SPR传感器阵列中微流通池的流场仿真

在自制表面等离子体激元共振(SPR)传感器阵列检测装置的微流通池的研制中,为了获得更好的检测精度和重复性,需要流经各检测点的流量尽可能一致。设计了几种不同的微流通池结构,并利用Comsol Multiphysics软件对其内部的流场进行仿真分析。首先,分析了常用梭状流通池的内部流场分布,并研究了不同长度过渡段对流场分布的影响。在此基础上,讨论了不同流通池形状,特别是流通池外形轮廓,如椭圆状流通池和边界流线型更好的流通池对流场分布的影响。此后,还研究了过渡区中微结构阵列设计对检测区流场分布的影响。仿真研究结果表明:适当延长过渡段长度,有利于改善流体流动的稳定性和均匀性,但这种改善将随长度的进一步增加逐渐减弱。而不同的边界外形轮廓对检测区域流场的影响不大;当采用一定的微结构阵列扰流设计后,流场均匀性一定程度上得到改善。     

EAST PCS异常事件处理算法

随着EAST实验装置的不断升级和等离子体放电参数的不断提高,获得的等离子体电流也在不断提高.高能量等离子体一旦破裂,将直接作用在实验装置的第一壁上面,带来毁灭性的灾害.本文研究了针对放电过程中出现的异常事件的处理,并且通过代码实现了控制等离子体电流在异常状态下斜率下降的算法.通过主动反馈控制的方法对发生异常的等离子体进行控制,达到对于电流的斜率下降.这个方法的有效性已经采用带圆等离子体模型的simulation server验证过了.     

飞行器翼型表面流场数据智能分区

飞行器翼型CFD仿真结果后处理分析自动化程度的提升能有效地提升产品设计效率,因此提出了一种翼型表面流场数据智能化分区方法,可有效得到翼型表面流场分区结果。首先,通过参数化批量修改气动外形得到翼型数据集,再利用数值模拟生成流场计算结果;然后,基于共形几何对流场数据进行降维并进行重采样和矩阵化,将其作为预测模型的标准输入;随后,构建卷积神经网络模型对流场数据进行训练和预测;最后,通过逆映射将分区结果重采样到翼型表面。实验表明,该方法可针对不同的物理量高效地对翼型表面流场数据进行分区,在测试数据集上的准确率在92%以上。     

基于低温等离子体的流动控制数值仿真研究

对于空气的流动控制问题,由于飞机在飞行中阻力较大,采用表面放电低温等离子体是较为有效的控制方法.采用格子玻尔兹曼模型对连续放置的两组等离子体发生器对边界层流动控制进行数值仿真.得到了壁面附近空气的速度分布.将等离子体发生器布置于NACA0015翼型表面,通过藕合计算得到的驱动速度,得到了等离子体作用前后翼型表面的流场分布以及升力系数、阻力系数曲线.数值仿真结果与实验结果较为吻合,等离子体可以有效的对翼型表面边界层进行控制,增加升力,降低阻力.     

运动模型式飞行器气动力参数测试系统研究

为了对飞行器进行气动布局、飞行性能评估,提出了一种运动模型式气动力参数测量方法,并研制了一套运动模型式测试系统,通过运动模型来模拟飞行器在空气中的运动,其构造主要包含载具、测试支架、六分量天平、风速测量模块、数据采集模块、控制模块、便携计算机等部分;该类测试系统在某些特定工况下能准确模拟飞行器的飞行状态,对流场参数和模型所受载荷进行实时测量,预测各运动状态下飞行器的气动特性;通过数值模拟分析和风洞实验对车载试验数据进行了对比验证,结果表明,该测试系统能够准确获得模型的气动力特性数据;该系统具有一定的拓展性和通用性,可使用火箭橇等其他载具开展高速运动模型式气动力参数测量试验。     

