RBCC发动机纯火箭模态性能特性仿真研究

凌文辉; 韦宝禧; 侯金丽; 姚达豪; 罗飞腾

doi:10.13224/j.cnki.jasp.20240433

RBCC发动机纯火箭模态性能特性仿真研究

doi: 10.13224/j.cnki.jasp.20240433

1.
北京动力机械研究所冲压发动机技术全国重点实验室，北京 100074
2.
华中科技大学航空航天学院，武汉 430074

基金项目: 国家自然科学基金面上项目（52076094）；冲压发动机技术重点实验室基金

详细信息

作者简介:
凌文辉（1967-），男，研究员，博士，研究方向为冲压发动机及组合推进技术

通讯作者:
韦宝禧（1983-），男，研究员，博士，研究方向为冲压发动机及组合推进技术。E-mail：weibaohuasky@163.com

中图分类号: V231.3
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出版历程
- 收稿日期: 2024-06-30
- 网络出版日期: 2025-03-02

Numerical study on pure rocket mode performance characteristics of a RBCC engine

1.
National Key Laboratory of Ramjet，Beijing Power Machinery Institute，Beijing 100074，China
2.
School of Aerospace Engineering，Huazhong University of Science and Technology，Wuhan 430074，China

摘要

摘要:
为掌握火箭基组合循环（RBCC）发动机的纯火箭模态性能特性及影响因素，设计构建了中心火箭布局的全轴对称特征RBCC发动机基准流道，以及相同设计膨胀比的半轴对称斜切喷管构型RBCC发动机、连续扩张喷管构型火箭发动机流道，开展了不同推力室室压条件下纯火箭模态三维膨胀流场仿真，获得了不同喷管构型、二次流、火箭推进剂方案下纯火箭模态内流工作特性、推力比冲性能特性。结果表明：纯火箭模态时非理想连续膨胀流动状态导致比冲性能损失，比冲效率在80%以下；大突扩膨胀后产生的强激波总压损失是性能下降的主要机制，全轴对称喷管构型相对于连续扩张喷管时比冲效率下降约8%，半轴对称斜切喷管可以减小膨胀总压损失，提高比冲效率；引入二次流可以调整膨胀-压缩波系结构以减小总压损失，提高二次流总温有利于提升整体性能。
- 火箭基组合循环（RBCC）发动机 /
- 纯火箭模态 /
- 斜切喷管 /
- 比冲效率 /
- 三维流场仿真
Abstract:
To gain a comprehensive understanding of the performance characteristics and influencing factors of the pure rocket mode in rocket-based combined cycle (RBCC) engines, a full-axisymmetric characteristic RBCC engine benchmark flow path with a central rocket layout was designed and constructed. Additionally, a half-axisymmetric scarfed nozzle configuration RBCC engine and a continuously expanding nozzle configuration rocket engine flow path with the same designed expansion ratio were developed. Three-dimensional expansion flow field simulations were conducted under different thrust chamber pressures in the pure rocket mode, yielding the internal flow characteristics and thrust specific impulse performance characteristics under various nozzle configurations, secondary flows, and rocket propellant schemes. The results indicated that in the pure rocket mode, the non-ideal continuous expansion flow state led to a loss of specific impulse performance, with the specific impulse efficiency being below 80%. The strong shock wave after the large sudden expansion led to a total pressure loss, which was the main mechanism for performance degradation. The overall axisymmetric nozzle configuration had a specific impulse efficiency decrease of about 8% compared with the continuous expansion nozzle. The semi-axisymmetric scarfed nozzle can reduce the total pressure loss of the expansion and improve the specific impulse efficiency. Introducing a secondary flow can adjust the expansion-compression wave system structure to a certain extent to reduce the total pressure loss. Increasing the total temperature of the secondary flow can be beneficial to enhancing overall performance.
- rocket-based combined cycle (RBCC) engine /
- pure rocket mode /
- scarfed nozzle /
- specific impulse efficiency /
- 3D flow field simulation

