As an important aerodynamic configuration of the new-generation UAV, the dorsal S-shaped inlet’s performance is affected by the complex coupling of inflow conditions and the boundary layer ingestion effect. To investigate the influence mechanisms of these factors on inlet performance, CFD based on the scale-adaptive simulation (SAS) turbulence model is used to systematically analyze the flow field and performance of a UAV dorsal S-shaped inlet within a typical flight envelope. It is found that with increasing Mach number (0.6–0.9), the exit total pressure recovery decreases significantly, while the circumferential distortion coefficient almost doubles. As the angle of attack varies from −10° to 10°, a slight decrease in total pressure recovery is observed, but distortion improves due to a relatively stable separation region. Changes in sideslip angle have minimal impact on overall performance but notably alter the symmetry of the vortex system, resulting in a decrease in distortion coefficient. Additionally, at a specific Mach number, back pressure correlates positively with inlet performance. The increase in back pressure can effectively inhibit the flow separation and enhance the total pressure recovery, while the distortion coefficient decreases. The research results provide an important theoretical basis for the design optimization of the new-generation UAV.

The unsteady buzz characteristics of a dorsal supersonic bump inlet for a flying-wing configuration are investigated using wind-tunnel tests and three-dimensional unsteady numerical simulations. This study focuses on off-design operation at a freestream Mach number of 1.8, with particular attention to the effects of angle of attack and downstream throttling on shock motion, pressure oscillation, and inlet stability. Wind-tunnel measurements show that the onset and development of buzz are highly sensitive to angle of attack. At high angles of attack, pressure oscillations first appear near the inlet compression surface and subsequently develop into large-amplitude fluctuations at the aerodynamic interface plane. The dominant experimental buzz frequency is approximately 60–70 Hz, and the numerical prediction of 71 Hz agrees well with the measured dominant frequency of 66 Hz. The simulations further reveal a strongly three-dimensional buzz cycle in which asymmetric separation over the bump, spanwise accumulation and discharge of low-energy flow, and alternating inlet blockage and recovery govern the large-amplitude shock excursion. The oscillatory flow field is dominated by shock–system expulsion and ingestion on the spanwise side with stronger back-pressure tolerance, accompanied by the formation of strong and weak shear layers during different stages of the cycle. These results provide insight into the buzz mechanism of dorsal bump inlets and support the assessment of starting performance and stable operating limits for supersonic inlets integrated with flying-wing configurations.

The S-shaped inlet is increasingly used in modern aviation for its compact layout and stealth benefits, but its complex geometry induces strong pressure gradients and secondary flows that impact performance. Existing studies on S-shaped inlets are mostly based on the rigid-wall assumption, neglecting deformation of lightweight structures under aerodynamic loads and their feedback effects on the flow field. This study investigates fluid–structure interaction (FSI) effects using a scale-adaptive simulation (SAS) with the Spalart–Allmaras turbulence model, coupled with a finite element structural solver via a bidirectional tightly coupled approach. Numerical simulations compare rigid and elastic S-shaped inlets, analyzing the influence of Mach number (0.2–0.8), angle of attack (−4° to 8°), and sideslip angle (0–10°). Results show that wall elasticity alters the internal flow field, delaying secondary flows and inhibiting vortex development. At higher Mach numbers (Ma ≥ 0.6), local supersonic regions and shock waves form in the bend, intensifying separation and increasing total pressure loss and distortion. Angle of attack has limited impact within 0–8°, while sideslip angle induces asymmetric streamwise vortices, redistributing outlet pressure with minimal effect on average performance. These findings offer theoretical guidance for designing S-shaped inlets that account for aeroelastic effects.

Boundary layer suction is a critical technique in hybrid laminar flow control (HLFC) for delaying transition and reducing drag. While the effectiveness of suction is well-established, systematic studies on the parametric optimization of suction hole diameter, location, and coefficient for two-dimensional airfoils remain scarce. This study addresses this gap through numerical investigations using the validated γ-${\widetilde{Re}}_{\theta t}$ transition model. The research systematically analyzes the synergistic effects of suction coefficient (C_q), location (5%, 10%, and 15% chord), and suction hole diameter (0.2 mm, 0.6 mm, and 1.0 mm) on transition characteristics and aerodynamic performance. The results reveal that suction location predominantly governs the viscous drag coefficient (C_Dv), whereas suction hole diameter primarily influences the pressure drag coefficient (C_Dp). Consequently, suction location selection proves more critical for drag reduction than suction hole diameter. The maximum drag reduction (11.9% decrease in C_D) and optimal transition delay (11.8% chord shift) are achieved using a small suction hole (0.2 mm) located at an aft position (15% chord) with a high suction coefficient. Furthermore, an optimal matching range exists between suction location and coefficient, which widens with decreasing suction hole diameter. Based on these findings, this study proposes an energy-efficient design strategy: employing small apertures across the suction region while gradually increasing suction rates toward the trailing edge to achieve significant drag reduction with minimal energy penalty.

To investigate the coupled internal–external flow and spillage characteristics of high-bypass-ratio turbofan nacelles, an improved compressor body-force model is developed. Starting from the three-dimensional Navier–Stokes equations, circumferentially averaged body-force source terms are extracted from baseline single-passage simulations. The body-force distribution is then normalized, concentrated circumferentially within blade regions using blade geometric data, and corrected radially using five-hole-probe measurements at the rotor inlet and outlet. By correlating the body force with two local aerodynamic parameters, namely the flow following angle and the axial velocity, the model responds to variations in the intake mass-flow rate and inflow angle. The model is first validated for NASA Rotor 67: the predicted characteristic curves agree with the single-passage RANS results within 0.5% over the entire operating range, and the spanwise distributions of total pressure, total temperature, and relative flow angle deviate by less than 1.5% over the 10–90% blade span, with larger deviations confined to the hub and tip regions. Coupled nacelle–fan simulations are then performed, and a powered nacelle spillage-drag test is conducted in a 2.4 m transonic wind tunnel. At Ma = 0.75, the numerical and experimental spillage-drag results show good agreement, and the predicted spillage-drag coefficient is approximately 0.026, with the nacelle frontal area as the reference area.