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高马赫数超燃冲压发动机技术研究进展

岳连捷 张旭 张启帆 陈科挺 李进平 陈昊 姚卫 仲峰泉 李飞 王春 陈宏

岳连捷, 张旭, 张启帆, 陈科挺, 李进平, 陈昊, 姚卫, 仲峰泉, 李飞, 王春, 陈宏. 高马赫数超燃冲压发动机技术研究进展. 力学学报, 2022, 54(2): 263-288 doi: 10.6052/0459-1879-21-547
引用本文: 岳连捷, 张旭, 张启帆, 陈科挺, 李进平, 陈昊, 姚卫, 仲峰泉, 李飞, 王春, 陈宏. 高马赫数超燃冲压发动机技术研究进展. 力学学报, 2022, 54(2): 263-288 doi: 10.6052/0459-1879-21-547
Yue Lianjie, Zhang Xu, Zhang Qifan, Chen Keting, Li Jinping, Chen Hao, Yao Wei, Zhong Fengquan, Li Fei, Wang Chun, Chen Hong. Research progress on high-Mach-number scramjet engine technologies. Chinese Journal of Theoretical and Applied Mechanics, 2022, 54(2): 263-288 doi: 10.6052/0459-1879-21-547
Citation: Yue Lianjie, Zhang Xu, Zhang Qifan, Chen Keting, Li Jinping, Chen Hao, Yao Wei, Zhong Fengquan, Li Fei, Wang Chun, Chen Hong. Research progress on high-Mach-number scramjet engine technologies. Chinese Journal of Theoretical and Applied Mechanics, 2022, 54(2): 263-288 doi: 10.6052/0459-1879-21-547

高马赫数超燃冲压发动机技术研究进展

doi: 10.6052/0459-1879-21-547
详细信息
    作者简介:

    岳连捷, 研究员, 主要研究方向: 超燃冲压发动机及组合循环发动机、高超声速气体动力学. E-mail: yuelj@imech.ac.cn

  • 中图分类号: O35

RESEARCH PROGRESS ON HIGH-MACH-NUMBER SCRAMJET ENGINE TECHNOLOGIES

  • 摘要: 吸气式高超声速飞行在空间运输和国家空天安全领域具有极高价值, 超燃冲压发动机是其核心动力装置. 目前飞行马赫数4.0 ~ 7.0超燃冲压发动机技术日趋成熟, 发展更高速的飞行动力技术成为今后临近空间竞争焦点之一. 本文对飞行马赫数8.0 ~ 10.0 的高马赫数超燃冲压发动机技术进行了分析和综述. 首先论述其亟待解决的关键问题和技术, 分别包括高焓离解与热化学非平衡效应、超高速气流燃料增混与燃烧强化技术、高超声速燃烧与进气压缩的匹配及工作模态、高焓低雷诺数边界层流动及其控制方法、高焓低密度流动/燃烧的热防护技术, 以及高马赫数发动机的地面试验风洞技术. 然后, 进一步介绍了国内外高焓激波风洞与驱动技术以及国内外典型的地面和飞行试验进展. 进而针对推进和热防护的总体性能评估、高马赫数发动机内凸显的高焓离解与热化学非平衡效应、超高速气流燃料增混和燃烧强化技术综述了相关研究进展及结论, 讨论了高马赫数超燃冲压发动机的可行性以及各关键技术的特点. 最后进行了总结并对后续研究提出了几点建议.

     

  • 图  1  不同飞行马赫数$M{a_{\text{f}}}$的自由来流总温$T_0^*$及滞止状态下空气中O和N组分摩尔分数$ {X_{\text{N}}} $$ {X_{\text{O}}} $

    Figure  1.  Freestream stagnation temperature $T_0^*$, and O and N mole fractions $ {X_{\text{N}}} $ and $ {X_{\text{O}}} $ of stagnant air under different flight Mach number $M{a_{\text{f}}}$

    图  2  不同飞行马赫数$M{a_{\text{f}}}$典型燃烧室入口马赫数$M{a_{{\text{in}}}}$条件下理论燃烧效率${\eta _{\text{b}}}$

    Figure  2.  Theoretical combustion efficiency ${\eta _{\text{b}}}$ under typical combustor inflow$M{a_{{\text{in}}}}$of different flight $M{a_{\text{f}}}$

