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An improved adaptive sampling and experiment design method for aerodynamic optimization

Chinese Journal of Aeronautics 2015年第5期28卷 1391-1399页

作者： Huang Jiangtao Gao Zhenghong Zhou Zhu Zhao Ke Computational aerodynamics institute China Aerodynamics Research and Development Center National Key Laboratory of Aerodynamic Design and Research Northwestern Polytechnical University

Experiment design method is a key to construct a highly reliable surrogate model for numerical optimization in large-scale project. Within the method, the experimental design criterion directly affects the accuracy of the surrogate model and the optimization efficient. According to the shortcomings of the traditional experimental design, an improved adaptive sampling method is proposed in this paper. The surrogate model is firstly constructed by basic sparse samples. Then the supplementary sampling position is detected according to the specified criteria, which introduces the energy function and curvature sampling criteria based on radial basis function （RBF） network. Sampling detection criteria considers both the uniformity of sample distribution and the description of hypersurface curvature so as to significantly improve the prediction accuracy of the surrogate model with much less samples. For the surrogate model constructed with sparse samples, the sample uniformity is an important factor to the interpolation accuracy in the initial stage of adaptive sam- pling and surrogate model training. Along with the improvement of uniformity, the curvature description of objective function surface gradually becomes more important. In consideration of these issues, crowdness enhance function and root mean square error （RMSE） feedback function are introduced in C criterion expression. Thus, a new sampling method called RMSE and crowd- ness enhance （RCE） adaptive sampling is established. The validity of RCE adaptive sampling method is studied through typical test function firstly and then the airfoil/wing aerodynamic opti- mization design problem, which has high-dimensional design space. The results show that RCE adaptive sampling method not only reduces the requirement for the number of samples, but also effectively improves the prediction accuracy of the surrogate model, which has a broad prospects for applications.

关键词： Aerodynamic optimization Crowdness enhance function RBF model RCE adaptive sampfing RMSE feedback

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Experimental investigation on aero-heating of rudder shaft within laminar/turbulent hypersonic boundary layers

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Chinese Journal of Aeronautics 2019年第5期32卷 1215-1221页

作者： Qiang LI Liang NIE Kouli ZHANG Yu LI Suyu CHEN Guangsheng ZHU Hypervelocity aerodynamics institute China Aerodynamics Research and Development Center Science and Technology on Space Physics Laboratory China Academy of Launch Vehicle Technology

The aero-heating of the rudder shaft region of a hypersonic vehicle is very harsh, as the peak heat flux in this region can be even higher than that at the stagnation point. Therefore, studying the aero-heating of the rudder shaft is of great significance for designing the thermal protection system of the hypersonic vehicle. In the wind tunnel test of the aero-heating effect, we find that with the increase of the angle of attack of the lifting body model, the increasement of the heat flux of the rudder shaft is larger under laminar flow conditions than that under turbulent flow conditions. To understand this, we design a wind tunnel experiment to study the effect of laminar/turbulent hypersonic boundary layers on the heat flux of the rudder shaft under the same wind tunnel freestream conditions. The experiment is carried out in the ?2 m shock tunnel(FD-14 A) affiliated to the China aerodynamics Research and Development Center(CARDC). The laminar boundary layer on the model is triggered to a turbulent one by using vortex generators, which are 2 mm-high diamonds. The aero-heating of the rudder shaft(with the rudder) and the protuberance(without the rudder) are studied in both hypersonic laminar and turbulent boundary layers under the same freestream condition. The nominal Mach numbers are 10 and 12, and the unit Reynolds numbers are2.4 × 10~6 m^(-1) and 2.1 × 10~6 m-1. The angle of attack of the model is 20°, and the deflection angle of the rudder and the protuberance is 10°. The heat flux on the model surface is measured by thin film heat flux sensors, and the heat flux distribution along the center line of the lifting body model suggests that forced transition is achieved in the upstream of the rudder. The test results of the rudder shaft and the protuberance show that the heat flux of the rudder shaft is lower in the turbulent flow than that in the laminar flow, but the heat flux of the protuberance is the other way around,i.e., lower in the laminar flow than in th

关键词： Heat flux Hypersonic boundary layer Rudder Shock tunnel Transition Vortex generator

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An application of space-filling curves to improve results of turbulent aerodynamics modeling with convolutional neural networks

