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.
Abstract This paper proposes a modular parametric framework for axisymmetric high-speed wind tunnel nozzle design, which innovatively introduces the Terminal Characteristic Line (TCL) theory to isolate upstream and downstream flow fields. It divides the wind tunnel nozzle into several modules for design. In the upstream initial expansion section module, this work improves the Method of Characteristics (MOC), which abandons empirical axial Mach number distributions and supports user-controlled length. In the downstream damping section module, it references the minimum-length nozzle (MLN) design method, serving as a case for modular design. The framework is developed in Python code and open-sourced on GitHub, which can design and output inviscid contours and boundary layer correction results based on user needs. Validation using Computational Fluid Dynamics (CFD), which covers multiple design cases from the framework, shows excellent outlet flow uniformity and demonstrates length-to-diameter ratio control capability.
Abstract With their growing prevalence, double-deck bridges can be classified by the relative width of upper and lower decks into three girder configurations: equal-width (Model A), inverted-trapezoidal (Model B), and lower-deck-cantilevered (Model C). This study combines wind tunnel tests and numerical simulations to quantify the influence of girder configuration, with a focus on measurements of global aerodynamics and load partitioning among subcomponents (upper/lower decks, windward/leeward trusses). Measurements indicate that Model B achieves the lowest drag and smallest absolute lift coefficients under most angles of attack (AoA), yet exhibits the highest moment coefficients. Model A yields peak drag coefficients at AoA = −6° to 2° with intermediate lift and moment values. Model C exhibits drag coefficients comparable to Model A at large AoAs and to Model B at small AoAs, while demonstrating the most pronounced AoA-dependence. Regarding load partitioning among subcomponents, the windward/leeward trusses carry over 50% of mean drag in all girders, with the magnitude sequence of Model C > Model A > Model B. For mean lift and moment, Models A and C distribute generally evenly between upper and lower decks, whereas Model B concentrates them on the upper deck. Furthermore, Dynamic Mode Decomposition (DMD) reveals three dominant vortex modes associated with lower-deck shedding, inter-truss vortex dynamics, and trailing-edge separation. Models A and C exhibit strong inter-deck interference and coherent vortices in the truss gap, resulting in higher drag and greater moment sensitivity, while Model B shows weaker coupling but intensified local separation near the upper deck. These flow characteristics, together with global force and load partitioning results, offer direct guidance for wind-resistant design: Model B is recommended for high-wind regions owing to its lowest drag; however, the concentration of aerodynamic loads on the upper deck necessitates structural reinforcement. Model A supports standardized design through balanced load partitioning between decks, while Model C exhibits relatively high drag and strong sensitivity to AoA, warranting careful aerodynamic mitigation in practical applications.
Abstract The influence of streamwise-traveling waves imposed at the walls on drag variation in compressible channel flow is investigated using direct numerical simulations. The variations in drag are mapped in the frequency-wavenumber $$({\omega }^{+},{k}^{+})$$ ( ω + , k + ) space for three bulk Mach numbers, $$M{a}_\mathrm{b}=0.3, 0.8$$ M a b = 0.3 , 0.8 and 1.5, at a fixed bulk Reynolds number $$R{e}_\mathrm{b}=3000$$ R e b = 3000 . The traveling waves significantly modify friction drag, showing similar trends across all $$M{a}_\mathrm{b}$$ M a b . Depending on the phase speed ( $${c}^{+}$$ c + ), the waves can induce either drag reduction ( $$DR$$ DR ) or drag increase ( $$DI$$ DI ). In the region of low $${c}^{+}$$ c + , the $$DR$$ DR is mostly around $$40\text{\%}$$ 40 \% , especially when $$M{a}_\mathrm{b}=1.5$$ M a b = 1.5 , where it reaches $$66.6\text{\%}$$ 66.6 \% . At moderate $${c}^{+}$$ c + , a substantial $$DI$$ DI occurs. At high $${c}^{+}$$ c + , $$DR$$ DR reappears. The flow statistics and visualizations reveal that the $$DR$$ DR mechanism resembles that of an oscillating wall, promoting larger coherent structures and suppressing ejection and sweep events. Particularly, when $${c}^{+}\approx 10$$ c + ≈ 10 , phase-locking is observed in both $$DR$$ DR and $$DI$$ DI cases, featuring streamwise streaks cyclically tilted by the wall forcing, which in turn influences the turbulence statistics. Additionally, low-frequency oscillations generate a periodic wave-like pattern in the streamwise direction, significantly modulating the streak structures, whereas high-frequency oscillations do not produce this effect.