低温等离子体降解偏二甲肼废水研究

该文研究了悬浮电极介质阻挡放电装置(FE-DBD)产生的低温等离子体对偏二甲肼废水的降 解效果,并对处理条件进行优化。首先,对比了低温等离子体装置、氙灯和紫外灯降解偏二甲肼废水的效果;然后,考察了低温等离子体装置的放电间隙、初始溶液 pH、工作时间和氢氧化钠加投量对偏 二甲肼降解的影响;最后,探究了低温等离子体对偏二甲肼废水 pH 值的影响。实验结果表明,不加入其他试剂的情况下,低温等离子体装置降解偏二甲肼效果好于氙灯及紫外灯;装置放电间隙从 4 mm缩短至 2 mm,偏二甲肼降解率增加 47.2%。随着等离子体处理时间的增加,偏二甲肼的含量降低,处理 20 min 即可降解 82.1% 偏二甲肼。同时,低温等离子体处理会引起偏二甲肼废水 pH 值下降,处理10 min 后废水 pH 从 10 下降至 6.9。废水初始 pH 在 2～10 时,偏二甲肼降解率随废水 pH 值的升高而增大 ：与 pH＝2 相比,初始 pH＝10 时偏二甲肼降解率增加 65.9%。低温等离子体处理 10 min 后,往废水中加入氢氧化钠溶液至终浓度为 1 mg/mL,再继续处理 10 min,可将偏二甲肼降解率提高至 95%。     

大气环境仿真中数据模型及数据库的设计与实现

大气环境对飞行器有着至关重要的影响,不仅影响着飞行器本身的飞行和作战性能,而且也影响着飞行器操纵人员的决策.所以,大气环境仿真是高逼真度飞行仿真器设计的一个的重要组成部分.该文基于数据场的概念建立了大气环境数据模型,然后根据这个大气环境数据模型设计了一个具有四叉树结构的大气环境数据库,很好地解决了多分辨率的问题.最后将大气环境数据库映射到SEDRIS表示.     

托卡马克等离子体的高精度快响应平衡控制系统的研究

为了实现高功率波加热下等离子体的水平位移平衡控制,针对HT-6M托卡马克 装置薄壁真空室的特点,成功地研制了由载流等离子体、薄壁真空室和反馈场磁体所组成的 平衡控制系统.用1MW大功率GT0直流斩波器作为反馈控制磁场电源,并采用Bang—Bang 控制模式的技术,使该系统响应快、控制精度高.在各种实验条件下,等离子体的实际水平位 置控制在±3mm以内,相应的控制精度小于1.5%.     

空天飞行器质量特性管理系统设计与实现

质量特性作为空天飞行器总体设计的关键控制参数,与飞行器总体性能和技术指标密切相关,其设计与控制贯穿空天飞行器研制全过程。基于数字化设计环境和飞行器数字样机,提出了全飞行器质量特性管理系统方案,并在此基础上开发了原型系统,对设计思路进行了验证。相比于传统质量特性的计算及管理方法,该系统可有效提高设计效率,实现飞行器质量特性全寿命周期的动态可控、可见和可追溯。该成果后续还可推广应用于其他航天器研制。     

底部结构对飞行器阻力特性的影响北大核心CSCD

为优化底部形状减小底部阻力,研究底部流场结构对超声速飞行器的阻力特性影响,采用计算流体动力学方法,建立了雷诺平均NS方程数值求解方法。通过与实验数据比较,验证了方法的可靠性,对含圆柱尾部、船尾的典型飞行器在不同马赫数和攻角下进行数值仿真。仿真结果表明:船尾设计具有减阻效果,且升阻比优于圆柱尾部。攻角变化对底部流场结构影响显著。飞行器的升力系数、阻力系数、升阻比及俯仰力矩系数,随攻角增加而增大。随马赫数增加,飞行器头部压力升高,底部压力降低,导致总阻力、底部阻力增大,然而,由于动压增大,使得相应的总阻力系数及底阻系数随之减小。计算分析为飞行器底部设计和优化提供了参考。     

