Transonic Rudder Buzz on Tailless Flying Wing UAV

Transonic rudder buzz responses based on the computational fluid dynamics or computational structural dynamics(CFD/CSD)loosely method are analyzed for a tailless flying wing unmanned aerial vehicle(UAV).The Reynolds-averaged Navier-Stokes(RANS)equations and finite element methods based on the detailed aerodynamic and structural model are established,in which the aerodynamic dynamic meshes adopt the unstructured dynamic meshes based on the combination of spring-based smoothing and local remeshing methods,and the lower-upper symmetric-Gauss-Seidel(LU-SGS)iteration and Harten-Lax-van Leer-Einfeldt-Wada(HLLEW)space discrete methods based on the shear stress transport(SST)turbulence model are used to calculate the aerodynamic force.The constraints of the rudder motions are fixed at the end of structural model of the flying wing UAV,and the structural geometric nonlinearities are also considered in the flying wing UAV with a high aspect ratio.The interfaces between structural and aerodynamic models are built with an exact match surface where load transferring is performed based on 3Dinterpolation.The flying wing UAV transonic buzz responses based on the aerodynamic structural coupling method are studied,and the rudder buzz responses and aileron,elevator and flap vibration responses caused by rudder motion are also investigated.The effects of attack,height,rotating angular frequency and Mach number under transonic conditions on the flying wing UAV rudder buzz responses are discussed.The results can be regarded as a reference for the flying wing UAV engineering vibration analysis.

Analysis of flow separation control using nanosecond-pulse discharge plasma actuators on a flying wing

Dielectric barrier discharge （DBD） plasma is one of most promising flow control method for its several advantages. The present work investigates the control authority of nanosecond pulse DBD plasma actuators on a flying wing model＇s aerodynamic characteristics. The aerodynamic forces and moments are studied by means of experiment and numerical simulation. The numerical simulation results are in good agreement with experiment results. Both results indicate that the NS-DBD plasma actuators have negligible effect on aerodynamic forces and moment at the angles of attack smaller than 16-. However, significant changes can be achieved with actuation when the model＇s angle of attack is larger than 16° where the flow separation occurs. The spatial flow field structure results from numerical simulation suggest that the volumetric heat produced by NS-DBD plasma actuator changes the local temperature and density and induces several vortex structures, which strengthen the mixing of the shear layer with the main flow and delay separation or even reattach the separated flow.

Experimental investigation of lift enhancement for flying wing aircraft using nanosecond DBD plasma actuators

he effects of the arrangement position and control parameters of nanosecond dielectric barrier discharge （NS-DBD） plasma actuators on lift enhancement for flying wing aircraft were investigated through wind tunnel experiments at a flow speed of 25 m s-1.The aerodynamic forces and moments were obtained by a six-component balance at angles of attack ranging from -4° to 28°.The lift,drag and pitching moment coefficients were compared for the cases with and without plasma control.The results revealed that the maximum control effect was achieved by placing the actuator at the leading edge of the inner and middle wing,for which the maximum lift coefficient increased by 37.8% and the stall angle of attack was postponed by 8° compared with the plasma-off case.The effects of modulation frequency and discharge voltage were also investigated.The results revealed that the lift enhancement effect of the NS-DBD plasma actuators was strongly influenced by the modulation frequency.Significant control effects were obtained atf =70 Hz,corresponding to F＋ ≈ 1.The result for the pitching moment coefficient demonstrated that the plasma actuator can induce the reattachment of the separation flows when it is actuated.However,the results indicated that the discharge voltage had a negligible influence on the lift enhancement effect.

Modeling and Motion Simulation for A Flying-Wing Underwater Glider with A Symmetrical Airfoil

The flying-wing underwater glider (UG), shaped as a blended wing body, is a new type of underwater vehicle and still requires further research. The shape layout and the configuration of the internal actuators of the flying-wing UG are different from those of "legacy gliders" which have revolving bodies, and these two factors strongly affect the dynamic performance of the vehicle. Considering these differences, we propose a new configuration of the internal actuators for the flying-wing UG and treat the flying-wing UG as a multi-body system when establishing its dynamic model. In this paper, a detailed dynamic model is presented using the Newton-Euler method for the flying-wing UG. Based on the full dynamic model, the effect of the internal actuators on the steady gliding motion of vehicle is studied theoretically, and the relationship between the state parameters of the steady gliding motion and the controlled variables is obtained by solving a set of equilibrium equations. Finally, the behaviors of two classical motion modes of the glider are analyzed based on the simulation. The simulation results demonstrate that the motion performance of the proposed flying-wing UG is satisfactory.

