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The efficient utilization of propeller slipstream energy is important for improving the ultra-short takeoff and landing capability of Distributed Electric Propulsion (DEP) aircraft. This paper presents a quasi-three-dimensional (2.5D) high-lift wing design approach considering the three-dimensional (3D) effects of slipstream for DEP aircraft, aiming at maximizing the comprehensive lift enhancement benefit of the airframe-propulsion coupling unit. A high-precision and efficient momentum source method is adopted to simulate the slipstream effects, and the distributed propellers are replaced by a rectangular actuator disk to reduce the difficulty of grid generation and improve the grid quality. A detailed comparison of the 2.5D and 3D configurations based on the X-57 Mod Ⅳ is performed in terms of flow characteristics and computational cost to demonstrate the rationality of the above design approach. The optimization results of the high-lift wing of the X-57 Mod Ⅳ show that the aerodynamic performance of the landing configuration is significantly improved, for instance, the lift coefficient increases by 0.094 at the angle of attack of 7°, and 0.097 at the angle of attack of 14°. This novel approach achieves efficient and effective design of high-lift wings under the influence of distributed slipstream, which has the potential to improve the design level of DEP aircraft.

In the context of increasing dimensionality of design variables and the complexity of constraints, the efficacy of Surrogate-Based Optimization (SBO) is limited. The traditional linear and nonlinear dimensionality reduction algorithms are mainly to decompose the mathematical matrix composed of design variables or objective functions in various forms, the smoothness of the design space cannot be guaranteed in the process, and additional constraint functions need to be added in the optimization, which increases the calculation cost. This study presents a new parameterization method to improve both problems of SBO. The new parameterization is addressed by decoupling affine transformations (dilation, rotation, shearing, and translation) within the Grassmannian submanifold, which enables a separate representation of the physical information of the airfoil in a high-dimensional space. Building upon this, Principal Geodesic Analysis (PGA) is employed to achieve geometric control, compress the design space, reduce the number of design variables, reduce the dimensions of design variables and enhance predictive performance during the surrogate optimization process. For comparison, a dimensionality reduction space is defined using 95% of the energy, and RAE 2822 for transonic conditions are used as demonstrations. This method significantly enhances the optimization efficiency of the surrogate model while effectively enabling geometric constraints. In three-dimensional problems, it enables simultaneous design of planar shapes for various components of the aircraft and high-order perturbation deformations. Optimization was applied to the ONERA M6 wing, achieving a lift-drag ratio of 18.09, representing a 27.25% improvement compared to the baseline configuration. In comparison to conventional surrogate model optimization methods, which only achieved a 17.97% improvement, this approach demonstrates its superiority.

Cancer-associated fibroblasts (CAFs) are known to promote angiogenesis in oral squamous cell carcinoma (OSCC). However, the epigenetic mechanisms through which CAFs facilitate angiogenesis within the tumor microenvironment are still poorly characterized. Nicotinamide N’-methyltransferase (NNMT), a member of the N-methyltransferase family, was found to be a key molecule in the activation of CAFs. This study shows that NNMT in fibroblasts contributes to angiogenesis and tumor growth through an epigenetic reprogramming-ETS2-VEGFA signaling axis in OSCC. Single-cell RNA Sequencing (scRNA-seq) analysis suggests that NNMT is mainly highly expressed in fibroblasts of head and neck squamous cell carcinoma (HNSCC). Moreover, analysis of the TCGA database and multiple immunohistochemical staining of clinical samples also identified a positive correlation between NNMT and tumor angiogenesis. This research further employed an assembled organoid model and a fibroblast-endothelial cell co-culture model to authenticate the proangiogenic ability of NNMT. At the molecular level, high expression of NNMT in CAFs was found to promote ETS2 expression by regulating H3K27 methylation level through mediating methylation deposition. Furthermore, ETS2 was verified to be an activating transcription factor of VEGFAin this study. Collectively, our findings delineate an epigenetic molecular regulatory network of angiogenesis and provide a theoretical basis for exploring new targets and clinical strategy in OSCC.

