Experimental investigation of counter-rotating four vortex aircraft wake

An experimental investigation on the wake vortex formation and evolution of a four vortex system of a generic model in the near field and extended near field as well as the behaviour and decay in the far field region has been conducted by means of hot-wire anemometry in a wind tunnel. The results were obtained during an experimental campaign as part of the EC project “FAR-Wake”. The model used consists of a wing–tail plane configuration with the wing producing positive lift and the tail plane negative lift. The circulation ratio of tail plane to wing is ?0.3 and the span ratio is 0.3. Thus, a four vortex system with counter-rotating neighboured vortices exists. The model set-up was chosen on the condition to create a most promising four vortex system with respect to accelerate wake vortex decay by optimal perturbations enhancing inherent instability mechanisms. The flow field has been investigated for a half plane of the entire wake up to a distance of 48 span dimensions downstream of the model. The results obtained at 1, 12, 24 and 48 span distances are shown as non-dimensional axial vorticity and vertical turbulence intensities. A significant decay in peak vorticity, swirl velocity and circulation is observable during the downward motion of the vortices. Spectral analysis of the unsteady velocity data reveals a peak in the power spectral density distributions indicating the presence of a dominating instability. Using two hot-wire probes cross spectral density distributions have also been evaluated, which highlight the co-operative instability leading to a rapid wake vortex decay within 30 span dimensions downstream.     

The turbulent decay of trailing vortex pairs in stably stratified environments

The decay of trailing vortex pairs in thermally stably stratified environments is investigated by means of large eddy simulations. Results of in-situ measurements in the wakes of different aircraft are used to find appropriate intitializations for the simulation of wake turbulence in the quiescent atmosphere. Furthermore, cases with weak atmospheric turbulence are investigated. It is shown that the early development of the vortices is not affected by turbulence and develops almost identically as in 2D simulations of wake vortices in stably stratified environments. In a quiescent atmosphere the subsequent vortex decay is controlled by the interaction of short-wave disturbances, owing to the aircraft induced turbulence, and baroclinic vorticity, owing to stable stratification. As a consequence, vertical vorticity streaks between the vortices are induced which are substantially intensified by vortex stretching and finally lead to rapid turbulent wake-vortex decay. When in addition atmospheric turbulence is also present, the long-wave instability is dominantly promoted. For very strong stratification (Fr<1) it is observed that wake vortices may rebound but lose most of their strength before reaching the flight level. Finally, the simulation results are compared to the predictive capabilities of Greene's approximate model.     

新概念机翼尾流特性实验

大型飞机常采用开启襟翼以增大机翼升力系数,实现较大迎角的起飞和降落,而机翼在大迎角状态下,翼尖会产生能量集中且自由消散时间长的飞机尾涡,严重影响后续起降飞机的安全。基于Rayleigh-Ludwieg不稳定性,提出一种新概念飞机襟翼布局,通过水槽实验发现:新概念布局的襟翼对翼尖涡的消散具有明显的促进作用,不同参数组合下襟翼涡对翼尖涡的运动特性和能量变化的影响均有不同。实验结果也为飞机尾流控制的研究提供了参考,在满足飞行力学设计的基础上,合理运用增升装置构建四涡系统可以有效促进飞机尾流的消散,提高机场飞机起降效率。     

Excitation of instabilities in the wake of an airfoil by means of active winglets

Strong wake vortices that develop behind every aircraft as a byproduct of lift production pose a threat where aircraft fly in close staggering such as in the vicinity of airports. One approach to alleviate these vortex wakes is the use of high lift systems or control surfaces of the wing to create an unstable vortex system. The inherent instability of this vortex system shall then lead to an accelerated decay of the vortex wake, triggered for example by a periodic motion of the control surfaces. In the work presented here a simple wing model with winglets able to produce a vortex system of up to six distinct vortices is investigated in towing tank experiments. Theoretical studies show that these vortex systems potentially have a high degree of instability. By means of active oscillation of rudders integrated into the winglets, these vortex systems are to be excited to initiate an accelerated decay of the vortices. It is shown that configurations exist which exhibit strong instabilities, that lead to a significantly lower hazard level behind the vortex generating wing, even when not actively excited. However, an additional oscillation does not seem to accelerate decay of these vortex systems in relation to the statical reference case.     