飞翼布局飞行器研究现状分析

对飞翼布局飞行器的发展脉络进行了梳理,涵盖了从其早期的进化史到现阶段的研究进展,并对未来的发展方向进行了思考和展望。同时,论述了飞翼布局飞行器的几个主要控制难点,具体包括:宽速域非线性状态下的姿态控制问题、多效应耦合下的高精度自主起降控制问题、气动模型近似及强扰动状态下的鲁棒控制难题,以及因多舵面冗余而引起的控制分配问题。此外,还深入讨论了当前重点研究的飞翼布局飞行器的流动控制问题,并对接下来主动流动控制技术的研究动向进行了阐述。

具有飞行包线限制的飞翼无人机鲁棒自适应容错姿态控制

为实现复杂环境下飞翼无人机姿态的精确跟踪控制,考虑参数不确定性、外部扰动、执行器故障及飞行包线限制的影响,提出一种基于Nussbaum增益的鲁棒自适应容错控制方法。在受扰动的飞翼无人机姿态运动学与动力学模型基础上考虑执行器故障与系统不确定的影响,建立面向控制的姿态控制模型。通过引入时变障碍Lyapunov函数,在保证飞行包线限制的同时确保姿态跟踪误差的瞬态与稳态性能。通过自适应的有界估计与Nussbaum增益,补偿总的不确定项与执行器故障的影响。通过稳定性分析严格证明新提出的控制方法的可行性。仿真结果表明,新的控制方法能够确保飞行包线限制,同时保证预设的瞬态与稳态性能,实现飞翼无人机高精度的姿态跟踪控制。

平直翼飞翼布局飞机的操稳特性

飞翼布局飞机气动效率高,常被太阳能飞机所采用,但操稳特性较差.平直翼飞翼布局飞机具有与主流后掠飞翼布局飞机不同的气动特性.使用涡格法计算了轻小型平直翼飞翼布局无人机的气动力系数与气动导数,分析了某验证机的俯仰静稳定性及不同迎角、爬升角下的纵向与横航向稳定性,讨论了重心位置低于机翼平面对其稳定性的影响,并制作航空模型试飞.研究发现,平直翼飞翼布局飞机的纵向稳定性与常规布局飞机不同,可以在无增稳条件下维持横航向稳定,荷兰滚倾向弱.重心位置降低可以改变配平迎角,使其在正迎角下俯仰更稳定,但会使其在负迎角或大爬升角下的俯仰稳定性恶化.飞机在飞行中表现出多种特殊行为,需要进一步研究.

环量控制飞翼翼型隐身与气动联合优化

飞翼无人机可利用主动流动控制技术取代传统舵面,避免舵面偏转产生的雷达回波,能够获得更好的隐身性能。本文面向装载环量控制器的大展弦比飞翼无人机,提出了一种兼顾气动和隐身性能的飞翼翼型联合优化方法。针对包含环量控制(Circulation control,CC)机构的二维翼型建立了基于代理模型的联合优化框架,并通过寻优求解,获得了Pareto前沿解集。选择Pareto前沿中的3个点对应的二维翼型进行了气动与隐身性能评估,数值计算结果证实了本文提出的联合优化方法的可行性。

基于高仿生形态布局的仿鸽扑翼飞行机器人系统设计

针对现有扑翼飞行机器人存在的飞行形态与实际鸟类相差较大,以及翅膀、尾翼布局和俯仰、转向控制方式仿生度较低的问题,提出一种形态布局与鸽子相仿的扑翼飞行机器人系统设计及实现方案.通过设计弧面-折翼-后掠翅膀、仿鸟扇形尾翼以及尾翼挨近翅膀后缘布置的布局方式,使扑翼机器人飞行形态更加接近真实鸟类,提高扑翼机器人的形态仿生度.在此基础上,设计结合下扑角调控无需尾翼大角度上翘的俯仰控制方式,以及不依赖于尾翼的翅膀收缩转向控制方式,在提高仿生度的同时保证飞行控制的有效性.在具体设计过程中,首先参考鸽子翅膀型式选择不同类型翅膀并进行风洞测试,确定出下扑角变化时仍能保持较优升推力性能的翅膀设计方案;其次,对各种尾翼型式进行分析和比较,结合鸽子尾翼特点进行仿鸽尾翼及俯仰、转向控制机构设计,并通过风洞测试验证;最后,设计飞控系统并装配整机,进行外场飞行测试,验证仿鸽扑翼飞行机器人平台的稳定性和可控性.

