ASYMMETRIC VORTICES FLOW OVER SLENDER BODY AND ITS ACTIVE CONTROL AT HIGH ANGLE OF ATTACK

The studies of asymmetric vortices flow over slender body and its active control at high angles of attack have significant importance for both academic field and engineering area.This paper attempts to provide an update state of art to the investigations on the fields of forebody asymmetric vortices.This review emphasizes the correlation between micro-perturbation on the model nose and its response and evolution behaviors of the asymmetric vortices.The critical issues are discussed, which include the formation and evolution mechanism of asymmetric multi-vortices;main behaviors of asymmetric vortices flow including its deterministic feature and vortices flow structure;the evolution and development of asymmetric vortices under the perturbation on the model nose;forebody vortex active control especially discussed micro-perturbation active control concept and technique in more detail.However present understanding in this area is still very limited and this paper tries to identify the key unknown problems in the concluding remarks.

Control of Asymmetric Flow Fields of Slender Bodies at High Angle of Attack

The wind tunnel experiments is conducted to get inspiration for understanding the mechanism of the asymmetric flow pattern and developing an innovative flow control technique for a slender body at high angle of attack. The bi-stable situation of the side forces is observed, which could be easily switched by a tiny disturbances either from coming flow or from artificial disturbances at nose tip （including manufacturing defect）. In turbulent flows the side forces switched randomly between positive and negative. There exists a hysteresis loop of side force with the rolling angle. A rod in front of the slender body is used to change the vortex pattern, which could be kept even the rod is moved out from the stream. A miniature strake attached to the nose tip of the model can be moved to different circumferential position. When the strake is stationary, the hysteresis loop disappears and the side force does not change with the turbulent fluctuation of coming flow. The results from dynamic measurements of section side force indicates that when the strake swung at lower frequency the side force can follow the cadence of the swinging strake. With increasing frequency, the magnitude of the side force decreases. At still high frequency, the side force diminishes to zero. If the strake is swinging, while the middle position can be changed to different circumferential angle Фs on either left or right side, the side forces can be changed proportionally with the angle Фs. On the basis of the experimental results, the mechanism of the asymmetry is discussed.

COMPUTATION OF FIELD STRUCTURE AND AERODYNAMIC CHARACTERISTICS OF DELTA WINGS AT HIGH ANGLES OF ATTACK

A numerical investigation of the structure of the vortical flowfield over delta wings at high angles of attack in longitudinal and with small sideslip angle is presented. Three-dimensional Navier-Stokes numerical simulations were carried out to predict the complex leeward-side flowfield characteristics that are dominated by the effect of the breakdown of the leading-edge vortices. The methods that analyze the flowfield structure quantitatively were given by using flowfield data from the computational results. In the region before the vortex breakdown, the vortex axes are approximated as being straight line. As the angle of attack increases, the vortex axes are closer to the root chord, and farther away from the wing surface. Along the vortex axes, as the adverse pressure gradients occur, the axial velocity decreases, that is, A is negativee, so the vortex is unstable, and it is possible to breakdown. The occurrence of the breakdown results in the instability of lateral motion for a delta wing, and the lateral moment diverges after a small perturbation occurs at high angles of attack. However, after a critical angle of attack is reached the vortices breakdown completely at the wing apex, and the instability resulting from the vortex breakdown disappears.

平均载荷系数0.5一级的双模式压气机设计研究与验证

为全面提升压气机的研制能力,本文对平均载荷系数0.5一级的带单双外涵两种模式(双模式)核心驱动风扇的四级压气机开展了设计技术研究与验证。针对其平均级载荷系数高、两个模式性能变化大等特点,开展了双模式高负荷压气机设计、低损失大流量范围导叶设计以及低损失大攻角范围可调静子设计等技术研究工作,并在此基础上完成双模式核心驱动风扇与四级压气机的设计,开展0.8~1.0相对转速的三维特性分析,全面研究单双外涵两种模式下各级的匹配情况。试验结果表明:单双外涵两种模式下,压气机流量、压比、效率达到设计指标。单外涵模式1.0相对转速,最高效率0.874,双外涵模式0.924相对转速,最高效率0.87,两种模式下压气机流量调节范围达到31.4%。各级匹配良好,突破了双模式核心驱动风扇与压气机一体化匹配设计和平均载荷系数0.5一级的高负荷压缩部件设计等关键技术,为下一代发动机的核心压缩部件设计奠定了基础。