Wake vortex characteristics of transport aircraft

The flow and flight physics of wake vortex systems has been intensively investigated concentrating on a large variety of aspects. This paper gives a brief overview on past and present wake vortex research activities such as early studies, integrated programs, model and flight tests, numerical investigations, fundamental physical aspects and alleviation strategies. Then, detailed results of the properties of the wake near field and extended near field are presented addressing typical length and time scales and especially turbulence quantities. Progressing from the near field to the far field wake instability mechanisms are explained along with their relevance for wake vortex decay. Characteristic quantities are given for the short and long wave instabilities associated with vortex merging and wakes consisting of two and four trailing vortices. A non-dimensional frequency parameter is introduced to classify the main instability types. Means for wake vortex alleviation are described aimed at influencing the wake vortex turbulence field or triggering and amplifying the inherent instabilities. The methods discussed include passive means such as the effects of spoilers, differential flap setting and four-vortex systems and active means using oscillating flaps or auxiliary devices.     

空中交通中尾流间隔的研究

分析了雷达尾流间隔标准的分类，通过尾流的形成机理和对后机影响的分析，建立了后机的危险遭遇模型和前机的尾流消散模型，给出了在确定的时间间隔下确定尾流危险遭遇概率的方法。     

Numerical study of aircraft wake vortex evolution near ground in stable atmospheric boundary layer

The evolutions of aircraft wake vortices near ground in stable atmospheric boundary layer are studied by Large Eddy Simulation(LES). The sensitivity of vortex evolution to the Monin-Obukhov(M-O) scale is studied for the first time. The results indicate that increasing stability leads to longer lifetimes of upwind vortices, while downwind vortices will decay faster due to a stronger crosswind shear under stable conditions. Based on these results, an empirical model of the vortex lifetime as a function of 10-m-high crosswind and the M-O scale is summarized. This model can provide an estimate of the upper boundary of the vortex lifetime according to the realtime crosswind and atmospheric stability. In addition, the lateral translation of vortices is also inspected. The results show that vortices can travel a furthest distance of 722 m in the currentlystudied parameter range. This result is meaningful to safety analysis of airports that have parallel runways.     

反弹自消散系统尾涡控制机制研究

为了研究加快尾涡消散的机理,以获得更合理的尾流间隔,进行了壁面反弹实验的数值模拟计算。实验采取在模拟机翼后缘安置反弹面的方法,诱发反弹二次涡与主涡相交出现不稳定性,以加快尾流衰减。数值计算运用N-S方程,对雷诺应力项采用Realizableκ-ε涡粘模型进行计算,分析经二次涡干涉后尾涡轴向涡量衰减、涡核下沉运动等参数的变化,以探究反弹二次涡对飞机主涡消散的演变影响。实验结果表明,反弹涡对主涡衰减有明显的促进作用。该实验为人工干预缩减机场尾流间隔的研究奠定了理论基础。     

Wake-vortex characteristics of military-type aircraft measured at airport oberpfaffenhofen using the DLR laser doppler anemometer

The DLR ground-based cw Doppler lidar was used for experimental investigations of wake vortices generated by military-type aircraft, including fixed-wing as well as rotor-wing aircraft. The measurements were mainly focused on the roll-up phase of the vortices. The experimental data has been analysed regarding vortex structure and circulation, partly also regarding their temporal development. The influence of flight parameters was measured during consecutive aircraft fly-bys. Small variations in the velocity profiles could be observed which were overlayed by the signatures of the variable wind and turbulence field. This problem can be solved on a statistical basis by measuring many cases under the same atmospheric conditions. The influence of aircraft design could be demonstrated by comparison of the vortex signatures of two B707-type aircraft showing pronounced design differences, namely a standard B707 and a AWACS E3A.     