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Chinese Journal of Aeronautics 2024年第2期37卷 81-92页

作者： Mikhail PETROV Sofia ZIMINA Moscow institute of Physics&Technology Dolgoprudny 141701Russia

When carrying out calculations for turbulent flow simulation,one inevitably has to face the choice between accuracy and speed of *** order to simultaneously obtain both a computationally efficient and more accurate model,a surrogate model can be built on the basis of some fast special model and knowledge of previous calculations obtained by more accurate base models from various test bases or some results of serial *** objective of this work is to construct a surrogate model which allows to improve the accuracy of turbulent calculations obtained by a special model on unstructured *** this purpose,we use 1D Convolutional Neural Network(CNN)of the encoder-decoder architecture and reduce the problem to a single dimension by applying space-filling *** an approach would have the benefit of being applicable to solutions obtained on unstructured *** this work,a non-local approach is applied where entire flow fields obtained by the special and base models are used as input and ground truth output ***-Allmaras(SA)model and Near-wall Domain Decomposition(NDD)method for SA are taken as the base and special models *** efficiency and accuracy of the obtained surrogate model are demonstrated in a case of supersonic flow over a compression corner with different values for angleαand Reynolds number *** conducted an investigation into interpolation and extrapolation by Re and also into interpolation byα.

关键词： Space-filling curves Convolutional neural network Domain decomposition Turbulent flows Unstructured mesh

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Effect of gas blowing on the head of thermal ablation vehicle

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Advances in aerodynamics 2022年第1期4卷 188-202页

作者： Qing Li Xianxu Yuan Jianqiang Chen Lin Bi State Key Laboratory of aerodynamics China Aerodynamics Research and Development CenterMianyang621000China Computational aerodynamics institute China Aerodynamics Research and Development CenterMianyang621000China

The reentry vehicle will encounter thermal ablation,especially at the stagnation point regime.A theoretical work has been done to analyze the thermal effect of gas blowing due to thermal ablation of surface material on the head of a general hypersonic *** deriving the formulation,research takes into account the effect of gas blowing on the thermal dynamics balance,and then solves them by numerical *** is found that gas blowing will increase the temperature and heat flux at the surface of stagnation point regime.

关键词： Stagnation point flow Viscous compression of fluid parcel Gas blowing

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Inverse design of low boom configurations using proper orthogonal decomposition and augmented Burgers equation

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Chinese Journal of Aeronautics 2019年第6期32卷 1380-1389页

作者： Yidian ZHANG Jiangtao HUANG Zhenghong GAO Chao WANG Bowen SHU School of Aeronautics Northwestern Polytechnical UniversityXi’an 710072China Computational aerodynamics institute China Aerodynamics Research and Development CenterMianyang 621000China

Mitigation of sonic boom to an acceptable stage is a key point for the next generation of supersonic transports. Meanwhile, designing a supersonic aircraft with an ideal ground signature is always the focus of research on sonic boom reduction. This paper presents an inverse design approach to optimize the near-field signature of an aircraft, making it close to the shaped ideal ground signature after the propagation in the atmosphere. Using the Proper Orthogonal Decomposition(POD) method, a guessed input of augmented Burgers equation is inversely achieved. By multiple POD iterations, the guessed ground signatures successively approach the target ground signature until the convergence criteria is reached. Finally, the corresponding equivalent area distribution is calculated from the optimal near-field signature through the classical Whitham F-function theory. To validate this method, an optimization example of Lockheed Martin 1021 is demonstrated. The modified configuration has a fully shaped ground signature and achieves a drop of perceived loudness by 7.94 PLdB. This improvement is achieved via shaping the original near-field signature into wiggles and damping it by atmospheric attenuation. At last, a nonphysical ground signature is set as the target to test the robustness of this inverse design method and shows that this method is robust enough for various inputs.

