火箭推进

为使冲压发动机性能最优并始终工作在安全状态,需要使其工作的喘振裕度最小且在喘振边界内。在喘振边界安装一种能够准确分辨超声速或亚声速流场状态的流场探测装置,控制进气道结尾激波位置。针对一维变截面流动控制方程,研究了流场探测装置的安装位置,以及激波越过流场探测装置后控制系统的减油规律。仿真研究结果表明,用试验数据修正理论仿真的方法可以准确地捕捉进气道结尾激波,同时根据某型冲压发动机的设计临界喘振裕度,确定了流场探测装置安装位置位于距进气道锥尖的距离; 进气道沿程压力跟随发动机燃油流量的变化而变化,压力波传播时间相对于燃油调节时间可以忽略; 由于进气道内激波前后运动存在明显的压力滞环现象,当激波越过喘振边界时,进气道出口压力会进一步上升,发动机喘振危险加大,应使用加速电磁阀快速减小燃油流量,控制激波回到安全区域。

In order to make the ramjet performance optimization and always work in a security state, the ramjet's buzz margin should be keep in minimum and within the buzz boundary. A flow field detecting device that can distinguish the supersonic or subsonic flow field status is installed in the buzz boundary to control shock wave location at the intake terminal. Aiming at the one-dimensional variable cross-section flow control equations, the installation position of the flow field detecting device and the fuel reduction law of the control system after the shock wave goes through flow field detecting device are studied in this paper. The simulation result indicates that the method to use the test data to correct theoretical simulation data can accurately catch the shock wave at intake terminal, and according to the critical buzz margin designed for a certain type ramjet, can determine the installation location of the shock exploring device. The pressure along intake changes with the variation of engine's fuel flow, pressure wave propagation time relative to the fuel regulating time can be ignored. As the existence of the pressure hysteresis phenomenon is obvious before and after the intake shock motion, when the shock goes through the buzz boundary, the pressure at intake exit will rise further, the engine buzz risk increases, and the acceleration electromagnetic valve should be adopted to quickly reduce fuel flow, so as to make the shock wave go back to a safe area.

引言
1 计算模型、控制方程及计算方法
2 仿真计算及结果分析
3 结论

图1 进气道喉部到出口截面的面积变化<br/>Fig.1 Area variation curve of the intake throat to exit

图1 进气道喉部到出口截面的面积变化
Fig.1 Area variation curve of the intake throat to exit

图2 稳态马赫数变化曲线<br/>Fig.2 Mach number variation curve in the steady state

图2 稳态马赫数变化曲线
Fig.2 Mach number variation curve in the steady state

图3 不同摩擦系数下的进气道马赫数分布<br/>Fig.3 Intake shock-position under different friction coefficients

图3 不同摩擦系数下的进气道马赫数分布
Fig.3 Intake shock-position under different friction coefficients

图4 临界喘振裕度时的马赫数分布<br/>Fig.4 Intake shock-position in the critical buzz margin

图4 临界喘振裕度时的马赫数分布
Fig.4 Intake shock-position in the critical buzz margin

图5 不同压力扰动幅值下进气道压力的阶跃响应<br/>Fig.5 Intake pressure response curve under different amplitudes of pressure disturbance

图5 不同压力扰动幅值下进气道压力的阶跃响应
Fig.5 Intake pressure response curve under different amplitudes of pressure disturbance

图6 试验燃油流量设定值与反馈值的变化规律<br/>Fig.6 Fuel flow control instruction and the actual regulation curve for a test

图6 试验燃油流量设定值与反馈值的变化规律
Fig.6 Fuel flow control instruction and the actual regulation curve for a test

图7 进气道出口截面积和结尾激波压力的响应曲线<br/>Fig.7 The pressure response curve of intake exit and after terminal shock wave

图7 进气道出口截面积和结尾激波压力的响应曲线
Fig.7 The pressure response curve of intake exit and after terminal shock wave

图8 进气道激波位置的变化<br/>Fig.8 Intake shock-position curve with time

图8 进气道激波位置的变化
Fig.8 Intake shock-position curve with time

图9 结尾激波前压力的响应曲线<br/>Fig.9 The pressure response curve of before terminal shock wave

图9 结尾激波前压力的响应曲线
Fig.9 The pressure response curve of before terminal shock wave

图10 进气道出口压力变化曲线<br/>Fig.10 Intake exit pressure curve

图10 进气道出口压力变化曲线
Fig.10 Intake exit pressure curve

图11 激波探测元件的压力变化<br/>Fig.11 Shock-exploring device pressure curve

图11 激波探测元件的压力变化
Fig.11 Shock-exploring device pressure curve

图12 不同燃油流量变化率时进气道出口压力的变化<br/>Fig.12 Intake exit pressure curves comparsion

图12 不同燃油流量变化率时进气道出口压力的变化
Fig.12 Intake exit pressure curves comparsion

图13 激波探测元件的压力响应曲线对比<br/>Fig.13 Shock-exploring device pressure curves comparsion

图13 激波探测元件的压力响应曲线对比
Fig.13 Shock-exploring device pressure curves comparsion

图14 不同变化幅值时进气道出口压力的响应特性<br/>Fig.14 Intake exit pressure curves of different deceased amplitudes

图14 不同变化幅值时进气道出口压力的响应特性
Fig.14 Intake exit pressure curves of different deceased amplitudes

图15 不同变化幅值激波探测装置处的压力变化曲线<br/>Fig.15 Shock-exploring device pressure curve of different deceased amplitudes

图15 不同变化幅值激波探测装置处的压力变化曲线
Fig.15 Shock-exploring device pressure curve of different deceased amplitudes

图16 不同变化幅值时激波位置曲线<br/>Fig.16 Intake shock-position curve of different deceased amplitudes

图16 不同变化幅值时激波位置曲线
Fig.16 Intake shock-position curve of different deceased amplitudes

图17 不同变化幅值进气道出口压力和激波位置曲线<br/>Fig.17 Intake exit pressure and shock-position curve of different deceased amplitudes

图17 不同变化幅值进气道出口压力和激波位置曲线
Fig.17 Intake exit pressure and shock-position curve of different deceased amplitudes

图18 不同变化幅值时激波探测装置处压力变化曲线<br/>Fig.18 Shock-exploring device pressure curve of different deceased amplitudes

图18 不同变化幅值时激波探测装置处压力变化曲线
Fig.18 Shock-exploring device pressure curve of different deceased amplitudes

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备注

引言

1 计算模型、控制方程及计算方法

2 仿真计算及结果分析

3 结论

期刊信息