超声速边界层燃烧壁面减阻特性与机理研究

The viscous drag generated due to the fully turbulent boundary layers on the surface of the flow-path in scramjet accounts for the most part of the viscous drag for hypersonic flight vehicles and can lead to significant loss of thrust. Recent studies have showed that the injection and combustion of a gaseous fuel in boundary layer could get great reductions in drag. However, under the influence of many factors and its complicated physical and chemical processes, the mechanism of drag reduction by boundary layer combustion has not been fully understood. In this project, the problems of fuel injection, ignition and flame-spreading in the supersonic boundary layer are investigated to reveal the mechanism of the skin friction drag reduction by boundary layer combustion. Experiments are carried out to study the macroscopic characteristics of the combustion flow field and its drag reduction performance in boundary layer. Through analysis of the pressure distribution, the temperature distribution and the flame structures in the flow-path, the change rule of drag reduction with the incoming flow and the fuel injection conditions can be obtained. A high-accuracy numerical method for the combustion of supersonic boundary layer is established. Then the numerical simulation for the supersonic boundary layer combustion is conducted. Through deeply analyzing the effect of combustion addition on the turbulent structure and the turbulent transport process in the boundary layer, the mechanism of the skin friction drag reduction by boundary layer combustion is determined. Lastly, based on both the experimental and numerical investigations, the predicting model is established. This study is of great significance for revealing the mechanism of drag reduction and further grasping its controlling principle for boundary layer combustion, which is benefit for enhancing the whole performance of scramjet.

超燃冲压发动机流道内部摩擦阻力占高超声速飞行器摩擦阻力的绝大部分，使推力遭受较大损失。采用边界层内燃烧的方式可大幅降低壁面摩擦力。然而其诸多的影响因素及内部复杂的物理化学过程，使其减阻机理至今仍未完全清楚。本项目针对超声速边界层内燃料喷注、点火与火焰传播问题开展深入研究，揭示超声速来流边界层燃烧减阻机理。开展超声速边界层燃烧减阻实验研究，结合流道压强、温度等参数分布及火焰结构特征分析，获得壁面摩擦阻力随来流及燃料喷注的变化规律。建立适用于超声速边界层燃烧的高精度数值计算方法，研究燃烧引入对边界层内部湍流结构及输运过程的影响。最后基于实验及数值模拟结果，揭示超声速边界层燃烧减阻机理并建立边界层燃烧减阻性能预示模型。该项研究不仅有助于揭示超声速边界层燃烧减阻机理，而且为进一步有效控制边界层燃烧流动，提升超燃冲压发动机性能提供理论依据。

超声速/高超声速下飞行器内流道阻力巨大，相对于其他减阻方式，边界层燃烧的减阻幅度更大且非常适合应用于超燃冲压发动机燃烧室内。本项目对边界层燃烧减阻机理及其在发动机内的应用开展研究。首先以超声速来流边界层基本模型为研究对象，对超声速来流下边界层喷注燃料、燃烧的基本规律进行探究。研究表明，边界层内燃烧释热导致边界层内气流密度的减小和厚度的增大，致使其法向速度梯度减小，进而降低摩擦阻力，这是超声速边界层燃烧减阻的主导宏观机理。随着空气/燃料温度比的增加，壁面摩擦阻力和传热均增加；空气中H2O的引入会抑制氧气与氢燃料的混合和燃烧，进而使减阻效果明显降低，但对壁面传热率影响较小；空气/燃料压力比增加时，燃料自点火位置提前，火焰更贴近壁面，从而使壁面传热量增大。基于形状因子H，发现燃料在边界层内的自点火所带来的“扰动”会诱发边界层转捩，致使摩擦阻力急剧增大。随着喷注角度的增加，转捩位置提前，但对减阻效果影响较小；燃料喷口直径的增大会导致转捩延迟，折射激波强度降低，进而使整体摩擦减阻效果较好；随后，设计了一种边界层燃烧装置，对边界层燃烧在超燃冲压发动机内的摩擦减阻特性进行研究，结果表明，在燃烧室中安装壁面射流装置时，用于产生推力的主燃料和用于减阻的壁面燃料的分配比为3:1时，可获得最佳的减阻幅度。由于主燃料和壁面喷气燃料燃烧的耦合，将壁面射流位置设置得太靠近上游或下游并不能获得最大的整体减阻效果。对不同燃烧室型面的影响研究表明，发动机下壁面采用扩张型面会抑制燃烧室内的边界层燃烧，不利于燃烧室内的壁面摩擦减阻，但促进了边界层燃烧火焰在尾喷管内的发展，尾喷管段的减阻效果更为明显；而当发动机流道下壁面向主流收缩时会增强燃烧室内的边界层燃烧，进而增大燃烧室内的减阻效果；相对于平直流道，燃烧室型面的收缩或扩张均在一定程度上降低了发动机壁面摩擦阻力，采用收缩型面构型的减阻效果最优。