跨声速机翼阻力分解的一种改进方法

提出黏性区域探测器的一种改进形式，并用于捕捉激波和翼梢涡的熵增阻力；给出尾迹平面的可压缩涡动力学诱导阻力表达式，并与基于热力学的诱导阻力对比。在跨声速来流状态下，对ONERA M6和某民用飞机巡航状态下的机翼阻力进行分解，同时分析该民用飞机机翼安装翼梢小翼前、后的远场阻力构成。结果表明：新的区域探测器合理可靠，黏性阻力与伪熵阻力的计算结果更加准确；2种诱导阻力计算方式的计算结果一致，但基于涡动力学的诱导阻力计算方法受积分平面位置的影响更小；安装翼梢小翼基本不影响整个流场的黏性阻力，减阻的主要效果体现为诱导阻力的减小。     

The implementation of a knowledge-based framework for the aerodynamic optimization of a morphing wing device

In the field of aerospace engineering currently a lot of research effort is directed towards the reduction of cruise drag of civil transport aircraft in order to reduce fuel burn, and hence environmental impact and costs. In order to reduce cruise drag, a promising method is under consideration by adjusting, or rather morphing the rear part of the aircraft’s wing during cruise flight. Given the premature state of knowledge of such a design implementation, a knowledge-based computational framework is developed. The purpose of this framework is to allow for an aerodynamic optimization of a section of the wing. The framework is set up in such a way that all relevant design knowledge generated in the process can be captured and used in a subsequent mechanical design process. In this fashion, the complex design process of a novel morphing wing device can be automated to a certain degree. This automation can be used to construct a large number of different feasible and optimized designs with varying boundary conditions of a complex experimental device.This article describes the initial 2-dimensional aerodynamic design step of the morphing device under consideration and how it is implemented in a knowledge-based optimization framework. It describes the initial stage of the development of this tool, as it will be expanded by a number of design steps that each adds more detail to the design in all relevant aspect fields (aerodynamic, structural, actuation, etc.). Ultimately, this tool will be used to obtain a thorough evaluation of a number of different proposed structural solutions and allow for a comparison between them.     

Coupled adjoint aerostructural wing optimization using quasi-three-dimensional aerodynamic analysis

This paper presents a method for wing aerostructural analysis and optimization, which needs much lower computational costs, while computes the wing drag and structural deformation with a level of accuracy comparable to the higher fidelity CFD and FEM tools. A quasi-three-dimensional aerodynamic solver is developed and connected to a finite beam element model for wing aerostructural optimization. In a quasi-three-dimensional approach an inviscid incompressible vortex lattice method is coupled with a viscous compressible airfoil analysis code for drag prediction of a three dimensional wing. The accuracy of the proposed method for wing drag prediction is validated by comparing its results with the results of a higher fidelity CFD analysis. The wing structural deformation as well as the stress distribution in the wingbox structure is computed using a finite beam element model. The Newton method is used to solve the coupled system. The sensitivities of the outputs, for example the wing drag, with respect to the inputs, for example the wing geometry, is computed by a combined use of the coupled adjoint method, automatic differentiation and the chain rule of differentiation. A gradient based optimization is performed using the proposed tool for minimizing the fuel weight of an A320 class aircraft. The optimization resulted in more than 10 % reduction in the aircraft fuel weight by optimizing the wing planform and airfoils shape as well as the wing internal structure.     

迷你翼梢小翼增升减阻效应

选取某窄体客机的翼梢小翼为研究对象,采用Spalart Allmaras模型对无翼梢小翼、全尺寸翼梢小翼和迷你翼梢小翼3种机翼构型进行数值模拟,通过流场分析和速度分解等手段,研究翼梢小翼的增升减阻机理。结果表明：迷你翼梢小翼有恢复涡核流速、减弱涡流掺混程度和梳理翼梢气流的作用;增升减阻的关键在于迷你翼梢小翼对气流方向的修正;翼梢小翼的局部流动差异会对整体机翼近场造成影响。由于尺寸较小,迷你翼梢小翼能在较大攻角范围内改善传统翼梢小翼的性能,具有一定的实践意义。     

乘波体组合高压捕获翼构型的性能分析

针对高速飞行器大容积、高升力、低阻力和高升阻比的设计需求,提出高压捕获翼(High pressure zone Capture Wing,HCW)的概念.在高速巡航条件下,合理配置HCW可以充分利用来流压缩产生的高压气体,从而提高飞行器升力;HCW采用与来流平行的薄板装置,其附加阻力较小,可以大幅提高升阻比.采用CFD分析工具,比较不同容积的乘波体构型与HCW组合前后的气动性能.结果表明,在不同容积构型下升阻比均有明显提高,最小提升量可达10%.此外,容积越大,升力和升阻比增加效果越明显.     

