High-Mach-Number Scramjet Engine Tests in JF12 Shock Tunnel
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摘要: 针对高Mach数（Ma ≥ 7）超燃冲压发动机高气动阻力下的燃烧组织问题，提出一种双突扩燃烧室结构方案.使用数值模拟方法考察了射流与双突扩燃烧室组合方式的混合燃烧特性.设计了双突扩超燃冲压发动机模型，在力学研究所JF12长试验时间激波风洞内，开展了Ma=7.0和Ma=9.5的氢燃料点火和燃烧试验对比.在风洞有效试验时间100 ms内，实现了Ma=7.0和Ma=9.5超燃冲压发动机的成功点火与稳定燃烧.在Ma=7.0情况下，进气道采用三维压缩，燃烧室入口设计Mach数Mac=2.5，壁面压力分布实验结果显示燃烧放热靠近燃烧室扩张段上游；在Ma=9.5情况下，进气道采用二维压缩，燃烧室入口设计Mach数Mac=3.5，由于燃烧室流动速度特别高，燃烧放热靠近燃烧室扩张段下游.Abstract: A schematic configuration of dual-step supersonic combustor was proposed to overcome the high gasdynamic drag of scramjet engine in high-Mach-number flight (Ma ≥ 7). Numerical simulations were conducted to evaluate the coupled mixing and combustion features of fuel jet and dual-step flowfield. A scramjet model with dual-step was designed and tested in the JF12 Shock Tunnel, which can provide the pure air medium for simulating high-Mach-number scramjet tests. During the 100 ms test duration of JF12 Shock Tunnel, successful ignition and stable combustion of hydrogen-fueled scramjet models were realized in both simulation conditions of Ma=7.0 and Ma=9.5 and the experimental results were compared. In the simulation condition of Mach 7, three-dimensional inlet was used and the entrance Mach number of supersonic chamber was about Mac=2.5. The wall pressure distributions showed that the combustion took place in the upstream of expansion section of duct. In the simulation condition of Mach 9.5, two-dimensional inlet was used and the entrance Mach number of supersonic chamber was about Mac=3.5. The combustion took place near the exit of inner duct of scramjet model.
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Key words:
- scramjet /
- high-Mach-number /
- shock tunnel /
- pure air /
- JF12 shock tunnel
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图 5 超燃冲压发动机下壁面压力测点分布
Figure 5. Distribution of pressure transducers on the lower wall of the test model
图 8 燃料喷孔上游压力测量曲线
Figure 8. Curve of pressure measured at upstream of fuel injection nozzle
图 10 Ma=7条件下冷态和热态实验发动机壁面压力分布
Figure 10. Wall pressure distributions of engine in cold and hot tests at Ma=7
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