飞翼无人机外形与进排气几何一体化参数化建模

飞翼布局是无人作战飞机的优先布局方案,其特点是机体与发动机进排气系统高度融合。针对飞翼式无人作战飞机概念方案快速几何建模需求,提出一种飞翼机体外形和进排气几何外形的一体化参数化建模方法。基于类函数/形函数方法,建立飞翼机体和进排气几何外形的参数化数学模型;确定关联的外形参数和关联控制规则,实现飞翼外形参数与进排气几何外形的匹配;将所建立的模型与CATIA二次开发方法相结合,实现飞翼式无人机概念方案三维几何模型的自动生成。结果表明:针对不同的进排气方案,本文方法能自动调节飞翼外形,匹配进排气系统的几何外形,有效地提高了飞翼式无人作战飞机概念设计的效率。

飞行汽车机翼折叠系统机构运动学分析

为延长飞行汽车的飞行时间及实现飞行汽车可以在空中飞行模式与陆地行驶模式下正常工作,在根据升力公式对飞行汽车机翼进行设计选型后,对机翼折叠、回收机构进行设计并基于复数矢量法建立机翼折叠、回收机构运动学模型,使用MATLAB分析计算了机构关键杆件的角位移、角速度、角加速度变化情况,并用Adams对机构进行运动仿真分析。分析结果表明:机翼折叠机构与回收机构的角位移、角速度、角加速度图像均未出现激增或骤降,机构结构设计合理,可实现平滑稳定运动,运动特性良好。其中,通过控制折叠机构驱动杆件由83.25°偏转至116.75°实现折叠外翼90°的偏转。通过控制回收机构驱动滑块位移516.6 mm实现机翼90°的展开。最终在30 s内完成机翼折叠及回收作业。

典型飞翼布局轰炸机气动与红外特性数值分析

以美国B2轰炸机作为研究对象,构建了进排气系统一体化整机仿真物理模型。计算了在不同飞行条件下轰炸机的气动/红外特性。结果表明:当飞行马赫数提高时,外流产生的对流换热加热蒙皮效果越发明显,跨声速飞行时,由于气动加热作用使机身前缘温度升高;水平方向上随着探测角度增加,3~5μm与8~14μm波段内轰炸机红外辐射强度降低;竖直方向探测时,3~5μm波段内轰炸机下方最大红外辐射强度值仅为上方最大值的20%,对规避来自地面及低空的红外探测具有重要意义;3~5μm波段内降低壁面发射率对轰炸机红外辐射的抑制效果在后方及上方探测角度内更为明显,最大降幅达60%;8~14μm波段内降低壁面发射率后轰炸机红外特征显著降低,最大降幅达70%。

面向扑翼飞行机器人的电子稳像算法设计

在扑翼飞行机器人的飞行过程中,由于其特有的驱动方式,机身存在周期性的俯仰和滚转运动,导致航拍视频出现抖动,影响成像效果的清晰度和稳定性.为了解决这一问题,本文提出了一种基于ORB和滑动均值滤波的电子稳像算法,用于在线消抖处理.首先,针对扑翼飞行机器人航拍图像的抖动周期与机翼扑动周期相一致的特性,设计了一种估计算法来根据图像特征估计机翼扑动周期.这一算法能够更准确地捕捉到抖动的周期性特征,为后续的稳像处理提供了重要参数.其次,提出了一种与扑动周期相关联的运动滤波算法,能够根据不同飞行工况对滤波参数自适应地进行动态调整.本文提出的算法优点在于能够根据扑翼飞行机器人实际飞行情况实时调整参数,从而更好地适应不同的飞行工况,进一步提高了稳像效果.最后,为了验证算法的可行性和稳定性,将视觉成像装置搭载在扑翼飞行机器人上进行了飞行实验.实验结果表明,相较于常用的电子稳像算法,本文所设计的算法在扑翼飞行机器人中表现出更好的稳像效果.最后,总结了本文所提出的算法优点,并对未来研究方向做出了展望.

翼身融合飞翼式水下滑翔机的水动力性能研究

本文依据翼身融合式飞行器的设计理念设计一种飞翼式水下滑翔机,及其主要设计参数。通过计算发现该飞翼式水下滑翔机较传统水下滑翔机拥有更大的升阻比,可达到15以上。流场分布结果显示飞翼式水下滑翔机在特定速度0.1 m/s、0.3 m/s和0.5 m/s下,机翼周围并未出现明显的流动分离,且机翼表面压力较大程度取决于攻角的大小。对比了不同雷诺数下的机翼表面涡脱落情况,发现随着攻角增大涡的脱落急剧增多,且翼梢小翼尾部涡脱落最为严重,极大影响滑翔机的水动力特性。

考虑强风干扰的固定翼飞行器“神经元”飞行气动建模

大气环境中的风是影响飞行器实际飞行的主要动态环境干扰因素之一,强风干扰下的复杂空气动力学是固定翼飞行器安全稳定飞行面临的严峻挑战。为提高强风干扰下固定翼飞行器的环境适应能力,发展了一种基于深度元学习的创新固定翼飞行器“神经元”飞行气动建模方法,该方法采用相对于地面坐标系的变量进行描述,根据多元函数的切比雪夫级数理论,将气动力和气动力矩函数分解为不同变量函数的乘积和,通过生成对抗网络技术构建强风干扰下飞行器空气动力模型的共性基函数模型,进而预测飞行器在飞行过程中受到的气动力(矩)。研究结果表明,文章建立的固定翼无人机空气动力共性基函数模型准确,可以较好地预测未知风况条件下飞行器的气动力与气动力矩,为实时空气动力学建模的迁移应用奠定良好基础。

