In this study, a one-way coupled point-particle mean flow simulation was utilized to investigate the effects and underlying mechanisms of near-wall particle clustering in channel flow. The mean velocity field was obtained from the turbulence database. Initially, the particles were uniformly distributed in the channel with their velocities following a Gaussian distribution. The results show that: 1） Inertial particles with a uniform initial distribution ultimately achieve equilibrium in the vertical direction and exhibit slight clustering near the wall. As the initial velocity fluctuations of particles gradually increase, the clustering effect becomes more pronounced. However, once the root mean square of the particles’ initial velocity fluctuations exceeds a certain threshold, further enhancement in near-wall particle clustering is no longer observed, and the particle concentration profile remains essentially unchanged. 2） As the particle Stokes number increases, the near-wall particle clustering also becomes more pronounced. 3） By modifying the particle-wall collision conditions, it was observed that the near-wall particle clustering increases substantially as the coefficient of restitution decreases. Finally, the particle stochastic drag model and the continuous random walk （CRW） model of fluid velocity fluctuation at the particle position were used to consider the effect of turbulent fluctuations on the near-wall particle clustering. It was found that when considering the turbulent fluctuations, the particle clustering in the near-wall region is significantly stronger than that when only the mean flow is considered, indicating that the turbulent fluctuation plays a dominant role in the near-wall particle clustering.
Aiming at the requirement of rapid assessment of aerodynamic thermal environment in the early design of hypersonic vehicle, a rapid prediction method of aerodynamic thermal integrating engineering algorithms with one-dimensional system simulation was designed. The Prandtl boundary layer theory was used to divide the inviscid/viscous flow field, the modified Newton theory was used to solve the surface pressure, the reference enthalpy method was used to determine the qualitative temperature of the boundary layer, and the equivalent cone method was introduced to deal with the angle of attack effect, and a complete aerodynamic heat calculation system was constructed. The calculation system was embedded into the Flowmaster platform controller element to realize the lightweight system simulation driven by the engineering semi-empirical models. The blunt cone in the NASA experimental report was used to verify the credibility of the fast prediction method. At the same time, the X-37B-like aircraft was used as an engineering case to calculate and analyze under cruise and ballistic conditions. The results show that the proposed rapid prediction method is efficient and reliable, and the calculation accuracy meets the engineering requirements in the initial design. It can effectively support the thermal environment assessment in the initial design of the aircraft and provide an efficient analysis tool for the thermal protection system design.
To address the need for enhancing wind tunnel testing efficiency, this study proposed an auxiliary analysis method for wind tunnel test data, based on matching and recommending similar aerodynamic patterns. A hybrid feature design was achieved by integrating expert-defined features with deep latent features extracted through a hybrid architecture combining a variational autoencoder and a temporal network. Leveraging this approach, unsupervised self-organizing mapping was employed to mine aerodynamic patterns from historical wind tunnel big data. An aerodynamic knowledge graph was then constructed based on similar aerodynamic pattern modes, establishing topological associations within wind tunnel data. The method intelligently matches and recommends historical data with analogous aerodynamic patterns, thereby facilitating auxiliary analysis of wind tunnel test data for novel aircraft configurations. Experimental results demonstrate significant pattern transferability in cross-configuration and cross-condition aerodynamic analysis. For the test set, the median Pearson correlation coefficient across six aerodynamic components exceeds 82%, validating the method’s effectiveness. This work forms a closed-loop technical framework of “data governance, pattern mining, knowledge evolution, and engineering empowerment”, providing a viable technical pathway for intelligent aerodynamics research.
This study focused on the critical technologies for wide-speed-range high-speed vehicle design, particularly the integration of aerodynamic configuration and adaptive morphing flight strategies. Firstly, an integrated design of a waverider forebody and an inward-turning adjustable inlet was developed, employing a longitudinal-inlet-to-lateral-exhaust configuration. The scramjet flow path was modified from a single central duct to dual bilateral ducts downstream of the throat, reducing turbine channel offset and nozzle height while maintaining a high total pressure recovery coefficient. Next, by combining the vehicle's center-of-mass motion equations with energy-maneuverability analysis, a rapid estimation method for fuel weight consumption during accelerated climb was established. Finally, a variable-geometry flight configuration was designed, and its performance was compared with that of a fixed-configuration approach in terms of climb time and fuel consumption. The results demonstrate that the variable-geometry configuration reduces climb duration by 21.4% and saves 5.4% in fuel weight.