不同入水攻角下高速射弹的流固耦合特性

高速入水时弹体结构响应(变形)大小关系射弹能否安全入水。当前,基于刚体模型的高速入水数值计算方法无法深入揭示初始扰动影响下射弹入水过程中复杂多相流、水动力与弹体结构响应等之间的相互耦合作用规律。为解决上述难题,基于流体力学和结构动力学,建立一种高速入水流固耦合数值计算方法,重点研究攻角对某型射弹高速入水过程的影响,分析攻角影响下空泡演化、冲击载荷、弹体运动与结构变形的相互耦合作用机理。研究结果表明:随攻角增大,弹体下表面与空泡壁面逐渐产生强烈撞击,迫使空泡壁面弯曲;对该型射弹,当攻角>3°时,射弹会因尾翼拍击液面出现第2次载荷峰值,此时射弹因沾湿产生弯矩,迫使射弹出现塑性应变,随入水深度增加,射弹形成明显的尾翼弯折,因沾湿产生的表面压力最大值超过10 MPa;攻角为2°时,射弹虽因弹头尾翼同时受力产生弯矩,但内部应力未超过弹体材料的屈服强度,未出现塑性变形,因此,对该型射弹,其安全入水攻角≤2°。

高速条件下细长旋成体背风区流动特性试验研究

导弹全方向攻击时会产生由旋涡主导的复杂流动,发生转捩和流动分离,对导弹的机动性和控制能力具有重大影响。为探索攻角及马赫数对弹体流动形态的影响,针对某细长旋成体在FD-12风洞开展高速粒子图像测速及荧光油流试验,来流Ma分别为0.4、0.6、0.8,攻角范围α为0°~180°,获取了细长体背风区的流场特性演化规律。研究结果表明:当攻角α<90°时,随着攻角的增大,模型背风区流场从附着流变化为分离流,且分离涡从对称变为非对称,最终演化为非定常流动;当攻角α>90°时,情况有所不同,α=100°及α=120°时背风区存在非定常涡脱落现象,时均流场具有明显的非对称性;当攻角α=150°时,迎风面分离区出现不对称偏移,初始分离区下边界已越过模型端面,在模型中后部形成分离线;当攻角α=180°时,时均流场没有明显的特征,在模型头部位置出现了环形分离区及再附区,表明由于底部扰动的影响,该截面流场呈现局部的非定常、非线性流动状态;在攻角相同情况下,增大流场的来流Ma,不但会使分离涡的影响范围变大,也会导致分离涡的位置抬高。

4 m×3 m风洞大迎角机构上位机软件

针对大迎角机构机械装置、控制硬件的改造升级,基于Labview框架平台设计编写上位机控制软件。优化更新基于TCP/IP协议的Socket通信,增设虚拟支杆设定、模型防碰撞识别等功能。编制的控制软件已成功应用于气动中心4 m×3 m低速风洞大迎角试验。结果表明,该软件有效提升了试验质量效率和设备运行安全。

面向战机大迎角机动过程的智能学习控制

针对战机大迎角动力学呈现的强非线性、气动不确定和通道耦合特性,提出了一种基于智能学习的自适应机动跟踪控制方法.通过将通道耦合视为集总扰动的一部分,把模型分解为迎角子系统、侧滑角子系统和滚转角速率子系统.采用神经网络估计不确定,设计跟踪误差反馈与集总干扰估计前馈相结合的控制器获取期望操纵力矩,并基于串接链分配方法求解气动舵偏角和推力矢量偏角.对于神经网络权重更新,构建预测误差表征集总干扰的估计性能,结合跟踪误差设计复合学习更新律.基于李雅普诺夫方法证明了闭环系统的一致最终有界稳定性.针对眼镜蛇机动和赫伯斯特机动指令进行了仿真验证和抗干扰参数拉偏测试,结果表明所提方法具有较高的机动指令跟踪精度和鲁棒性能.