The development of the Urban Air Mobility concept has proposed stringent requirements for the fixed-wing vehicle’s Short/Vertical Take-Off and Landing (S/VTOL) performance. With its lift-enhancing impact, the channel wing, can improve aircraft low-speed performance and Short Take-Off and Landing (STOL) capability. It is critical for the performance analysis and flow mechanism study of a channel wing that considers the influence of propeller slipstream to guide the design of S/VTOL aircraft. The law and mechanism for the effect of the enclosing angle of the arc wing, the phase angle of the propeller, the tip clearance, the rotational speed and the chordwise position of the propeller on the channel wing are explored utilizing the quasi-steady multi reference frame method. A channel wing with a larger enclosing angle has a better ability to enhance lift and reduce drag using propeller slipstream. The effect of propeller phase angle on channel wing aerodynamic forces is periodic and weak. Increasing propeller rotational speed is helpful to enhance lift and resist flow separation for channel wing. It can reduce drag for tractor configuration but increase drag for pusher configuration. However, the nose-down pitching moment of a channel wing will grow dramatically, making longitudinal trimming in aircraft layout design challenging.

It is a challenging work to design micro aerial vehicle with great aerodynamic performance because the tiny wingspan at low-Reynolds-number cannot provide lift efficiently. The aerodynamic configuration of a classic delta-wing paper airplane is investigated in the present work with numerical method to discover its potential for micro aerial vehicle designs. Furthermore, the effect of the ventral gap on the aerodynamic characteristics of the paper airplane is investigated herein. The stall angles of attack reach 37.5° and 40°, respectively, for ventral opened configuration and the closed one, and the maximum lift coefficient reaches 1.49 and 1.46. The ventral-opened configuration has negative pitching moment coefficient (−0.01431) even at 37.5° while the closed one has a positive coefficient (0.01402). The reason may be the gap leads to a strong back-flow vortex before the trailing edge in the ventral gap which produces a strong nose-down moment. Generally, the ventral gap improves lift and dramatically influences the longitudinal stability compared with the one without it.

To improve specific stiffness and strength of cylinder type shell, it is essential to make use of free space inside the cylinder to generate internal supported structure for reinforcement. In this article, an available method is introduced to design this complex structure based on the advantages of additive manufacturing, which include the subprocess of design domain generation, topology optimization, reconstruction, and analysis. Firstly, the design domain inside cylinder is created by removing swept volume from cylindrical container space. The initial structure is obtained by topology optimization and reconstructed by considering the constraints of manufacturing. Then, mechanical properties of the refined structure are checked by finite element method to give a final model. A cylindrical thin shell with slider is chosen as an example to show the detailed work process of the proposed design method, and the designed structure is generated by selective laser melting technology finally. It is proved that the proposed method is suitable for designing additive manufacturing cylinder shell with internal support. Furthermore, a traditional cylindrical thin shell is selected as a benchmark to be compared with the designed model by mechanical properties. Through the comparisons, it is concluded that internal supported cylindrical thin shell has greater bearing capacity and less weight cost than the classical cylindrical structure.

Natural laminar flow technology can significantly reduce aircraft aerodynamic drag and has excellent technical appeal for transport aircraft development with high aerodynamic efficiency. Accurately and efficiently predicting the laminar-to-turbulent transition and revealing the maintenance mechanism of laminar flow in a transport aircraft’s flight environment are significant for developing natural laminar flow wings. In this research, we carry out natural laminar flow flight experiments with different Reynolds numbers and angles of attack. The critical N-factor is calibrated as 9.0 using flight experimental data and linear stability theory from a statistical perspective, which makes sure that the relative error of transition location is within 5%. We then implement a simplified eNtransition prediction method with a similar accuracy compared with linear stability theory. We compute the sensitivity information for the simplified eNmethod with an adjoint-based method, using the automatic differentiation technique (ADjoint). The impact of Reynolds numbers and pressure distributions on TS waves is analyzed using the sensitivity information. Through the sensitivity analysis, we find that: favorable pressure gradients not only suppress the development of TS waves but also decrease their sensitivity to Reynolds numbers; there exist three special regions which are very sensitive to the pressure distribution, and the sensitivity decreases as the local favorable pressure gradient increases. The proposed sensitivity analysis method enables robust natural laminar flow wings design.

