UAV flight test of plasma slats and ailerons with microsecond dielectric barrier discharge

Plasma flow control（PFC） is a promising active flow control method with its unique advantages including the absence of moving components, fast response, easy implementation, and stable operation. The effectiveness of plasma flow control by microsecond dielectric barrier discharge（μs-DBD）, and by nanosecond dielectric barrier discharge（NS-DBD） are compared through the wind tunnel tests, showing a similar performance between μs-DBD and NS-DBD. Furthermore, theμs-DBD is implemented on an unmanned aerial vehicle（UAV）, which is a scaled model of a newly developed amphibious plane. The wingspan of the model is 2.87 m, and the airspeed is no less than 30 m/s. The flight data, static pressure data,and Tufts images are recorded and analyzed in detail. Results of the flight test show that the μs-DBD works well on board without affecting the normal operation of the UAV model. When the actuators are turned on, the stall angle and maximum lift coefficient can be improved by 1.3° and 10.4%, and the static pressure at the leading edge of the wing can be reduced effectively in a proper range of angle of attack, which shows the ability of μs-DBD to act as plasma slats. The rolling moment produced by left-side μs-DBD actuation is greater than that produced by the maximum deflection of ailerons,which indicates the potential of μs-DBD to act as plasma ailerons. The results verify the feasibility and efficacy of μs-DBD plasma flow control in a real flight and lay the foundation for the full-sized airplane application.

Fault diagnosis and accommodation for linear parameter-varying in flight control systems

The fault diagnosis and accommodation strategy for a class of linear parameter-varying (LPV) systems were investigated. A fast adaptive fault estimation (FAFE) algorithm for LPV systems module, based on an adaptive observer, proposed to enhance the performance of fault estimation including rapidity and accuracy. Then, the obtained fault estimate was used to construct the fault tolerant control (FTC) law. The design method was formulated as a convex linear matrix inequalities (LMIs) optimization problem. Once the faults are estimated, the fault tolerant controller is implemented as a dynamic output feedback controller. This controller can compensate for the effect of the faults by stabilizing the closed-loop systems. Finally, a helicopter model in a vertical flight with actuator fault was used to the effectiveness of the proposed approach.

Microjet-Based Active Flow Control on a Fixed Wing UAV

A fixed wing, remote controlled, unmanned aerial vehicle (UAV), has been retrofitted with a system of microjet-based actuators to test microjet efficacy for flow control during flight. This allows one to evaluate the flow control system in a more complex and uncontrolled environment relative to the “clean” laboratory flowfield. The system is composed of off-the-shelf components and provides some appreciation of the challenges associated with implementing such an active control scheme in more “practical” configurations. A wing section with actuators was first tested in a low-speed wind tunnel to characterize microjet control efficacy in the laboratory using surface flow visualizations and particle image velocimetry (PIV). The laboratory results show that the microjet actuators are an effective means of controlling separation with fairly low supply pressures and flow rates. Only relatively robust and straightforward diagnostics can be used to determine the flow conditions on the UAV during flight. As such, tufts are installed on the aircraft’s wing to serve as a qualitative way of measuring control efficacy. Results from the flight tests confirm that this flow control system is capable of delaying flow separation in the complex flows occurring during flight.

基于符合性证据链的航电系统工程试验方法

为形成逻辑严密、可追溯的适航符合性验证资料,提出一种基于符合性证据链的航电系统工程试验方法,并对实现中的关键问题进行分析研究。以自动飞行控制系统的工程试验为例,设计与实现了基于自动飞行控制系统适航验证需求的验证平台。在验证方案设计中,针对复杂验证环境中部分参数不确定性描述问题,提出一种基于分类概率多场景分析的验证方案设计方法,基于适航验证需求,对飞行数据进行筛选,通过对筛选后的飞行数据进行统计特性分析、随机抽样与合并缩减,生成包含发生概率的确定性场景,描述验证环境中部分参数的不确定性。在验证数据分析中,针对多场景多参数条件下的适航符合性判断问题,提出基于加权Dempster-Shafer证据理论的适航符合性评估方法,以确定性场景发生概率为权重,进行证据融合,避免部分小概率场景对融合结果的干扰。基于实际飞行数据,自动飞行控制系统的自动飞行模式工程试验结果表明,所提方法有效可行。

