高超声速飞机是指飞行速度达到或者超过5马赫的一类飞行器，因其具备远程快速运输能力，近年来已成为国际上研究的热点问题。由于该类飞行器的飞行速度涵盖亚、跨、超和高超声速，这就要求在气动布局设计上需兼顾全速域下的气动性能，包括：高超声速巡航阶段的高升阻比需求，水平起降阶段的高升力需求和全速域范围内的稳定性匹配需求等，因此气动布局设计的难度较现有飞行器而言大幅增加，是高超声速飞机的核心关键技术之一。

高压捕获翼新型气动布局基于有益气动干扰，通过在飞行器上方合理增加一个上置翼，可在高超条件下同时获得高容积率、高升力和高升阻比，此外，相比传统单翼布局，该布局的双升力面设计可使起降阶段的升力获得明显提升，为宽速域飞行器的气动布局设计提供了一条新思路。而对于该类新型气动布局，由于外形复杂，飞行器各部件之间存在强烈的流动耦合作用，且在不同飞行速域和飞行姿态条件下，气动干扰类型和流场结构将呈现明显差异，直接影响到飞行器的气动性能。基于上述背景，本文采用数值模拟方法，对高压捕获翼气动布局在宽速域下的流动耦合特性及其对气动性能的影响机理开展了系统性研究。首先，重点围绕机体与捕获翼间的流动耦合特性，以一种单翼原理性构型开展研究。然后，在此基础上进一步考虑下翼面的影响，以一种双翼原理型构型开展研究。

本文主要研究结果和结论如下：

1)研究并揭示了机体与捕获翼间的流场拓扑结构随来流马赫数变化的演化过程，总结获得了不同速域内的主要气动干扰形式。具体如下：亚声速来流条件下，机体中后段的圆台上表面由于逆压梯度的存在出现流动分离，且分离区范围随来流马赫数增加而逐渐增大。当来流速度继续增大进入跨声速域时，在机体拐点后产生激波并且与分离区相互作用，在下游产生二次激波，导致捕获翼下表面出现明显的压力波动。随来流马赫数逐渐增大，机体与捕获翼之间的超声速范围增大，激波出现位置后移，分离区逐渐缩小，二次激波基本消失。在超声速来流条件下，捕获翼前缘激波作用于机体上表面并且发生反射，随来流马赫数的不断增加，在机体与捕获翼之间以此出现单次反射、二次反射和多次反射等不同的激波反射形式。此外，研究结果还表明高压捕获翼和机体之间的支撑结构在宽速域内对流场特性均未产生本质影响。

2)研究了机体尾截面形状变化对宽速域气动特性的影响。首先，机体尾截面展向适当变宽可在一定程度上可改善宽速域气动性能：亚声速来流条件下，机体中后段的圆台上表面逆压梯度减弱，流动分离得到明显抑制。随来流攻角增大，圆台上表面逆压梯度减弱，分离区逐渐消失；当攻角进一步增大时，受捕获翼垂向抑制的影响，机体尾截面展向变宽可加快圆台背风面分离涡消失进程，并且可延缓机体背风面横向绕流发展。跨声速来流条件下，随机体尾截面展向变宽，来流经过机体拐点后产生的激波逐渐减弱，分离区范围明显缩小，并且随攻角增大，机体尾截面展向变宽可加速分离区消失进程；超声速来流条件下，随机体尾截面展向变宽，机体拐点处产生的膨胀波逐渐减弱，捕获翼下表面压力系数逐渐增大，有益于提高升力特性。此外与椭圆形尾截面相比，采用椭圆顶圆角矩形尾截面使整机的升力系数和阻力系数进一步增大，亚声速和超声速来流条件下对捕获翼的气动特性和机体与捕获翼间的流场结构无显著影响；跨声速来流条件下，当攻角较大时，受到机体三维效应的影响，机体圆台上表面会出现局部的流动分离。

3)在参数化设计的基础上对双翼原理性构型开展了宽速域气动特性分析，研究了设计参数对气动性能的影响，结果表明，捕获翼与下翼面间距对于整机宽速域气动特性的影响最为显著。当翼间距较小时，在亚声速来流条件下，下翼面前缘诱发的边条涡强度随攻角增大而增强，逐渐向后发展并影响捕获翼下表面，削弱捕获翼的增升效果；跨声速来流条件下，捕获翼与下翼面之间的激波将同时作用与捕获翼下表面和下翼面上表面，同样削弱增升效果；超声速来流条件下，捕获翼前缘激波作用在下翼面上表面，随来流马赫数增大，激波面逐渐扫过下翼面。导致与参考构型相比，双翼布局的高升力系数优势无法体现。当将捕获翼和下翼面的装配攻角适当增加，等效于两个翼面间距适当增大，两翼面之间的气动干扰逐渐减弱，捕获翼的增升效果改善，双翼原理性构型的高升力优势明显。

