Introducing active flow control into the design of flapping wing is an effective way to enhance its aerodynamic performance.In this paper,a novel active flow control technology called Co-Flow Jet(CFJ)is applied to fla...Introducing active flow control into the design of flapping wing is an effective way to enhance its aerodynamic performance.In this paper,a novel active flow control technology called Co-Flow Jet(CFJ)is applied to flapping airfoils.The effect of CFJ on aerodynamic performance of flapping airfoils at low Reynolds number is numerically investigated using Unsteady Reynolds Averaged Navier-Stokes(URANS)simulation with Spalart-Allmaras(SA)turbulence model.Numerical methods are validated by a NACA6415-based CFJ airfoil case and a S809 pitching airfoil case.Then NACA6415 baseline airfoil and NACA6415-based CFJ airfoil with jet-off and jet-on are simulated in flapping motion,with Reynolds number 70,000 and reduced frequency 0.2.As a result,CFJ airfoils with jet-on generally have better lift and thrust characteristics than baseline airfoils and jet-off airfoil when Cμgreater than 0.04,which results from the CFJ effect of reducing flow separation by injecting high-energy fluid into boundary layer.Besides,typical kinematic and geometric parameters,including the reduced frequency and the positions of the suction and injection slot,are systematically studied to figure out their influence on aerodynamic performance of the CFJ airfoil.And a variable Cμjet control strategy is proposed to further improve effective propulsive efficiency.Compared with using constant Cμ,an increase of effective propulsive efficiency by22.6%has been achieved by using prescribed variable CμNACA6415-based CFJ airfoil at frequency 0.2.This study may provide some guidance to performance enhancement for Flapping wing Micro Air Vehicles(FMAV).展开更多
The flapping motion has a great impact on the aerodynamic performance of flapping wings. In this paper, a surging motion is added to an airfoil performing pitching-plunging combined motion to figure out how it influen...The flapping motion has a great impact on the aerodynamic performance of flapping wings. In this paper, a surging motion is added to an airfoil performing pitching-plunging combined motion to figure out how it influences the lift performance and flow pattern of flapping airfoils.Firstly, the numerical methods are validated by a NACA0012 airfoil pitching case and a NACA0012 airfoil plunging case. Then, the E377m airfoil which has typical geometric characteristics of the bird-like airfoil is selected as the calculation model to study how phase differences φ1 between surging motion and plunging motion affect the aerodynamic performance of flapping airfoils. The results show that the airfoil with surging motion has comprehensively better lift performance and thrust performance than the airfoil without surging motion when 15°< φ1< 90°. It is demonstrated that surging motion has a powerful ability to improve the aerodynamic performance of flapping airfoil by adjusting φ1. Finally, to further explore how flapping airfoil improves lift performance by considering surging motion, the flapping motions of E377m airfoil with the highest lift coefficient and lift efficiency are obtained through trajectory optimization. The surging motion is removed in the highest lift case and highest lift efficiency case respectively, and the mechanism that surging motion adjusts the aerodynamic force is analyzed in detail by comparing the vortex structure and kinematic parameters. The results of this paper help reveal the aerodynamic mechanism of bird flight and guide the design of Flapping wing Micro Air Vehicles(FMAV).展开更多
A variable-fidelity method can remarkably improve the efficiency of a design optimization based on a high-fidelity and expensive numerical simulation,with assistance of lower-fidelity and cheaper simulation(s).However...A variable-fidelity method can remarkably improve the efficiency of a design optimization based on a high-fidelity and expensive numerical simulation,with assistance of lower-fidelity and cheaper simulation(s).However,most existing works only incorporate‘‘two"levels of fidelity,and thus efficiency improvement is very limited.In order to reduce the number of high-fidelity simulations as many as possible,there is a strong need to extend it to three or more fidelities.This article proposes a novel variable-fidelity optimization approach with application to aerodynamic design.Its key ingredient is the theory and algorithm of a Multi-level Hierarchical Kriging(MHK),which is referred to as a surrogate model that can incorporate simulation data with arbitrary levels of fidelity.The high-fidelity model is defined as a CFD simulation using a fine grid and the lower-fidelity models are defined as the same CFD model but with coarser grids,which are determined through a grid convergence study.First,sampling shapes are selected for each level of fidelity via technique of Design of Experiments(DoE).Then,CFD simulations are conducted and the output data of varying fidelity is used to build initial MHK models for objective(e.g.C_D)and constraint(e.g.C_L,C_m)functions.Next,new samples are selected through infillsampling criteria and the surrogate models are repetitively updated until a global optimum is found.The proposed method is validated by analytical test cases and applied to aerodynamic shape optimization of a NACA0012 airfoil and an ONERA M6 wing in transonic flows.The results confirm that the proposed method can significantly improve the optimization efficiency and apparently outperforms the existing single-fidelity or two-level-fidelity method.展开更多
Rotor noise is one of the most important reasons for restricting helicopter development;hence,the optimization design of rotor blade considering aeroacoustic and aerodynamic performance at the same time has always bee...Rotor noise is one of the most important reasons for restricting helicopter development;hence,the optimization design of rotor blade considering aeroacoustic and aerodynamic performance at the same time has always been the focus of research attention.For complex rotor design problems with a large number of design variables,the efficiency of the traditional Kriging model needs to be improved.Thus,Hierarchical Kriging(HK)model is employed in this study for rotor optimization design.By using the validated RANS solver and acoustic method based on the FWHpds equation,an efficient aerodynamic/aeroacoustic optimization method for high-dimensional problem of rotors in hover based on HK model is developed.By using present HK model and new infill-sampling criteria,the number of design variables is increased from less than 20-53.Results of two analytical function test cases show that the HK model is efficient and accurate in calculation.Subsequently,the helicopter rotor blade is optimally designed for aerodynamic/aeroacoustic performance in hover based on the HK model with high dimensional design variables.The objective function is adopted to improve the rotational noise characteristics by reducing the absolute peak of the acoustic pressure.In addition,the constraints of thrust,hover efficiency,solidity,and airfoils thickness are strictly satisfied.Optimization results show that the Kriging model finds the objective of reducing the noise by 2.87 dB after 248 iterations while the HK model does it only after 164 iterations.The optimization efficiency of the HK model is significantly higher than that of the traditional Kriging model.In the case analyzed,the HK model saves 35%of the time used by the Kriging model.展开更多