To overcome the limitations of existing hot jet generation approaches in hot jet wind tunnel testing, this paper presents a wind tunnel experimental methodemploying methane/air high-temperature combustion gas for hot jet simulation. A design methodology for a methane/air high-temperature gas generator compatible with wind tunnel testing was established. The gas generator realized sustained stable combustion over extended durations, with the combustion chamber pressure, temperature, and nozzle exit species composition parameters systematically calibrated. Pressure measurement tests on the surface of a Ma = 3 flat-plate hot jet model were completed. The test results exhibited good repeatability, and the numerical results were in good agreement with the experimental data. Force measurement tests on a Ma = 5 axisymmetric divert-control jet model were also carried out, yielding stable and highly repeatable force and moment data. It is demonstrated that the wind tunnel test method using methane/air combustion to simulate hot jets is stable and reliable, providing a new solution for hot jet wind tunnel tests of flight vehicles.
Traditional circulation control method can enhance the lift and increase the lift-to-drag ratio of airfoils effectively. However, its failure at high angles of attack and the additional drag caused by airfoil geometry modification in circulation-control-off situations suppress the method in further augmenting lift and lift-to-drag ratio of airfoils. Based on the understanding of the mechanism of circulation control, a leading-edge circulation control （LECC） method was proposed and an LECC airfoil was conceived and designed. Numerical simulation was carried out to study the effects of jet momentum coefficient and angle of attack on the aerodynamic characteristics of the LECC airfoil. The study reveals the following items. Firstly, the geometry modification of LECC can lead to lift increase and drag reduction. Secondly, with the increase of the jet momentum coefficient, the lift of the airfoil increases, and there exists a critical value of jet momentum coefficient beyond which the control effect of LECC is weakened slightly. Thirdly, as the jet momentum coefficient increases, the drag coefficient decreases almost linearly and also experiences a critical value of the jet momentum coefficient at which the drag coefficient changes from positive to negative, in other words, a net thrust is generated on the airfoil. Fourthly, LECC can delay the onset of airfoil stall, widen the range of available angles of attack a great deal, increase the lift curve slope, and achieve excellent lift characteristics. Lastly, LECC can generate negative drag over a wide ranges of angle of attack.
Guo et al.[1] proposed using implicit treatment with asynchronous iterations in Reynolds-averaged Navier-Stokes （RANS） simulations for training cases[2]. They recommended selecting from a series of intermediate iteration steps in the generated flow fields to obtain multiple step mappings between the flow fields at steps n and n + 1, which enhances data for model training. Based on Guo et al.’s approach, ITAI-RANS was employed for training cases to produce a series of converged fields rather than intermediate fields. A selection algorithm was applied to these converged fields, expanding the training set and improving data augmentation. Using two-dimensional periodic hill flows with various geometries and Reynolds numbers as training-prediction cases, multiple tests were conducted by varying the selection of these cases. The results demonstrate that the model trained with augmented data significantly outperforms the original RANS model, the original tensor basis neural network （TBNN） model, and Guo et al.’s model[1].
Free-stream temperature is an important physical parameter in high-speed wind tunnel, which may have an impact on boundary layer transition. This study investigated the effect of free-stream temperature on stationary Görtler vortices in hypersonic flow using direct numerical simulations （DNS), two-dimensional spatial eigenvalue analysis （BiGlobal) and plane-marching parabolized stability equations （PSE3D). In the simulations, Görtler vortices were excited with spanwise wavelengths of 3 mm. The flow visualization of the DNS data reveals prominent sinuous perturbations in the transition process. When the free-stream temperature increases, the location where Görtler streaks break down moves upstream. With the increase in free-stream temperature within an appropriate range, the swing of Görtler streaks occurs earlier. Subsequent stability analyses based on BiGlobal and PSE3D confirm that sinuous secondary instability modes are the most unstable, responsible for the breakdown of the Görtler vortices, and the growth rate of the dominant sinuous mode increases significantly with an increase in the free-stream temperature. Further analysis indicates that increasing the free-stream temperature remarkably promotes the growth of the Görtler vortices, thus increasing the spanwise gradient of the streamwise velocity, which results in the increase of energy production of spanwise velocity shear. Therefore, the sinuous secondary instability is destabilized, leading to the promotion of boundary layer transition. In summary, the free-stream temperature has large effect on Görtler vortices.
