Exploration of Experimental Techniques on Ma=10 Scramjets in FD-21 High Enthalpy Shock Tunnel
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摘要: 针对Mach数8以上（Ma>8）冲压发动机地面试验能力不足问题，基于FD-21高能脉冲风洞，开展了吸气式推进试验技术探索，提升了FD-21风洞的重活塞驱动能力，获得了总压18.66 MPa、总温3 950 K、Ma=9.62、静压436.6 Pa、速度3 km/s的高焓大动压模拟流场，同时发展了高时间分辨率吸收光谱测量技术和基于重模型自由飞原理的发动机推阻测量方法.在此基础上，设计了弯曲激波压缩二元发动机，构建了燃料在线供应与喷注控制、模型悬挂与瞬态释放及相关测量一体的试验系统，在所建立的Ma=9.62风洞模拟环境中进行了集成验证试验，定量测得了有/无氢气射流与空气/氮气超声速气流作用下二元发动机的壁面压力、吸收光谱峰值吸收率、轴向力等数据，并利用纹影观测到了进气道唇口与燃烧室部位的波系特征.多次试验所得的壁面压力、峰值吸收率、轴向力随时间变化曲线均存在2 ms以上的平台，表明二元发动机建立了准定常流动.冷热态及氮气对照组对应的壁面压力分布、峰值吸收率、轴向力等数据呈现出了明显不同，且二者规律近似一致，一方面说明所建立的模拟流场、燃烧诊断技术、发动机推阻测量技术是有效的，另一方面也表明二元发动机实现了点火燃烧、获得有效热功转换，为后续相关研究奠定了良好的基础.
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关键词:
- FD-21高焓激波风洞 /
- Ma＝10超燃冲压发动机 /
- 自由射流试验 /
- 重模型自由飞技术 /
- 吸收光谱测量技术
Abstract: Multiple experimental techniques were explored to test and verify the scramjet performance at Mach number above 8 in FD-21 shock tunnel. The piston driver ability was further improved to obtain the stagnation simulation conditions of 18.66 MPa and 3 950 K, together with the freestream conditions of Mach number of 9.62, static pressure of 436.6 Pa and velocity of 3 km/s. Furthermore, two key experimental techniques were established, involving the tunable diode laser absorption spectroscopy (TDLAS) for combustion diagnosis and the large-scale free-flight force measurement technique for scramjet testing. And subsequently, a two-dimensional curved shock compression scramjet model (2DSM) was designed, together with a set-up of the free-flight system integrated with hydrogen on-board supply/injection control and data acquisition. A series of free-jet experiments were performed at the above-mentioned test conditions, including airflow with/without hydrogen injection and the cross check shots with nitrogen test gas. Those data were quantitatively obtained including wall pressure, peak absorption, and axial force, as well as schlieren images at inlet and combustor. A plateau lasted over 2 ms exists in the time distributions of wall pressure, peak absorption, axial force, indicating the quasi-steady flow establishment of 2DSM. The consistency of those data was observed over shots, verifying that these techniques were effective for scramjet performance measurement. An inner net thrust as well as a remarkable increase in wall pressure was achieved between these hot and cold shots, where the hot shots denote the cases of airflow with hydrogen injection and the cold shots denote those of the airflow without hydrogen injection or nitrogen-flow with hydrogen injection. These phenomena observed indicate the occurrence of hydrogen igniting and burning, together with obvious conversion of heat into power. The present achievements are helpful for experimental investigation of high Mach number scramjets. -
图 2 入射激波运动Mach数在激波管内的变化
Figure 2. Incident shock Mach number along shock tube axial location
图 3 距离激波管末端1.17 m(S9)处壁面压力变化