This article presents a novel approach for predicting transition locations over airfoils,which are used to activate turbulence model in a Reynolds-averaged Navier-Stokes flow solver.This approach combines Dynamic Mode...This article presents a novel approach for predicting transition locations over airfoils,which are used to activate turbulence model in a Reynolds-averaged Navier-Stokes flow solver.This approach combines Dynamic Mode Decomposition(DMD)with e^Ncriterion.The core idea is to use a spatial DMD analysis to extract the modes of unstable perturbations from a steady flowfield and substitute the local Linear Stability Theory(LST)analysis to quantify the spatial growth of Tollmien–Schlichting(TS)waves.Transition is assumed to take place at the stream-wise location where the most amplified mode’s N-factor reaches a prescribed threshold and a turbulence model is activated thereafter.To improve robustness,the high-order version of DMD technique(known as HODMD)is employed.A theoretical derivation is conducted to interpret how a spatial highorder DMD analysis can extract the growth rate of the unsteady perturbations.The new method is validated by transition predictions of flows over a low-speed Natural-Laminar-Flow(NLF)airfoil NLF0416 at various angles of attack and a transonic NLF airfoil NPU-LSC-72613.The transition locations predicted by our HODMD/e^Nmethod agree well with experimental data and compare favorably to those obtained by some existing methods■.It is shown that the proposed method is able to predict transition locations for flows over different types of airfoils and offers the potential for application to 3D wings as well as more complex configurations.展开更多
Accurate prediction of sonic boom is one of key challenges for the design of a low-boom supersonic aircraft. For most of available far-field prediction methods, the effect of atmospheric turbulence appearing in the pl...Accurate prediction of sonic boom is one of key challenges for the design of a low-boom supersonic aircraft. For most of available far-field prediction methods, the effect of atmospheric turbulence appearing in the planetary boundary layer cannot be considered, which results in remarkable inaccuracy of predicting ground-level sonic boom waveform. Although some efforts have been made to overcome the shortcoming, the turbulence effects are not yet well described so far. This article proposes an improved method by extending the two-dimensional Heterogeneous One-Way Approximation for the Resolution of Diffraction(HOWARD) equation to account for the axial and transverse convections of wind fluctuation as well as the effect of temperature fluctuation. The proposed method is validated by comparing the predictions with the flight-test data of JAXA D-SEND#1 LBM, which shows that the result of the proposed method is in better agreement with the flight-test data than that of the method without considering atmospheric turbulence effects.Then, distortion mechanism of sonic boom waveforms caused by atmospheric turbulence is analyzed by using the proposed method. It is indicated that the effect of turbulent convection makes uniform sonic-boom wavefronts irregular, which creates the condition of diffraction effect to perturb waveforms. Finally, the proposed method is applied to investigate the behavior of two types of waveforms given by the sonic boom minimization theory. Results show that a far-field waveform with a weaker initial shock is more beneficial for low-boom design of a supersonic aircraft.展开更多
Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor.A new high-order scheme(we call it WENO-K)proposed by our research group is employed to minimize numerical d...Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor.A new high-order scheme(we call it WENO-K)proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors.WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction,and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities.Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid.To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid,a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid.Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation.Numerical results for Caradonna-Tung(C-T)rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%.Compared with WENO-JS scheme,WENO-K scheme achieves about 40%improvement on accuracy of predicting rotor thrust with only 4.1%extra computational cost.More importantly,WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270°compared to WENO-JS scheme.展开更多
After the last flight of the Concorde in 2003,sonic boom has been one of the obstacles to the return of a supersonic transport aircraft to service.To reduce the sonic boom intensity to an acceptable level,it is of gre...After the last flight of the Concorde in 2003,sonic boom has been one of the obstacles to the return of a supersonic transport aircraft to service.To reduce the sonic boom intensity to an acceptable level,it is of great significance to study the effect of lift distribution on far-field sonic boom,since lift is one of the most important contributors to an intense sonic boom.Existing studies on the longitudinal lift distribution used low-fidelity methods,such as Whitham theory,and in turn,only preliminary conclusions were obtained,such as that extending the lift distribution can reduce sonic boom.This paper uses a newly developed high-fidelity prediction method to quantitatively study the effect of longitudinal lift distribution on the sonic boom of a Canard-Wing-Stabilator-Body(CWSB)configuration.This high-fidelity prediction method combines near-feld CFD simulation with far-field propagation by solving the augmented Burgers equation.A multipole analysis method is employed for the extraction of near-field waveform in order to reduce computational cost.Seven configurations with the same total lift but different distributions are studied,and the quantitative relationship between the longitudinal lift distribution and far-field sonic boom intensity is investigated.It is observed that a small lift generated by the stabilator can prevent aft-stabilator and aft-fuselage shocks from merging,while the balanced lift generated by the canard and wing can effectively keep the corresponding shocks further apart,which is beneficial for reducing both the on-track and off-track sonic boom.In turn,the acoustic level perceived at the ground can be reduced by 5.9 PLdB on-track and 5.4 PLdB off-track,on average.展开更多
