EXPERIMENTAL STUDY OF TURBULENT WINGBODY JUNCTION AND WAKE FLOW

The experiments on wing-body junction and wake flow were carried cul in alow-turbulence level wind tunnel. Intensive measurements of various flow parameters,such as the distributions of the pressure on the surfaces of the airfoil and the plate wall,the mean and the fluctuating velocities, as well as the turbulent kinetic energy, u'v' andu'w'etc., were performed. The results indicate that the secondary flow entrains the highvelocity low-turbulence fluid into the boundary layer and the wake vortex dominates thewake flow. Some regions of negative eddy viscosity are also found in the wake flow.

SEPARATION CONTROL FOR THE OUTWING OF A STRAKE－WING BY ROTATING CONE PLACED NEAR THE LEADING EDGE

Based on the research result on the strake-wing, when the size of a strakeis not large, there is a separation zone near the leading edge of the outwing of thestrake-wing at middle angles of attack. So the idea on separation control by rotating acone placed near the leading edge is presented. The cone surface consists of the part ofthe wing. The effect of rotating the cone on aerodynamic characteristics of thestrake-wing is investigated. The results show that a rotating surface could play an important role in controlling the flow separation for a 3-dimensional wing. For example,the relative increment in maximum lift coefficient attains 30%. The separation zone issuppressed to a certain extent.

EXPERIMENTAL RESEARCH OF THE CONFIGURATION OF WING－TIP DRAG REDUCTION FOR LIGHT AIRCRAFT

hree kinds of devices of drag reduction are presented swept wingtip,stage by stage swept wingtip and downbend wingtip. The effects of changing geometryparameters of the swept wingtip on the drag reducing are also presented. Wind-tunnelexperiment results indicate that a properly designed swept wingtip results in an incre-ment in induced efficiency of 4%～ 7% and that swept wingtip can increase longitudinalstatic-stability. Water-tunnel experiment results indicate that the reason for drag re-ducing of swept wingtip is that when the angle of attack is not zero, the strong end vor-tex of the wing is weakened by the combined effect of the leading edge and trailing edgevortices of the swept wingtip.

Structural Design and Control of Variable Camber Wing Driven by Ultrasonic Motors

A novel variable camber wing driven by ultrasonic motors is proposed.Key techniques of distributed layout of drive mechanisms,coordination control of distributed ultrasonic motors as well as novel flexible skin undergoing one-dimensional morphing are studied.The system integration of small variable camber wing is achieved.Distributed layout of parallelogram linkages driven by geared ultrasonic motors is adopted for morphing,aimed at reducing the load for each motor and producing various aerodynamic configurations suitable for different flying states.Programmable system-on-chip(PSoC)is used to realize the coordination control of the distributed ultrasonic motors.All the morphing driving systems are assembled in the interior of the wing.The wing surface is covered with a novel smooth flexible skin in order to maintain wing shape and decrease the aerodynamic drag during morphing.Wind tunnel test shows that the variable camber wing can realize morphing under low speed flight condition.Lift and drag characteristics and aerodynamic efficiency of the wing are improved.Appropriate configurations can be selected to satisfy aerodynamic requirements of different flight conditions.The study provides a practical application of piezoelectric precision driving technology in flow control.

Proportional fuzzy feed-forward architecture control validation by wind tunnel tests of a morphing wing

In aircraft wing design,engineers aim to provide the best possible aerodynamic performance under cruise flight conditions in terms of lift-to-drag ratio.Conventional control surfaces such as flaps,ailerons,variable wing sweep and spoilers are used to trim the aircraft for other flight conditions.The appearance of the morphing wing concept launched a new challenge in the area of overall wing and aircraft performance improvement during different flight segments by locally altering the flow over the aircraft＇s wings.This paper describes the development and application of a control system for an actuation mechanism integrated in a new morphing wing structure.The controlled actuation system includes four similar miniature electromechanical actuators disposed in two parallel actuation lines.The experimental model of the morphing wing is based on a full-scale portion of an aircraft wing,which is equipped with an aileron.The upper surface of the wing is a flexible one,being closed to the wing tip;the flexible skin is made of light composite materials.The four actuators are controlled in unison to change the flexible upper surface to improve the flow quality on the upper surface by delaying or advancing the transition point from laminar to turbulent regime.The actuators transform the torque into vertical forces.Their bases are fixed on the wing ribs and their top link arms are attached to supporting plates fixed onto the flexible skin with screws.The actuators push or pull the flexible skin using the necessary torque until the desired vertical displacement of each actuator is achieved.The four vertical displacements of the actuators,correlated with the new shape of the wing,are provided by a database obtained through a preliminary aerodynamic optimization for specific flight conditions.The control system is designed to control the positions of the actuators in real time in order to obtain and to maintain the desired shape of the wing for a specified flight condition.The feasibility and effectiveness of the developed control system by use of a proportional fuzzy feed-forward methodology are demonstrated experimentally through bench and wind tunnel tests of the morphing wing model.