Figure 3. Time variation of wall pressure located at 1.17 m(S9)from the shock tube downstream end
图 4 距离激波管末端1.17 m(S9), 0.02 m(S10)处壁面压力变化
Figure 4. Time variation of wall pressure located at 1.17 m(S9) and 0.02 m(S10) from shock tube downstream end
图 5 弯曲激波压缩二元发动机模型(2DSM)
Figure 5. Two-dimensional curved shock compression scramjet model(2DSM)
图 7 微小位移光学追踪方法
Figure 7. A nonlinear method to extract the tiny displacement from image
图 8 二元发动机的TDLAS光路布置及其保护结构模型(2DSM-B)
Figure 8. Four light-path configurations with protecting structure of TDLAS on 2DSM-B
图 10 有无氢气喷注下二元发动机各测点压力随时间变化
Figure 10. Time variation of wall pressure on 2DSM for the shots with hydrogen injection and no injection
图 11 不同工况下发动机沿程壁面静压分布
Figure 11. Streamwise pressure distributions on 2DSM wall for different shots
图 12 氢气喷注当量比近似一致、不同试验介质对应的TDLAS测得的V1峰值吸收率随时间变化
Figure 12. Peak absorbance of different test-gas shots from TDLAS at the approximately identical equivalence ratio
图 13 二元发动机进气道与燃烧室纹影图
Figure 13. Flow pattern at inlet and flow pattern on combustor of 2DSM from high speed schlieren
图 14 氢气喷注时发动机轴向加速度随时间变化(2021车次, ϕ=0.416)
Figure 14. Time variation of axial acceleration and wall pressure on 2DSM-A (No. 2021, ϕ=0.416)
表 1 FD-21风洞总压18.66 MPa、总温3 950 K模拟条件下名义Ma=10喷管出口参数
Table 1. Freestream parameters of the simulated Ma=10 conditions in FD-21 shock tunnel
P∞/Pa T∞/K Tvib∞/K U∞/(m/s) Ma∞ CN2, ∞/(%) CO2, ∞/(%) CNO, ∞/(%) CO, ∞/(%) 436.6 257.5 1 478 3 000 9.62 74.32 20.80 4.85 0.03 
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表 2 基于二元发动机模型的验证试验概况
Table 2. Test information of 2DSM
Shot No. test gas model equivalence
ratiomeasure methods 2018 air 2DSM-A 0 piezoresistive sensors 2019 air 2DSM-A 0 2020 air 2DSM-A 0 piezoresistive sensors, high-speed photography, accelerometer 2021 air 2DSM-A 0.416 2022 air 2DSM-A 0.358 2023 air 2DSM-B 0.346 piezoresistive sensors, high-speed Schlieren, accelerometer, TDLAS 2024 air 2DSM-B 0.360 2025 nitrogen 2DSM-B 0.347 2026 air 2DSM-B 0
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表 3 TDLAS测得的温度与水蒸气(H2O)分压(2023车次, ϕ=0.346)
Table 3. Time-averaged temperature and partial pressure of water vapor from TDLAS(shot No. 2023, ϕ=0.346)
light-path time-averaged temperature/K time-averaged partial pressure of H2O/Pa V1 1 215.8 203.7 V2 1 194.8 207.9 H1 1 195.6 174.7 H2 1 193.4 196.1 mean value 1 199.9 195.6 max error 1.3% 10.7%
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表 4 基于自由飞原理测得的不同工况下二元矩形发动机模型的轴向力数据
Table 4. Axial force on 2DSM-A and 2DSM-B measured by large-scale free-flight techniques for different shots
Shot No. test gas equivalence ratio model axial accelera-tion/(m/s2) dimensionless aixal force dimensionless inner net force between hot and cold shots 2020 air 0 2DSM-A 14.08 65.70 0 2021 air 0.416 2DSM-A 16.80 91.17 -25.47 2022 air 0.358 2DSM-A 11.44 52.39 13.31 2023 air 0.346 2DSM-B 21.28 105.44 14.16 2024 air 0.360 2DSM-B 21.82 113.88 5.72 2025 nitrogen 0.347 2DSM-B 24.34 119.60 0
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