星载高精度偏振扫描仪热控设计与验证

偏振扫描仪安装于卫星对地板,周围有其他载荷的遮挡,热环境复杂,是卫星热控设计的难点。文章根据扫描仪不同部组件的控温要求设计热控方案,包括:主光学组件与框架隔热并包覆多层隔热组件;红外探测器热沉、头部电路盒分别设计独立的散热路径,通过外贴热管连接辅助散热板散热;结合电加热主动控温;在主要部位安装隔热垫。该热控方案已通过热平衡试验在地面的考核验证,在轨遥测数据显示:主光学组件温度控制在(16±1)℃,头部电路盒温度控制在6~8℃,红外探测器热沉温度控制在(-22±1)℃,满足扫描仪各项温度指标要求。以上表明该热控方案合理有效,对同类型光学载荷的热控设计及优化具有借鉴意义。

受控制律影响的电传运输机飞行品质试飞技术

针对电传运输机控制律设计特点,该文分析保护功能、自动配平功能等飞控功能对飞行品质试飞科目的影响,梳理受影响的关键试飞科目,并根据影响情况,适应性提出受控制律影响的运输机失速特性、抖振边界、空中最小操纵速度、高速特性等科目改进的试飞方法,总结试飞过程中的问题和经验,突破受电传控制律影响的飞机飞行品质关键科目试飞技术,为后续电传运输机飞行品质试飞提供参考。

基于动力学特性优化的直升机航向操纵系统减振效果验证

基于结构动力学基本原理及其影响因素,对直升机在飞行过程中出现的航向操纵拉杆支座断裂以及后舱脚蹬振动较大的问题,通过航向操纵系统增加配重和部分拉杆更换为球柔钢索这两种方式,优化航向操纵系统的安装动特性。试飞验证结果表明,两种方式均能有效降低拉杆振动,进而明显改善支座受载和脚蹬处的振动。

MBSE在航空飞行控制系统的应用研究

现代飞机系统的复杂性不断增长,传统的开发方法在管理和维护方面变得越来越具有挑战性。首先,对航空飞控系统研发过程在当前系统工程实践方面进行了评估,阐述了传统开发方式的不足和基于模型的系统工程(MBSE)的核心原则和优点,以及在实现中常用的工具和方法;然后,展示了MBSE如何在飞行控制系统研发时使用模型来支持复杂系统的规范、设计、分析、验证和确认的过程;最后,结合先进的MBSE工具和方法的发展,MBSE结合敏捷的开发前景,以及提高MBSE实践的互操作性和标准化的需求对未来的研究方向进行了展望。本文研究为利用MBSE方法论和工具来改进飞行控制系统的设计、开发和性能感兴趣的航空工业专业人员、系统工程师提供资源和解决思路。

主动噪声控制技术在涡扇飞机飞行试验中的应用研究

为研究主动噪声控制技术在涡扇飞机真实飞行环境下的作用效果,基于自适应滤波算法设计了一套主动降噪系统,并在与航线飞机舱内环境完全相同的试飞机上集成安装,挑选航线爬升和不同巡航高度典型飞行试验工况开展试飞试验,通过对试飞数据的分析结果表明:在飞机复杂的飞行环境及舱内声场环境下,主动降噪系统在目标降噪区域仍有明显的降噪效果,且降噪效果与基础噪声有较大的关系,通常是往空间噪声声压级趋于一致的方向调整,基础噪声越大,降噪效果越好;且对低频的噪声控制效果更为稳定,效果更好;并对主动降噪系统在涡扇飞机上的布局方案形成指导建议。

Initial virtual flight test for a dynamically similar aircraft model with control augmentation system

