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为了探索边界层非强迫转捩对进气道性能的影响,采用数值计算的方法开展了边界层转捩对轴对称混压式高超声速进气道流场特性的研究。研究表明:随着进气道中心锥锥尖钝化半径增大,边界层转捩先推迟。当锥尖钝度大到一定程度时,边界层转捩位置前移。随着钝化半径进一步增大,边界层转捩再次推迟,转捩位置逐渐后移。来流湍流度越大,边界层越不稳定,边界层转捩越易发生。与湍流边界层相比,考虑边界层转捩时进气道的总压恢复系数及流量系数较高、热载荷及阻力系数较小,Ma=6.5时喉道处总压恢复系数最高上升17.3%,进气道阻力最大下降17.4%。边界层转捩对壁面热流密度分布影响较大,但对壁面压力分布影响较小。钝化影响进气道的自起动性能,随着钝化半径增大,自起动马赫数升高,而边界层转捩对进气道自起动性能影响较小。
In order to explore the influence of the non-forced transition of the boundary layer on the performance of the inlet duct, the numerical simulation is used to study the characteristics of the flow field of the axisymmetric hybrid hypersonic inlet along the boundary layer. The research shows that with the increase of the passivation cone tip radius, the boundary layer transition is delayed first. When the cone tip sharpness to a certain extent, the boundary layer transition 捩 position forward. As the passivation radius further increases, the boundary layer transition is postponed again, and the transition position is gradually shifted backward. The greater the turbulence of the incoming flow, the more unstable the boundary layer and the more likely the boundary layer transition occurs. Compared with the turbulent boundary layer, the total pressure recovery factor and the flow coefficient of the inlet are higher when the boundary layer transition is considered, and the thermal load and drag coefficient are smaller. The maximum pressure recovery factor at the throat of Ma = 6.5 increases by 17.3% , The maximum air intake resistance decreased by 17.4%. The boundary layer transition has a great influence on the wall heat flux density distribution, but has little effect on the wall pressure distribution. Passivation affects the self-starting performance of the inlet. As the passivation radius increases, the self-starting Mach number increases, while the boundary layer transition has little effect on the self-starting performance of the inlet.