Boundary-Layer-Ingesting (BLI) propulsion technology uses a rear mounted engine to draw in the boundary layer on the fuselage, reducing the airflow velocity at the inlet of the intake duct and the outlet of the propulsion system, thereby achieving efficient flight of the aircraft. Compared to traditional propulsion systems, BLI propulsion systems have the potential to significantly reduce fuel consumption, flight resistance, and pollution emissions, and are a key development direction in the field of aircraft and engine integration in the future. As a specific practitioner of boundary layer ingesting, the performance of the BLI inlet directly affects the comprehensive benefits of the entire BLI propulsion system compared to traditional propulsion systems. Since its emergence, it has attracted widespread attention from scholars and research institutions at home and abroad. Based on this, this article provides an overview of the research background and progress of BLI intake duct technology, with a focus on discussing the aerodynamic design methods, research methods, internal flow mechanism, and internal flow control technology of BLI intake ducts. Finally, based on the analysis and summary of the current research status at home and abroad, the future development of BLI intake duct is discussed.

To induce the near-wall low-momentum fluids moving laterally, an aerodynamic design method of hypersonic forebody/compression surface with the controlled lateral surface pressure is proposed. The basic principle is that the lateral pressure distribution on one section after the external conical flow field is prescribed, and the spatial coordinates of the section can be derived inversely through the coordinate transform, then the forebody/compression surface can be obtained by the stream tracing method. The numerical results demonstrate that the lateral pressure gradient dominates the motion of the near-wall low-momentum fluids on the forebody/compression surface. For the conventional forebody/compression surface, it has strong lateral pressure gradient on the 1^st stage of the forebody, which can induce a lateral flow with a deflection angle below 3° at the design point (Ma = 7.0 and H = 28 km). However, it almost has no lateral pressure gradient on the subsequent compression surfaces, and the lateral flow is also fairly weak. The controlled lateral pressure distribution forebody can intensify the lateral pressure gradient about 7 times within the sector-angle ranging from 0° to 40°, the deflection angle increases about 5° on the 1^st stage of the forebody, and the lateral pressure gradient increases significantly with the deflection angle increased over 7° on the 2^st and 3^rd stages of the forebody. Consequently, the boundary layer thickness decreases about 20%, and the total-pressure recovery coefficient in the sector region of the inlet increases about 1.56%.

A flow control method using multiple micro-vane vortex generators is proposed to improve the fullness of boundary layer profiles over the forebody ramp of hypersonic vehicles and to reduce the potential risk of flow separation in the inlet. The flowfield characteristics and mixing mechanism at an incoming Mach number of 7.0 are numerically analyzed. The effect of the installation angle of vortex generators on the flowfield is studied as well. Results show that micro-vane vortex generators can generate a local large-sideslip-angle and low-pressure flow in the near-wall region. At the lateral sides of mocro vanes, glancing shocks and expansion waves are generated, which induce the lateral mixing of low-momentum fluid with the mainstream. When installation angles are positive, with the increase of installation angle, the mixing effect gets stronger but the pressure loss increases. Micro vanes with negative installation angles lead to the strongest mixing effect, however, they also bring significant total pressure loss. Compared with the one without flow control, the shape factor of the boundary layer gets smaller by using vortex generators, among which those with an installation angle of 15 degrees lead to the smallest shape factor and consequently the fullest boundary layer profile, which has better anti-separation capability under reverse pressure gradients.

Cowl-induced incident Shock Wave/Boundary Layer Interactions (SWBLI) under the influence of gradual expansion waves are frequently observed in supersonic inlets. However, the analysis and prediction of interaction lengths have not been sufficiently investigated. First, this study presents a theoretical scaling analysis and validates it through wind tunnel experiments. It conducts detailed control volume analysis of mass conservation, considering the differences between inviscid and viscous cases. Then, three models for analysing interaction length under gradual expansion waves are derived. Related experiments using schlieren photography are conducted to validate the models in a Mach 2.73 flow. The interaction scales are captured at various relative distances between the shock impingement location and the expansion regions with wedge angles ranging from 12° to 15° and expansion angles of 9°, 12°, and 15°. Three trend lines are plotted based on different expansion angles to depict the relationship between normalised interaction length and normalised interaction strength metric. In addition, the relationship between the coefficients of the trend line and the expansion angles is introduced to predict the interaction length influenced by gradual expansion waves. Finally, the estimation of normalised interaction length is derived for various coefficients within a unified form.