A Dielectric Barrier Discharge(DBD) plasma actuator can create a body force which locally accelerates the base flow leading to an attenuation of broadband disturbance to delay the transition. In this study, numerical ...A Dielectric Barrier Discharge(DBD) plasma actuator can create a body force which locally accelerates the base flow leading to an attenuation of broadband disturbance to delay the transition. In this study, numerical simulation on an NLF0416 airfoil is conducted to investigate transition delay and drag reduction by a DBD plasma actuator. To simulate plasma’s effect more accurately, boundary-layer data is acquired from Reynolds Averaged Navier Stocks(RANS) equations instead of laminar boundary layer equations, although RANS equations need a much finer boundary-layer grid, and the linear stability analysis method is used to analyze the boundary layer and get the transition point. In this study, the influences of different actuation intensities and positions are investigated, and results show that if the actuation intensity is stronger and the actuation position is closer to the base transition point, more drag reduction can be obtained. However, the efficiency of plasma transition delay is really low. For example, when the actuation voltage is 16 k V,the actuation frequency is 1 k Hz, and the main Mach number is 0.1, the saved power due to drag reduction is about 5.09 W, but the power consumed is about 32.61 W, and the efficiency is just15.6%.展开更多
To meet the challenge of drag reduction for next-generation supersonic transport aircraft,increasing attention has been focused on Natural Laminar Flow(NLF)technology.However,the highly swept wings and high-Reynolds-n...To meet the challenge of drag reduction for next-generation supersonic transport aircraft,increasing attention has been focused on Natural Laminar Flow(NLF)technology.However,the highly swept wings and high-Reynolds-number conditions of such aircraft dramatically amplify Crossflow(CF)instabilities inside boundary layers,making it difficult to maintain a large laminar flow region.To explore novel NLF designs on supersonic wings,this article investigates the mechanisms underlying the attenuation of Tollmien-Schlichting(TS)and CF instabilities by modifying pressure distributions.The evolution of TS and CF instabilities are evaluated under typical pressure distributions with different leading-edge flow acceleration region lengths,pressure coefficient slopes and pressure coefficient deviations.The results show that shortening the leading-edge flow acceleration region and using a flat pressure distribution are favorable for suppressing CF instabilities,and keeping a balance of disturbance growth between positive and negative wave angles is favorable for attenuating TS instabilities.Based on the uncovered mechanisms,a strategy of supersonic NLF design is proposed.Examination of the proposed strategy at a 60°sweep angle and Ma=2 presents potential to exceed the conventional NLF limit and achieve a transition Reynolds number of 17.6million,which can provide guidance for NLF design on supersonic highly swept wings.展开更多
High-resolution numerical simulations for wake vortical flows have long been a challenge in rotor aerodynamics.A novel spectrum-optimized sixth-order Weighted Essentially NonOscillatory(WENO)scheme is proposed to disc...High-resolution numerical simulations for wake vortical flows have long been a challenge in rotor aerodynamics.A novel spectrum-optimized sixth-order Weighted Essentially NonOscillatory(WENO)scheme is proposed to discretize inviscid fluxes on moving overset grids,and the Improved Delayed Detached Eddy Simulation(IDDES)is employed to resolve turbulent vortices.The integration of these methods facilitates a comprehensive numerical investigation into the unsteady vortical flows over coaxial rotors in hover.The results highlight the substantial improvement in numerical resolution,in terms of both spatial structure and temporal evolution of unsteady multiscale wake vortices.Coaxial rotors in hover manifest three primary scales of wake vortex structures:(A)the helical evolution of primary blade tip vortices and the periodic occurrence of strong Blade-Vortex Interactions(BVI);(B)the continuous shedding of small-scale horseshoeshaped vortices from the trailing edges of rotor blades,forming the vortex sheets;(C)the emergence of small-scale secondary vortex braids induced by interactions between rotor tip vortices and the vortex sheets.These vortex structures and their interactions cause high-frequency oscillations in rotor disk loads and induce unsteady perturbations in the local flow field.Interactions among these primary vortices,coupled with the generation of secondary vortices,result in the dissipation,distortion,and breakup of the rotor tip vortices,ultimately forming a vortex soup.Notably,a substantial quantity of seemingly weak small-scale secondary vortex braids significantly contribute to energy dissipation during the evolution of wake vortices for coaxial rotors in hover.ó2024 Chinese Society of Aeronautics and Astronautics.Production and hosting by Elsevier Ltd.This is an open access article under the CC BY-NC-ND license(http://creativecommons.org/licenses/by-nc-nd/4.0/).展开更多