基于Transformer的飞机状态预测

在非定常气动力下,为防止飞机进入危险状态,通过建模进行状态预测,是保障飞行安全的重要手段,传统方法建模过程复杂、工程化难度大且普适性不强。为更好解决大迎角下飞行状态预测,使用基于深度学习的时序序列预测方法,推测飞机的飞行状态,达到最大限度发掘飞机性能、保障飞行安全的目的。提出一种多任务Transformer模型,同时完成飞行状态参数回归和飞行状态分类。实验结果表明,相比于同类模型,该模型的预测性能有大幅提升。

横风环境下高速列车气动力模拟与分析

为探究不同风攻角、风向角下高速列车所受气动力特性,通过数值模拟的方法对高速列车所在风场进行研究。采用时间序列法模拟得到列车移动点的脉动风速时程,根据准定常假设下风速与气动力之间的关系计算得到列车所受静风力和抖振风力,分析了不同参数对计算结果的影响。结果表明,在同一模拟点下,相较于顺风向脉动风,竖向脉动风的脉动量更小;同一风攻角,风向角越大,静风力越小;同一风向角,风攻角越接近3°,静横向力越小;风攻角越接近0°,静升力越小;随着风攻角由-3°过渡到3°,同一车速下,脉动横向力逐渐减小,而脉动升力先减小后增大;同一风攻角下,随着风向角增大,脉动横向力标准差先增大后减小;脉动升力除风攻角为-3°且呈现逐渐增大的趋势外,其余两种均为先减小后增大;气动力的离散程度即变异系数与风攻角关系不大,主要与风向角有关,风向角越大变异系数越大。

升力翼列车通过隧道的气动效应研究

升力翼列车是一种通过增加升力翼来提升气动升力的新概念列车,可等效降低自身重力,有效减少轮轨磨损。本研究基于RNG k−ε湍流模型,采用滑移网格模拟方法,研究了不同攻角升力翼列车通过隧道的气动效应,并通过动模型实验数据对数值计算方法的精度进行验证。研究结果表明:升力翼列车进入隧道后列车升力增大,相较于明线,隧道内平均升力增加了33.3%;在进入隧道时攻角由12.5°减小为7.5°,可以较好地减小进入隧道时的升力波动,同时也可减小列车和隧道表面的压力峰值,有利于列车平稳通过隧道。通过对比有、无升力翼的列车可发现,车体前端主要受到升力翼增加车隧阻塞比的影响,而压力上升;车体后端主要受到升力翼尾流的影响,压力降低。本研究结果可为升力翼列车平稳通过隧道提供技术支持。

高航速流动不分离外形带攻角流动特性分析

边界层发生分离的两个必要条件是流体粘性和正压梯度,设计外形使其表面具有尽量大的负压梯度区域,延缓边界层转捩和流动分离,从而达到减阻目的。本文采用基于细长体理论的流动不分离外形设计方法,设计临界速度为100 m/s的航行体外形,运用数值仿真分析其在不同速度和攻角情况下的流动特性。研究发现,0°攻角下仿真得到的压强分布与理论计算结果一致,摩擦阻力与估算结果一致,证明可以通过外形设计使航行体边界层在高雷诺数条件下保持层流流动不分离状态,以达到大幅减阻的目的;小攻角不会破坏航行体表面流体附着状态,但攻角会使流动出现横向漩涡,边界层发生转捩,使阻力明显增加,但减阻效果仍然存在。

A physical model of asymmetric vortices flow structure in regular state over slender body at high angle of attack

In subcritical Reyolds number flow region, the repeatable and determinate asymmetric vortices flow at regular state can be obtained by manually setting mini-perturbation on the nose of a pointed ogive-cylinder model at high angle of attack and zero side slip. Test results of this study involve surface pressure distributions, sectional-side-force distributions and flow visualizations. The analyses of these results revealed a complicated multi-vortices system at regular state inwhich asymmetric twin vortices with inception region and fully developed region, asymmetric triple vortices, four vortices region, five vortices region and Karman-vortex-street-like flow region aredeveloped along the slender body. The correlation between multi-vortices system structures and corresponding sectional-side-force distribution is given. The behaviors of multi-vortices flow struc-ture at the peculiar points of sectional-side-force distributions and characteristics of corresponding pressure distribution are analysed. Finally, a physical model of asymmetric multi-vortices flowstructure in regular state over slender body is developed.

