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

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

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^{*, †}

*.
中国科学院力学研究所高温气体动力学国家重点实验室, 北京100190
†.
中国科学院大学工程科学学院, 北京100049

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
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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

RESEARCH PROGRESS ON HIGH-MACH-NUMBER SCRAMJET ENGINE TECHNOLOGIES

Yue Lianjie^{*, †
,},
Zhang Xu^{*, †},
Zhang Qifan^*,
Chen Keting^{*, †},
Li Jinping^{*, †},
Chen Hao^*,
Yao Wei^{*, †},
Zhong Fengquan^{*, †},
Li Fei^{*, †},
Wang Chun^{*, †},
Chen Hong^{*, †}

*.
State Key Laboratory of High Temperature Gas Dynamics, Institute of Mechanics, Chinese Academy of Sciences, Beijing 100190, China
†.
School of Engineering Science, University of Chinese Academy of Sciences, Beijing 100049, China

摘要

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.
- 高马赫数 /
- 超燃冲压发动机 /
- 热化学非平衡 /
- 超声速燃烧 /
- 低雷诺数流动 /
- 激波风洞 /
- 飞行试验
Abstract: Hypersonic airbreathing flights are highly valued in both the fields of space transportation and national aerospace safety, and the scramjet engines are pivotal propulsion devices for these flights. The scramjet engines for flight Mach numbers within the range between 4.0 and 7.0 have been extensively studied and well developed in recent years, and the extension to the scramjet engines for higher flight Mach numbers within the range between 8.0 and 10.0 or even higher are sure to be a competing focus for near-space competitions in the following decades. The current paper analyzes and summarizes the recent research advances of scramjet engines with flight Mach numbers within the range between 8.0 and 10.0+ . First of all, the key scientific problems and technologies of the higher Mach number scramjet engines are highlighted, including the high-temperature dissociation and thermochemical nonequilibrium effects, mixing and combustion enhancement technologies in ultra-high-speed flows, the matching of hypersonic combustion and inflow compression and the operating modes, the high-enthalpy low Reynolds number boundary-layer flows and the boundary-layer flow control methods, the thermal protection technologies of high-enthalpy low-density combustion inflows, and the ground test facility technologies for high-Mach number scramjet engines, respectively. Second, the experimental apparatus related to high-enthalpy shock tunnels and the shock tunnel driving technologies and typical ground and flight experiments of the high-Mach number scramjet engines home and aboard in recent years are introduced. Third, research advances including overall performance analyses of thrusts and thermal protections, the prominent high-enthalpy dissociation and thermochemical nonequilibrium effects in high-Mach-number scramjet engines, and mixing and combustion enhancement technologies in the ultra-high-speed flows are reviewed, so as to assess the feasibilities of high-Mach-number scramjet engines, and to discuss the features of engines’ key technologies. Finally, the summary is presented and several suggestions are proposed for further studies of the higher Mach number scramjet engines.
- high Mach number /
- scramjet engine /
- thermal and chemical nonequilibrium /
- supersonic combustion /
- low Reynolds number flow /
- shock tunnel /
- flight experiment

HTML全文

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}}} $和$ {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 $M{a_{\text{f}}}$

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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 $M{a_{\text{f}}}$

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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

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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 $M{a_{{\text{in}}}}$

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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$M{a_{{\text{in}}}}$

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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 ${H_{\text{f}}}$

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图 7 不同$M{a_{\text{f}}}$和${H_{\text{f}}}$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 ${H_{\text{f}}}$

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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 $q$

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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 $T_0^*$

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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 ${T_{{\text{air}}}}$

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图 11 激波风洞结构及运行原理

Figure 11. Structure and operating principle of shock tunnel

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图 12 JF24激波风洞示意图

Figure 12. Schematic of JF24 shock tunnel

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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

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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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]

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图 32 典型OH组分质量分数分布^[11]

Figure 32. Typical OH mass fraction distribution^[11]

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

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

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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]

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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]

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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]

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图 37 进气道OH质量分数分布^[14]

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

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图 38 燃烧室和喷管温度分布^[14]

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

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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]

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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]

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图 41 激波诱导燃烧示意图^[101]

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

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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]

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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]

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图 44 定制的非对称喷孔排布^[106]

Figure 44. Tailored asymmetric fuel injection holes^[106]

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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]

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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]

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图 47 超混合型支板的典型构型^{[61, 118]}

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

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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]

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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]

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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
8	30	2619	18.7
	35	2710	9.2
	40	2831	4.8
10	30	3688	114.7
	35	3790	57.8
	40	3917	31.0
12	30	4915	575.0
	35	5025	296.2
	40	5168	163.6

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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 type		Operating time	Test medium	$T_0^*$/K	$M{a_{\text{f}}}$	Typical wind tunnel^[26]
continuous		minutes	pure air	<1000	<4	AEDC tunnels
intermittent	heat reservoir	seconds	pure air	<2100	<7	HTF glenn
intermittent	combustion heater	minutes	vitiated air	<2500	<8	8-ft HTT langley
impulse		milliseconds	pure air	<10000	<20	LENS

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表 3 典型高焓激波风洞

Table 3. Typical high-enthalpy shock tunnels

Tunnel name	Affiliation	Driver type	Total length/m	Operating time/ms	Stagnation pressure/MPa	Ma
T4^[35]	University of Queensland	free piston	36	1 ~ 2	<30	4 ~ 10
FD21^[36]	Chinese Academy of Aerospace Aerodynamics	free piston	109	2	<11	>8
LENS I^[28]	Calspan	heated lightweight gas	26	5 ~ 15	<80	6 ~ 15
Hypulse^[28]	GASL	lightweight gas and detonation	27	0.5 ~ 7	<30	5 ~ 12
JF24^[37]	Institute of Mechanics	gaseous detonation	23	5 ~ 10	<20	8 ~ 12

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表 4 T4激波风洞试验来流参数^[90]

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

	TECN	TNCN
$Ma$	7.53	8.42
$p$/Pa	3750	1877.9
${T_{\text{t}}}$/K	315.4	260.6
$ {T_{\text{v}}} $/K	315.4	640.3
${Y_{{\text{NO}}}}$	0.04789	0.01677
${Y_{\text{O}}}$	0.00016	$5.75 \times {10^{ - 8}}$

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