关键词： Aeroacoustics Augmented Burgers equation Low boom configuration Optimization Supersonic aerodynamics

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Characteristics of reattached boundary layer in shock wave and turbulent boundary layer interaction

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Chinese Journal of Aeronautics 2022年第6期35卷 172-185页

作者： Fulin TONG Junyi DUAN Xinliang LI State Key Laboratory of aerodynamics China Aerodynamics Research and Development CenterMianyang 621000China Computational aerodynamics institute China Aerodynamics Research and Development CenterMianyang 621000China LHD Institute of MechanicsChinese Academy of SciencesBeijing 100190China School of Engineering Science University of Chinese Academy of SciencesBeijing 100049China

The reattached boundary layer in the interaction of an oblique shock wave with a flatplate turbulent boundary layer at Mach number 2.25 is studied by means of Direct Numerical Simulation(DNS).The numerical results are carefully compared with available experimental and DNS data in terms of turbulence statistics,wall pressure and skin *** coherent vortex structures are significantly enhanced due to the shock interaction,and the reattached boundary layer is characterized by large-scale structures in the outer *** space-time correlation of fluctuating wall shear stress and streamwise velocity fluctuation reveals that the structural inclination angle exhibits a gradual decrease during the recovery *** scale interactions are analyzed by using a twopoint amplitude modulation correlation.A possible mechanism is proposed to account for the strong amplitude modulation in the downstream ***,the mean skin-friction is decomposed to understand the physically informed *** the upstream Turbulent Boundary Layer(TBL),the contribution associated with the Turbulence Kinetic Energy(TKE)production is greatly amplified,while the spatial growth contribution induced by the pressure gradient largely inhibits skin-friction *** on bidimensional empirical mode decomposition,the turbulence kinetic energy production contribution is further split into different terms with specific spanwise length scales.

关键词： Amplitude modulation Shock wave Skin friction decomposition Structural inclination angle Turbulent boundary layer

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Installation error calibration for MEMS angular measurement system in hypersonic wind tunnel

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Chinese Journal of Aeronautics 2023年第5期36卷 175-186页

作者： Xiong WANG Xueling ZHAO Bin ZHOU Qi WEI Rong ZHANG Hypervelocity aerodynamics institute China Aerodynamics Research and Development CenterMianyang 621000China Department of Precision Instrument Tsinghua UniversityBeijing 100084China

The accuracy of model attitude measurement has an important impact on wind tunnel test results. Microelectromechanical System Inertial Measurement Unit(MEMS IMU) provides a feasible way to measure model attitudes with high accuracy. However, the installation error between MEMS IMU coordinate system and the body coordinate system of test models can make the accuracy of the model attitude measurement decrease. In wind tunnel tests, the installation error depends on the relationship between the IMU and the model mechanism before tests. Therefore, infield calibration in wind tunnel tests is necessary to reduce installation errors. To improve attitude measurement accuracy, the least squares quaternion calibration method based on MEMS IMU and six-position calibration procedure are proposed. High-precision three-axis turntable tests are performed. The pitch accuracy after calibration is higher than that before calibration in the angle of attack sweeping tests. The Root-Mean-Square Errors(RMSE) in the roll and yaw are within0.01°, which are smaller than those before calibration. In the roll sweeping tests, RMSE of three attitude angles decrease significantly. In hypersonic wind tunnel tests, the pitch errors before and after calibration are within 0.05° and 0.02° in the angle of attack sweeping tests without wind. In five angle of attack sweeping tests with wind, the deviation between the mean of the pitch and the pitch after the elastic angle correction is within 0.03° and the standard deviation of five tests is within 0.01°. The proposed method is confirmed to enhance the accuracy of attitude measurement effectively, which is convenient for engineering applications.

关键词： Calibration Installation error Least squares MEMS Model attitude measurement Quaternion Wind tunnel test

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Direct numerical simulation of impinging shock wave and turbulent boundary layer interaction over a wavy-wall

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Chinese Journal of Aeronautics 2021年第5期34卷 350-363页

作者： Fulin TONG Dong SUN Xinliang LI LHD Institute of MechanicsChinese Academy of SciencesBeijing 100190China State Key Laboratory of aerodynamics China Aerodynamics Research and Development CenterMianvang 621000China Computational aerodynamics institute China Aerodynamics Research and Development CenterMicmyang 621000China School of Engineering Science University of Chinese Academy of SciencesBeijing 100049China

The interaction of an impinging oblique shock wave with an angle of 30°and a supersonic turbulent boundary layer at Ma_(∞)=2.9 and Re_(θ)=2400 over a wavy-wall is investigated through direct numerical simulation and compared with the interaction on a flat-plate under the same flow conditions.A sinusoidal wave with amplitude to wavelength ratio of 0.26 moves in the streamwise direction and is uniformly distributed across the spanwise *** influences of the wavy-wall on the interaction,including the characterization of the flow field,the skin-friction,pressure and the budget of turbulence kinetic energy,are systematically *** region of separation grows slightly and decomposes into four *** peaks of skin-friction are observed at the rear part of the interaction *** low-frequency shock motion can be seen in the wall pressure *** of the turbulence kinetic energy budget indicate that both diffusion and transport significantly increase near the crests,balanced by an amplified dissipation in the near-wall *** orthogonal decomposition analyses show that the most energetic structures are associated with the separated shock and the shear layer over the *** the bubbles in the first two troughs are dominated by a low-frequency enlargement or shrinkage.

