Performance and Process for Scramjet Combustor[J].Journal of Propulsion Technology,2013,34(10).)
[10]Ikui T,Matsuo K,Nagai M.The Mechanism of PseudoShock Waves[J].Bulletin of JSME,1974,17(108):731-739.
[11]O’Brien T F,Starkey R P,Lewis M J,et al.QuasiOne-Dimensional High-Speed Engine Model with Finite-Rate Chemistry[J].Journal of Propulsion and Power,2001,17(6).
[12]Theodore Brooke Smith.Development and Ground Testing of Direct Measuring Skin Friction Gages for High Enthalpy Supersonic Flight Tests[D].Blacksburg:Virginia Polytechnic Institute and State University,2001.
[13]Holger Babinsky,John K Harvery.Shock Wave Boundary-Layer Interaction[M].白菡尘译,长沙:国防工业出版社,2015:15-18.
[14]李大鹏.煤油双模态冲压发动机燃烧室工作过程研究[D].长沙:国防科技大学,2006,10.
[15]孙波,张堃元,金志光,等.流线追踪Busemann进气道设计参数的选择[J].推进技术,2007,28(1).(SUN Bo,ZHANG Kun-yuan,JIN Zhi-guang,et al.Selection of Design Parameters for Stream Traced Hypersonic Busemann Inlets[J].Journal of Propulsion Technology,2007,28(1).)
[16]李淑艳.高超声速进气道设计及性能研究[D].西安:西北工业大学,2006.
[17]南向军,张堃元.采用新型基准流场的高超声速内收缩进气道性能分析[J].宇航学报,2012,33(2).
[18]贺旭照,周正,倪鸿礼.密切内锥乘波前体进气道一体化设计和性能分析[J].推进技术,2012,33(4).(HE Xu-zhao,ZHOU Zheng,NI Hong-li.Integrated Design Methods and Performance Analyses of Osculating Inward Turning Cone Waverider Forebody Inlet(OICWI)[J].Journal of Propulsion Technology,2012,33(4).)
[19]王振国,梁剑寒,丁猛,等.高超声速飞行器动力系统研究进展[J].力学进展,2009,39(6).
[2]罗文雷,潘宇,谭建国,等.模型冲压发动机低压条件下燃烧效率试验[J].推进技术,2010,31(3).(LUO Wen-lei,PAN Yu,TAN Jian-guo,et al.Experimental Investigation on Combustion Efficiency of the Ramjet Model at Low Pressure[J].Journal of Propulsion Technology,2010,31(3).)
[3]Heiser W H,Pratt D T.Hypersonic Airbreathing Propulsion[M].Washington DC:American Institute of Aeronautics and Astronautics,Inc,1994.
[4]徐旭,陈兵,徐大军.冲压发动机原理及技术[M].北京:北京航空航天大学出版社,2014.
[5]邓远灏.超燃冲压发动机隔离段流动特性研究[D].西安:西北工业大学,2005.
[6]董贤蒙.超燃冲压发动机建模与仿真研究[D].长沙:国防科学技术大学,2011.
[7]陈军,白菡尘,柳森.燃烧室工作过程对冲压发动机性能潜力的影响研究[J].实验流体力学,2013,(3).
[8]Chen Jun,Bai Hanchen.Study of the Combustion Organization Method’s Effects on Performance of Ramjet[C].Shanghai:Sino-Russia Hypersonic Flow Conference,2011.
[9]陈军,白菡尘,柳森.化学平衡效应对冲压燃烧室性能潜力与工作过程关系的影响分析[J].推进技术,2013,34(10).(CHEN Jun,BAI Han-chen,LIU Sen.Effects of Chemical Equilibrium on Relation Between Performance and Process for Scramjet Combustor[J].Journal of Propulsion Technology,2013,34(10).)
[10]Ikui T,Matsuo K,Nagai M.The Mechanism of PseudoShock Waves[J].Bulletin of JSME,1974,17(108):731-739.
[11]O’Brien T F,Starkey R P,Lewis M J,et al.QuasiOne-Dimensional High-Speed Engine Model with Finite-Rate Chemistry[J].Journal of Propulsion and Power,2001,17(6).
[12]Theodore Brooke Smith.Development and Ground Testing of Direct Measuring Skin Friction Gages for High Enthalpy Supersonic Flight Tests[D].Blacksburg:Virginia Polytechnic Institute and State University,2001.
[13]Holger Babinsky,John K Harvery.Shock Wave Boundary-Layer Interaction[M].白菡尘译,长沙:国防工业出版社,2015:15-18.
[14]李大鹏.煤油双模态冲压发动机燃烧室工作过程研究[D].长沙:国防科技大学,2006,10.
[15]孙波,张堃元,金志光,等.流线追踪Busemann进气道设计参数的选择[J].推进技术,2007,28(1).(SUN Bo,ZHANG Kun-yuan,JIN Zhi-guang,et al.Selection of Design Parameters for Stream Traced Hypersonic Busemann Inlets[J].Journal of Propulsion Technology,2007,28(1).)
[16]李淑艳.高超声速进气道设计及性能研究[D].西安:西北工业大学,2006.
[17]南向军,张堃元.采用新型基准流场的高超声速内收缩进气道性能分析[J].宇航学报,2012,33(2).
[18]贺旭照,周正,倪鸿礼.密切内锥乘波前体进气道一体化设计和性能分析[J].推进技术,2012,33(4).(HE Xu-zhao,ZHOU Zheng,NI Hong-li.Integrated Design Methods and Performance Analyses of Osculating Inward Turning Cone Waverider Forebody Inlet(OICWI)[J].Journal of Propulsion Technology,2012,33(4).)
[19]王振国,梁剑寒,丁猛,等.高超声速飞行器动力系统研究进展[J].力学进展,2009,39(6).