Unsteady aerodynamic modeling at high angles of attack using support vector machines

Abstract Accurate aerodynamic models are the basis of flight simulation and control law design. Mathematically modeling unsteady aerodynamics at high angles of attack bears great difficulties in model structure determination and parameter estimation due to little understanding of the flow mechanism. Support vector machines （SVMs） based on statistical learning theory provide a novel tool for nonlinear system modeling. The work presented here examines the feasibility of applying SVMs to high angle.-of-attack unsteady aerodynamic modeling field. Mainly, after a review of SVMs, several issues associated with unsteady aerodynamic modeling by use of SVMs are discussed in detail, such as sele, ction of input variables, selection of output variables and determination of SVM parameters. The least squares SVM （LS-SVM） models are set up from certain dynamic wind tunnel test data of a delta wing and an aircraft configuration, and then used to predict the aerodynamic responses in other tests. The predictions are in good agreement with the test data, which indicates the satisfving learning and generalization performance of LS-SVMs.

Study on flow behavior and structure over chined fuselage at high angle of attack

A study of leeward vortex structure over chined fuselage and the effects of micro tip perturbation on its vortex flow have been carried out in wind tunnel experiments at Reynolds numbers from 1.26×105 to 5.04×105 with PIV and pressure measurement techniques.Firstly,the experiment results have proved that micro tip perturbation has no effects on the vortex flow and its aerodynamic characteristics over chined fuselage at high angle of attack,in which there are not any non-deterministic flow behaviors.Secondly,the evolution of leeward vortex structure over chined fuselage along the axis of model can be divided into four flow regimes:linear conical developed regime,decay regime of leeward vortex intensity,asymmetric leeward vortex break down regime and completely break down regime.And a correlation between leeward vortex structure and sectional aerodynamic force was also revealed in the present paper.Thirdly,the experiment results show the behavior of leeward vortex core trajectories and zonal characteristics of leeward vortex structure with angles of attack.Finally,the experiment results of Reynolds number effect on the leeward vortex flow have further confirmed research conclusions from previous studies:the flows over chined fuselage at high angles of attack are insensitive to variation of Reynolds number,and there is a little effect on the secondary boundary layer separation and the suction peak induced by leeward vortex.

HB-2 high-velocity correlation model at high angles of attack in supersonic wind tunnel tests

Responding to a need for experimental data on a standard wind tunnel model at high angles of attack in the supersonic speed range, and in the absence of suitable reference data, a series of tests of two HB-2 standard models of different sizes was performed in the T-38 trisonic wind tunnel of Vojnotehnickˇi Institut(VTI), in the Mach number range 1.5–4.0, at angles of attack up to+30°. Tests were performed at relatively high Reynolds numbers of 2.2 millions to 4.5 millions(based on model forebody diameter). Results were compared with available low angle of attack data from other facilities, and, as a good agreement was found, it was assumed that, by implication, the obtained high angle of attack results were valid as well. Therefore, the results can be used as a reference database for the HB-2 model at high angles of attack in the supersonic speed range, which was not available before. The results are presented in comparison with available reference data, but also contain data for some Mach numbers not given in other publications.