To satisfy the validation requirements of flight control law for advanced aircraft,a wind tunnel based virtual flight testing has been implemented in a low speed wind tunnel.A 3-degree-offreedom gimbal,ventrally installed in the model,was used in conjunction with an actively controlled dynamically similar model of aircraft,which was equipped with the inertial measurement unit,attitude and heading reference system,embedded computer and servo-actuators.The model,which could be rotated around its center of gravity freely by the aerodynamic moments,together with the flow field,operator and real time control system made up the closed-loop testing circuit.The model is statically unstable in longitudinal direction,and it can fly stably in wind tunnel with the function of control augmentation of the flight control laws.The experimental results indicate that the model responds well to the operator＇s instructions.The response of the model in the tests shows reasonable agreement with the simulation results.The difference of response of angle of attack is less than 0.5°.The effect of stability augmentation and attitude control law was validated in the test,meanwhile the feasibility of virtual flight test technique treated as preliminary evaluation tool for advanced flight vehicle configuration research was also verified.

Two freedom linear parameter varying μ synthesis control for flight environment testbed

To solve the problem of robust servo performance of Flight Environment Testbed(FET)of Altitude Ground Test Facilities(AGTF) over the whole operational envelope, a two-degree-offreedom μ synthesis method based on Linear Parameter Varying(LPV) schematic is proposed, and meanwhile a new structure frame of μ synthesis control on two degrees of freedom with double integral and weighting functions is presented, which constitutes a core support part of the paper. Aimed at the problem of reference command's rapid change, one freedom feed forward is adopted, while another freedom output feedback is used to meet good servo tracking as well as disturbance and noise rejection; furthermore, to overcome the overshoot problem and acquire dynamic tuning,the integral is introduced in inner loop, and another integral controller is used in outer loop in order to guarantee steady errors; additionally, two performance weighting functions are designed to achieve robust specialty and control energy limit considering the uncertainties in system. As the schedule parameters change over large flight envelope, the stability of closed-loop LPV system is proved using Lyapunov inequalities. The simulation results show that the relative tracking errors of temperature and pressure are less than 0.5% with LPV μ synthesis controller. Meanwhile, compared with non-LPV μ synthesis controller in large uncertainty range, the proposed approach in this research can ensure robust servo performance of FET over the whole operational envelope.

Virtual flight test of pitch and roll attitude control based on circulation control of tailless flying wing aircraft without rudders

Circulation Control(CC) realizes rudderless flight control by driving compressed air jet to generate a virtual rudder surface, which significantly improves low detectability. The layout plan of combined control rudder surface is proposed based on the tailless flying wing aircraft. The closed-loop jet actuator system and stepless rudder surface switching control strategy are used to quantitatively study the control characteristics of circulation actuator for pitch and roll attitude through 3-DOF virtual flight test in a wind tunnel with a powered model at wind speed of 40 m/s. The results show that the combined use of circulation actuators can achieve bidirectional continuous and stable control of the aircraft’s pitch and roll attitude, with the maximum pitch rate of 12.3(°)/s and the maximum roll rate of 21.5(°)/s;the response time of attitude angular rate varying with the jet pressure ratio is less than 0.02 s, which can satisfy the control response requirements of aircraft motion stability for the control system;the jet rudder surface has a strong moment control ability, and the pitch moment of the jet elevator with a pressure ratio of 1.28 is the same as that of the mechanical elevator with 28° rudder deflection, which can expand the flight control boundary.