关键词： Direct numerical simulation Low-frequency unsteadiness Shock waves Turbulent boundary layer Wavy-wall

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Hypersonic shock wave and turbulent boundary layer interaction in a sharp cone/flare model

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Chinese Journal of Aeronautics 2023年第3期36卷 80-95页

作者： Fulin TONG Junyi DUAN Jiang LAI Dong SUN Xianxu YUAN State Key Laboratory of aerodynamics China Aerodynamics Research and Development CenterMianyang 621000China Computational aerodynamics institute China Aerodynamics Research and Development CenterMianyang 621000China LHD Institute of MechanicsChinese Academy of SciencesBeijing 100190China School of Engineering Science University of Chinese Academy of SciencesBeijing 100049China

A direct numerical simulation of hypersonic Shock wave and Turbulent Boundary Layer Interaction(STBLI)at Mach 6.0 on a sharp ***-angle circular cone/flare configuration at zero angle of attack is *** flare angle is *** the momentum thickness Reynolds number based on the incoming turbulent boundary layer on the sharp circular cone is Re θ=*** is found that the mean flow is separated and the separation bubble occurring near the corner exhibits *** Reynolds analogy factor changes dramatically across the interaction,and varies between 1.06 and 1.27 in the downstream region,while the QP85 scaling factor has a nearly constant value of 0.5 across the *** evolution of the reattached boundary layer is characterized in terms of the mean profiles,the Reynolds stress components,the anisotropy tensor and the turbulence kinetic *** is argued that the recovery is incomplete and the near-wall asymptotic behavior does not occur for the hypersonic *** addition,mean skin friction decomposition in an axisymmetric turbulent boundary layer is carried out for the first *** of the interaction,the contributions of transverse curvature and body divergence are negligible,whereas the positive contribution associated with the turbulence kinetic energy production and the negative spatial-growth contribution are *** on scale decomposition,the positive contribution is further divided into terms with different spanwise length *** negative contribution is analyzed by comparing the convective term,the streamwise-heterogeneity term and the pressure gradient term.

关键词： Shock wave Hypersonic turbulent boundary layer Heat transfer Skin friction decomposition Direct numerical simulation

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Aerodynamic configuration optimization by the integration of aerodynamics, aerothermodynamics and trajectory for hypersonic vehicles

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Chinese Science Bulletin 2014年第33期59卷 4608-4615页

作者： Yi Feng Wei Tang Yewei Gui School of Aerospace Tsinghua University Computational aerodynamics institute of China aerodynamics Research and Development Center

The problem of aerodynamic configuration design optimization is a multidisciplinary design optimization(MDO) problem, and recently the MDO method is widely adopted in the field of hypersonic vehicle configuration design. From the aerodynamic point of view, the aerodynamics, aerothermodynamics and trajectory are considered in this paper. Generally speaking, the aerodynamic characteristics, aerodynamic heating and trajectory are determined by the aerodynamic configuration and the design of flight trajectory. The design method considering these three disciplines is proposed. The parametric geometrical configurations are proposed, and the aerodynamic characteristics are predicted by the rapid and effective engineering method. The optimization of aerodynamic configuration considering the integration of aerodynamics,aerothermodynamics and trajectory is investigated based on the parametric geometrical configuration. Maximum lift-to-drag ratio, maximum range of the trajectory and minimum total heat load of the stagnation point are chosen as the three optimal goals. The detailed research indicates that the optimal configurations and trajectories with different weighting factors can be obtained by the optimization, and there are obvious differences between them. The optimal configuration and flight trajectory obtained by the optimization can be used as the feasible schemes in the future work.

关键词：高超声速飞行器空气动力学气动热力学优化配置整合优化多学科设计优化空气动力特性 MDO方法

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