Neng Xiong

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.

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.

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 data processing and uncertainty evaluation method of the five-hole probe, considering compressibility, is studied to meet the application requirements of the five-hole probe in a high-speed wind tunnel. The data processing method that considers compressibility is theoretically derived in detail. Based on interpolation fitting and an iterative solution, a data analysis method is established. On this basis, a Monte Carlo simulation method is further established to evaluate the uncertainty of the parameters of the flow field to be measured. It provides a theoretical basis and technical means for the calibration and application of a five-hole probe, considering compressibility in a high-speed wind tunnel.

In order to meet the need of evaluating the effectiveness of airfoil drag reduction design, the calculation and uncertainty analysis method for airfoil drag coefficient considering compressibility is studied. Based on the momentum conservation equation, the method of solving the drag coefficient of a two-dimensional airfoil is theoretically derived in detail. The drag coefficient integral equation is established and optimized by converting the integrated variable. On this basis, the accurate measurement test method and uncertainty analysis method for airfoil drag coefficient are established, which provides a theoretical basis and technical means for accurate measurement and uncertainty analysis of airfoil drag coefficient in high-speed wind tunnels.

To clarify the influence of the serpentine nozzle configurations on the flow characteristics and aerodynamic performance of aircraft, the flow features and aerodynamic performances of the double-ducted serpentine nozzles with different aspect ratios (AR), length–diameter ratios (LDR) and shielding ratios (SR) are numerically investigated. The results show that the asymmetric nozzle flow occurs due to the curved profile of serpentine nozzles, and a local accelerating effect exists at the S-bend, causing the increase in wall shear stress. The unilateral unsymmetrical expansion of the tail jet in the upward direction interacts with the separated external flow of the afterbody, forming an obvious cross-shock wave and shear layer structure. The surface pressure of the afterbody increases along the external flow direction, and decreases sharply in the separation point of the boundary layer. With the increase in AR and LDR, the local accelerating effect of the nozzle flow weakens, while with the increase in SR, the accelerating effect increases. The total pressure recovery coefficient, flow coefficient and axial thrust coefficient all decrease with the increase in AR, LDR, and SR. The thrust vector angle decreases with the increase in AR but is less affected by LDR and SR.

The aircraft wake vortex is an important factor affecting flight safety; as an important part of the aircraft, the powered nacelle will inevitably have an important impact on the aircraft wake vortex, so it is of great practical significance to research it. The present study focused on the numerical simulation of the wake flow of large aircraft (as the front aircraft) and the comparative analysis of the influence of engine jets on the wake flow. In order to meet the accuracy requirements and control the consumption of computing resources, LES and RANS methods were compared, and the RANS method was finally selected for subsequent calculation. The dynamic effect of jet flow was simulated by simplifying the boundary conditions of the inlet fan and outlet bypass as the mass flow boundary condition. The simulation results showed that the engine nacelle will have a significant impact on the morphology of the aircraft wake flow (position and strength of the main vortex in the wake flow system), which is caused by the vortices formed under the shear flow and separated flow of the nacelle. However, the nacelle will not significantly change the total strength of the wake vortex (half-plane circulation). The engine jet intensity causes additional turbulent mixing, which will accelerate the fusion of the nacelle vortex and ultimately change the intensity ratio of the inner wing vortex and the wingtip vortex, affecting the trajectory of the wake of the mean vortex. The study provides a corresponding reference for the following research on a wake vortex by a powered nacelle.

To study the aerodynamic characteristics of the convex structure of a surface-monitoring device on a high-speed train and to evaluate its impact on the aerodynamic performance of the high-speed train, numerical simulation research was conducted on three different layouts of the monitoring device. The computational fluid dynamics (CFD) method was used for the simulation study, and the unsteady compressible NS equation was used as the control equation. Hexagonal grid technology was used to reduce the demand for the grid quantity. The rationality of the grid size and layout was verified through grid independence research. To increase the accuracy of the numerical simulation, the γ-Re_θ transition model and improved delayed detached eddy simulation (IDDES) method were coupled for the simulation research. The aerodynamic characteristics of the different operation directions and configurations were compared and analyzed. The research results showed that the windward side of the single pantograph detection device experienced positive pressure, and the sideline and leeward sides experienced negative pressure. Increasing the fillet radius of the sideline could appropriately reduce the aerodynamic resistance. When the speed was about 110 m/s, the drag force coefficient of the detection device was 210~410 N, and the lateral force was small, which means that it had little impact on the overall aerodynamic force of the train. According to the results of the unsteady analysis of the layout with a large space, the resistance during forward travel was greater than that during negative travel. The streamlined upwind surface was conducive to reducing the scope of the leeward separation zone and the amplitude of the pressure fluctuation in the leeward zone, and it thus reduced the resistance. For the running trains, a vortex was formed on their leeward surface. The pressure monitoring results showed that the separated airflow had no dominant frequency or energy peak. The possibility of the following train top and other components experiencing resonance damage is low.

Cryogenic wind tunnels provide the for possibility aerodynamic tests to take place over high Reynolds numbers by operating at a low gas temperature to meet the real flight simulation requirements, especially for state-of-the-art large transport aircrafts. However, undesirable temperature gradients between the test model and the surroundings will be caused by the thermal non-equilibrium, changing the boundary layer characteristics and resulting in test errors. To study the nonadiabatic wall effects on the aerodynamic characteristics of the model in cryogenic wind tunnels, a numerical study was carried out for the CHN-T1 standard model under different wall temperature gradients. A code with a finite volume method and γ-Re_θt transition model were used. The analysis concluded that the change in wall temperature significantly affects the surface pressure distribution, transition position and skin-friction coefficient of the model, thus varying the lift and drag coefficients of the aircraft. The influences on the flow characteristics of both laminar and turbulent boundary layers by the wall temperature gradient were also investigated.

Lossy dielectric bars have potential applications in high power microwave (HPM) radiation-effect experiments performed in wind tunnels. Under plane excitation, beat wave patterns are observed along lossy and lossless dielectric bars in 3-D electromagnetic simulation results. Mode analysis is performed to clarify the two main modes in the beat waves—the leaky and guided modes. Comparisons are made between the theoretical and simulated results and acceptable agreements are obtained. The absorbing array of lossy dielectric bars is further investigated. The array can attenuate 90% of the microwave power at 10 GHz when a ten-meter series of such arrays is arranged. It is expected that the proposed absorbing array can achieve the required attenuation, considering that the testing object under microwave radiation is also an absorbing structure itself.