Since proposed, the self-similarity variables based genuinely multidimensional Riemann solver is attracting more attentions due to its high resolution in multidimensional complex flows. However, it needs numerous logical operations in supersonic cases, which limit the method’s applicability in engineering problems greatly. In order to overcome this defect, a hybrid multidimensional Riemann solver, called HMTHS (Hybrid of MulTv and multidimensional HLL scheme based on Self-similar structures), is proposed. It simulates the strongly interacting zone by adopting the MHLLES (Multidimensional Harten-Lax-van Leer-Eifeldt scheme based on Self-similar structures) scheme at subsonic speeds, which is with a high resolution by considering the second moment in the similarity variables. Also, it adopts the MULTV (Multidimensional Toro and Vasquez) scheme, which is with a high resolution in capturing discontinuities, to simulate the flux at supersonic speeds. Systematic numerical experiments, including both one-dimensional cases and two-dimensional cases, are conducted. One-dimensional moving contact discontinuity case and sod shock tube case suggest that HMTHS can accurately capture one-dimensional expansion waves, shock waves, and linear contact discontinuities. Two-dimensional cases, such as the double Mach reflection case, the supersonic shock / boundary layer interaction case, the hypersonic flow over the cylinder case, and the hypersonic viscous flow over the double-ellipsoid case, indicate that the HMTHS scheme is with a high resolution in simulating multidimensional complex flows. Therefore, it is promising to be widely applied in both scholar and engineering areas.

Transition prediction is a hot research topic of fluid mechanics. For subsonic and transonic aerodynamic flows, eNmethod based on Linear Stability Theory (LST) is usually adopted reliably to predict transition. In 2013, Coder and Maughmer established a transport equation for Tollmien-Schlichting (T-S) instability so that the eNmethod can be applied to general Reynolds-Average-Navier-Stokes (RANS) solvers conveniently. However, this equation focuses on T-S instability, and is invalid for crossflow instability induced transition which plays a crucial role in flow instability of three-dimensional boundary layers. Subsequently, a transport equation for crossflow instability was developed in 2016, which is restricted to wing-like geometries. Then, in 2019, this model was extended to arbitrarily shaped geometries based on local variables. However, there are too many tedious functions and parameters in this version, and it can only be used for incompressible flows. Hence, in this paper, after a large amount of LST analyses and parameter optimization, an improved version for subsonic and transonic boundary layers is built. The present improved model is more robust and more concise, and it can be applied widely in aeronautical flows, which has great engineering application value and significance. An extensive validation study for this improved transition model will be performed.

It is of great significance to develop a high-efficiency and low-noise propeller optimization method for new-generation propeller aircraft design. Coupled with free form deformation method, dynamic mesh interpolation technology, optimization algorithm, surrogate model, aerodynamic calculation and aeroacoustic prediction model module, the integrated aerodynamic and aeroacoustic design method of propeller is built. The optimization design for the six-blade propeller is carried out. The non-reduction in efficiency, thrust coefficient and the minimum of aerodynamic noise is treated as the optimization design objective. The spatial vorticity distribution of the propeller before and after the design is also analyzed by using unsteady computational fluid dynamics method. The results show that the optimized propeller can effectively reduce the aerodynamic noise level. The maximum total sound pressure level can be reduced by 5 dB without reducing its aerodynamic performance. The developed method has good application potential in low-noise optimization design of propeller and other rotating machinery.

Aerial refueling technology has been widely applied in various fields and it is one of the hotspots but difficulties for the aeronautical technologies. DLR-F6 WBNP model is used as a tanker and a fighter model is used as a receiver. The flow field of Probe–Drogue refueling and Flying Boom refueling is numerically simulated using the Reynolds-averaged Navier–Stokes equations, and the effects of the jet flow and the aerodynamic characteristics of the receiver are taken into consideration. The results indicate that the effect of downwash of the tanker reduces the lift coefficient and decreases the pitching moment coefficient of the receiver. The jet flow of tanker increases the dynamic pressure while decreases the local angle of attack, which increases the pressure difference between the upper and lower surfaces of receiver. Compared with the results without jet, the jet flow can increase the lift and the drag of the receiver and reduces the pitching moment, and even cause the change of rolling moment direction. Therefore, engine jet is an important factor when simulating aerial refueling.