Abstract Structural evolution of the porous surface caused by aerodynamic heating and ablation is crucial for evaluating the ablative performance of light-weight thermal protection materials for high speed vehicles. However, the numerical coupling prediction method of the flow, heat transfer, thermochemical reactions and the pore-scale structure evolution still remains limited. In this study, a hybrid micro-continuum scale approach is employed to predict this complicated coupling phenomenon for three-dimensional (3D) porous media. The results indicate that in the simulation of the oxidation ablation process of a typical carbon fibrous porous medium, the proposed method can successfully reproduce the needle-like morphology observed experimentally, revealing the gradual recession and surface roughness formation caused by surface heterogeneity. Moreover, under high-temperature flow conditions where the reaction rate is much larger than the mass-transfer rate due to convection and diffusion, the incoming oxygen gas can be consumed rapidly at the flow/porous media interface. This fast consumption is found to be dominant, resulting in a low oxygen supply condition at the gas–solid interface and a non-linear ablation recession rate. The recession rate is found to reach its maximum value of 13.05 μm/s, followed by a gradual decline to 8.35 μm/s. The proposed hybrid micro-continuum multiscale modeling approach can potentially offer valuable pore-scale insights into the ablation behavior of porous media, thereby enhancing the predictive accuracy for thermal protection systems.
Abstract This study presents an optimization framework for improving the aerodynamic performance of a square cylinder by optimizing the placement of a pair of dielectric barrier discharge (DBD) plasma actuators using Kriging-enhanced Genetic Algorithm (GA). High-fidelity Navier-Stokes simulations at Reynolds numbers ranging from 1 to 300 provide baseline insights into flow separation, vortex shedding, and aerodynamic force fluctuations. Plasma actuators are modeled as body forces, with configurations explored through Latin Hypercube Sampling (LHS) within a Design of Experiments (DoE) framework. The optimization process is refined by using adaptive infill sampling, guided by Expected Improvement (EI), to efficiently identify optimal actuator placements. The comparative analysis of predicted and actual force coefficients validates the effectiveness of the Kriging model, GA, and infill sampling for optimizing the location of plasma actuators. Results highlight that ( $$y/h = \pm 1$$ y / h = ± 1 ) is the most optimized location for steady plasma actuation, where it achieved 55.73% reduction in drag and 98.08% significant reduction in lift oscillations, relative to the baseline condition ( $$C_{d,\text {avg}} = 1.44$$ C d , avg = 1.44 , $$C_{l,\text {rms}} = 0.453$$ C l , rms = 0.453 ). Comprehensive flow physics analysis, including flow streamlines, velocity profiles, power spectral density, and turbulence intensity, combined with comparisons to other potential actuator locations ( $$y/h = \pm 0.8, \pm 0.7, \pm 0.6$$ y / h = ± 0.8 , ± 0.7 , ± 0.6 ), further confirms the superiority of steady plasma actuation at the suggested optimized location. These findings underscore the robustness of Kriging-assisted GA optimization in plasma-based flow control, offering an energy-efficient solution for drag reduction and vortex suppression in aerospace, automotive, and civil engineering applications.
This study introduces a machine learning (ML) framework for efficient aero-structural characterization of wing planform shapes, addressing the computational challenges posed by traditional CFD and FEA methods. The goal is to develop an ML-based aero-structural optimization framework that replaces time intensive computational tasks with faster yet reliable approach. Leveraging advanced parameterization, data normalization, and reduced-order modeling (ROM), two regressor chain-based surrogate models are developed to predict aerodynamic and structural responses from identical shape parameters. For CFD buffet predictions, a Proper Orthogonal Decomposition-based surrogate model (POD-ML) is employed, while for stress predictions, the data is first clustered, and then POD is applied within each cluster, forming the K-means ML model. Using the ONERA M6 wing as a baseline, the surrogate models' performance is evaluated across various shape parameters like sweep, dihedral, and twists. The models effectively predict aerodynamic responses, such as pressure coefficients (Cp), with a maximum Mean Squared Error (MSE) of 0.05, requiring only 1,000 training samples — significantly fewer than conventional neural network models. The models also handle complex structural predictions with good accuracy, particularly for stress fields, achieving a maximum MSE of 0.10. Comparisons reveal that K-means ML outperforms POD-ML in stress field prediction, though it performs slightly worse for aerodynamic responses. Together, these methods provide substantial computational savings, facilitating rapid design iterations and advancing the optimization of wing architectures.