This study aims to investigate the intricate dynamic characteristics of the high-speed duct during the over-under Turbine-Based Combined Cycle (TBCC) inlet mode transition process while operating in an off-design state under throttled conditions. A typical over-under TBCC inlet, designed for a working Mach number range of 0–6 with a transition Mach number of 3.5, is examined through experimental studies in a supersonic wind tunnel with a freestream Mach number of 2.9. The investigation focuses on the complex oscillatory flow and unique hysteresis observed in the mode transition process of the high-speed duct under the mildly throttled condition, utilizing high-speed schlieren and dynamic pressure acquisition system. The findings reveal that the high-speed duct undergoes four distinct oscillation stages akin to those in a higher throttled state during the mode transition, albeit with smaller dominant frequency and energy. Moreover, an irregular alternating “big/little buzz” mode is observed in the early stage of the large oscillation stage. Notably, the mildly throttled state exhibits three intriguing hysteresis properties compared to the unthrottled and higher throttled states. Firstly, hysteresis is observed in the shock train motion stage in the duct before unstart, along with the corresponding inverse process. Subsequently, hysteresis is noted in the unstart and restart of the high-speed duct, with a smaller hysteresis interval than in the unthrottled state. Finally, the hysteresis characteristics of oscillation mode switching and the corresponding inverse process are explored. Based on the analysis, the first two hysteresis phenomena are associated with the formation and dissipation of the separation bubble. The significant adverse pressure gradient constrains the cross-sectional capacity of the channel, rendering the disappearance of the separation bubble more challenging. The hysteresis in oscillation mode switching is linked to not only the channel cross-sectional capacity but also the state of the incoming boundary layer.

The study presents an experimental exploration into the mode transition of an over-under TBCC (Turbine-Based Combined Cycle) inlet, with a specific emphasis on the flow characteristics at off-design transition Mach number. A systematic investigation was undertaken into the mode transition characteristics in both unthrottled and throttled conditions within a high-speed duct, employing high speed Schlieren and dynamic pressure acquisition systems. The results show that the high-speed duct faced flow oscillations primarily dictated by the separation bubble near the duct entrance during the downward rotation of splitter, leading to the duct’s unstart under the unthrottled condition. During the splitter’s reverse rotation, a notable hysteresis of unstart/restart of the high-speed duct was observed. Conversely, hysteresis vanishes when the initial flowfield nears the critical state owing to downstream throttling. Moreover, the oscillatory diversity, a distinctive characteristic of the high-speed duct, was firstly observed during the mode transition induced by throttling. The flow evolution was divided into four stages: an initial instability stage characterized by low-frequency oscillations below 255 Hz induced by shock train self-excitation oscillation and high-frequency oscillations around 1367 Hz caused by the movement of separation bubble. This stage is succeeded by the “big buzz” phase, comprised of pressure accumulation/release within the overflow-free duct and shock motion outside the duct to retain dynamic flow balance. The dominant frequency escalated with the increase of the internal contraction ratio in the range of 280 Hz to 400 Hz. This was followed by a high-frequency oscillation stage around 453 Hz dominated by a large internal contraction ratio with low pulsating energy, accompanied by a continuous supersonic overflow. Lastly, as the splitter gradually intersected the boundary layer of the first-stage compression surface, the capture area and the turbulence intensity of the incoming flow underwent a sudden shift, leading to a more diverse flow oscillation within the duct, manifested as various forms of mixed buzz.

