Zhengyin Ye

This study employs direct numerical simulations to investigate the influence of a vibrating wall panel on the oblique shock wave/turbulent boundary layer interaction flows at incoming free-stream Mach number 2.9. We examine two distinct low-frequency vibration modes, defined by their streamwise deformation patterns: a full-panel (symmetric) mode and a half-panel (antisymmetric) mode. By analysing phase-average flow quantities, we demonstrate that the antisymmetric mode primarily forces a streamwise oscillation of the separation bubble, whereas the symmetric mode amplifies its inherent expansion–contraction (‘breathing’) motion. Both panel vibrating modes enhance quasi-periodic shock oscillation and modulate the plateau in the mean wall pressure distributions. Spectral analysis confirms that low-frequency pressure fluctuations are amplified not only at the mean separation point but also in the mid-region of the panel. Through correlation analysis and dynamic mode decomposition, we establish that these vibrations lead to strong spatiotemporal coherence, associating upstream fluctuations near the separation to downstream propagation. Critically, a comparative analysis reveals that the symmetric mode facilitates significantly longer downstream propagation of pressure disturbances than the antisymmetric mode. This highlights how the spatial pattern of panel deformation, not just the vibration frequency, modulates the unsteady flow characteristics.

Lightweight structural design improves vehicle performance but reduces stiffness, making prescribed wall motion relevant to shock-reflection behavior. This study employs computational fluid dynamics simulations with a prescribed harmonic motion of the wedge to investigate how the vibration amplitude and frequency influence asymmetric shock reflections. The flow fields of both Mach reflection (MR) and regular reflection (RR) are shown to exhibit four evolution forms under the vibrating boundary. The corresponding distribution domains and evolution trajectories with respect to vibration frequency are summarized. Further analysis shows that increasing the vibration to upper or lower limits delays MR → RR transitions and advances RR → MR transitions within a vibration cycle, whereas increasing frequency postpones both transitions. Based on this, we examine the variation of MR–RR transition boundaries with amplitude and frequency. From the perspective of flow physics, the instantaneous MR–RR transition in the present unsteady cases is shown to be consistent with the sonic criterion. It is also found that the flow field exhibits a notable lag relative to the prescribed wedge motion, which is identified as a key mechanism contributing to the MR–RR transition-boundary variation and hysteresis behavior in the present prescribed-wedge-motion cases. These findings clarify the mechanisms of vibration-induced asymmetric shock-reflection evolution and provide physical insight into high-speed configurations involving moving or vibrating walls.

Laminar–turbulent transition in boundary layers is of fundamental significance in the design of high-speed vehicles. Subcritical transition occurring upstream of the linear instability region is frequently observed in high-speed blunt wedge flows. To identify the nonmodal perturbations which may lead to subcritical transition, the boundary layer instability of a 500 mm-length blunt-nose wedge with Mach 7.0 is investigated by the nonmodal stability method, the BiGlobal stability analysis method, and direct numerical simulation. The nonmodal stability method is developed by solving the harmonic linearized Navier–Stokes equations and the corresponding adjoint system. This method is capable of finding the most dangerous nonmodal perturbations (optimal perturbations). The optimal perturbations are shown to be temporally steady (with zero frequency), streamwise growth, and spanwise periodic. These perturbations can evolve into streamwise streaks, which are prone to secondary instabilities. The BiGlobal results indicate that the most unstable secondary instability mode exhibits a high growth rate, with the N factor exceeding more than 14, which is sufficient to trigger the boundary layer transition. Mode transformation from subcritical sinuous modes into most unstable sinuous modes is verified. During the nonlinear evolution, higher-frequency harmonic modes are excited. The nonlinear evolution of the secondary unstable modes exhibits a relatively uniform saturation amplitude, which is subjected to the mean flow distortion (MFD). The MFD plays a major role during the nonlinear process. This study provides the physical mechanisms behind nonlinear secondary instability modes in high-speed boundary layers.

An efficient and accurate aerodynamic heating analysis method is critical for thermal protection system (TPS) design in hypersonic vehicles. A computational fluid dynamics (CFD) coupled Newtonian cooling (NC) framework, which was proposed to predict aerothermal and drive thermal conduction efficiently, is first validated on four representative hypersonic configurations, covering flow complexities from axisymmetric bow shocks to shock/shock interactions. Compared with high-fidelity CFD coupled computational thermal dynamics, the CFD/NC approach maintains structural temperature response prediction errors at corresponding nodes within 5% while significantly reducing computational cost, especially for the three-dimensional cases. Moreover, a segmented coupling strategy (SCS) is proposed, which periodically updates heat transfer coefficients with CFD data, thereby extending the robustness and applicability of the CFD/NC approach to long-duration cases. This study advances the CFD/NC framework by providing systematic validation on diverse hypersonic configurations and by introducing SCS that extends its reliability for long-duration heating. These developments yield a balanced and scalable approach for rapid TPS design optimization and also offer potential for integration into aerothermoelastic analyses of hypersonic aircraft.