Background:Metastasis to the infraclavicular and supraclavicular lymph nodes(ISLNs)is an important factor that predicts poor survival in patients with breast cancer;however,pathological nodal staging does not traditio...Background:Metastasis to the infraclavicular and supraclavicular lymph nodes(ISLNs)is an important factor that predicts poor survival in patients with breast cancer;however,pathological nodal staging does not traditionally include ISLNs because of their non-routine surgical dissection.This study aimed to evaluate the prognostic impact of ISLN metastasis and propose a refined nodal staging system tailored for patients undergoing neoadjuvant chemotherapy(NAC).Methods:We retrospectively reviewed 1,072 patients with breast cancer with or without ISLN metastasis who received NAC at two institutions(Fujian cohort and Hebei cohort)from 2010 to 2022.We conducted detailed survival analysis to evaluate the diagnostic consistency and prognostic significance of ISLNs.Results:There were no survival differences among patients with ISLN involvement across different assay method-ologies and patient cohorts.Among 887 patients in the Fujian cohort,238 patients(26.8%)with positive ISLNs had significantly inferior 3-year progression-free survival(PFS,75.9%vs.90.4%,P<0.001)and overall survival(OS,90.6%vs.95.9%,P<0.001).After adjusting for potential confounders,ISLN involvement persisted as an independent predictor of both PFS and OS.We propose a refined axillary classification that combines pathologi-cal axillary staging post-NAC with ISLN involvement,revealing 3-year PFS rates of 95.3%,87.6%,73.4%,and 64.5%for the respective four groups defined by this refined classification combining axillary stage and ISLN status.Conclusions:Involvement of the ISLNs was associated with a worse prognosis,underscoring their prognostic value.This finding highlights the potential of ISLN status to influence decisions regarding adjuvant treatment in patients with breast cancer.展开更多
The present study introduces a Gauss-Seidel fluid-structure interaction(FSI)method including the flow solver,structural statics solver and a fast data transfer technique,for the research of structural deformation and ...The present study introduces a Gauss-Seidel fluid-structure interaction(FSI)method including the flow solver,structural statics solver and a fast data transfer technique,for the research of structural deformation and flow field variation of rotor blades under the combined influence of steady aerodynamic and centrifugal forces.The FSI method is illustrated and validated by the static aeroelasticity analysis of a transonic compressor rotor blade,NASA Rotor 37.An improved local interpolation with data reduction(LIWDR)algorithm is introduced for fast data transfer on the fluid-solid interface of blade.The results of FSI calculation of NASA Rotor 37 show that when compared with the radial basis function(RBF)based interpolation algorithm,LIWDR meets the interpolation accuracy requirements,while the calculation cost can be greatly improved.The data transmission time is only about 1%of that of RBF.Moreover,the iteration step of steady flow computation within one single FSI has little impact on the converged aerodynamic and structural results.The aerodynamic load-caused deformation accounts for nearly 50%of the total.The effects of blade deformation on the variations of aerodynamic performance are given,demonstrating that when static aeroelasticity is taken into account,the choke mass flow rate increases and the peak adiabatic efficiency slightly decreases.The impact mechanisms on performance variations are presented in detail.展开更多
To increase the efficiency and robustness of stability-based transition prediction in flow simulations, simplified methods are introduced to substitute direct stability analyses for rapid disturbance growth prediction...To increase the efficiency and robustness of stability-based transition prediction in flow simulations, simplified methods are introduced to substitute direct stability analyses for rapid disturbance growth prediction. For low-speed boundary layers, these methods are mainly established based on self-similar assumptions, which are not applicable to non-similar boundary layers in hypersonic flows. The objective of this article is to investigate the application of surrogate models to stability analysis of non-similar flows over blunt cones, focused on parameterization of boundary-layer (BL) profiles. Firstly, correlations between BL edge and profile parameters are analyzed, along with self-similar flow parameters and discrete points on BL profiles, which present four groups of BL characteristic parameters. Secondly, using these parameters as inputs, surrogate models are built for disturbance growth prediction over an MF-1 blunt cone. Results show that, surrogate models using four BL edge parameters and a BL shape factor {Ue, Te, ρe, ηe, H12} for stability analysis can achieve comparable accuracy with those using 16 discrete BL profile parameters, which are more precise than those using merely self-similar parameters or BL edge parameters. Thirdly, the established surrogate models are validated by stability analysis and transition prediction over the MF-1 blunt cone in flight experiments at the instants of t = 17 s ~ 22 s. Compared with direct linear stability analyses, the mean relative error of predicted disturbance growth rates by surrogate models is 8.0% and the maximum relative error of N factor envelopes is 6.6%, which indicates feasible applications of surrogate models to stability analysis and transition prediction of non-similar boundary layers in hypersonic flows.展开更多
基金co-supported by the National Key Research and Development Program of China(No.:2017YFB1300102)the National Natural Science Foundation of China(No.:11872314)。
文摘Introducing active flow control into the design of flapping wing is an effective way to enhance its aerodynamic performance.In this paper,a novel active flow control technology called Co-Flow Jet(CFJ)is applied to flapping airfoils.The effect of CFJ on aerodynamic performance of flapping airfoils at low Reynolds number is numerically investigated using Unsteady Reynolds Averaged Navier-Stokes(URANS)simulation with Spalart-Allmaras(SA)turbulence model.Numerical methods are validated by a NACA6415-based CFJ airfoil case and a S809 pitching airfoil case.Then NACA6415 baseline airfoil and NACA6415-based CFJ airfoil with jet-off and jet-on are simulated in flapping motion,with Reynolds number 70,000 and reduced frequency 0.2.As a result,CFJ airfoils with jet-on generally have better lift and thrust characteristics than baseline airfoils and jet-off airfoil when Cμgreater than 0.04,which results from the CFJ effect of reducing flow separation by injecting high-energy fluid into boundary layer.Besides,typical kinematic and geometric parameters,including the reduced frequency and the positions of the suction and injection slot,are systematically studied to figure out their influence on aerodynamic performance of the CFJ airfoil.And a variable Cμjet control strategy is proposed to further improve effective propulsive efficiency.Compared with using constant Cμ,an increase of effective propulsive efficiency by22.6%has been achieved by using prescribed variable CμNACA6415-based CFJ airfoil at frequency 0.2.This study may provide some guidance to performance enhancement for Flapping wing Micro Air Vehicles(FMAV).