Shale gas is becoming an increasingly promising alternative energy resource because of its high efficiency and environment-friendly characteristic. The amount of adsorbed gas on the shale matrix surfaces and dissolved gas in the shale matrix bulk is the dominant factor in the long-term productivity of shale reservoir. Although experimental measurements have been extensively carried out to investigate the gas adsorption and diffusion properties in the shale matrix, they cannot provide the detailed information on the microscopic transport mechanism of shale gas during the gas production process. Molecular simulation can accurately visualize the gas adsorption/desorption and diffusion processes in the shale matrix. In the present study, the recent research advances of molecular simulation on gas adsorption/desorption and diffusion in the shale matrix are reviewed. Firstly, the density functional theory (DFT) for shale gas molecule desorption/adsorption on the surface of the matrix crystal is illustrated. Then, the grand canonical Monte Carlo (GCMC) method predicting the amount of shale gas desorption/adsorption in the shale matrix crystal is introduced. Finally, molecular dynamics simulation (MD) for gas diffusion in the shale matrix is elucidated. Further developments of the molecular simulation method in shale gas production are also discussed.

The Vortex Particle Method (VPM) is a meshless Lagrangian vortex method. Its low numerical dissipation is exceptionally suitable for wake simulation. Nevertheless, the inadequate numerical stability of VPM prevents its widespread application in high Reynolds number flow and shear turbulence. To better simulate these flows, this paper proposes the stability-enhanced VPM based on a Reformulated VPM (RVPM) constrained by conservation of angular momentum, integrating a relaxation scheme to suppress the divergence of the vorticity field, and further coupling the Sub-Grid Scale (SGS) model to account for the turbulence dissipation caused by vortex advection and vortex stretching. The validity of the RVPM is confirmed by simulating an isolated vortex ring’s evolution. The results also demonstrate that the relaxation scheme of vorticity enhances the numerical stability of the VPM by mitigating the divergence of the vorticity field. The leapfrogging vortex rings simulation demonstrates that the RVPM with the present SGS model can more precisely feature the leapfrog and fusion of vortex rings and has improved numerical stability in high Reynolds number flows. The round turbulent jet simulation confirms that the stability-enhanced VPM can stably simulate shear turbulence and accurately resolve fluctuating components and Reynolds stresses in the turbulence.

A Non-Intrusive Reduced-Order Model (NIROM) based on Proper Orthogonal Decomposition (POD) has been proposed for predicting the flow fields of transonic airfoils with geometry parameters. To provide a better reduced-order subspace to approximate the real flow field, a domain decomposition method has been used to separate the hard-to-predict regions from the full field and POD has been adopted in the regions individually. An Artificial Neural Network (ANN) has replaced the Radial Basis Function (RBF) to interpolate the coefficients of the POD modes, aiming at improving the approximation accuracy of the NIROM for non-samples. When predicting the flow fields of transonic airfoils, the proposed NIROM has demonstrated a high performance.

This paper builds an aerodynamic optimization design system by coupling the free form deformation parameterization technique, the dynamic mesh technique based on inverse distance weighting interpolation method and particle swarm optimization algorithm, Kriging surrogate model, with the computational fluid dynamics solver based on Reynolds-averaged Navier–Stokes equations. And it carries out aerodynamic optimization designs with a supersonic transport aircraft configuration. The wing shape is parameterized by defining 60 design variables and the wing twist is by four design variables. Meanwhile, the wing thickness constraints are considered. Afterwards, a single-point optimization design aiming at minimum of the supersonic cruise drag is conducted, and in order to consider the transonic cruise condition, a multipoint aerodynamic optimization design aiming at minimum of a weighted drag for the supersonic cruise drag and transonic cruise drag is performed. The analyses of optimization results illustrate the specific drag reduction, the characteristics of pressure distributions for both the supersonic cruise condition and transonic cruise condition, and how the shock wave region and strength have improved on the supersonic cruise condition. The optimization design this paper presents realizing the drag reduction is quite practical and effective, which is instructive in supersonic aircraft design.