As a typical flow structure in hypersonic vehicles, the shock wave/boundary layer interaction in compression corners leads to increasingly prominent structural vibration issues due to aerodynamic heating and lightweight design requirements. In this study, the influences of panel vibration with different amplitudes and frequencies on the supersonic compression corner flow characteristics are numerically studied. The results show that in all calculation cases, as the panel amplitude increases, the size of separation bubble at the corner gradually decreases during moments when no separation bubble is generated in the panel region. When the panel amplitude increases to a certain value, a separation bubble is generated in the panel region during the period of large panel concavity. This causes the size of the separation bubble at the corner to significantly increase and the separation shock to significantly move upstream and then leads the high-frequency proportion in the pressure fluctuation in the panel region to increase and the skin friction coefficient to decrease significantly. As the panel vibration frequency increases, the generation of the separation bubble in the panel region is gradually delayed. When the panel vibration frequency increases to a certain value, the separation bubble will not be generated in the panel region. As the panel vibration frequency further increases, the separation bubble at the corner almost disappears during the period of small panel concavity. This study can enhance the understanding of the influences of panel vibration on the compression corner flow characteristics.

Swept shock wave/boundary layer interaction (SSBLI) is a typical flow phenomenon in hypersonic vehicles. Under the influence of lightweight structural design, the stiffness of the panel will be significantly reduced, which can easily lead to fluid–structure interaction (FSI) issues. Based on an in-house FSI program, the dynamical characteristics of the flexible panel and the influence of the FSI on flow characteristics was investigated by varying the area of a plate immersed in the separation zone induced by SSBLI. The results show that as the immersion area increases, the dominant mode of the panel changes from low-order to high-order, and the pressure and displacement response transition from limit cycle oscillation to multi-frequency irregular oscillation. The evolution of pressure pulsations at the line of surface-flow coalescence and the line of surface-flow divergence in the flow field is consistent with that of the panel. Furthermore, the degree of destruction of the quasi-conical characteristics increases with the increase in the immersed area. Moreover, as the displacement of the panel increases, the “depression” of the separation shock wave becomes more pronounced, and the expansion wave gradually strengthens and moves farther away from the sharp fin. This study provides an important reference for the refined design of aircraft under complex hypersonic flows.

The uncertain factors in wind tunnel experiments, especially the interference effects, will cause the flow around the test model to differ from that under the free flow condition. The wind tunnel wall interference effects are challenging to be eliminated, especially in transonic flow regions where classical linear correction methods are ineffective. To address this issue, an efficient wind tunnel wall interference correction framework based on data assimilation is proposed for steady and shock buffet flows around the airfoils in the transonic region in this work. The ensemble Kalman filter (EnKF) is used to find the combination of inflow conditions, which minimizes the differences between the pressure coefficients measured in the wind tunnel experiment and those estimated by computational fluid dynamics (CFD). The polynomial chaos expansion method is used to propagate uncertainty in the forecast step of the EnKF in order to improve efficiency. Typical wind tunnel wall interference correction problems are studied, including two steady cases: the transonic flow around the Royal Aircraft Establishment 2822 (RAE2822) and the National Aerospace Laboratory 7301 (NLR7301) airfoils, and one unsteady case: the transonic shock buffet on the National Aeronautics and Space Administration supercritical (phase 2)-0714 (NASA SC(2)-0714) airfoil. It is demonstrated that with the corrected Mach number and angle of attack, the estimated CFD results are in better agreement with the measured experimental data than traditional linear methods and the computational efficiency is improved compared to the original EnKF.

Compared to the shock-free condition, the weak shock impingement stabilizes the flexible panel, while the strong shock impingement leads to the early onset of panel flutter with a significant increase in flutter amplitude and frequency. However, the reason for this change by shock impingement remains unclear. The current research examines the mechanism of this change by an in-house code where the von Kármán’s large deflection plate theory is coupled with two-dimensional Euler equations. Compared to the shock-free condition, the oblique shock impingement leads to the change of local dynamic pressure on the panel as well as the static pressure differential across the panel. The analysis on the influence of these changes indicates that, on the one hand, the average dynamic pressure on the panel becomes larger than the shock-free condition, accelerating the onset of panel flutter. On the other hand, the change of the static pressure differential across the panel alters the coupling characteristic between different natural frequencies (modes) of the panel structure. The dynamic response of panel flutter under shock impingements is dominated by the coupling between the second and third modes instead of the first two modes for panel flutter under the shock-free condition. The combined effect of these two changes leads to the change of flutter characteristics of the panel under shock impingement. These findings provide valuable insights into the mechanism of shock-induced panel flutter.

The traditional method for obtaining aerodynamic parameters of airfoils by solving Navier–Stokes equations is a time-consuming computing task. In this article, a novel data-driven deep attention network (DAN) is proposed for reconstruction of incompressible steady flow fields around airfoils. To extract the geometric representation of the input airfoils, the grayscale image of the airfoil is divided into a set of patches, and these are input into the transformer encoder by embedding. The geometric parameters extracted from the transformer encoder, together with the Reynolds number, angle of attack, flow field coordinates, and distance field, are input into a multilayer perceptron to predict the flow field of the airfoil. Through analysis of a large number of qualitative and quantitative experimental results, it is concluded that the proposed DAN can improve the interpretability of the model while obtaining good prediction accuracy and generalization capability for different airfoils and flow-field states.