    图  3  压力50.66 kPa和506.6 kPa条件下O2和N2分子振动能松弛时间${\tau _{\text{t}}}$随温度$T$的变化

    Figure  3.  O2 and N2 vibration energy relaxation time vs. temperature T under pressures of 50.66 kPa and 506.6 kPa

    图  4  不同飞行马赫数$M{a_{\text{f}}}$典型燃烧室入口马赫数$M{a_{{\text{in}}}}$条件下射流穿透深度${y_{\text{p}}}$与驻留时间${{\tau }_{{\rm{res}}}}$

    Figure  4.  Jet penetration depth ${y_{\text{p}}}$ and residual time ${{\tau }_{{\rm{res}}}}$ vs. different $M{a_{\text{f}}}$ under typical $M{a_{{\text{in}}}}$

    图  5  不同燃烧室入口马赫数$M{a_{{\text{in}}}}$条件下理论燃烧效率${\eta _{\text{b}}}$和比冲${I_{{\text{sp}}}}$

    Figure  5.  Theoretical combustion efficiency ${\eta _{\text{b}}}$ and specific impulse ${I_{{\text{sp}}}}$ vs. combustor inflow$M{a_{{\text{in}}}}$

    图  6  不同飞行马赫数$M{a_{\text{f}}}$和高度${H_{\text{f}}}$条件下自由来流动压$q$和单位雷诺数$R{e_L}$的等值线

    Figure  6.  Free stream dynamic pressure $q$ and Reynolds number $R{e_L}$ isolines under different $M{a_{\text{f}}}$ and ${H_{\text{f}}}$

    图  7  不同$M{a_{\text{f}}}$${H_{\text{f}}}$条件下自由来流动压$q$和边界层自然转捩长度${L_T}$的等值线

    Figure  7.  Free stream dynamic pressure $q$ and boundary layer natural transition length ${L_T}$ isolines under different $M{a_{\text{f}}}$ and ${H_{\text{f}}}$

    图  8  不同飞行马赫数$M{a_{\text{f}}}$等动压$q$条件下同一位置的无量纲化层流边界层厚度${\delta _{\text{L}}}$

    Figure  8.  Normalized laminar boundary layer thickness ${\delta _{\text{L}}}$ at the same location vs. $M{a_{\text{f}}}$ under a constant flight dynamic pressure $q$

    图  9  烧氢补氧风洞不同总温$T_0^*$试验气流组分摩尔分数

    Figure  9.  Test inflow mole fractions provided by H2-O2-air combustion tunnel under different $T_0^*$

    图  10  烧氢补氧风洞不同空气初始温度${T_{{\text{air}}}}$下试验气流中$ {{\text{H}}_{\text{2}}}{\text{O}} $摩尔分数

    Figure  10.  Test inflow $ {{\text{H}}_{\text{2}}}{\text{O}} $ mole fractions of H2-O2-air combustion tunnel under different initial air temperature ${T_{{\text{air}}}}$

    图  11  激波风洞结构及运行原理

    Figure  11.  Structure and operating principle of shock tunnel

    图  12  JF24激波风洞示意图

    Figure  12.  Schematic of JF24 shock tunnel

    图  13  JF24激波风洞典型的来流总压-时间曲线

    Figure  13.  Typical stagnation pressure time histories of the test inflow by JF24 shock tunnel

    图  14  X-43 A马赫数9.6飞行试验示意图[38]

    Figure  14.  X-43 A Mach 9.6 flight mission profile[38]

    图  15  X-43 A马赫数9.6飞行(F3)与地面发动机试验壁面压力分布对比[40]

    Figure  15.  Comparison of X-43 A engine ground test and Mach 9.6 flight (F3) pressure distribution[40]

    图  16  HiFire-7全尺寸发动机模型T4激波风洞试验典型的沿程压力[50]

    Figure  16.  Typical wall-pressure distributions of HiFire-7 full-scale engine tests in T4 shock tunnel[50]

    图  17  SCRAMSPACE I HEG风洞试验与CFD沿程压力分布对比[58]

    Figure  17.  Typical wall-pressure distributions of SCRAMSPACE I by HEG tunnel test and CFD data[58]