To address the issues in traditional double-swept waverider generation methods, where the primary sweep angle is difficult to control precisely and a transitional segment exists between the first and second sweep angles, this paper proposed a transition-free double-swept waverider generation method. Based on the conventional double-swept waverider generation method that incorporates a transitional segment, the proposed approach introduced geometric constraints to eliminate this segment, thereby achieving a seamless connection between the two swept leading edges. This design strategy ensures complete controllability of the waverider’s leading edge during the design phase. Furthermore, a geometric modification approach based on cubic spline interpolation was employed to resolve the geometric singularities on the lower surface that arise from removing the transitional segment. Finally, the designed double-swept waverider was analyzed and validated. The modification analysis indicates that the proposed method not only ensures first-derivative continuity on the lower surface but also maintains sufficiently small geometric errors. The computational fluid dynamics results indicate that the proposed waverider exhibits pronounced waveriding characteristics, which align well with the design expectations.
In response to the demand for high-precision plasma flow field simulation, a three-temperature model for electronic energy nonequilibrium was developed. Numerical simulation of various high-speed flow fields was carried out, and the formation mechanism of plasma flow around the aircraft was analyzed. Numerical simulations were performed for four operating conditions of the RAM-C Ⅱ flight test, and it was verified that the prediction performance of the eleven-species three-temperature model for the flow field plasma was better than that of the seven-species two-temperature model and the eleven-species two-temperature model. Numerical simulations were performed for the flow field under various flight conditions of the ball-head model. The research shows that the electronic temperature is an important variable to control the process of ionization reaction. Gas-phase electronic excitation is closely related to the formation of surrounding plasma. In a certain temperature range, the reduction of the characteristic time of electronic energy relaxation and the rise of electronic temperature create positive feedback. Consequently, the plasma distribution in the flow field will undergo abrupt changes as flow velocity increases.
In blast impact research, accurately measuring the pressure-time history is critical for damage assessment, yet conventional pressure measurement techniques struggle to fully characterize the spatiotemporal pressure changes on the surfaces of test models. The fast pressure-sensitive paint （PSP） technique offers a promising solution to this challenge. This paper developed a pressure field measurement technique for blast impact research based on fast PSP, established a corresponding optical measurement system, and conducted validation experiments in a shock tube. Firstly, to address the time delay in the coating’s dynamic response and the time-averaging effect of camera exposure, a temporal deconvolution algorithm was proposed, which effectively corrected the signal lag of PSP measurements during transient pressure surges in the explosion shock process. Secondly, to address the variation in the gas composition of explosion products, a frequency filtering method combined with proper orthogonal decomposition （POD） was developed, which effectively decoupled the effects of gas composition changes from pressure variations in the PSP signal, thereby mitigating oxygen-concentration interference in pressure measurements. Experimental results indicate that the fast PSP technology can achieve high-precision full-field dynamic pressure measurements under blast conditions, promising to provide robust technical support for experimental research in related fields.
High-speed vehicles often suffer severe aerodynamic heating during cruise, leading to surface ablation phenomena. The ablation effect significantly influences the perturbation evolution and laminar-turbulent transition in boundary layers. This paper first reviewed the local scattering theory that characterizes the ablation effect. To address the impact of distributed ablation morphology on transition, a numerical prediction method based on the harmonic linearized Navier-Stokes （HLNS） framework was introduced. Additionally, a wind tunnel experimental strategy was designed to measure the influence of ablation morphology on transition. Finally, the HLNS method was employed to predict the effect of ablation morphology on transition under wind tunnel experimental conditions, and the numerical results show good agreement with experimental data. This work provided a theoretical foundation and methodological support for establishing an engineering transition prediction method based on local scattering mechanisms.