Optimization and design of an aircraft's morphing wing-tip demonstrator for drag reduction at low speeds, Part II- Experimental validation using Infra-Red transition measurement from Wind Tunnel tests

In the present paper, an ‘in-house＇ genetic algorithm was numerically and experimentally validated. The genetic algorithm was applied to an optimization problem for improving the aerodynamic performances of an aircraft wing tip through upper surface morphing. The optimization was performed for 16 flight cases expressed in terms of various combinations of speeds, angles of attack and aileron deflections. The displacements resulted from the optimization were used during the wind tunnel tests of the wing tip demonstrator for the actuators control to change the upper surface shape of the wing. The results of the optimization of the flow behavior for the airfoil morphing upper-surface problem were validated with wind tunnel experimental transition results obtained with infra-red Thermography on the wing-tip demonstrator. The validation proved that the 2D numerical optimization using the ‘in-house＇ genetic algorithm was an appropriate tool in improving various aspects of a wing＇s aerodynamic performances.

基于NSGA-Ⅱ算法的机翼风洞模型静强度试验加载方案优化设计

针对机翼风洞模型的静强度试验,提出一种基于NSGA-Ⅱ算法试验加载方案的优化设计方法,包括加载布局的设计和试验载荷的多目标优化设计。加载布局的设计中引入控制剖面的设置,并在此基础上建立试验载荷的多目标优化设计模型,以控制剖面内的剪力、弯矩和转矩误差为优化目标,以静力等效原则和误差最小要求为约束条件,并以CHN-T1标模为例,基于NSGA-Ⅱ算法进行多目标优化模型的求解,算法采用可行性法则处理约束条件。最终求解出最优的试验加载方案并进行方案的有限元仿真验证。结果表明:试验方案与流固耦合下的理想试验的最大等效应力、应变值和变形位移的误差均小于6%,经优化设计的加载方案较好地模拟了气动载荷。

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.

Gust load alleviation wind tunnel tests of a large-aspect-ratio flexible wing with piezoelectric control

An active control technique utilizing piezoelectric actuators to alleviate gust-response loads of a large-aspect-ratio flexible wing is investigated. Piezoelectric materials have been extensively used for active vibration control of engineering structures. In this paper, piezoelectric materials further attempt to suppress the vibration of the aeroelastic wing caused by gust. The motion equation of the flexible wing with piezoelectric patches is obtained by Hamilton＇s principle with the modal approach, and then numerical gust responses are analyzed, based on which a gust load alleviation（GLA） control system is proposed. The gust load alleviation system employs classic propor tional-integral-derivative（PID） controllers which treat piezoelectric patches as control actuators and acceleration as the feedback signal. By a numerical method, the control mechanism that piezoelectric actuators can be used to alleviate gust-response loads is also analyzed qualitatively. Furthermore, through low-speed wind tunnel tests, the effectiveness of the gust load alleviation active control technology is validated. The test results agree well with the numerical results. Test results show that at a certain frequency range, the control scheme can effectively alleviate the z and x wingtip accelerations and the root bending moment of the wing to a certain extent. The control system gives satisfying gust load alleviation efficacy with the reduction rate being generally over 20%.