    图  18  HIEST激波风洞试验典型的沿程压力分布[63]

    Figure  18.  Typical wall-pressure distributions based on tests in HIEST shock tunnel[63]

    图  19  高马赫数燃烧室激波风洞试验典型火焰图[66]

    Figure  19.  Typical flame image in high Mach number circular combustor by shock tunnel test[66]

    图  20  不同${k_{AR}}$条件下最大飞行马赫数${(M{a_0})_{\max }}$[68]

    Figure  20.  ${(M{a_0})_{\max }}$ at different ${k_{AR}}$[68]

    图  21  采用传统等压加热和等压-等温混合加热循环的发动机最大推力[74]

    Figure  21.  Thrust vs. cruise Mach number using traditional and compound thermodynamic cycles, respectively[74]

    图  22  隔热涂层与再生冷却组合热防护示意图[79]

    Figure  22.  Schematic diagram of combined active and passive thermal protection systems[79]

    图  23  不同当量比$\varPhi$燃烧室主体的最大壁温${T_{{\rm{wa}}}}$[79]

    Figure  23.  Maximum temperature of combustor main structure ${T_{{\rm{wa}}}}$ at different $\varPhi$[79]

    图  24  不同燃烧室长径比$L/D$条件下${(M{a_0})_{\max }}$[68]

    Figure  24.  Maximum ${(M{a_0})_{\max }}$ at different $L/D$[68]

    图  25  耦合涡轮泵燃料供给系统的二次冷却循环[83]

    Figure  25.  Schematic of recooling cycle with turbine pump fuel supply system[83]

    图  26  再生冷却与二次冷却典型燃烧室沿程壁温[83]

    Figure  26.  Comparison of gas-side wall temperature distributions in recooling cycle and regenerative cooling cycle, respectively[83]

    图  27  不同气体模型二元进气道典型沿程壁温[85]

    Figure  27.  Typical wall-temperature distributions of 2-D inlet using different gas models[85]

    图  28  热力学平衡和非平衡假设下典型的隔离段马赫数云图[13]

    Figure  28.  Typical Mach number contours assuming thermal equilibrium and nonequilibrium, respectively[13]

    图  29  热力学平衡和非平衡隔离段典型静压云图[13]

    Figure  29.  Typical pressure contours assuming thermal equilibrium and nonequilibrium, respectively[13]

    图  30  进气道内首个高温分离区横截面的有效温度分布[10]

    Figure  30.  Typical cross-sectional effective temperature distributions of the first hot pocket in the inlet[10]

    图  31  高马赫数超燃冲压发动机典型的无燃料流场数值纹影[90]

    Figure  31.  Typical numerical schlieren images of scramjet engine without fuel[90]

    图  32  典型OH组分质量分数分布[11]

    Figure  32.  Typical OH mass fraction distribution[11]

    图  33  热力学非平衡燃烧场典型${T_{\text{t}}}$${T_{\text{v}}}$分布[11]

    Figure  33.  Typical ${T_{\text{t}}}$ and ${T_{\text{v}}}$ contours of thermal-nonequilibrium combustion flow[11]

    图  34  稳定燃烧典型的OH组分质量分数分布[12]

    Figure  34.  Typical OH mass fraction distributions of stabilized combustion[12]

    图  35  热力学非平衡燃烧场喷孔附近${T_{\text{t}}}$分布[12]

    Figure  35.  Typical ${T_{\text{t}}}$ contour of thermal-nonequilibrium combustion flow[12]

    图  36  喷管内水组分质量分数与入口值之比云图[7]

    Figure  36.  ${{\text{H}}_{\text{2}}}{\text{O}}$ mass fraction contour of the nozzle normalized by the inlet value[7]

    图  37  进气道OH质量分数分布[14]

    Figure  37.  OH mass fraction contour in the inlet[14]

    图  38  燃烧室和喷管温度分布[14]

    Figure  38.  T contour of combustor and nozzle[14]

    图  39  基于DZFM模型的典型高马赫数发动机流场OH质量分数分布[96]

    Figure  39.  Typical OH mass fraction distribution of high Mach number scramjet based on DZFM model[96]

    图  40  进气道喷注典型OH组分质量分数分布[100]

    Figure  40.  Typical OH mass fraction contour by inlet injection[100]