Drag of a flexible fiber in a 2D moving viscous fluid

This paper reports the numerical study of the drag of a flexible elastic fiber immersed in a two-dimensional viscous flow using the immersed boundary (IB) method. We found drag reduction of a flexible fiber compared to a stiff one and the drag coefficient decreases with respect to the dimensionless fiber length within a certain range. The results are a starting point for the understanding of the role of flexibility in biological organisms in fluid flows.     

Multi-fidelity wing aerostructural optimization using a trust region filter-SQP algorithm

A trust region filter-SQP method is used for wing multi-fidelity aerostructural optimization. Filter method eliminates the need for a penalty function, and subsequently a penalty parameter. Besides, it can easily be modified to be used for multi-fidelity optimization. A low fidelity aerostructural analysis tool is presented, that computes the drag, weight and structural deformation of lifting surfaces as well as their sensitivities with respect to the design variables using analytical methods. That tool is used for a mono-fidelity wing aerostructral optimization using a trust region filter-SQP method. In addition to that, a multi-fidelity aerostructural optimization has been performed, using a higher fidelity CFD code to calibrate the results of the lower fidelity model. In that case, the lower fidelity tool is used to compute the objective function, constraints and their derivatives to construct the quadratic programming subproblem. The high fidelity model is used to compute the objective function and the constraints used to generate the filter. The results of the high fidelity analysis are also used to calibrate the results of the lower fidelity tool during the optimization. This method is applied to optimize the wing of an A320 like aircraft for minimum fuel burn. The results showed about 9 % reduction in the aircraft mission fuel burn.     

短波地空通信自动化系统

针对目前短波地空通信系统中的设备操作复杂、可靠性差等缺点,研究了原有系统的设备性能、管理方法和操作流程,设计了一套改进的自动化系统。采用系统设计的方法,确定系统总体结构,进行设备选型和连接,实现了系统集成;设计了接口设备——选呼控制器,使地空通信呼叫过程由手动操作转化为自动控制,实现了流程改进。系统简化了操作流程,提高了功能可靠性,同时扩大了远程控制距离。测试与应用表明:系统的呼叫范围可达6000 km,单次选呼成功率提高了20个百分点,操作台的远程控制距离由原有的几十米提高到几十甚至上百公里。     

基于等离子流动控制的翼型减阻技术

为研究基于等离子流动控制的减阻技术,基于Langtry-Menter转捩模型提出边界层转捩数值模拟技术.该技术可有效结合转捩模型与湍流模型,用标准模型验证其精确性,为采用等离子流动控制抑制边界层分离和转捩研究提供数值模拟平台.采用基于现象学模型的等离子流动控制数值模拟技术,对流动分离以及边界层转捩抑制进行数值模拟,为基于等离子流动控制的翼型减阻技术提供参考.     

Design of airfoils in incompressible viscous flows by numerical optimization

A method is outlined for the design of airfoils in incompressible viscous flows by numerical optimization wherein a reduced number of design coordinates are used to define the airfoil shape. The optimization problem is formulated as a nongradient search in a finite constrained parameter space. The approach is to define the airfoil as a linear combination of basic shapes which may be analytically or numerically defined. The design problem is to determine the participation of each of these basic shapes in defining the optimum airfoil. The aerodynamic analysis program is specially developed to fit the requirements of the optimization program and is based on the vortex singularity method for inviscid flow analysis and the momentum integral method for boundary layer analysis. Four examples have been worked out to illustrate the proposed design method. In these, modifications to four different airfoil geometries are made to achieve either a minimum drag coefficient or a minimum pitching moment coefficient under prescribed constraints. The results show that significant drag or pitching moment reduction is possible through shape manipulation alone.