乘波前体两侧高超声速内收缩进气道一体化设计

Integrated Design of Waverider Forebody and Lateral Hypersonic Inward Turning Inlets

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 study the flow field and aerodynamics performance of inlet system with two inlets on each side, an in- let system for a hypersonic aircraft with a waverider forebody and hypersonic inward turning inlets with controlled pressure rise law is designed and numerically simulated, whose flow field and characteristics of speed, angle of attack and sideslip angle are obtained. The results indicate that, at the design point, the external flow filed of the forebody is independent of the inlet internal flow filed, but at take-over Mach number the leading shock waves by the forebody and the inlet are coupling. Because the boundary layer does not enter the inlets, the interaction of the shock wave and the boundary layer is weak, and so there is no separation in the flow field. With the increase of the freestream Mach number, the total pressure recovery co- efficient decreases, the compression ratio increases greatly, but the lift-drag ratio almost does not change. With the increase of the angle of attack, the mass capture ratio and compression ratio increase greatly, the total pressure recovery coefficient decreases slightly, and the lift-drag ratio increases. With the increase of the sideslip angle, the overall performance of the two inlets deceases, and the inlet at downwind decreases more than the other at upwind.