Aerospace vehicles are expected to play extremely important roles in military and civilian applications. One of the key technologies is the airbreathing combined-cycle engine that provides propulsion within the atmosphere. As a key component of the air-breathing combined-cycle engine, the air inlet urgently requires solutions to broad geometric/aerodynamic adjustment requirements, the intricate organization and prediction needs of wide-area internal flow, and the smooth transition demands between multiple engine modes. In response to these three major needs, the variable air inlet system necessitates further research in fixed geometric/variable geometric adjustment technology, shock-dominated flow physics and flow control technology, and the dynamic evolution characteristics of the flow field during mode transition. Regarding the current state of research, rotation is the primary adjustment method for rectangular inlets; translation is the principal adjustment implementation method for axisymmetric inlets; deformable modulation holds a significant complementary and auxiliary value for both rectangular and axis-symmetric adjustable inlets. Researchers, both domestically and internationally, have gained a deep understanding of shock wave/boundary layer interactions, inlet buzz mechanisms, and unstart/restart hysteresis mechanisms, and made considerable progress in the modelling and prediction of the unstart/restart boundaries. Boundary layer bleeding remains an effective control measure against various detrimental phenomena in shock-dominated flows. In terms of mode transition, external over-under inlets exhibit relatively weak inter-flowpath coupling interaction, whereas internal over-under inlets display more significant inter-flowpath coupling effects. In the future, emphasis should be placed on developing efficient and lightweight adjustment technology, establishing a rapid design system for wide-envelope adjustable inlets, strengthening research on complex internal flow mechanisms during mode transition, and actively intersecting with engine control disciplines to break through intelligent inlet control technology, thereby supporting the development of China’s aerospace vehicles to a new level.

The hysteresis during the throat regulation process of a supersonic variable inlet is unconducive to restart. Hence, detailed experimental studies of such a hysteresis and its control are necessary. A throat variable supersonic inlet was designed at a shock-on-lip Mach number of 4.0 and an Internal Contraction Ratio (ICR) ranging over 1.21–2.94. Meanwhile, a distributed bleed system was proposed to suppress the hysteresis. The wind tunnel tests were conducted at Mach number 2.9. The throat regulation processes were recorded using a high-speed schlieren and dynamic pressure acquisition system. The results indicate that the unstart and restart ICRs during the uncontrolled inlet’s throat regulation process were 1.95 and 1.48, respectively, demonstrating an unstart-restart hysteresis. Four typical flowfields were summarized during the uncontrolled inlet’s restart process. The proposed bleed control increased the unstart and restart ICRs to 2.06 and 1.75, respectively, and the inlet realized the designed state as the ICR was further decreased to 1.67. The controlled inlet’s hysteresis loop was decreased compared to the uncontrolled inlet. Finally, the mechanism of the hysteresis, dominated by the entrance separation-induced wave system, was clarified. The mechanisms of the bleed control to broaden the unstart and restart boundaries and suppress the hysteresis were elucidated.

The inlet with scavenge duct is an important part of turboprop aircraft engine. This type of inlet normally has a complex shape, of which the design is challenging and directly affects the flow field quality of the engine entrance and thus the engine performance. In this paper, the parametric design method of a turboprop aircraft inlet with scavenge duct is established by extracting and controlling the transition law of the critical characteristic parameters. The inlet’s performance and internal flow characteristics are examined by wind-tunnel experiment and numerical simulation. The results indicate that a flow tendency of winding up on both sides is formed due to the induction of the inlet profile, as well as a vortex pair on the back side of the power output shaft. The vortex pair dominates the pressure distortion index on the Aerodynamic Interface Plane (AIP). In addition, with the increase of freestream angle of attack, the total-pressure recovery coefficient of the inlet increases gradually while the total pressure distortion index decreases slightly. On the basis of the experimental results under different working conditions, the parametric design method proposed in this paper is feasible.

The variable geometry supersonic inlet tends to decrease the throat area to reduce the Mach number upstream of the terminal shock, so as to reduce the flow loss. However, excessive Internal Contraction Ratio (ICR) exposes the inlet to a greater risk of unstart, which inevitably results in a process of increasing the throat area to aid the inlet restart. In the above throat regulation process, the inlet undergoes the start, unstart, and restart states in turn. In order to reveal the flow structure and mechanism of this process, a two-dimensional unsteady numerical simulation combined with a dynamic mesh technique were employed. The shock-on-lip Mach number of the studied inlet is 4.0 and the flight angle of attack is + 6°. Analysis was focused on the state with a freestream Mach number of 3.0. The results clearly show that the flow response hysteresis appears, and restart is only realized when the throat area is obviously increased as compared to that of unstart due to the historical unstart flow structure. In addition, three typical flow fields were analyzed, and it is found that the separation ahead of the inlet was the key factor affecting the hysteresis. Finally, unstart and restart boundaries of the inlet were discussed, and the factors influencing its deviation from the typical boundaries of dual-solution area were analyzed. The newly predicted unstart and restart boundaries are much closer to the CFD results.