喷流对飞机尾流涡影响的试验研究

飞机产生的尾流涡,特别是大尺度的翼尖涡,对尾随其后的飞行器是非常有害的,本文旨在探索利用飞机发动机产生的喷流加速尾流涡消亡的方法。试验采用简化的飞机模型(有尾翼),建立了包含一对翼尖涡及一对反向旋转的尾翼涡(通过以负迎角安装尾翼得到)的4涡尾流系统。在无外来扰动的情况下,不同的尾翼设置下得到的尾翼涡对翼尖涡的作用效果不同,有的能导致翼尖涡提前消亡,有的则不能。考察了不同强度的喷流对不同4涡尾流系统的影响,且作为对比,对无尾翼(2涡系统)及无喷流下的各种情况也分别作了观测。试验在拖曳水槽中进行,运用体视粒子图像测速(SPIV)技术,观测了与模型拖曳方向垂直的、从机翼后缘到下游约45翼展间均布的一系列切面。结果表明:当喷流直接作用于涡时,其效果主要取决于两者之间的初始距离及相对强度;而当喷流作用于整个4涡尾流系统时,其引入的扰动对不同的系统均能起到一定程度的改善作用,这种作用的关键在于利用喷流优化对翼尖涡进行扰动的机制,而不仅仅取决于喷流的强度。     

基于线性稳定性分析的翼尖涡摇摆机制

在涡不稳定性特征的影响下，翼尖涡会在尾迹中发生摇摆运动。为了揭示翼尖涡摇摆的本质原因以及发展机理，采用体视粒子图像测速（SPIV）技术和线性稳定性分析方法对不同雷诺数和迎角下NACA0015等直翼产生的翼尖涡在尾迹区的不稳定性特征及发展进行研究。结果表明：在1~6倍弦长的尾迹区内，翼尖涡存在摇摆现象，摇摆幅值随流向放大，且摇摆运动沿流向逐渐呈现出各向异性特征；在大迎角条件下，翼尖涡摇摆幅值随流向增长更快。采用线性稳定性分析方法，定量化分析翼尖涡的稳定性、空间/时间不稳定性放大率和扰动频率随流向的发展过程。结果显示，在雷诺数2.1×10⁵~3.5×10⁵范围内，翼尖涡均处于临界稳定状态，扰动频率为3~5 Hz。基于线性稳定性分析结果，发现在大迎角条件下翼尖涡时间/空间不稳定性放大率更大，解释了当迎角增大时翼尖涡摇摆幅值随流向增长更快的现象。另外，由线性稳定性分析得到的最不稳定模态显示翼尖涡的横向速度扰动具有明显的方向性，从而诱导翼尖涡产生摇摆运动；速度扰动方向的周期性变化则使翼尖涡摇摆区别于一维的随机振荡，而是表现为在各方向均含有分量且具有主频的摇摆运动。这种由不稳定性导致的速度扰动是翼尖涡摇摆的内在机制，其不稳定性放大率控制着摇摆幅值的增长速率，而其横向速度扰动的方向性与周期性则决定了翼尖涡的摇摆特征。     

飞机尾流的快速建模方法

尾流的强度主要由飞机的飞机重量、飞行速度和机翼形状所决定和仿真尾流的保守被动模型可很好的描述尾流系统中水蒸汽、位温等保守被动量的运动演化规律,但这种参数仿真时间长、对计算机要求高,不能实时预测任一机型所产生的尾流的状态分布特性.为此提出了一种尾流的快速建模方法,很好的解决了以往尾流实时仿真时的缺点,为飞机飞行过程中实时预测前机尾流的影响区域提供理论依据,从而减少尾流事故的发生.      