4)研究了双翼原理性构型中翼面几何形状对整机宽速域气动特性的影响。结果表明，随下翼面后掠角增大，捕获翼受到的遮挡效应减弱，捕获翼贡献的升/阻力系数系数增大，但整体的升/阻力系数和升阻比均减小。下翼面后掠角增大，亚声速时，下翼面产生的边条涡强度降低，对捕获翼的干扰减弱；跨声速时，随下翼面后掠角较大时，捕获翼前缘激波对下翼面无直接影响；超声速时，捕获翼前缘激波对下翼面的干扰区范围缩小，但干扰区内物面压力逐渐增大。随捕获翼后掠角增大，翼身组合体贡献的升/阻力系数略有增大，但整体的升/阻力系数略有减小升阻比基本保持不变。捕获翼后掠角增大，亚声速时，对下翼面无明显的影响；跨声速时，捕获翼前缘激波强度减弱位置后移，并且翼间流道内马赫数减小；超声速时，捕获翼前缘激波随捕获翼后掠角增大而减弱，在下翼面上的作用点逐渐后移，干扰区逐渐缩小，加速了激波面扫过下翼面的干扰进程。此外整机升/阻力系数与总俯视投影面积正相关，并且俯视投影面积变化量相同时，翼身组合体俯视投影面积变化对气动力的影响较捕获翼更加敏感。

本文的研究工作为高超声速飞机气动构型设计提供了一条新思路，相关结果可对工程设计提供理论指导和数据支撑。

Hypersonic aircraft refers to a class of aircraft with a flight speed of up to or exceeding Mach 5. Due to its long-range rapid transportation capability, it has recently become a focus of international research. Because the flight speed of this kind of aircraft covers subsonic, transonic, supersonic, and hypersonic, it is necessary to take into account the aerodynamic performance in the full speed range in the aerodynamic configuration design, including the high lift-to-drag ratio requirement in the hypersonic cruise phase, the high lift requirement in the horizontal take-off and landing phase, and the stability matching requirement in the full speed range. Therefore, the difficulty of aerodynamic layout design is greatly increased compared with existing aircraft, which is one of the core key technologies of hypersonic aircraft.

Based on beneficial aerodynamic interference, a new type of High-pressure Capturing Wing(HCW) aerodynamic configuration can achieve high volume ratio, high lift, and high lift-to-drag ratio under hypersonic conditions by reasonably adding an upper wing above the aircraft. In addition, compared to traditional single wing layouts, the dual lift surface design of this layout can significantly improve the lift during takeoff and landing, providing a new idea for the aerodynamic configuration design of wide-speed domain aircraft. Due to the complex shape, there is a strong flow coupling between the components of the aircraft, and under different flight speed and flight attitude conditions, the types of aerodynamic interference and flow field structure will show obvious differences, which directly affect the aerodynamic performance of the aircraft. Based on the above background, this dissertation systematically studies the flow coupling characteristics of the High-pressure Capturing Wing aerodynamic configuration in a wide speed range and its influence mechanism on the aerodynamic performance by numerical simulation. Firstly, focusing on the flow coupling characteristics between the fuselage and HCW, a single wing principle configuration was studied. Then, based on this, further consideration is given to the impact of the lower wing surface, and the study is conducted with a dual wing principle configuration.

The main research results and conclusions are as follows,

1) The evolution process of the flow field topology between the fuselage and HCW with the change of the incoming Mach number is studied and revealed, and the main aerodynamic interference forms in different speed domains are summarized. The details are as follows: Under subsonic inflow, the flow separation occurs on the upper surface of the round platform in the fuselage due to the existence of adverse pressure gradient, and the range of the separation zone gradually increases with the increase of the inflow Mach number. When the flow velocity continues to increase into the transonic region, a shock wave is generated after the inflection point of the fuselage and interacts with the separation zone, and a secondary shock wave is generated downstream, resulting in obvious pressure fluctuations on the lower surface of HCW. With the increase of the free stream Mach number, the supersonic range between the fuselage and HCW increases, the position of the shock wave moves backward, the separation zone gradually shrinks, and the secondary shock wave basically disappears. Under the condition of supersonic flow, the shock wave of the leading edge of the HCW acts on the upper surface of the fuselage and reflects. With the increase of the Mach number of the flow, there are different forms of shock wave reflection such as single reflection, secondary reflection, and multiple reflection between the fuselage and HCW until the hypersonic design state is reached. Furthermore, the results show that the strut structure has no significant effect on the flow field structure over a wide speed range.