基金supported by the National Natural Science Foundation of China(No.11872314)the Key R&D Program in Shaanxi Province of China(No.2020GY-154)。
文摘The flapping motion has a great impact on the aerodynamic performance of flapping wings. In this paper, a surging motion is added to an airfoil performing pitching-plunging combined motion to figure out how it influences the lift performance and flow pattern of flapping airfoils.Firstly, the numerical methods are validated by a NACA0012 airfoil pitching case and a NACA0012 airfoil plunging case. Then, the E377m airfoil which has typical geometric characteristics of the bird-like airfoil is selected as the calculation model to study how phase differences φ1 between surging motion and plunging motion affect the aerodynamic performance of flapping airfoils. The results show that the airfoil with surging motion has comprehensively better lift performance and thrust performance than the airfoil without surging motion when 15°< φ1< 90°. It is demonstrated that surging motion has a powerful ability to improve the aerodynamic performance of flapping airfoil by adjusting φ1. Finally, to further explore how flapping airfoil improves lift performance by considering surging motion, the flapping motions of E377m airfoil with the highest lift coefficient and lift efficiency are obtained through trajectory optimization. The surging motion is removed in the highest lift case and highest lift efficiency case respectively, and the mechanism that surging motion adjusts the aerodynamic force is analyzed in detail by comparing the vortex structure and kinematic parameters. The results of this paper help reveal the aerodynamic mechanism of bird flight and guide the design of Flapping wing Micro Air Vehicles(FMAV).
基金sponsored by the National Natural Science Foundation of China(Nos.11772261 and 11972305)Aeronautical Science Foundation of China(No.2016ZA53011)Foundation of National Key Laboratory(No.JCKYS2019607005).
文摘A variable-fidelity method can remarkably improve the efficiency of a design optimization based on a high-fidelity and expensive numerical simulation,with assistance of lower-fidelity and cheaper simulation(s).However,most existing works only incorporate‘‘two"levels of fidelity,and thus efficiency improvement is very limited.In order to reduce the number of high-fidelity simulations as many as possible,there is a strong need to extend it to three or more fidelities.This article proposes a novel variable-fidelity optimization approach with application to aerodynamic design.Its key ingredient is the theory and algorithm of a Multi-level Hierarchical Kriging(MHK),which is referred to as a surrogate model that can incorporate simulation data with arbitrary levels of fidelity.The high-fidelity model is defined as a CFD simulation using a fine grid and the lower-fidelity models are defined as the same CFD model but with coarser grids,which are determined through a grid convergence study.First,sampling shapes are selected for each level of fidelity via technique of Design of Experiments(DoE).Then,CFD simulations are conducted and the output data of varying fidelity is used to build initial MHK models for objective(e.g.C_D)and constraint(e.g.C_L,C_m)functions.Next,new samples are selected through infillsampling criteria and the surrogate models are repetitively updated until a global optimum is found.The proposed method is validated by analytical test cases and applied to aerodynamic shape optimization of a NACA0012 airfoil and an ONERA M6 wing in transonic flows.The results confirm that the proposed method can significantly improve the optimization efficiency and apparently outperforms the existing single-fidelity or two-level-fidelity method.
基金sponsored by the National Natural Science Foundation of China(Nos:11772261,11972305)"111"Project of China(No.:B17037).
文摘Rotor noise is one of the most important reasons for restricting helicopter development;hence,the optimization design of rotor blade considering aeroacoustic and aerodynamic performance at the same time has always been the focus of research attention.For complex rotor design problems with a large number of design variables,the efficiency of the traditional Kriging model needs to be improved.Thus,Hierarchical Kriging(HK)model is employed in this study for rotor optimization design.By using the validated RANS solver and acoustic method based on the FWHpds equation,an efficient aerodynamic/aeroacoustic optimization method for high-dimensional problem of rotors in hover based on HK model is developed.By using present HK model and new infill-sampling criteria,the number of design variables is increased from less than 20-53.Results of two analytical function test cases show that the HK model is efficient and accurate in calculation.Subsequently,the helicopter rotor blade is optimally designed for aerodynamic/aeroacoustic performance in hover based on the HK model with high dimensional design variables.The objective function is adopted to improve the rotational noise characteristics by reducing the absolute peak of the acoustic pressure.In addition,the constraints of thrust,hover efficiency,solidity,and airfoils thickness are strictly satisfied.Optimization results show that the Kriging model finds the objective of reducing the noise by 2.87 dB after 248 iterations while the HK model does it only after 164 iterations.The optimization efficiency of the HK model is significantly higher than that of the traditional Kriging model.In the case analyzed,the HK model saves 35%of the time used by the Kriging model.
