提高高马赫数超燃冲压发动机推力的理论方法

韩信; 刘云峰; 张子健; 张文硕; 马凯夫

doi:10.6052/0459-1879-21-350

提高高马赫数超燃冲压发动机推力的理论方法

doi: 10.6052/0459-1879-21-350

韩信^{*, †},
刘云峰^{*, †,},
张子健^**,
张文硕^{*, †},
马凯夫^{*, †}

*.
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
**.
香港理工大学航空及民航工程学系, 香港九龙 999077

基金项目: 国家自然科学基金资助项目(11672312)

详细信息

作者简介:
刘云峰, 高级工程师, 主要研究方向: 激波与爆轰物理. E-mail: liuyunfeng@imech.ac.cn

中图分类号: V231.3
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- 被引次数: 0
出版历程
- 收稿日期: 2021-07-23
- 录用日期: 2021-12-21
- 网络出版日期: 2021-12-22
- 刊出日期: 2022-03-18

THE THEORETICAL METHOD TO INCREASE THE THRUST OF HIGH MACH NUMBER SCRAMJETS

Han Xin^{*, †},
Liu Yunfeng^{*, †
,},
Zhang Zijian^**,
Zhang Wenshuo^{*, †},
Ma Kaifu^{*, †}

*.
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
**.
Department of Aeronautical and Aviation Engineering, The Hong Kong Polytechnic University, Kowloon 999077, Hong Kong, China

摘要

摘要: 斜爆轰发动机和激波诱导燃烧冲压发动机在高马赫数吸气式发动机中具有重要应用前景, 但是斜爆轰发动机是否具有足够大的净推力, 还是一个未知的问题, 因此需要对高马赫数冲压发动机的推进性能以及提高推力的方法进行理论研究. 本文主要分为3部分. 第1部分理论研究了超燃冲压发动机中的爆燃波和爆轰波的传播特性. 保证发动机稳定燃烧是提高推力的前提. 通过对爆燃波和爆轰波传播特性研究, 得到了影响发动机燃烧稳定性的关键参数和物理规律. 第2部分研究了发动机处于热壅塞临界状态下的燃烧规律和推力特性. 在临界状态下, 燃烧室入口气流速度正好等于爆轰波传播速度, 二者处于平衡状态, 这是发动机推进性能的理论上限. 第3部分研究了提高高马赫数超燃冲压发动机推力的理论方法. 对于高马赫数冲压发动机, 燃烧室入口气流速度远远大于爆轰波的传播速度, 这部分速度差就是提高推力的理论空间. 对于马赫数Ma ≥ 12的超燃冲压发动机, 理论上燃烧产生的爆燃波或激波不会引起发动机不起动, 因此可以通过进一步添加燃料和氧化剂的方法来提高其推力. 理论分析结果表明, 对于高马赫数超燃冲压发动机, 不但燃烧流场是容易稳定的, 而且可以有很多方法来进一步提高推力.
- 高马赫数超燃冲压发动机 /
- 斜爆轰发动机 /
- 激波诱导燃烧冲压发动机 /
- 超声速燃烧 /
- 提高推力的方法
Abstract: The oblique detonation engines and shock-induced combustion ramjets have been proved to be practicable for high flight Mach number air-breathing engines in recent years. However, whether the oblique detonation engines and shock-induced combustion ramjets have enough thrust or not is unknown yet. In this paper, the combustion characteristics and propulsive performance of scramjets are discussed theoretically. Firstly, the mechanism of engine unstart of scramjets is discussed from the point of view of shock/shock interaction and deflagration-to-detonation transition. The results show that engine unstart process is very similar to the deflagration-to-detonation transition process. In the combustor of scramjets, the maximum velocity of the deflagration wave is very close to the detonation velocity. Therefore, the C-J detonation velocity is defined as the stable operation boundary of scramjets. Secondly, the formula of thrust produced by the divergent nozzle is put forth and key parameters influencing thrust are obtained. According to the thrust formula, supersonic combustion is beneficial for increasing the thrust. The main way to increase the thrust is to increase the pressure of combustion products. The propulsive performance of scramjets is theoretically analyzed by using C-J detonation theory, which is the critical condition when the engine is thermally choked. Finally, the theoretical method to increase the thrust is discussed. For high flight Mach number scramjets Ma ≥ 12, the velocity in the isolator is much faster than the C-J detonation velocity in the combustor and the problem of engine unstart disappears. Therefore, extra fuel and oxidizer can be injected into the combustor to increase the thrust further as long as the shock wave generated by the high pressure combustion products is slower than the air velocity in the isolator. The theoretical results agree well with the existing experimental and numerical results, which can be used as a baseline for the development of high Mach number scramjets.
- high flight Mach number scramjets /
- oblique detonation engines /
- shock-induced combustion ramjets /
- supersonic combustion /
- the method to increase the thrust

