Abstract Opposing jet has been widely investigated as an effective aerodynamic drag reduction method in recent years. To study the impact of freestream angle of attack on the drag reduction effect of opposing jet, experimental and numerical approaches are employed to analyze the effects of jet pressure ratio, angle of attack, and jet deflection angle on the flow field structure and the underlying mechanism to drag reduction of the opposing jet in a Mach 6 freestream. The results indicate that, within the range of this study, drag reduction increases with higher jet pressure ratios, reaching a maximum reduction of 32.67% at a pressure ratio of 0.60. The drag reduction decreases as the angle of attack increases, with a 5.44% decline observed at 6° compared to 0°. This deterioration stems from the strong compression of the incoming flow on the windward side, which enhances the reattachment shock and increases the peak pressure dramatically. Additionally, at higher angles of attack, the adoption of an appropriate jet deflection angle can significantly enhance the effectiveness of drag reduction, achieving an improvement of up to 2.65%. Three typical flow fields are systematically summarized and analyzed, which have rarely been compared together in previous studies, namely the completely axisymmetric case, freestream with angle of attack but no jet deflection, and freestream with angle of attack while the jet undergoes deflection, respectively. Among all conditions, the position and distance away from the body of the intersection point of the bow shock and reattachment shock wave play a crucial role in drag reduction effect, shifting this point downstream and away from the surface significantly enhances the drag reduction effect.
This paper presents the development and experimental validation of a novel cable-driven parallel suspension mechanism (CDPSM) for investigating unstable, high-angle-of-attack aerodynamics in low-speed wind tunnel. Designed to overcome the limitations of conventional testing in simulating post-stall flight dynamics, the six-degree-of-freedom CDPSM enables precise execution of highly nonlinear motions, specifically amplitude-frequency sweep maneuvers. The system achieves this through a geometric mapping between aircraft pose commands and cable lengths, governed by a terminal sliding mode controller that integrates an extended state observer for high-fidelity motion control. A series of unsteady aerodynamic tests, incorporating coupled frequency-amplitude variations, were conducted with forces measured via an internal six-component balance. The results confirm the CDPSM’s capability to accurately simulate prescribed unstable motions and capture the ensuing aerodynamic responses. Critically, the measured forces revealed significantly more complex hysteresis loops and variation patterns than those obtainable from traditional sinusoidal motion tests, highlighting the unique capabilities of the proposed system for investigating advanced post-stall flight dynamics.
Laminar flow separation caused by shock wave-boundary layer interaction is a critical issue for supersonic vehicles, which significantly affects their aerodynamic performance. In-depth studies of the flow separation characteristics and accurate predictions of the separation zone dimensions are essential for optimizing flow control strategies and determining the location of the peak heat flux. This study focuses on the incident shock wave flat plate model and employs direct numerical simulation to investigate the separation flow caused by an incident shock wave-laminar boundary layer under varying shock intensities and wall temperatures. The research results show that with increasing shock intensity, a secondary separation zone appears within the separation region, and the emergence of this secondary zone increases the normal pressure gradient of the primary separation zone. The size of the secondary separation zone first increases, then decreases, and ultimately vanishes as the wall temperature increases. Additionally, based on the Katzer prediction formula, the separation bubble dimension prediction formula was optimized by considering the effects of wall temperature. Experimental and numerical validations demonstrate that the optimized formula can predict the scale of the primary separation zone in the Mach number range of 1.6–15 and wall temperature ratios of 0.2–1. The formula exhibits excellent predictive accuracy under both adiabatic and isothermal wall conditions and remains reliable in accurately predicting the primary separation bubble dimension even in the presence of secondary separation.
This work numerically investigates the interaction of shock waves with sinusoidal fuel-air interfaces using an enhanced detonationFoam solver, which employs a detailed nine-species, twenty-one-step H _2 -O _2 chemical mechanism. Effects of shock strength, interface amplitude, and interface thickness on the evolution of shocked interface are examined. Two distinct ignition modes are identified: weak ignition at Ma = 2.0 where transverse waves create localized ignition sites at spike and bubble tips, and strong ignition at Ma = 2.7 where intense shock compression triggers rapid ignition across the interface. Combustion generally exerts negligible influence on interface evolution in both regimes, with dimensionless amplitude evolution collapsing onto a single curve regardless of reaction activity. Large-amplitude interfaces exhibit more complex ignition behavior with multiple triple points and Mach stems inducing multi-site ignition. A modified VMG model (Vandenboomgaerde, et al. Phys Rev E 58(2):1874–1882, 1998) incorporating a reduction factor successfully predicts the linear growth rate across different amplitudes, confirming that hydrodynamic effects dominate instability evolution even with enhanced combustion complexity. Interface diffusion-layer thickness introduces regime-dependent effects on ignition. In the weak-ignition regime, a non-monotonic behavior emerges: moderate thickening advances ignition by reducing scalar dissipation rate, while excessive thickening delays ignition by attenuating shock-induced compression. In contrast, under strong-ignition conditions, interface thickening consistently delays ignition as it weakens the abrupt thermal compression. For thick interfaces at high Mach numbers, combustion-enhanced viscosity plays a significant role, and the instability evolution is governed by the combined effects of baroclinic torque and viscous vorticity production. The heat release rate peaks briefly after ignition and rapidly decays as mixing becomes insufficient, explaining the limited overall influence of combustion on instability growth in non-premixed configurations.
In this work, an improved adaptive approach for large-eddy simulations (LES) is introduced based on a dynamic adaptive mesh refinement (AMR) algorithm and a corresponding adaptive subgrid model embedded in such AMR process. The objective of the present work is to develop a turbulent-flow simulation method that dynamically allocates meshes of varying resolutions, which thereby significantly reduces the required computational resources of traditional LES methods. To achieve this, the local velocity fluctuation intensity is proposed as a criterion to determine whether the grid resolution needs to be changed. With this, the grid distribution is adjusted according to the motions of the local flow structures, enabling a more precise and cost-effective way to capture the multiscale evolution in turbulent flow fields. Meanwhile, with the updated position and velocity information, an adaptive subgrid model is constructed within the AMR framework to determine the behaviors of the unresolved structures smaller than the finest resolution of the AMR mesh. As a result, the developed dynamic remeshing and the corresponding adaptive subgrid model enhance the performance and efficiency of LES models. Through a number of validation examples, the accuracy, capability, and efficiency of the proposed adaptive LES methodology have been demonstrated.
