Flight Condition Achievement of Mach Number 8 in a New Shock Tunnel of CAAA and its Scramjet Experimental Investigation
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摘要: 针对高Mach数超燃冲压发动机实验能力空缺问题，基于航天十一院新建的FD-21高能脉冲风洞，进行了Ma=8超燃飞行条件的模拟能力设计与调试，获得了总焓2.9 MJ/kg、总压11.01 MPa实验条件，实现了Ma=8、高度31 km飞行条件的风洞模拟.在此基础上，研发了匹配的氢燃料供应及喷注时序控制系统，设计了超燃冲压发动机模型，开展了超燃冲压发动机模型自由射流应用性风洞实验，获得了氢气燃料与空气、氮气超声速气流耦合流动作用下的实验模型壁面压力数据.在当量比近似一致条件下，空气来流对应的燃烧室壁面压力明显高于氮气来流情况，表明氢气在1 ms有效实验时间内完成了与超声速空气来流的混合、点火与燃烧，获得燃烧释热特性，确认了在FD-21高能脉冲风洞开展高Mach数超燃实验是切实可行的，为后续研究奠定了良好的基础.
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关键词:
- 高能脉冲风洞 /
- Ma=8飞行条件复现 /
- 超燃实验 /
- 燃料供应系统 /
- 高Mach数自由射流实验
Abstract: Aimed at the lack of experimental conditions of high-Mach-number scramjet, the ability to simulate flight conditions of Mach number 8 was explored in FD-21 high-energy shock tunnel of CAAA. The simulation in total enthalpy of 2.9 MJ/kg and total pressure of 11.01 MPa was obtained, duplicating a flight Mach number of 8 at the altitude of 31 km. Furthermore, experimental techniques of hydrogen combustion were developed. A hydrogen fuel supply system with the time sequencer was equipped. A semi-free jet test model was designed to examine the possibility of scramjet test in FD-21 shock tunnel and check the performance of fuel supply system. Several commissioning tests were conducted at the Ma=8 simulation conditions. A comparison of wall pressure distributions between different test gas experiments with the almost identical hydrogen injection mass flow rate was made. It is shown that a pronounced pressure increment is observed in the downstream of the hydrogen injection location, while only a slight difference exists in the upstream of the hydrogen injection location. These phenomena indicate the achievement of hydrogen mixing, igniting and burning in supersonic airflow, which verifies the feasibility of scramjet test in FD-21 shock tunnel. The present achievements are helpful for experimental investigation of high-Mach-number scramjet. -
图 4 压缩管及激波管上的测点位置示意图
Figure 4. Measure porthole locations on compression and shock tubes(not scale)
图 5 激波管各测点压力随时间变化情况
Figure 5. Variation of pressure distributed in shock tube with time
图 9 实验模型及压力、热电偶测点
Figure 9. Scramjet test article with pressure and thermocouple porthole
图 12 基于Tannehill平衡空气模型及k-ωSST湍流模型, 轴对称数值计算所得Φ=1.2 m的喷管出口参数分布
Figure 12. Axial parameters distributions at the FD-21 Φ= 1.2 m nozzle exit from CFD with Tannehill equilibrium air model and k-ω SST turbulence model
图 13 空气来流条件下, 总压、模型壁面静压及氢气喷注前室压力数据曲线(65车次, 总压Pt=10.96 MPa, PH2 -inj=0.698 MPa, 当量比ϕ=0.71)
Figure 13. Variation of total pressure, model wall pressure and hydrogen injection pressure with time for shot 65 of air inflow, where Pt=10.96 MPa, PH2 -inj=0.698 MPa, ϕ=0.71
图 14 氮气来流条件下, 总压、模型壁面静压及氢气喷注前室压力数据曲线(67车次, 总压Pt=10.36 MPa, PH2-inj=0.625 MPa, 当量比ϕ=0.69)
Figure 14. Variation of total pressure, model wall pressure and hydrogen injection pressure with time for shot 67 of nitrogen inflow, where Pt=10.36 MPa, PH2-inj=0.625 MPa, ϕ=0.69
图 15 空气(65车次)/氮气(67车次)来流条件下, 实验及数值所得模型壁面静压沿程分布
Figure 15. Time-average wall pressure distributions along the centre line of injection-side wall under the condition of air (slot 65) and nitrogen(slot 67), together with numerical results
表 1 世界主要高焓脉冲风洞设备参数
Table 1. Simulation parameters of worldwide high-enthalpy impulse wind tunnels
wind tunnel affiliation driver technique driver tube driven tube nozzle exit diameter/m Vmax/
(km/s)Masimulated tef/ms hmax/
(MJ/kg)L/m D/m L/m D/m T4 UQ free-piston 26 0.229 10 0.076 0.388 ~5.5 4~10 ≥1 15 T5 Caltech 30 0.3 12 0.09 0.314 ~6.3 4~7 ≥1 20 HIEST JAXA 42 0.6 17 0.18 0.88 ~7 8~16 ≥2 25 HEG DLR 33 0.55 17 0.15 1.2 ~7 6~10 1~6 25 FD-21 CAAA 75 0.668 34 0.29 1.2/2.0 ~7 10~16 2~10 25 LENS Ⅰ GASL heated light gas 7.62 0.28 18.3 0.203 1.5 ~4.6 6~22 5~18 10 LENS Ⅱ 18.3 0.61 30.5 0.61 1.8 ~2.7 3~10 30~100 4 JF-12 IMCAS detonation 99 0.4 89 0.72 2.5 ~3 5~9 ~100 3.5 
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表 2 模拟弹道点Ma=8, H=31 km对应的风洞运行参数
Table 2. Operation parameters of FD-21 shock tunnel to simulate flight conditions of Ma=8 and H=31 km
compression tube with helium and argon shock tube with air ωHe ωAr P4i/kPa T4i/K P1/kPa T1/K 0.1 0.9 30 300 61 300
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表 3 激波管上压电传感器安装位置
Table 3. Locations of piezoelectric sensors on shock tube
marks location/m S0 0.72 S1 4 S2 8 S4 16 S6 23.99 S7 27.99 S8 31.99 S9 33.19 S10 33.96
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表 4 激波管末端入射激波Mach数及模拟总压
Table 4. Shock Mach number and simulated total pressure at the end of shock tube
test No. Mas between S7 and S8 error PtS10/MPa error 69 5.13 3.62% 9.64 -4.48% 70 4.90 -1.09% 9.96 -1.30% 71 4.90 -1.05% 10.54 4.44% 72 5.12 3.52% 10.49 3.95% 73 4.74 -4.26% 9.7 -3.88% 74 4.91 -0.73% 10.22 1.27% average 4.95 10.09
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