基金supported by the National Natural Science Foundation of China (No. 11772261)the Aeronautical Science Foundation of China (No. 2016ZA53011)+1 种基金the ATCFD Project (No. 2015-F-016)the 111 Project of China (No. B17037)
文摘This article presents a novel approach for predicting transition locations over airfoils,which are used to activate turbulence model in a Reynolds-averaged Navier-Stokes flow solver.This approach combines Dynamic Mode Decomposition(DMD)with e^Ncriterion.The core idea is to use a spatial DMD analysis to extract the modes of unstable perturbations from a steady flowfield and substitute the local Linear Stability Theory(LST)analysis to quantify the spatial growth of Tollmien–Schlichting(TS)waves.Transition is assumed to take place at the stream-wise location where the most amplified mode’s N-factor reaches a prescribed threshold and a turbulence model is activated thereafter.To improve robustness,the high-order version of DMD technique(known as HODMD)is employed.A theoretical derivation is conducted to interpret how a spatial highorder DMD analysis can extract the growth rate of the unsteady perturbations.The new method is validated by transition predictions of flows over a low-speed Natural-Laminar-Flow(NLF)airfoil NLF0416 at various angles of attack and a transonic NLF airfoil NPU-LSC-72613.The transition locations predicted by our HODMD/e^Nmethod agree well with experimental data and compare favorably to those obtained by some existing methods■.It is shown that the proposed method is able to predict transition locations for flows over different types of airfoils and offers the potential for application to 3D wings as well as more complex configurations.
基金supported by the National Natural Science Foundation of China(Nos.U20B2007,11972305)the Aeronautical Science Foundation of China(No.2019ZA053004)+1 种基金the Shaanxi Science Fund for Distinguished Young Scholars(No.2020JC-13)the“111”Project of China(No.B17037)。
文摘Accurate prediction of sonic boom is one of key challenges for the design of a low-boom supersonic aircraft. For most of available far-field prediction methods, the effect of atmospheric turbulence appearing in the planetary boundary layer cannot be considered, which results in remarkable inaccuracy of predicting ground-level sonic boom waveform. Although some efforts have been made to overcome the shortcoming, the turbulence effects are not yet well described so far. This article proposes an improved method by extending the two-dimensional Heterogeneous One-Way Approximation for the Resolution of Diffraction(HOWARD) equation to account for the axial and transverse convections of wind fluctuation as well as the effect of temperature fluctuation. The proposed method is validated by comparing the predictions with the flight-test data of JAXA D-SEND#1 LBM, which shows that the result of the proposed method is in better agreement with the flight-test data than that of the method without considering atmospheric turbulence effects.Then, distortion mechanism of sonic boom waveforms caused by atmospheric turbulence is analyzed by using the proposed method. It is indicated that the effect of turbulent convection makes uniform sonic-boom wavefronts irregular, which creates the condition of diffraction effect to perturb waveforms. Finally, the proposed method is applied to investigate the behavior of two types of waveforms given by the sonic boom minimization theory. Results show that a far-field waveform with a weaker initial shock is more beneficial for low-boom design of a supersonic aircraft.
基金co-supported by the National Natural Science Foundation of China(No.12072285)Shaanxi Science foundation for Distinguished Young Scholars,China(No.2020JC-13)。
文摘Accurate prediction of tip vortices is crucial for predicting the hovering performance of a helicopter rotor.A new high-order scheme(we call it WENO-K)proposed by our research group is employed to minimize numerical dissipation and extended to numerical simulation of unsteady compressible viscous flows dominated by tip vortices over hovering rotors.WENO-K is referred to as an adaptively optimized WENO scheme with Gauss-Kriging reconstruction,and its advantage is to reduce dissipation in smooth regions of flow while preserving high-resolution around discontinuities.Here WENO-K scheme is adopted to reconstruct left and right state values within the Roe Riemann solver updating the inviscid fluxes on a structured dynamic overset grid.To minimize the accuracy loss for high-order reconstruction on artificial boundaries of overset grid,a method of multilayer fringes is proposed to carry out interpolation between background grid and blade grid.Massively parallel computing considering automatic load balance on averagely partitioned overset grid is developed to reduce the wall-clock time of an unsteady simulation.Numerical results for Caradonna-Tung(C-T)rotor in hover at the conditions of subsonic and transonic tip Mach numbers show that the thrust coefficient error for the result of WENO-K scheme is no more than 3%.Compared with WENO-JS scheme,WENO-K scheme achieves about 40%improvement on accuracy of predicting rotor thrust with only 4.1%extra computational cost.More importantly,WENO-K scheme can capture more sophisticated unsteady flow structures and resolve tip vortices to a larger wake age with an increment of about 270°compared to WENO-JS scheme.
基金sponsored by the National Natural Science Foundation of China(Nos.12072285,U20B2007)the Shaanxi Science Fund for Distinguished Young Scholars,China(No.2020JC-13)the Natural Science Funding of Shaanxi Province,China(No.2020JM-127).