Although high-order unstructured grid finite volume methods based on variational reconstruction offer advantages like high accuracy and computational efficiency, their iterative convergence robustness and speed significantly lag behind second-order methods when solving steady-state aerodynamic problems using implicit schemes. This paper presented an efficient iterative acceleration strategy: constructing an incomplete LU factorization （ILU） based on a low-order approximate Jacobian matrix, and defining the ILU（k） fill-in pattern utilizing graph distance. This ILU scheme can be employed as a standalone iterative solver or as a preconditioner for the GMRES method. Furthermore, the loop ordering of the ILU algorithm was optimized for the compressed row storage block sparse matrix format, effectively reducing matrix storage requirements. Computational practice demonstrates that the ILU preconditioner requires only modest additional storage when using fourth-order variational reconstruction. Comparative tests on classical two-dimensional steady-state cases reveal that the ILU method and its combination with GMRES （GMRES+ILU） generally accelerate convergence by a factor of 2 to 3, and can overcome convergence failures encountered with the LU-SGS iterative method on certain problems. Finally, the effectiveness of the proposed method was validated in large-scale steady-state computations for a three-dimensional wing-body configuration, demonstrating rapid convergence and satisfactory accuracy. Overall, the iterative strategy presented in this paper significantly enhances both the convergence robustness and computational efficiency of unstructured grid variational reconstruction finite volume methods for solving steady-state problems.
The accurate prediction of mean flow profiles in compressible wall-bounded turbulence is crucial for engineering applications. However, high-fidelity simulations remain computationally prohibitive, whereas traditional Reynolds-averaged Navier-Stokes （RANS） methods often suffer from limited accuracy. A novel modular iterative framework based on inverse velocity transformation for accurate and efficient prediction of mean profiles in compressible turbulent channel flows （CTCFs） was presented. The key innovation lies in replacing the conventional algebraic mean temperature-velocity （TV） relations, which are based on Reynolds analogy, with a recently proposed integral relation, thereby removing the strong dependence on empirical scalings for centerline mean temperature. This advancement not only avoids the ill-posedness arising from repeatedly imposing symmetry constraints in prior approaches but also significantly enhances the generality and robustness of the framework. Comprehensive numerical evaluations demonstrate that the proposed method performs excellently across a wide range of Mach and Reynolds numbers. It predicted mean velocity and temperature profiles with maximum relative errors below 3% and 2%, respectively, when compared with benchmark direct numerical simulation （DNS） data. Meanwhile, it reliably predicted wall skin friction and wall heat flux, achieving absolute relative errors below 2% and 1% in moderate- to high-Reynolds-number regimes. Compared with existing approaches, the present framework preserves high predictive accuracy while markedly reducing the reliance on prior empirical knowledge during the iteration, offering a powerful tool for rapid prediction and engineering-oriented modeling of compressible wall-bounded turbulence.
Plasma actuators have received extensive attention for active flow control due to their advantages of quick response and operational flexibility.The aerodynamic,thermal and chemical effects of the plasma play a dominant role in flow control.Closed-loop feedback control is expected to further improve plasma control efficiency and facilitate its practical application.A numerical study was carried out to evaluate the energy efficiency of the thermal effects of arc plasma for shock wave control,using experimental data from a Mach 2.5 flow over a 7° compression ramp.Quantitative regulation of shock wave strength was achieved using proportional-integral closed-loop feedback control.The effects of arc length and power on the energy efficiency of arc plasma were comparatively analyzed.The results show that 60%～80%of the total arc energy is utilized for shock wave control through thermal effects,and this percentage increases with arc length.
Detonation instability is an unresolved fundamental problem of detonation theory. In this study, the convective flux analysis methodology was put forth for the first time to explain the underlying physics of unstable detonation. One-dimensional hydrogen/air detonation was numerically simulated with Euler equations and one-step overall detonation model. The convective flux was extracted at each time step. The numerical results show that the flux of compression wave and heat release at the detonation front has the Λ shape, while the flux of rarefaction wave has the negative Λ shape. During the unstable detonation propagation process, the convective flux varies by more than an order of magnitude. The combination of these fluxes leads to different detonation propagation modes, including CJ detonation, weakly unstable detonation, periodic detonation, pulsating detonation and overdriven detonation.