Experimental investigation of influence of strake wings on self-induced roll motion at high angles of attack

The modern high performance air vehicles are required to have extreme maneuverability,which includes the ability of controlled maneuvers at high angle of attack. However, the nonlinear and unsteady aerodynamic phenomena, such as flow separation, vortices interaction, and vortices breaking down, will occur during the flight at high angle of attack, which could induce the uncommanded motions for the air vehicles. For the high maneuverable and agile air missile, the nonlinear roll motions would occur at the high angle of attack. The present work is focused on the selfinduced nonlinear roll motion for a missile configuration and discusses the influence of the strake wings on the roll motion according to the results from free-to-roll test and PIV measurement using the models assembled with different strake wings at a = 60°. The free-to-roll results show that the model with whole strake wings(baseline), the model assembled with three strake wings(Case A)and the model assembled with two opposite strake wings(Case C) experience the spinning, while the model assembled with two adjacent strake wings(Case B), the model assembled with one strake wing(Case D) and the model with no strake wing(Case E) trim or slightly vibrate at a certain "×"rolling angle, which mean that the rolling stability can be improved by dismantling certain strake wings. The flow field results from PIV measurement show that the leeward asymmetric vortices are induced by the windward strake wings. The vortices would interact the strake wings and induce crossflow on the downstream fins to degrade the rolling stability of the model. This could be the main reason for the self-induced roll motion of the model at a = 60°.

High AOA decoupling control for aircraft based on ADRC

In this paper, a practical decoupling control scheme for fighter aircraft is proposed to achieve high angle of attack(AOA)tracking and super maneuver action by utilizing the thrust vector technology. Firstly, a six degree-of-freedom(DOF) nonlinear model with 12 variables is given. Due to low sufficiency of the aerodynamic actuators at high AOA, a thrust vector model with rotatable engine nozzles is derived. Secondly, the active disturbance rejection control(ADRC) is used to realize a three-channel decoupling control such that a strong coupling between different channels can be treated as total disturbance, which is estimated by the designed extended state observer. The control surface allocation is implemented by the traditional daisy chain method. Finally,the effectiveness of the presented control strategy is demonstrated by some numerical simulation results.

Bifurcation analysis of polynomial models for longitudinal motion at high angle of attack

To investigate the longitudinal motion stability of aircraft maneuvers conveniently, a new stability analysis approach is presented in this paper. Based on describing longitudinal aerodynamics at high angle-of-attack （a ＆lt; 50 ） motion by polynomials, a union structure of two-order differential equation is suggested. By means of nonlinear theory and method, analytical and global bifurcation analyses of the polynomial differential systems are provided for the study of the nonlinear phenomena of high angle-of-attack flight. Applying the theories of bifurcations, many kinds of bifurcations, such as equilibrium, Hopf, homoclinic （heteroclinic） orbit and double limit cycle bifurcations are discussed and the existence conditions for these bifurcations as well as formulas for calculating bifurcation curves are derived. The bifurcation curves divide the parameter plane into several regions; moreover, the complete bifurcation diagrams and phase portraits in different regions are obtained. Finally, our conclusions are applied to analyzing the stability and bifurcations of a practical example of a high angle-of-attack flight as well as the effects of elevator deflection on the asymptotic stability regions of equilibrium. The model and analytical methods presented in this paper can be used to study the nonlinear flight dynamic of longitudinal stall at high angle of attack.

Time delay compensation in lateral-directional flight control systems at high angles of attack

The previous studies of time delay compensation in flight control systems are all based on the conventional aerodynamic derivative model and conducted in longitudinal motions at low angles of attack.In this investigation,the effects of time delay on the lateral-directional stability augmentation system in high-a regime are discussed based on theβmodel,which is proposed in our previous work and proved as a more accurate aerodynamic model to reveal the lateraldirectional unsteady aerodynamic characteristics at high angles of attack.Both theβmodel and the quasi-steady model are used for simulating the effects of time delay on the flying qualities in high-a maneuvers.The comparison between the simulation results shows that the flying qualities are much more sensitive to the mismatch of feedback gains than the state errors caused by time delay.Then a typical adaptive controller based on the conventional dynamic derivative model and a gain-prediction compensator based onβmodel are designed to address the time delay in different maneuvers.The simulation results show that the gain-prediction compensator is much simpler and more efficient at high angles of attack.Finally,the gain-prediction compensator is combined with a linearizedβmodel reference adaptive controller to compensate the adverse effects of very large time delay,which exhibits excellent performance when addressing the extreme conditions at high angles of attack.