文摘After the last flight of the Concorde in 2003,sonic boom has been one of the obstacles to the return of a supersonic transport aircraft to service.To reduce the sonic boom intensity to an acceptable level,it is of great significance to study the effect of lift distribution on far-field sonic boom,since lift is one of the most important contributors to an intense sonic boom.Existing studies on the longitudinal lift distribution used low-fidelity methods,such as Whitham theory,and in turn,only preliminary conclusions were obtained,such as that extending the lift distribution can reduce sonic boom.This paper uses a newly developed high-fidelity prediction method to quantitatively study the effect of longitudinal lift distribution on the sonic boom of a Canard-Wing-Stabilator-Body(CWSB)configuration.This high-fidelity prediction method combines near-feld CFD simulation with far-field propagation by solving the augmented Burgers equation.A multipole analysis method is employed for the extraction of near-field waveform in order to reduce computational cost.Seven configurations with the same total lift but different distributions are studied,and the quantitative relationship between the longitudinal lift distribution and far-field sonic boom intensity is investigated.It is observed that a small lift generated by the stabilator can prevent aft-stabilator and aft-fuselage shocks from merging,while the balanced lift generated by the canard and wing can effectively keep the corresponding shocks further apart,which is beneficial for reducing both the on-track and off-track sonic boom.In turn,the acoustic level perceived at the ground can be reduced by 5.9 PLdB on-track and 5.4 PLdB off-track,on average.
基金supported by the National Numerical Wind Tunnel Project (No. NNW2018-ZT3B08)。
文摘A Dielectric Barrier Discharge(DBD) plasma actuator can create a body force which locally accelerates the base flow leading to an attenuation of broadband disturbance to delay the transition. In this study, numerical simulation on an NLF0416 airfoil is conducted to investigate transition delay and drag reduction by a DBD plasma actuator. To simulate plasma’s effect more accurately, boundary-layer data is acquired from Reynolds Averaged Navier Stocks(RANS) equations instead of laminar boundary layer equations, although RANS equations need a much finer boundary-layer grid, and the linear stability analysis method is used to analyze the boundary layer and get the transition point. In this study, the influences of different actuation intensities and positions are investigated, and results show that if the actuation intensity is stronger and the actuation position is closer to the base transition point, more drag reduction can be obtained. However, the efficiency of plasma transition delay is really low. For example, when the actuation voltage is 16 k V,the actuation frequency is 1 k Hz, and the main Mach number is 0.1, the saved power due to drag reduction is about 5.09 W, but the power consumed is about 32.61 W, and the efficiency is just15.6%.
基金supported by the National Natural Science Foundation of China(No.12072285)the National Key Research and Development Program of China(No.2023YFB3002800)the Youth Innovation Team of Shaanxi Universities,China。
文摘To meet the challenge of drag reduction for next-generation supersonic transport aircraft,increasing attention has been focused on Natural Laminar Flow(NLF)technology.However,the highly swept wings and high-Reynolds-number conditions of such aircraft dramatically amplify Crossflow(CF)instabilities inside boundary layers,making it difficult to maintain a large laminar flow region.To explore novel NLF designs on supersonic wings,this article investigates the mechanisms underlying the attenuation of Tollmien-Schlichting(TS)and CF instabilities by modifying pressure distributions.The evolution of TS and CF instabilities are evaluated under typical pressure distributions with different leading-edge flow acceleration region lengths,pressure coefficient slopes and pressure coefficient deviations.The results show that shortening the leading-edge flow acceleration region and using a flat pressure distribution are favorable for suppressing CF instabilities,and keeping a balance of disturbance growth between positive and negative wave angles is favorable for attenuating TS instabilities.Based on the uncovered mechanisms,a strategy of supersonic NLF design is proposed.Examination of the proposed strategy at a 60°sweep angle and Ma=2 presents potential to exceed the conventional NLF limit and achieve a transition Reynolds number of 17.6million,which can provide guidance for NLF design on supersonic highly swept wings.
文摘High-resolution numerical simulations for wake vortical flows have long been a challenge in rotor aerodynamics.A novel spectrum-optimized sixth-order Weighted Essentially NonOscillatory(WENO)scheme is proposed to discretize inviscid fluxes on moving overset grids,and the Improved Delayed Detached Eddy Simulation(IDDES)is employed to resolve turbulent vortices.The integration of these methods facilitates a comprehensive numerical investigation into the unsteady vortical flows over coaxial rotors in hover.The results highlight the substantial improvement in numerical resolution,in terms of both spatial structure and temporal evolution of unsteady multiscale wake vortices.Coaxial rotors in hover manifest three primary scales of wake vortex structures:(A)the helical evolution of primary blade tip vortices and the periodic occurrence of strong Blade-Vortex Interactions(BVI);(B)the continuous shedding of small-scale horseshoeshaped vortices from the trailing edges of rotor blades,forming the vortex sheets;(C)the emergence of small-scale secondary vortex braids induced by interactions between rotor tip vortices and the vortex sheets.These vortex structures and their interactions cause high-frequency oscillations in rotor disk loads and induce unsteady perturbations in the local flow field.Interactions among these primary vortices,coupled with the generation of secondary vortices,result in the dissipation,distortion,and breakup of the rotor tip vortices,ultimately forming a vortex soup.Notably,a substantial quantity of seemingly weak small-scale secondary vortex braids significantly contribute to energy dissipation during the evolution of wake vortices for coaxial rotors in hover.ó2024 Chinese Society of Aeronautics and Astronautics.Production and hosting by Elsevier Ltd.This is an open access article under the CC BY-NC-ND license(http://creativecommons.org/licenses/by-nc-nd/4.0/).