    图  41  激波诱导燃烧示意图[101]

    Figure  41.  Schematic of shock-induced combustion[101]

    图  42  进气道喷注、燃烧室喷注和组合喷注3种方式在不同当量比$\phi $条件下的推力系数${C_{\rm{T}}}$[103]

    Figure  42.  Thrust coefficient ${C_{\rm{T}}}$ vs. ER $\phi $ by inlet/combustor/combined injection schemes[103]

    图  43  进气道喷注典型进气道出口组分分布[106]

    Figure  43.  Typical inlet outlet species mass fraction contour of inlet injection[106]

    图  44  定制的非对称喷孔排布[106]

    Figure  44.  Tailored asymmetric fuel injection holes[106]

    图  45  多孔介质喷孔和离散喷孔的进气道化学冻结流静温$T$和静压$p$云图[101]

    Figure  45.  $T$ and $p$ contours of chemically frozen flow with porous and porthole inlet injections, respectively[101]

    图  46  非预混和预混补氧喷注示意图[111]

    Figure  46.  Schematic of nonpremixed and premixed oxygen enrichment[111]

    图  47  超混合型支板的典型构型[61, 118]

    Figure  47.  Typical hypermixer strut configurations[61, 118]

    图  48  超混合型支板后缘产生流向涡示意图[119]

    Figure  48.  Schematic diagram of streamwise vortex generation by hypermixer strut[119]

    图  49  抗分离型支板燃烧室的典型质量分数云图[118]

    Figure  49.  Typical contours of (a) schlieren and (b) H2O mass fraction using the separation-resistant strut[118]

    表  1  不同飞行马赫数$M{a_{\text{f}}}$和高度${H_{\text{f}}}$条件下气流的总温$T_0^*$和总压$p_0^*$

    Table  1.   Freestream $T_0^*$ and $p_0^*$ under different flight $M{a_{\text{f}}}$ and ${H_{\text{f}}}$

    $M{a_{\text{f}}}$${H_{\text{f}}}$/km$T_0^*$/K$p_0^*$/MPa
    830261918.7
    3527109.2
    4028314.8
    10303688114.7
    35379057.8
    40391731.0
    12304915575.0
    355025296.2
    405168163.6
    下载: 导出CSV

    表  2  超/高超声速风洞的主要类型和特性

    Table  2.   Typical types and characteristics of supersonic/hypersonic test facilities

    Wind tunnel typeOperating timeTest medium$T_0^*$/K$M{a_{\text{f}}}$Typical wind tunnel[26]
    continuousminutespure air<1000<4AEDC tunnels
    intermittentheat reservoirsecondspure air<2100<7HTF glenn
    combustion heaterminutesvitiated air<2500<88-ft HTT langley
    impulsemillisecondspure air<10000<20LENS
    下载: 导出CSV

    表  3  典型高焓激波风洞

    Table  3.   Typical high-enthalpy shock tunnels

    Tunnel nameAffiliationDriver typeTotal length/mOperating time/msStagnation pressure/MPaMa
    T4[35]University of Queenslandfree piston361 ~ 2<304 ~ 10
    FD21[36]Chinese Academy of Aerospace Aerodynamicsfree piston1092<11>8
    LENS I[28]Calspanheated lightweight gas265 ~ 15<806 ~ 15
    Hypulse[28]GASLlightweight gas and detonation270.5 ~ 7<305 ~ 12
    JF24[37]Institute of Mechanicsgaseous detonation235 ~ 10<208 ~ 12
    下载: 导出CSV

    表  4  T4激波风洞试验来流参数[90]

    Table  4.   T4 shock tunnel test inflow parameters[90]

    TECNTNCN
    $Ma$7.538.42
    $p$/Pa37501877.9
    ${T_{\text{t}}}$/K315.4260.6
    $ {T_{\text{v}}} $/K315.4640.3
    ${Y_{{\text{NO}}}}$0.047890.01677
    ${Y_{\text{O}}}$0.00016$5.75 \times {10^{ - 8}}$
    下载: 导出CSV
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出版历程
  • 收稿日期:  2021-10-25
  • 录用日期:  2021-12-22
  • 网络出版日期:  2021-12-23
  • 刊出日期:  2022-02-17

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