HTML全文

图 1 发动机燃烧室物理模型

Figure 1. Physical model of a combustor

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图 2 不同当量比下的C-J爆轰波速度、C-J爆燃波速度和声速的比较

Figure 2. Comparison of C-J detonation velocity, C-J deflagration velocity and sound velocity of H₂/Air mixture at 0.1 MPa and 300 K

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图 3 不同初始温度T₁下的C-J爆轰波速度和C-J爆燃波速度的比较

Figure 3. Comparison of C-J detonation and C-J deflagration velocity of H₂/Air mixture at different initial temperature T₁

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4 不同初始温度T₁下的C-J爆轰波速度和C-J爆燃波速度的数值模拟结果

4. Numerical results of C-J deflagration velocity of H₂/Air mixture at different initial temperature T₁

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图 4 不同初始温度T₁下的C-J爆轰波速度和C-J爆燃波速度的数值模拟结果 (续)

Figure 4. Numerical results of C-J deflagration velocity of H₂/Air mixture at different initial temperature T₁ (continued)

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图 5 无量纲推力与喷管入口马赫数的关系

Figure 5. Relationship between dimensionless thrust and inlet Mach number of nozzles

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图 6 超燃冲压发动机、C-J爆轰发动机和斜爆轰发动机示意图

Figure 6. Schematic of scramjets, C-J detonation engine and oblique detonation engine

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图 7 初始温度300 K下不同当量比的C-J爆轰波传播速度

Figure 7. C-J detonation velocity under different equivalence ratio at 300 K

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图 8 不同初始温度下C-J爆轰波传播速度

Figure 8. C-J detonation velocity under different static temperature at ER = 1.0

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图 9 不同初始温度下的C-J爆轰波的压比

Figure 9. Pressure ratio under different static temperature at ER = 1.0

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图 10 不同初始温度下的C-J爆轰波压比

Figure 10. Pressure ratio under different ER at 1000 K and 1500 K

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图 11 斜爆轰波的温度云图

Figure 11. Temperature contours of oblique detonation waves

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图 12 激波管入射激波马赫数M_s与压比的关系

Figure 12. Relationship between incident shock wave Mach number M_s and driver pressure ratio

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图 13 燃烧室和喷管中的OH质量分数云图

Figure 13. The contours of OH mass fraction in the combustor and nozzle

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图 14 沿壁面压力分布 (1 atm = 101.3 kPa)

Figure 14. The pressure distribution along the wall (1 atm = 101.3 kPa)

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表 1 国内外超燃冲压发动机试验结果汇总

Table 1. Summary of some typical scramjets experimental results

Cases	Fuel	Velocity in isolator/(m·s⁻¹)	Unstart equivalence ratio	Detonation velocity/(m·s⁻¹)	References
1	H₂	1720	0.5	1635	[7-10]
2	H₂	1000	0.10	985	[11]
3	C₂H₄	1000	0.32	1133	[12-13]
4	C₂H₄	1060	0.39	1434	[15-16]
5	C₂H₄	900	0.21	1139	[19]
6	H₂	1750	0.48	1612	[20-22]
7	H₂	2500	1.26 (start)	2039	[23]

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表 2 不同飞行马赫数下斜爆轰发动机参数

Table 2. Parameters of oblique detonation under different flight Mach numbers

Ma_∞	Ma₁	T₁/K	p₂/p₁	Ma₂	β_ODW/(°)
9	4.41	618.1	8.16	1.86	47.6
10	4.68	683.3	7.94	2.15	41.9
11	4.93	752.9	7.86	2.38	37.8
12	5.52	826.6	7.87	2.61	34.8

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表 3 不同飞行马赫数等熵压缩后的参数

Table 3. Parameters behind isentropic compression for different flight Mach numbers

Ma_∞	T_∞/K	T₁/K	Ma₁
8	225	1250	2.69
9	225	1250	3.22
10	225	1250	3.68
11	225	1250	4.12
12	225	1250	4.63

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