Aimed at the computational aeroacoustics multi-scale problem of complex configurations discretized with multi-size mesh, the flux reconstruction method based on modified Weight Essentially Non-Oscillatory (WENO) scheme is proposed at the interfaces of multi-block grids. With the idea of Dispersion-Relation-Preserving (DRP) scheme, different weight coefficients are obtained by optimization, so that it is in WENO schemes with various characteristics of dispersion and dissipation. On the basis, hybrid flux vector splitting method is utilized to intelligently judge the amplitude of the gap between grid interfaces. After the simulation and analysis of 1D convection equation with different initial conditions, modified WENO scheme is proved to be able to independently distinguish the gap amplitude and generate corresponding dissipation according to the grid resolution. Using the idea of flux reconstruction at grid interfaces, modified WENO scheme with increasing dissipation is applied at grid points, while DRP scheme with low dispersion and dissipation is applied at the inner part of grids. Moreover, Gauss impulse spread and periodic point sound source flow among three cylinders with multi-scale grids are carried out. The results show that the flux reconstruction method at grid interfaces is capable of dealing with Computational AeroAcoustics (CAA) multi-scale problems.

In aerodynamic optimization, global optimization methods such as genetic algorithms are preferred in many cases because of their advantage on reaching global optimum. However, for complex problems in which large number of design variables are needed, the computational cost becomes prohibitive, and thus original global optimization strategies are required. To address this need, data dimensionality reduction method is combined with global optimization methods, thus forming a new global optimization system, aiming to improve the efficiency of conventional global optimization. The new optimization system involves applying Proper Orthogonal Decomposition (POD) in dimensionality reduction of design space while maintaining the generality of original design space. Besides, an acceleration approach for samples calculation in surrogate modeling is applied to reduce the computational time while providing sufficient accuracy. The optimizations of a transonic airfoil RAE2822 and the transonic wing ONERA M6 are performed to demonstrate the effectiveness of the proposed new optimization system. In both cases, we manage to reduce the number of design variables from 20 to 10 and from 42 to 20 respectively. The new design optimization system converges faster and it takes 1/3 of the total time of traditional optimization to converge to a better design, thus significantly reducing the overall optimization time and improving the efficiency of conventional global design optimization method.

As boundary layer transition plays an important role in aerodynamic drag prediction, the proposal and study of transition prediction methods simulating the complex flow phenomena are prerequisite for aerodynamic design. In this paper, with the application of the linear stability theory based on amplification factor transport transition equations on the two-equation shear stress transport (SST) eddy-viscosity model, a new method, the SST-NTS-NCFmodel, is yielded. The new amplification factor transport equation for the crossflow instability induced transition is proposed to add to the NTSequation proposed by Coder, which simulates Tollmien–Schlichting wave transition. The turbulent kinetic energy equation is modified by introducing a new source term that simulates the transition process without the intermittency factor equation. Finally, coupled with these two amplification factor transport equations and SST turbulence model, a four-equation transition turbulence model is built. Comparisons between predictions using the new model and wind-tunnel experiments of NACA64(2)A015, NLF(2)-0415 and ONERA-D infinite swept wing and ONERA-M6 swept wing validate the predictive quality of the new SST-NTS-NCFmodel.

A one-equation turbulence model which relies on the turbulent kinetic energy transport equation has been developed to predict the flow properties of the recirculating flows. The turbulent eddy-viscosity coefficient is computed from a recalibrated Bradshaw’s assumption that the constant a1= 0.31 is recalibrated to a function based on a set of direct numerical simulation (DNS) data. The values of dissipation of turbulent kinetic energy consist of the near-wall part and isotropic part, and the isotropic part involves the von Karman length scale as the turbulent length scale. The performance of the new model is evaluated by the results from DNS for fully developed turbulence channel flow with a wide range of Reynolds numbers. However, the computed result of the recirculating flow at the separated bubble of NACA4412 demonstrates that an increase is needed on the turbulent dissipation, and this leads to an advanced tuning on the self-adjusted function. The improved model predicts better results in both the non-equilibrium and equilibrium flows, e.g. channel flows, backward-facing step flow and hump in a channel.