Abstract An overview is given of investigations on a 70° slender delta wing applying pressure-sensitive paint (PSP) and particle image velocimetry (PIV) in parallel. Experiments are conducted within the range of Reynolds number from 0.6 $$\times$$ × 106 to 1.2 $$\times$$ × 106 and angle of attack from 15° to 30°. The focus is on systematically studying the correlation between leading-edge vortex (LEV) and surface pressure distribution under different conditions. It is found that the suction peak does not exactly coincide with the vortex center. The shift of the suction peak is more pronounced at higher angles of attack or Reynolds numbers, reflecting the effect of the asymmetry of the corresponding velocity gradient distribution on the pressure distribution. Additionally, under certain conditions, the reduction of streamwise variation may weaken the three-dimensionality of the LEVs, which has a significant impact on their pressure distribution.
In regard to the development of more efficient and silent wind turbines, the present study consists of experimentally assessing the potential aero-acoustic benefits offered by bio-inspired, serrated blades. To this end, a conventional wind turbine blade (NREL Phase VI) is modified such that its leading edge exhibits serrations of various designs (amplitude and wavelength). The resulting serrated blades are then characterized from both an aerodynamic and an acoustic perspective, which is achieved through dual aero-acoustic tests. Involving a small-scale, fully instrumented wind turbine rig, these aero-acoustic tests are performed within both a closed-vein aerodynamic wind tunnel and an anechoic chamber, thereby allowing for measurement of the aerodynamic performance and the noise signatures entailed by the blades’ serrations, depending on their designs. From an aerodynamic standpoint, the serrated blades exhibit superior performance to their baseline counterpart, which translate into significantly higher spinning rates and power coefficients. These benefits are robust, being relatively insensitive to the inflow conditions (yaw, pitch, upstream turbulence) and, to a lesser extent, the serration design considered. From an acoustic perspective, the serrated blades exhibit more diverse noise signatures, depending on their serration design, the rotational regime and, in some cases, the radiative direction considered. Whereas no clear trend can be drawn at lower spinning rates, the higher rotational regime reveals more consistent patterns, e.g., the blades with smaller (resp. larger) serrations systematically radiate less (resp. more) noise than their baseline counterpart. All in all, these findings indicate that leading-edge serrations may offer an efficient passive flow control solution to improve the aero-acoustic characteristics of wind turbine blades. They nevertheless call for extra caution when designing serrated blades, owing to their sensitivity towards their serration design.
Abstract Active flow control aircraft (AFCA) represent a significant development direction in next-generation aircraft design. Among them, rudderless flight control (RFC) and short takeoff and landing (STOL) are gaining widespread attentions and are gradually applied as disruptive technologies of next-generation aircraft, displaying marked strategic and engineering value. In 2020, DARPA launched “Control of Revolutionary Aircraft with Novel Effectors” (CRANE) project, where active flow control (AFC) was incorporated as a critical dimension of overall design with the goal of developing the next-generation AFCA. The full-scale demonstrator of CRANE was officially named X-65 in 2023. Two of the critical technologies being validated are RFC and STOL. This article reviews the developing process of typical AFCA adopting above technologies and provides an outlook on the futural development tendency. AFCA adopting STOL developed earlier, and comprehensive evaluations and flight tests have already been conducted on full-scale commercial and military aircraft, demonstrating a high level of technological maturity. In contrast, AFCA adopting RFC developed later and have so far undergone flight tests only on small-scale UAVs, needing further evaluation on full-scale aircraft. Future research hotspots and challenges of AFCA will contain AFC implementation scheme combining the bleed air system and synthetic jet actuators, full-envelope high-efficiency AFC technology, coordination control of multiple AFC technologies, flight dynamics modeling of AFCA, redundant high-reliability flight control, high-stealth heavy-payload wide-speed-range large-airspace flight platform, and aircraft overall design integrating AFC.
As the fundamental component of rotor blades, the airfoil dominates the aerodynamic performance of helicopter rotors. The aerodynamic shape optimization of rotor airfoils is thus an effective means of improving their aerodynamic performance. Given that the aerodynamic requirements for the airfoil vary across different states of flight, such as hovering, maneuvering, and forward flight, multi-objective optimization of its aerodynamic design is necessary. In this study, we propose a scheme that simultaneously considers the optimization of the tip and root airfoils to design rotor blades that deliver satisfactory aerodynamic performance. We combine a data-driven surrogate model with a multi-island genetic algorithm to efficiently and globally optimize the rotor airfoils by using the Helishape-7A rotor as the benchmark. We first used Latin hypercube sampling to establish a database of airfoils, and then applied deep neural networks to train surrogate models and obtain the aerodynamic coefficients of the tip and root airfoils. Following this, we optimized the aerodynamic shape of the tip and root airfoils using the multi-island genetic algorithm while considering multiple objectives, including the states of hovering, maneuvering, and forward flight of the helicopter. Finally, the baseline Helishape-7A rotor's aerodynamic performance was compared with that of the optimized rotor. The outcomes demonstrated that, in comparison to the baseline, the optimized rotor produced noticeably better aerodynamic performance.