2) The effect of changing the shape of the tail section on the aerodynamic characteristics of a wide speed range is studied. Firstly, the aerodynamic performance in the wide speed range can be improved to a certain extent by widening the spanwise dimension of the tail section of the fuselage. Under subsonic inflow, the reverse pressure gradient on the upper surface of the round platform is weakened, which has a significant inhibitory effect on the flow separation. As the angle gradually increases, the adverse pressure gradient on the upper surface of the cone weakens, and the separation zone gradually disappears. When the angle of attack further increases, the widening of the spanwise dimension can accelerate the disappearance process of the separation vortex on the leeward surface of the circular platform and delay the development of the transverse flow around the leeward surface of the fuselage. Under transonic incoming flow, with the spanwise widening of the tail section of the fuselage, the shock wave generated by the incoming flow passing through the inflection point of the fuselage gradually weakens, and the range of the separation zone is obviously reduced. With the increased angle of attack, the spanwise widening of the tail section of the fuselage can accelerate the disappearance of the separation zone. With the spanwise widening of the fuselage tail section under supersonic inflow, the expansion wave generated at the inflection point of the body gradually weakens, and the pressure coefficient of the lower surface of HCW gradually increases, which is beneficial to improving the lift characteristics. In addition, compared with the elliptical tail section, the lift coefficient and drag coefficient of the total aircraft are further increased by using the elliptical top fillet rectangular tail section, and the aerodynamic characteristics of HCW and the flow field structure between the fuselage and HCW are not significantly affected under subsonic and supersonic flow conditions. Under transonic flow, at high angle of attack, due to the influence of the three-dimensional effect of the fuselage, local flow separation occurs on the upper surface of the round platform.

3) A wide speed domain flow characteristic analysis for a dual wing principle configuration is performed using parametric design, and the influence of different design parameters on aerodynamic performance is studied. The results show that the distance between HCW and the lower wing surface has the most significant influence on the aerodynamic characteristics of the wide speed range. When the vertical distance is small, the strength of the strake vortex generated by the lower wing increases with the angle of attack and gradually develops backward, affecting the lower surface of HCW and weakening its effect. Under transonic inflow, the shock wave acts on the HCW and the lower wing surface at the same time, and the interaction between the two wings will also weakens the lift enhancement effect. Under supersonic inflow, the shock wave of the leading edge of HCW acts on the surface of the lower wing. As the incoming flow Mach number increases, the shock wave surface gradually sweeps over the lower wing. Affected by this, compared with the reference configuration without HCW, the high lift coefficient advantage of the dual wing configuration cannot be embodied out. When the wing installation angle of HCW and the lower wing is appropriately increased, the aerodynamic interference between the two wings is gradually weakened, the lift increase effect of HCW is improved, and the high lift advantage of the dual wing principle configuration is obvious.

4) Based on the previous work, the influence of the sweep angle of the lower wing and HCW on the aerodynamic characteristics of the whole aircraft in the wide speed range is further studied. The results show that with the increase in the sweep angle of the lower wing, the shielding effect of HCW is weakened, and the lift/drag coefficient contributed by HCW increases, but the overall lift/drag coefficient and lift-drag ratio decrease. The sweep angle of the lower wing surface increases, the intensity of the side vortex generated by the lower wing surface decreases at subsonic speed, and the interference to HCW is weakened. At transonic speed, the interference of the leading edge shock wave of the HCW on the lower wing surface is gradually weakened. When the sweep angle of the lower wing surface is large, the shock wave has no direct effect on the lower wing surface. At supersonic speed, the range of the interaction zone created by the leading edge shock wave of the HCW on the lower wing surface is reduced, but the surface pressure in the interaction zone is gradually increased. As the sweep angle of HCW increases, the lift/drag coefficient contributed by the wing-body combination increases slightly, but the overall lift/drag coefficient decreases slightly and the lift/drag ratio remains basically unchanged. When the sweep angle of HCW increases, there is no obvious effect on the lower wing surface at subsonic speed. At transonic speed, the shock wave intensity at the leading edge of the capture wing weakens and the action position moves backward. At supersonic speed, the shock wave at the leading edge of HCW weakens with the increase of the sweep angle of HCW, the action point on the lower wing surface gradually moves backward, and the interference zone gradually shrinks, which accelerates the interference process of the shock wave surface sweeping through the lower wing surface. In addition, the lift/drag coefficient of the total aircraft is positively correlated with the total top view projection area, and the influence of the change of the overhead projection area of the wing-body combination on the aerodynamic force is more sensitive than that of HCW.

The research work in this dissertation provides a new idea for the aerodynamic configuration design of hypersonic aircrafts, and the relevant results can be used as a reference for engineering design.