It is an inherent uncertainty problem that the application of laminar flow technology to the wing of large passenger aircraft is affected by flight conditions. In order to seek a more robust natural laminar flow control effect, it is necessary to develop an effective optimization design method. Meanwhile, attention must be given to the impact of crossflow (CF) instability brought on by the sweep angle. This paper constructs a robust optimization design framework based on discrete adjoint methods and non-intrusive polynomial chaos. Transition prediction is implemented by coupled Reynolds-Averaged Navier-Stokes (RANS) and simplified eNmethod, which can consider both Tollmien-Schlichting (TS) wave and crossflow vortex instability. We have performed gradient enhancement processing on the general Polynomial Chaos Expansion (PCE), which is advantageous to reduce the computational cost of single uncertainty propagation. This processing takes advantage of the gradient information obtained by solving the coupled adjoint equations considering transition. The statistical moment gradient solution used for the robust optimization design also uses the derivatives of coupled adjoint equations. The framework is applied to the robust design of a 25° swept wing with infinite span in transonic flow. The uncertainty quantification and sensitivity analysis on the baseline wing shows that the uncertainty quantification method in this paper has high accuracy, and qualitatively reveals the factors that dominate in different flow field regions. By the robust optimization design, the mean and standard deviation of the drag coefficient can be reduced by 29% and 45%, respectively, and compared with the deterministic optimization design results, there is less possibility of forming shock waves under flight condition uncertainties. Robust optimization results illustrate the trade-off between the transition delay and the wave drag reduction.

In this paper a new flow field prediction method which is independent of the governing equations, is developed to predict stationary flow fields of variable physical domain. Predicted flow fields come from linear superposition of selected basis modes generated by proper orthogonal decomposition (POD). Instead of traditional projection methods, kriging surrogate model is used to calculate the superposition coefficients through building approximate function relationships between profile geometry parameters of physical domain and these coefficients. In this context, the problem which troubles the traditional POD-projection method due to viscosity and compressibility has been avoided in the whole process. Moreover, there are no constraints for the inner product form, so two forms of simple ones are applied to improving computational efficiency and cope with variable physical domain problem. An iterative algorithm is developed to determine how many basis modes ranking front should be used in the prediction. Testing results prove the feasibility of this new method for subsonic flow field, but also prove that it is not proper for transonic flow field because of the poor predicted shock waves.

Due to the strong empirical demand of setting the design space variation in optimization design process which strongly affect the design result and efficiency, the adaptive aerodynamic optimization design method based on design variables space reconstruction concept has been established by the spatial statistical analysis of the design variables spatial distribution in the optimization process, which resolves the issue of design variables spatial distribution selection and lead to a better flexibility and convergence capability of the optimization model. As the design variables distribution has been statistically analyzed in the optimization process with the concept of clustering level, on the one hand the design variables spatial variation is reconstructed, on the other hand the design variables of a part of samples is adjusted in the reconstructed design space, both of which result in a better population diversity and a faster convergence with reservation of good genetic information. The NACA 0012 airfoil and NLF(1) 0416 airfoil have been optimized by this method, the results of which have been compared and analyzed with that of optimization of fixed design space variation. The design method approached has been proved of good feasibility with the optimization results, which results in an optimum solution in a larger variables distribution scale and shows better optimization efficiency.

In the originally published article version the family name of first author WANG Hui was misspellt “WAND”. The original article has been corrected.

In the design of a hypersonic airliner that can considerably shorten the flight time, how to accurately predict the vehicle’s aerodynamic heating loads is of great significance. In this study, a new shock-stable flux scheme called the simple low dissipation advection upwind splitting method (SLAU)-M1 is proposed for the prediction of hypersonic aerodynamic heating load. Based on the construction of the SLAU scheme for low-speed simulations, SLAU-M1 improves the robustness of the mass flux against shock instability. After validating the code employed, several numerical test cases are conducted. The one-dimensional (1D) sod shock tube case and the two-dimensional (2D) inviscid NACA0012 airfoil case show that SLAU-M1 features a high level of accuracy at both low and high speeds. To simulate the hypersonic viscous flow over a blunt cone, we adopt different aspect ratios (ARs) of cells near the shock. The results suggest that SLAU-M1 is much less sensitive to the AR of cells near the shock in predicting hypersonic aerodynamic heating loads. Moreover, the findings show that the theoretical value is considerably better than that of the other schemes. The hypersonic viscous flow over a 2D double ellipsoid case and that over the Hypersonic Flight Experiment vehicle case also indicate that SLAU-M1 exhibits a considerably high level of accuracy in hypersonic heating predictions. These properties suggest that SLAU-M1 promises to be widely used in the accurate prediction of the aerodynamic heating loads of hypersonic airliners.