连接翼布局气动特性研究

在一个小型低速风洞中进行了五种不同布局形式的连接翼方案实验研究。利用油流法研究了三种连接翼的流谱,初步分析了具有连接翼飞机的气流流动机理。为比较,同时对三角翼常规布局方案进行了实验,所有方案使用相类似的隐身布局机身。实验结果表明,连接翼布局有其特有的流型:翼面分前翼、后翼及外翼三部分,其流型受前翼涡、后翼涡、翼端涡、机身边条涡以及它们互相缠绕形成的新涡的控制。这些涡的产生、发展、离体和破裂的情况不同,形成不同方案气动特性的差别。连接翼布局气动特性优于常规翼布局,特别是最大升阻比可达12以上,失速迎角超过30°。通过前后翼后缘操纵面的有利组合,可以达到提高升阻比,满足纵、横向稳定性和操纵性要求的目的。结果显示,具有扁平机身的连接翼方案是一个有潜力的无人机布局形式。

飞机偏航-滚转耦合运动非定常空气动力实验

在3m低速风洞中设计制造了一套动态实验系统,不仅能模拟飞机单自由度机动飞行运动,还能实现模型绕体轴的偏航-滚转耦合运动。通过选择运动参数,可实现绕两个轴运动角速度之间的匹配,满足飞机典型机动飞行所需的绕速度轴的无侧滑偏航-滚转运动。试验测量了BJ-1飞机模型在不同迎角下单独滚转、单独偏航和偏航-滚转耦合运动时的非定常气动特性。结果表明,飞机机动飞行时多自由度运动的气动特性比单自由度运动气动特性复杂,耦合运动时的气动特性与两个单自由度运动的气动特性的叠加结果相比有很大差别。

大型飞机高速气动力关键问题解决的技术手段及途径

大型飞机采用超临界机翼,并具有尺度大、飞行雷诺数高等特点,其研制中必须解决好高升阻比机翼、翼身组合体设计,推进系统/机体一体化设计,抖振特性、静气动弹性特性预测及超临界机翼流动控制等高速气动力问题。要解决这些关键气动力问题,必须进行一系列相关的大型高速风洞试验,以及解决相应的试验技术问题。

柔性翼微型飞行器气动特性的实验研究

柔性翼有望提高微型飞行器(MAV)的抗风能力。为进一步了解柔性翼的气动性能,建立MAV数学模型,并为飞行仿真和飞行控制设计做准备,在低雷诺数风洞中对柔性翼微型飞行器进行了风洞试验,同时采用翼型、机翼平面形状和尺寸大小均相同的刚性翼进行了风洞对比试验。对比结果表明:柔性翼相比于对应的刚性翼,失速迎角较大;柔性翼的最大升力系数较大,但是柔性翼的变形在提高升力的同时也增大了阻力,升阻比的情况较为复杂;在较低雷诺数情况下,柔性翼的纵向静稳定性略优于刚性翼;柔性翼的长周期和短周期模态的衰减特性和阻尼特性略优于刚性翼。

飞翼布局低速支撑干扰及局部外形畸变影响数值模拟

飞翼布局具有气动效率高、隐身性能好的优点,是未来军民用飞机的重要发展方向。该类布局模型风洞试验尾撑、腹撑及背撑等支撑形式的干扰量及由此带来的模型局部外形畸变影响较为复杂,目前还没有通用的支撑方案和试验修正方法。采用CFD数值模拟方法,分别对某小展弦比飞翼布局标模低速尾撑支杆干扰和尾部外形畸变影响进行了研究,结果表明:在常用角度范围内,所使用的数值模拟方法是可靠的,可用于风洞试验支撑方案的评估及支撑干扰的修正;对纵向特性,尾撑支杆干扰量和尾部外形畸变影响量相对全量较小;对横航向特性,尾撑支杆干扰量基本可忽略,尾部外形畸变影响量与全量相当。

微型扑翼飞行器升力特性研究

为了对微型扑翼飞行器升力机理进行探索性研究,本文利用西北工业大学微型飞行器专用风洞进行了扑动翼频率、风速、迎角、扑动翼弯度对其升力特性影响的研究。并且利用示波器对扑动翼扑动相位角和产生升力大小之间的关系进行了研究。试验结果表明升力特性的研究为微型扑翼飞行器总体设计和气动布局设计提供了一定的工程指导。