动态尾涡间隔计算工具的开发与实现

基于建立的尾涡消散模型、运动模型、遭遇模型,以Delphi 7.0为平台,采用面向对象的设计思想,开发了动态尾涡间隔计算工具,包括前机诱导尾涡流场计算、后机遭遇响应过程分析、安全间隔计算等功能模块。根据输入的飞机参数、气象数据、机场数据,实现对尾涡间隔的实时、高效计算。基于开发的计算工具,以B777-200和B737-800为前后机进行了验证计算。     

Flight safety,aircraft vortex wake and airport operation capacity

One of the major problems that challenge today's aeronautics is the problem of improving flight safety. A zone of increased hazard is the aerospace in the vicinity of an airport. Here, one of aircraft accidents' causes is wake turbulence generated by aircraft. The encountering of an aircraft on take-off or landing with the vortex wake of a preceding aircraft can lead to an accident especially dangerous near the ground.     

飞机尾涡系Rayleigh-Ludwieg不稳定性实验研究

以飞机起降过程中主翼和尾翼产生反向涡系存在相互作用的事实为背景,设计了一套反向双漩涡发生装置。通过改变两涡的位置关系与初始涡强度比值,采用流动显示与粒子成像测速(PIV)技术,对涡系相交不稳定性的作用特性进行了研究。结果表明:小涡的引入改变了主涡原有运动轨迹,合理地引入小涡的位置与小涡的强度,对主涡能量的衰减有明显的促进作用,但它们之间不呈现明显的线性关系;涡空间运动轨迹的分析,对未来完善机场起降控制模型有一定借鉴意义;实验结果也为飞机整体设计提供了一定参考依据,在满足飞行力学的设计基础上,优化整体气动布局对降低飞机尾流强度有显著的影响。     

涡破裂诱导的垂尾抖振气动弹性分析

通过试验方法分析了三角翼前缘分离涡与垂尾抖振之间的关系,深入研究了尾迹流动对垂尾抖振各阶模态的激励作用。计算得到了垂尾模型固有频率及各阶模态。在风洞试验中,应用激光片光烟流场显示技术,得到了三角翼模型在风速为30 m/s下,各迎角的涡结构;使用加速度传感器测量了垂尾翼根和翼梢的抖振响应;使用热线风速仪测量了垂尾翼根和翼梢位置的脉动速度分量。结果表明:前缘涡破裂后产生的高湍流度的尾迹是垂尾抖振的直接原因,抖振边界与涡破裂的强度和位置有关;涡破裂后尾迹与垂尾产生共振,使得抖振加速度响应频率与垂尾固有频率一致;涡破裂后,在较小迎角下,尾迹对垂尾的高频振动模态的激励较为明显,在较大迎角下,涡破裂流动对垂尾低频振动模态的激励加强了。     

Full-flight simulator study for wake vortex hazard area investigation

Simplified hazard areas (SHA) can help to overcome the wake vortex problem. Avoiding these zones allows safe and undisturbed flight operations. The definition of the hazard areas is based on the nominal required roll control power. The nominal limit for this value must be chosen carefully to ensure that outside the hazard area all wake vortex effects are noncritical. For the determination of such a limit offline simulations and full flight simulator studies were conducted. For a given roll control limit the hazard area dimensions are established with the “simplified hazard area prediction” method (SHAPe). Parameterization of the input parameters makes this method universally applicable to different aircraft types. SHAPe represents an essential element of the wake vortex prediction and observation system within the DLR project Wirbelschleppe II, for the generation of safe dynamic wake vortex separations.     

鸭式布局机翼的数值计算

本文将计算单独机翼的非线性离散涡法和集中涡核法以及二者的结合发展应用于鸭式布局机翼的气动特性计算。在集涡核的计算模型中除了对机翼附涡系和自由尾涡的模拟外，还包括了对机翼前缘涡的自由涡面、集中涡核及涡核卷吸作用的模拟。本文给出了三种不同形状鸭式布局的算例，结果表明，本方法的计算结果与实验结果和其他计算结果吻合得很好或接近。