一种悬浮扑翼机器人的视觉自主避障方法研究

提出了一种基于视觉感知的悬浮扑翼机器人自主避障方案,以满足室内障碍物环境中低速、安全和长航时飞行要求。首先,借鉴鸟类视觉避障原理,提出了一种基于光流视觉检测的避障方法,用于识别动、静态障碍物;其次,设计了一种氦气球与扑翼结合的飞行机器人,利用氦气浮力提供主要升力,扑翼运动提供推进力和次要升力,并设计了飞行控制方法以满足稳定飞行及避障的要求;最后,制作了机器人样机,测试了其飞行性能和自主避障能力,并与其他障碍物检测算法进行了对比。结果表明,样机的直飞航向偏移为5.52°,转弯速度可达23°/s,并以6.1 ms的单帧检测时间达到77%的避障成功率,证明其能够实现低速悬浮飞行和自主避障,为其开展室内探测任务奠定了基础。

苏联作家彼得·巴甫连科在民国时期的译介

彼得·巴甫连科是苏联“苏维埃年代杰出的作家”。初登文坛便引起中国作家的关注。鲁迅、郭沫若、茅盾等文学大家,均曾与之发生过文字上的交集。民国时期对巴甫连科的译介,由瞿秋白发轫,至抗战时期,形成高潮。其《复仇的火焰》《红翼东飞》《日苏未来大战记》等中文译作,既极大鼓舞了中国军民的斗志,也为中国的抗战提供了鉴照。而小说《复仇的火焰》版本甚众,其中的差异与讹误,尤待甄别。

A Study on the Wing Characteristics of Flies

This paper is concerned with the aerodynamic functions of fly wings. The free and tethered flight analyses were performed by using a digital high-speed video camera system. A liquid droplet impacting with a wing surface of fly was conducted to examine the wing characteristics. Microscopic observation of fly＇s wings were also conducted by using a laser beam microscope. The results of a series of observation and measurement revealed the flight characteristics of flies, such as the wing tip velocity, wing path, wing flexibility, wing structure, resistance to rain drops, and so forth.

Experimental Investigation on Aerodynamic Control of a Wing with Distributed Plasma Actuators

Experimental investigation of active flow control on the aerodynamic performance of a flying wing is conducted. Subsonic wind tunnel tests are performed using a model of a 35° swept flying wing with an nanosecond dielectric barrier discharge （NS-DBD） plasma actuator, which is installed symmetrically on the wing leading edge. The lift and drag coefficient, lift-to- drag ratio and pitching moment coefficient are tested by a six-component force balance for a range of angles of attack. The results indicate that a 44.5% increase in the lift coefficient, a 34.2% decrease in the drag coefficient and a 22.4% increase in the maximum lift-to-drag ratio can be achieved as compared with the baseline case. The effects of several actuation parameters are also investigated, and the results show that control efficiency demonstrates a strong dependence on actuation location and frequency. Furthermore, we highlight the use of distributed plasma actuators at the leading edge to enhance the aerodynamic performance, giving insight into the different mechanism of separation control and vortex control, which shows tremendous potential in practical flow control for a broad range of angles of attack.

Wing kinematics measurement and aerodynamics of free-flight maneuvers in drone-flies

The time courses of wing and body kinematics of two free-flying drone-flies, as they performed saccades, were measured using 3D high-speed video, and the morpho- logical parameters of the wings and body of the insects were also measured. The measured wing kinematics was used in a Navier-Stokes solver to compute the aerodynamic forces and moments acting on the insects. The main results are as following. （1） The turn is mainly a 90° change of heading. It is made in about 10 wingbeats （about 55 ms）. It is of interest to note that the number of wingbeats taken to make the turn is approximately the same as and the turning time is only a little different from that of fruitflies measured recently by the same approach, even if the weight of the droneflies is more than 100 times larger than that of the fruitflies. The long axis of body is about 40° from the horizontal during the maneuver. （2） Although the body rotation is mainly about a vertical axis, a relatively large moment around the yaw axis （axis perpendicular to the long axis of body）, called as yaw moment, is mainly needed for the turn, because moment of inertial of the body about the yaw axis is much larger than that about the long axis. （3） The yaw moment is mainly pro- duced by changes in wing angles of attack： in a right turn, for example, the dronefly lets its right wing to have a rather large angle of attack in the downstroke （generally larger than 50°） and a small one in the upstroke to start the turn, and lets its left wing to do so to stop the turn, unlike the fruitflies who generate the yaw moment mainly by changes in the stroke plane and stroke amplitude.