飞翼布局飞行器等离子体激励滚转操控试验

飞翼布局飞行器采用多个气动舵面共同作用来控制飞行,常规气动舵面的结构复杂,在大迎角时由于流动分离,舵面操纵效率显著降低。等离子体激励器具有结构简单、重量轻和响应快等优势,常被用在流动控制上。本文利用激励器抑制单侧翼面流动分离产生不对称的气动力,对飞翼布局飞行器滚转通道的控制进行了试验研究,得出了激励器在飞行器上的最优布置位置和最佳控制参数,并和常规副翼舵面滚转操控效果进行了对比。结果表明:布置于内翼、中翼前缘的等离子体激励器能够获得最佳的滚转控制效果;激励器调制频率对飞行器滚转控制效果的影响较大,而激励电压对滚转控制效果的影响较小;与常规副翼相比,等离子体激励器在大迎角时对滚转通道的操控效果优于副翼。

微型扑翼的推进特性实验研究

为了研究微型扑翼的推力和推进效率特性,研制了基于虚拟仪器(VI)的微型扑翼风洞实验系统。该系统以西北工业大学现有微型风洞为基础,采用了高精度六分量天平、角度传感器、新型机翼扑动机构、可编程电源、工控机等设备,利用基于虚拟仪器(VI)的LabView软件对上述设备进行集中控制与数据处理。利用上述实验系统研究了风速、扑动频率、展弦比、根稍比及机翼扑动幅度对微型扑翼推力和推进效率的影响,总结出了微型扑翼推力及推进效率的基本规律,为微型扑翼飞行器机翼设计及飞行控制方式提供了参考。

飞翼模型纵向气动特性等离子体流动控制试验

为了改善大展弦比飞翼模型纵向操纵性和稳定性,在低速风洞中开展了等离子体流动控制技术的试验研究.采用粒子图像测速技术获取了等离子体对翼面流场的影响.采用静态测力技术获取了等离子体对模型气动力和升降舵舵效的影响.采用虚拟飞行试验技术获取了等离子体对俯仰角和俯仰角速度时间历程的影响.通过对粒子图像测速和测力试验结果的分析表明,等离子体能够抑制翼面流动分离,阻止气动中心前移,改善模型的大迎角纵向气动特性.通过分析不同舵偏角的测力数据,来流风速V=50 m/s时等离子体能够改善飞翼模型大迎角的升降舵舵效,在不同舵偏角时均使模型的最大升力系数提高约0.1、失速迎角推迟4°以上.通过分析虚拟飞行试验结果,等离子体能够将模型的临界俯仰角提高3.6°,能够改善飞翼模型的纵向飞行稳定性和操纵性.

三种静气动弹性发散方法的一致性分析和验证

介绍了 3种静气弹发散方法 (柔度法、模态法、颤振法 ) ,并对这 3种方法的一致性从理论上进行了分析和证明。使用这 3种方法对 5个平板机翼模型的发散速度进行了计算 ,其实际发散速度也从低速风洞试验中获得。通过对比计算结果和各模型风洞试验结果 ,进一步验证了这 3种方法的一致性。另外 ,还提出了在模态法和颤振法中基于发散特征向量的结构模态选取原则。

一种柔性微型扑翼设计及其气动力特性的试验研究

微型扑翼机以其优良的机动性、低噪音、低成本、执行任务的多功能性以及在现代高科技局部战争中的潜在应用价值而得到各科技先进国家的重视,投入了大量精力开展研究工作。在我国也已经初步开展对微型扑翼机研究。但由于与微型扑翼飞行相关的低雷诺数空气动力学研究还处于萌芽阶段,对微型扑翼的设计研究工作尚无可供指导的成熟理论和借鉴的经验,目前微型扑翼的设计研究基本上是以试验为主。本文参照动物的扑翼飞行实例和固定翼飞行器机翼设计的经验,在对扑翼飞行基本原理分析研究的基础上,设计制作了一种微型扑翼及扑翼的驱动机构,采用直流电动机作为动力,通过调节驱动电压来控制扑动的频率。并对该微型扑翼进行了空气动力特性试验,得到了微型扑翼的空气动力特性,并从中得出了影响扑动升力的一些因素。实验结果表明,具有柔性翼面的扑翼,其产生的升力和推力较高而且比较稳定。