Robust and accurate numerical methods for numerical simulations of compressible fluid flows featuring discontinuities around complex geometries are challenging and of increasing practical interest. In this paper, a third-order weighted compact nonlinear scheme (WCNS) for solving inviscid compressible flows on Cartesian grids and a third-order constrained weighted least squares immersed boundary (CWLS-IB) method for dealing with irregular boundaries are developed. The numerical fluxes at the cell edges are calculated by a nonlinear combination of an optimal high-order polynomial on a global stencil and two low-order polynomials on two sub-stencils. A new global smoothness indicator of the global stencil is introduced to improve the accuracy of the scheme at both the first- and second-order critical points. The third-order CWLS method is employed to extrapolate the values at ghost cells outside the physical domain and a robust WENO-type extrapolation is adopted to eliminate oscillations when shocks occur near boundaries by introducing a nonlinear weight that can switch automatically between the high-order and the low-order extrapolation. Several numerical experiments are provided to verify the performance of the present WCNS scheme and the CWLS-IB method in terms of the accuracy, robustness and spatial resolution.
A numerical analysis of high-speed rarefied gas flow over a cavity has been performed using the direct simulation Monte Carlo (DSMC) method. The simulation was carried out over a wide range of gas rarefaction, from free-molecular flow to nearly continuum flow, for various values of the angle of attack, the free-stream Mach number, and the geometric size of the cavity. Flow fields, streamlines, vertical 1D distributions of macroscopic quantities, and the dimensionless particle flux density to the cavity floor were calculated with high accuracy. It was found that the flow pattern in the simulated system strongly depends on all flow parameters. At a certain distance from the cavity, the gas flow passes through a detached shock, where the flow regime changes from supersonic to subsonic. In close proximity to the cavity inlet, a complex flow structure can form, such as a flow separation zone, including a recirculation zone and a reverse flow zone. Inside the cavity, a circulating motion of the gas is formed, the pattern of which is significantly influenced by the geometric size of the cavity.
Shock buffet is one of the common flow instability problems in aviation engineering, which can affect flight handling quality and limit flight envelope. In this study, we propose a physics-guided aerodynamic optimization design framework to improve the buffet onset in the transonic region. The buffet instability mechanism correlates strongly with both the extent of flow separation and the spatial evolution of shock waves in the underlying steady flow. Therefore, this optimization framework is based on the deep deterministic policy gradient (DDPG) algorithm, which focuses on optimizing the shock wave position and separation area from the stable flow, while comprehensively considering various aerodynamic constraints for airfoil optimization. The proposed method is applied to two airfoil optimization parameter spaces: 5
Organizing supersonic fuel films near the wall offers a promising solution to simultaneously reduce friction and thermal loads in the internal flow passages of hypersonic vehicles. This paper presents a thorough review of the supersonic fuel film cooling from a numerical perspective. First, the foundational concepts and features of supersonic fuel films are outlined. Second, several critical issues in successfully modeling their flow and combustion processes are discussed in detail. Progress in numerical studies pertinent to supersonic fuel films is then summarized, mainly focusing on slot films, hole films, and shock-film-combustion interactions. Existing studies have demonstrated the feasibility of improving thermal protection and friction reduction performance of supersonic fuel films by boundary layer combustion, and provided preliminary insights into the underlying mechanisms. Nonetheless, current research status remains insufficient to achieve a comprehensive understanding of the effects induced by boundary layer combustion on fuel films. Key knowledge gaps and future research priorities are highlighted in the conclusion.
The one-equation Spalart–Allmaras (SA) model is employed to close the Reynolds averaged Navier–Stokes equations (RANS equations), and research on turbulent numerical algorithm is carried out. Using the coordinate transformation based on the characteristic line, we can obtain RANS equations without the convection terms. The intermediate velocity and correct velocity terms are introduced to split the momentum equations. The Runge–Kutta method is used to discretize it in time, and the Taylor expansion along the characteristic line is used to overcome the mesh update difficulty due to the coordinate transformation. Finally, based on the Galerkin space discretization, a third-order characteristic-based-split finite element algorithm is proposed. Numerical simulation of turbulent flow past a square cylinder is carried out to verify the effectiveness of the proposed algorithm. The prosposed algorithm is applied to simulate turbulent flow past a single conductor and 4-bundled conductors with crescent-shaped ice, the aerodynamic coefficients for different ice thicknesses and wind attack angles are obtained, and the galloping stability of crescent-shaped ice-covered conductor is discussed.