HTML

图 1 中心火箭布局的全轴对称RBCC发动机基准流道

Figure 1. Benchmark flowpath configuration of a axisymmetric RBCC engine with center rocket

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图 2 不同喷管构型RBCC发动机流道及连续扩张钟型喷管火箭发动机

Figure 2. RBCC engine flow channels with different nozzle configurations and continuously expanded bell-shaped nozzle rocket engine

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图 3 3种发动机流道构型的计算网格

Figure 3. Computational meshes of three engine flow channel configurations

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图 4 不同网格的沿程壁面静压对比

Figure 4. Comparison of longitudinal wall static pressures on different meshes

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图 5 煤油/液氧火箭发动机理想比冲随室压变化特性

Figure 5. Variation of ideal specific impulse of kerosene/LO₂ rocket engine with chamber pressure

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图 6 液氢/液氧火箭发动机理想比冲随室压变化特性

Figure 6. Variation of ideal specific impulse of LH₂/LO₂ rocket engine with chamber pressure

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图 7 不同喷管构型的推力性能对比

Figure 7. Comparison of thrust performance of different nozzle configurations

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图 8 不同喷管构型的比冲效率对比

Figure 8. Comparison of specific impulse efficiency among different nozzle configurations

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图 9 3种构型下对称面上的马赫数分布云图对比

Figure 9. Comparison of Mach number distribution contours on the symmetry plane for three configurations

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图 10 不同火箭构型的沿程总压和马赫数对比

Figure 10. Comparison of total pressure and Mach number along the path for different rocket configurations

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图 11 不同火箭室压对称面上的流场结构

Figure 11. Flow field structure on the symmetry plane for different rocket chamber pressures

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图 12 不同二次流流量比下对称面上参数分布

Figure 12. Parameters distribution on the symmetry plane for different secondary flow rate ratios

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图 13 二次流流量比对发动机比冲和推力的影响

Figure 13. Impact of secondary flow rate ratio on engine specific impulse and thrust

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图 14 不同二次流流量比下对称面马赫数分布

Figure 14. Mach number distribution on the symmetry plane for different secondary flow rate ratios

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图 15 不同二次流流量比下沿程总压对比（空气）

Figure 15. Comparison of total pressure along the path for different secondary flow rate ratios （air）

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图 16 二次流流量比对比冲和推力的影响（燃气）

Figure 16. Impact of secondary flow rate ratio on engine specific impulse and thrust （gas）

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图 17 不同二次流流量比下对称面上马赫数分布（燃气）

Figure 17. Mach number distribution on the symmetry plane for different secondary flow rate ratios （gas）

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图 18 不同二次流流量比下沿程总压对比（燃气）

Figure 18. Comparison of total pressure along the path for different secondary flow rate ratios （gas）

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图 19 二次流流量比对比冲和推力的影响（斜切喷管）

Figure 19. Impact of secondary flow rate ratio on engine specific impulse and thrus （scarfed nozzle）

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图 20 不同二次流流量比下对称面马赫数分布（斜切喷管）

Figure 20. Mach number distribution on the symmetry plane for different secondary flow rate ratios （scarfed nozzle）

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图 21 不同二次流流量比下沿程总压对比（斜切喷管）

Figure 21. Comparison of total pressure along the path for different secondary flow rate ratios （scarfed nozzle）

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图 22 采用液氢/液氧推进剂时的不同喷管构型发动机比冲性能特性

Figure 22. Specific impulse characteristics of engines with different nozzle configurations when using LH₂/LO₂ propellants

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表 1 不同喷管构型的发动机流道关键尺寸对比

Table 1. Comparison of key dimensions of engine flow channels with different nozzle configurations

流道构型	火箭入口半径/mm	火箭喉部半径/mm	内喷管出口半径/mm	发动机出口半径/mm	收缩比 A_in/A_t	内部膨胀比 A_out/A_t	总膨胀比 A_e/A_t
全轴对称喷管火箭	14	5.207	11.684	63.5	7.229	5.035	148.72
斜切喷管火箭	14	5.207	14.004	63.5	7.229	7.233	148.72
连续扩张喷管火箭	14	5.207		63.5	7.229		148.72