To investigate the regulatory mechanism of Reynolds number on soliton-like coherent structures （SCS） in a circular free jet, a series of direct numerical simulations were conducted over a Re range of 2000 to 6600. The results show that while variations in Re did not alter the overall time-averaged structure of the jet, they significantly affected its momentum transport: at lower Re, the centerline velocity decayed faster, but the radial spreading was slower. Under all conditions, SCS emerged within alternating high- and low-speed regions, with their core soliton-like structure remaining stable across different Re. Furthermore, the closer to the jet centerline, the earlier the SCS formed, with larger amplitude and greater quantity. As Re increased, the inception locations of SCS generally shifted upstream, and their number increased. Lagrangian material surface tracking and correlation analysis confirmed that SCS exhibit soliton-like properties and induce vortex generation in the flanking shear layers during their development, with earlier destabilization and breakdown occurring at higher Re. This study demonstrates that while Re did not alter the core soliton-like morphology of SCS, it systematically modulated key evolutionary parameters — such as their inception location, quantity, and stability — by adjusting the balance between inertial effects and viscous dissipation, thereby shaping the transition pathway of the free jet.
Transported probability density function(TPDF)methods are attractive for modeling turbulent combustion,as the nonlinear chemical reaction terms in the PDF transport equation can be treated exactly.In this paper,a TPDF method was applied to model the Burgers turbulence governed by the one-dimensional Burgers equation with a stochastic external force and random initial data.Firstly,an exact one-point velocity PDF transport equation was derived from the Burgers equation with a nonlinear stochastic external force,and then the conditional advection and diffusion terms were closed by using the respective models so as to obtain the modeled PDF transport equation.Secondly,three numerical methods were employed to solve the modeled PDF transport equation:the mesh-based Lagrangian particle Monte Carlo(MC)method,the original quadrature method of moments(QMOM),and the linear QMOM(LQMOM)developed in this work.The numerical tests show that the present PDF method is suitable for studying statistic properties of Burgers turbulence and the LQMOM and QMOM can yield results comparable to the DNS solution with the LQMOM being more efficient than the QMOM.However,the MC method produces unsatisfactory results for cases with stochastic forcing terms,which is probably due to numerical perturbations introduced by frequently adding and deleting particles in order to keep the particle number per grid cell unchanged.
In order to achieve long-duration flight in the upper atmosphere,inlet performance and aerodynamic drag were analyzed by employing the direct simulation Monte Carlo method.Effects of incomplete accommodation in the normal and tangential directions for the gas-surface interaction(GSI)were explored,and the feasibility of long-duration flight was evaluated from the perspective of thrust-drag balance.Results show that,in comparison with the complete-accommodation case,incomplete accommodation in the normal or tangential direction can raise the compression ratio,and incomplete accommodation in the normal direction makes the high-pressure region move upstream while incomplete accommodation in the tangential direction causes the high-pressure region to move downstream.Internal flows inside the inlet and external flows past the solar panel are mainly responsible for the aerodynamic drag of the spacecraft.Incomplete accommodation in the normal direction increases the drag while incomplete accommodation in the tangential direction can reduce the drag.Under some reasonable assumptions,the spacecraft considered can achieve long-duration flight in the upper atmosphere with incomplete accommodation in the tangential direction,provided that the thrust-to-power ratio reaches 18～20 mN/kW and the ionization efficiency attains 40%～50%.
Dredging thermal protection,as a semi-active thermal protection method,is characterized by its simple structure and the absence of the need to carry additional coolants,making it suitable for the leading edge thermal protection of hypersonic vehicles.This paper established a transient start-up model applicable to wedge-shaped dredging structures and conducted cold-start heat transfer calculations for dredging components under various flight conditions,structural parameters,and materials.The start-up time of the dredging components and the spatiotemporal distribution of physical parameters under different scenarios were investigated.Furthermore,the mechanism underlying the commonly observed thermal shock phenomenon during testing was explored.The results indicate that the dredging structure encounters neither freezing start limits nor sonic limits.However,under high heat loads,the pressure drop in the vapor channel is significant.During the incomplete start-up phase,the maximum pressure drop can be up to 40%higher than the steady-state value,potentially inducing the capillary limit.This is identified as the core mechanism leading to thermal shock burnout of the dredging component.Reducing the thickness of the non-dredging region can decrease the temperature difference across the dredging component,while increasing the height of the vapor channel reduces the vapor pressure drop,thereby avoiding the capillary limit.Dredging components made of nickel-based,niobium-based,and molybdenum-based alloys can withstand incoming flows at Mach 7,Mach 9,and Mach 11,respectively,at an altitude of 30 km without ablation.