基金supported by the Fujian Key Laboratory of Intelligent Imaging and Precision Radiotherapy for Tumors(Fujian Medical University)and Clinical Research Center for Radiology and Radiotherapy of Fujian Province(Digestive,Hematological and Breast Malignancies).
文摘Background:Metastasis to the infraclavicular and supraclavicular lymph nodes(ISLNs)is an important factor that predicts poor survival in patients with breast cancer;however,pathological nodal staging does not traditionally include ISLNs because of their non-routine surgical dissection.This study aimed to evaluate the prognostic impact of ISLN metastasis and propose a refined nodal staging system tailored for patients undergoing neoadjuvant chemotherapy(NAC).Methods:We retrospectively reviewed 1,072 patients with breast cancer with or without ISLN metastasis who received NAC at two institutions(Fujian cohort and Hebei cohort)from 2010 to 2022.We conducted detailed survival analysis to evaluate the diagnostic consistency and prognostic significance of ISLNs.Results:There were no survival differences among patients with ISLN involvement across different assay method-ologies and patient cohorts.Among 887 patients in the Fujian cohort,238 patients(26.8%)with positive ISLNs had significantly inferior 3-year progression-free survival(PFS,75.9%vs.90.4%,P<0.001)and overall survival(OS,90.6%vs.95.9%,P<0.001).After adjusting for potential confounders,ISLN involvement persisted as an independent predictor of both PFS and OS.We propose a refined axillary classification that combines pathologi-cal axillary staging post-NAC with ISLN involvement,revealing 3-year PFS rates of 95.3%,87.6%,73.4%,and 64.5%for the respective four groups defined by this refined classification combining axillary stage and ISLN status.Conclusions:Involvement of the ISLNs was associated with a worse prognosis,underscoring their prognostic value.This finding highlights the potential of ISLN status to influence decisions regarding adjuvant treatment in patients with breast cancer.
基金the Zhejiang Provincial Natural Science Foundation of China(Grant no.LXR22E060001)the National Science and Technology Major Project of China(Grant no.2017-II-0006-0020)the National Natural Science Foundation of China(Grant no.92152202).
文摘The present study introduces a Gauss-Seidel fluid-structure interaction(FSI)method including the flow solver,structural statics solver and a fast data transfer technique,for the research of structural deformation and flow field variation of rotor blades under the combined influence of steady aerodynamic and centrifugal forces.The FSI method is illustrated and validated by the static aeroelasticity analysis of a transonic compressor rotor blade,NASA Rotor 37.An improved local interpolation with data reduction(LIWDR)algorithm is introduced for fast data transfer on the fluid-solid interface of blade.The results of FSI calculation of NASA Rotor 37 show that when compared with the radial basis function(RBF)based interpolation algorithm,LIWDR meets the interpolation accuracy requirements,while the calculation cost can be greatly improved.The data transmission time is only about 1%of that of RBF.Moreover,the iteration step of steady flow computation within one single FSI has little impact on the converged aerodynamic and structural results.The aerodynamic load-caused deformation accounts for nearly 50%of the total.The effects of blade deformation on the variations of aerodynamic performance are given,demonstrating that when static aeroelasticity is taken into account,the choke mass flow rate increases and the peak adiabatic efficiency slightly decreases.The impact mechanisms on performance variations are presented in detail.
基金National Numerical Wind Tunnel Project(No.NNW2018-ZT1A03)National Natural Science Foundation of China(No.12072285 and No.11972305).
文摘To increase the efficiency and robustness of stability-based transition prediction in flow simulations, simplified methods are introduced to substitute direct stability analyses for rapid disturbance growth prediction. For low-speed boundary layers, these methods are mainly established based on self-similar assumptions, which are not applicable to non-similar boundary layers in hypersonic flows. The objective of this article is to investigate the application of surrogate models to stability analysis of non-similar flows over blunt cones, focused on parameterization of boundary-layer (BL) profiles. Firstly, correlations between BL edge and profile parameters are analyzed, along with self-similar flow parameters and discrete points on BL profiles, which present four groups of BL characteristic parameters. Secondly, using these parameters as inputs, surrogate models are built for disturbance growth prediction over an MF-1 blunt cone. Results show that, surrogate models using four BL edge parameters and a BL shape factor {Ue, Te, ρe, ηe, H12} for stability analysis can achieve comparable accuracy with those using 16 discrete BL profile parameters, which are more precise than those using merely self-similar parameters or BL edge parameters. Thirdly, the established surrogate models are validated by stability analysis and transition prediction over the MF-1 blunt cone in flight experiments at the instants of t = 17 s ~ 22 s. Compared with direct linear stability analyses, the mean relative error of predicted disturbance growth rates by surrogate models is 8.0% and the maximum relative error of N factor envelopes is 6.6%, which indicates feasible applications of surrogate models to stability analysis and transition prediction of non-similar boundary layers in hypersonic flows.