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表 2 火箭推力室的模拟工况条件

Table 2. Simulated conditions of rocket thruster

燃料	氧化剂	氧燃比	推力室室压 p_c/MPa	燃气总温 T_t/K
煤油	液氧	2.6	3	3640
			6	3640
			9	3698
液氢	液氧	7.0	3	3456
			6	3546
			9	3615

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表 3 仿真网格无关性验证

Table 3. Mesh independence study of simulation

网格方案	网格数量/10⁴	燃烧室出口总压/Pa	燃烧室出口静压/Pa
粗网格	120	956472	5588
中等网格	250	961994	5543
细网格	450	962617	5536

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表 4 基于APL-10C发动机的典型试验状态工况

Table 4. Typical test conditions for the APL-10C engine

试验工况	氧燃比	火箭室压/MPa	环境压力/MPa	燃气总温/K	工作模态
E02	11.58	3.597	0.1006	3242	纯火箭模态
E04	11.67	3.507	0.1006	3244
E05	11.05	3.362	0.0413	3207
106	11.34	3.514	0.1013	3227	单独推力室
108	10.96	3.369	0.1013	3202	单独推力室

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表 5 仿真计算与试验数据的比冲对比

Table 5. Comparison of impulse between the simulation results and experiments data

工况	试验比冲/s	CFD比冲/s	相对偏差/%
E02	220.9	239.8	8.58
E04	220.0	238.5	8.41
E05	251.5	253.7	0.88
106	264.3	261.7	−0.99
108	265.3	261.3	−1.5

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表 6 空气二次流入口参数

Table 6. Parameters of the secondary flow inlet （air）

火箭室压/ MPa	流量比/%	二次流温度/K	入口动量/ （kg·m/s）	推力贡献/N
3	2	300	0.243	2.349
3	4	300	0.702	2.741
3	8	300	2.024	1.760
6	2	300	0.601	3.637
6	4	300	1.403	5.860
6	8	300	4.039	3.883
9	2	300	0.708	7.862
9	4	300	2.071	9.009
9	8	300	5.979	6.084

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表 7 燃气二次流入口参数

Table 7. Parameters of the secondary flow inlet （gas）

火箭室压/ MPa	流量比/ %	二次流总温/K	入口动量/ （kg·m/s）	推力贡献/N
6	2	900	1.432	3.539
6	3	900	2.777	2.855
6	4	900	4.354	2.070
6	2	2 000	2.749	2.930
6	3	2 000	5.446	1.530
6	4	2 000	7.842	0.073

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表 8 斜切喷管构型二次流入口参数

Table 8. Parameters of the secondary flow inlet （scarfed nozzle）

火箭室压/ MPa	流量比/ %	二次流温度/K	入口动量/ （kg·m/s）	推力贡献/N
6	2	900	2.739	0.301
6	4	900	6.667	0.603
6	8	900	28.487	6.919

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表 9 采用液氢/液氧推进剂时二次流入口参数

Table 9. Parameters of the secondary flow inlet（LH₂/LO₂ propellant）

火箭室压/ MPa	流量比/%	二次流温度/K	入口动量/ （kg·m/s）	推力贡献/N
6	2	300	0.291	5.935
6	4	300	0.889	7.011
6	8	300	2.534	7.440

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表 10 液氢/液氧与煤油/液氧推力性能推进剂方案的比冲性能特性对比

Table 10. Comparison of specific impulse characteristics between LH₂/LO₂ and kerosene/LO₂ propellants

二次流流量比/%	煤油/液氧		液氢/液氧
二次流流量比/%	比冲/s	比冲效率/%	比冲/s	比冲效率/%
0	286.38	76.32	359.56	76.52
2	285.02	75.99	358.29	76.25
4	283.34	75.52	355.27	75.61
8	277.36	73.92	348.79	74.23

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