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为了探索两侧进气系统的流场结构及气动性能,采用吻切锥乘波前体、压升规律可控的一种高超声速内收缩进气道设计了两侧进气布局的高超声速飞行器一体化进气系统,并进行了数值模拟,研究了进气系统的流场结构、速度特性、攻角特性以及侧滑角特性等。结果表明,设计点前体外流场和进气道内流场相互独立,接力点前体前缘激波和进气道前缘激波相互耦合。由于未吞入前体附面层,因而进气道内激波附面层相互作用较弱,没有产生分离;随来流马赫数增大,进气道总压恢复系数减小,增压比增大显著,升阻比几乎不变;随攻角增大,流量系数增大明显,总压恢复系数略有减小,增压比增大明显,升阻比逐渐增大;随侧滑角增大,进气道总体性能逐渐减小,迎风侧进气道性能下降较小,背风侧进气道性能下降明显。
In order to explore the flow field structure and aerodynamic performance of air intake system on both sides, a hypersonic air-intake system with kiss-cut cone wave precursors and controlled pressure rise is designed to design a hypersonic vehicle Integrated air intake system, and numerical simulation of the intake system flow field structure, speed characteristics, angle of attack characteristics and side slip angle characteristics. The results show that the external flow field before the design point and the inlet flow field are independent of each other, and the leading shock and the leading shock of the inlet port couple with each other. Because of the non-swallowing of the precursors, the interaction of the shockwaves in the inlets is weak and there is no separation. As the flow Mach number increases, the total pressure recovery coefficient of the inlet decreases, With the increase of the angle of attack, the flow coefficient increases obviously, the total pressure recovery coefficient decreases slightly, the supercharging pressure ratio increases obviously, and the ratio of resistance to drag increases gradually. With the increase of side slip angle The overall performance of the inlet and the inlet decreases gradually, the performance of the inlet on the windward side decreases little, and the performance of the inlet on the leeward side decreases obviously.