A shock wave/boundary layer interaction (SWBLI) is a common phenomenon in supersonic inlet flow, which can significantly degrade the aerodynamic performance of the inlet by inducing boundary layer separation. To address this issue, in this paper, we propose the use of a dynamic vortex generator to control the SWBLI in a typical supersonic inlet. The unsteady simulation method based on dynamic grid technology was employed to verify the effectiveness of the proposed method of control and investigate its mechanism. The results showed that, in a duct of finite width at the inlet, the SWBLI generated complex three-dimensional (3D) flow structures with remarkable swirling properties. At the same time, vortex pairs were generated close to the side wall as a result of its presence, and this led to the intensification of transverse flow and, in turn, the formation of a complex 3D structure of the flow of the separation bubble. The dynamic vortex generator induced oscillations of variable intensity in the vortex system in the supersonic boundary layer that enhanced the mixing between the boundary layer flow and the mainstream. Meanwhile, the unique effects of “extrusion” and “suction” in the oscillation process continued to charge the airflow, and the distribution of velocity in the boundary layer significantly improved. As the oscillation frequency of the vortex generator increased, its charging effect on low-velocity flow in the boundary layer increased, and its control effect on the flow field of the SWBLI became more pronounced. The proposed method of control reduced the length of the separation bubble by 31.76% and increased the total pressure recovery coefficient at the inlet by 6.4% compared to the values in the absence of control.

Interactions between the turbulent boundary layer and incident compression waves (ICWs), sufficiently strong to induce boundary layer separation, are investigated numerically and analytically. This paper proposes two typical interaction modes of ICWs and the boundary layer: one where the convergence point of the ICWs is located on a flat plate and another where the ICWs impinge on the flat plate in a dispersed manner. For the two typical interaction modes, corresponding aerodynamic profiles for ICW generators are established. The macroscopic flow structures evolve with different ICW ranges on the flat plate, and can be classified into three categories. First, when the ICWs cover a smaller range, flow structures resemble that of the interaction between a single incident shock wave and a boundary layer. Second, when the ICWs cover a moderate range, the reattachment point moves upstream to the impingement point of the terminal compression wave. In this case, the reattachment process still features a reattachment shock. Third, when the ICWs cover a large range, the reattachment shock wave degrades from a reattachment shock wave to a series of compression waves. Inviscid models of the above wave structures were established. Separation scales in two interaction modes are also investigated. Additionally, based on mass conservation within control volumes established for the interaction zones, the dimensionless interaction intensity and length for two modes are defined, and it is found that within the Mach number scope of this study, two modes exhibit different relationships between “the dimensionless interaction intensity” and “the dimensionless interaction length.”

This paper presents the findings of a wind tunnel experiment aimed at investigating the mode transition process of a two-dimensional over-under turbine-based combined cycle inlet at an incoming flow Mach number of 2.9. The study utilized high-speed schlieren and dynamic pressure acquisition systems to examine the evolution process of the shock-dominated flow structure of the high-speed duct during the mode transition process. Additionally, the impact of mode transition speed on the unstart/restart characteristics of the high-speed duct was analyzed. The results indicate that, during the forward mode transition process, the increasing captured airflow of the high-speed duct leads to a higher number of shock reflections and the shock train moves forward in the duct, ultimately resulting in unstart. The unstarting flow field exhibits a small oscillation characteristic dominated by the separation bubble located at the entrance. However, evident hysteresis characteristics were observed in the restart process during the reverse transition. Furthermore, a higher mode transition speed delays the unstart and restart of the high-speed duct, consequently increasing the hysteresis interval. Theoretical analysis suggests that a larger mode transition speed leads to lower mass accumulation efficiency in the high-speed duct, thereby slowing the pressure response and causing the shock train to lag forward, resulting in delayed unstart. The delay in the restart process is attributed to the relative slip motion of the separation bubble with the upper surface of the splitter, in addition to its forced motion.

Many UAVs today have an S-bend inlet for the sake of stealth; however, the majority of them have a relatively gentle transition of the flow channel. This study presents an experimental investigation of the aerodynamic performance and swirl flow characteristics of a UAV inlet with double 90° bends, which is also integrated with an aircraft fuselage as well as a volute. The influences of angle of attack, sideslip angle, AIP Mach number and freestream speed are explored in detail. The influences of the deflectors installed ahead of the first 90° bend of the inlet and the baffle installed at the bottom of the volute are revealed. It is found that both the deflectors and the baffle are beneficial in enhancing the aerodynamic performance of the inlet and alleviating the intensity of the swirl flow inside the volute.

In the vertical take-off and landing (VTOL) state of tiltrotor aircraft, the inlet entrance encounters the incoming airflow at a 90° attack angle, resulting in highly complex internal flow characteristics that are extremely susceptible to gusts. Meanwhile, the flow quality at the inlet exit directly affects the performance of the aircraft’s engine. This work made use of an unsteady numerical simulation method based on sliding meshes to investigate the internal flow characteristics of the inlet during the hover state of a typical tiltrotor aircraft and the effects of head-on gusts on the inlet’s aerodynamic characteristics. The results show that during the hover state, the tiltrotor aircraft inlet features three pairs of transverse vortices and one streamwise vortex at the aerodynamic interface plane (AIP). The transverse vortices generated due to the rotational motion of the air have the largest scale and exert the strongest influence on the inlet’s performance, which is characterized by pronounced unsteady features. Additionally, strong unsteady characteristics are present within the inlet. Head-on gusts mainly affect the mechanical energy and non-uniformity of the air sucked into the inlet by influencing the direction of the rotor’s induced slipstream, thereby impacting the performance of the inlet. The larger head-on gusts have beneficial effects on the performance of the inlet. When the gust velocity reaches 12 m/s, there is a 1.01% increase in the total pressure recovery (σ) of the inlet, a 25.72% decrease in the circumferential distortion index (DC60), and a reduction of 62.84% in the area where the swirl angle |α| exceeds 15°. Conversely, when the gust velocity of head-on gusts reaches 12 m/s in the opposite direction, the inlet’s total pressure recovery decreases by 1.13%, the circumferential distortion index increases by 14.57%, and the area where the swirl angle exceeds 15° increases by 69.59%, adversely affecting the performance of the inlet. Additionally, the presence of gusts alters the unsteady characteristics within the inlet.