一种无人机飞行控制软件自动测试方法

针对机载飞行控制软件测试依赖硬件系统、测试周期长、手动测试方式效率低、重用性差、易出错和维护成本高等问题,研究了全数字仿真环境的关键技术,提出了一种机载飞行控制软件的自动测试方法;该测试方法在全数字仿真环境中进行测试,并对传统的测试数据生成算法进行改进,使用AETG-SA算法生成测试数据,将测试结果的反馈引入到算法中,动态调整算法参数,获得最优测试集合,提高了测试覆盖率;在全数字仿真环境中执行测试用例,减少了嵌入式软件测试过程中对硬件的依赖,对系统功能测试和故障模拟测试的覆盖更加全面;工程实践表明,基于全数字仿真环境的自动测试方法相较于传统全实物和半实物的测试方法测试充分性提升了4%,测试时间缩短了44%。

人工神经网络在飞机下沉速度控制中的应用

下沉速度是指飞机着陆接地时刻其重心的垂向速度,它直接关系到飞机起落装置在接地时所受到的载荷大小。在考核飞机起落架强度、刚度的着陆试验中,国军标对飞机的下沉速度有具体的要求,但在实际操作中,由于各种外部因素的干扰,飞行员很难利用现有的手段精准操纵飞机达到标准要求的下沉速度。本文通过合理分析,选取了几个影响陆基飞机下降速度的飞行参数,并将这些参数的实测飞行数据作为MATLAB人工神经网络模型的输入与输出,得到了较好的预测结果,探索了控制陆基飞机下沉速度的新思路与新方法。

基于STM32F4与FreeRTOS的无人机飞行控制系统

飞行控制系统多以高性能CPU为处理器和Vxworks为设计内核,存在成本高、体积大、内核源码不公开等问题,因此设计一款低成本、高可靠性的小型无人机飞行控制系统。基于STM32F4芯片开发一套由微型控制器、各种传感器和执行机构构成的飞行控制硬件系统,实现对无人机的姿态和舵机位置的控制;引入源码公开的FreeRTOS嵌入式实时系统作为软件平台开发飞行控制软件,实现飞行控制、自主导航、飞行管理等功能;通过固定翼航模的飞行试验对该飞行控制系统进行验证。结果表明:该系统具有良好的稳定性、实时性,控制精度高。

自适应增广控制频域分析与飞行试验

本文聚焦运载火箭自适应增广控制频域分析问题。首先,通过对控制器结构的简化,降低其非线性,同时构造单输入单输出环节,便于频域分析;之后,利用描述函数和常微分方程近似解析解的思想,对自适应增广控制进行了频域分析,获得了频域模型;最后,基于可重复使用验证飞行器对自适应增广控制器进行了飞行测试。飞行结果显示频域模型正确且精度较高,能够获得更好的控制效果。

某大型涡桨飞机副翼振动与控制研究

根据适航规章要求,研究了飞机操纵面振动设计要求和方法,论述了飞机操纵面作动器动刚度和自由间隙极限环振荡计算方法。针对某大型涡桨飞机在试飞中发现的副翼振动问题,论述了副翼异常振动产生的原因,建立了结构动力学有限元模型,开展副翼操纵刚度、阻尼变参的频率响应分析,研究副翼的振动特性。提出了副翼助力器活塞打孔增加阻尼的临时改进方法,及根据操纵面振动设计、自由间隙控制要求的副翼结构和系统优化方法。副翼改进设计后,试飞数据分析表明,副翼的振动响应幅值降幅明显,振动收敛性较好,保障了型号试飞的继续开展。

自动飞行系统飞行模式操作验证方法研究

现代民用飞机的自动飞行控制系统是主要的人机交互系统之一,其飞行模式操作控制功能由两大物理部分组成:FMCP硬件与自动飞行控制软件。对于复杂的人机交互功能的验证,需要采用面向开发过程的模型验证方法,首先由系统功能层级到功能模块层级逐级分解飞行模式操作控制功能开发架构,进而从功能模块层级到系统功能层级自下而上设计飞行模式操作控制功能验证架构,最后针对功能验证架构中的集成模型测试阶段,以FMCP速度视窗功能为例,通过对所有相关的人机操作进行分类定义,分别定义了速度目标值的状态转换与动作响应,进而针对定义梳理出完整的速度目标值状态转换与动作响应矩阵,矩阵中的每一个表格均对应着速度目标值的状态转换或者动作操作及其对应的响应情况。为自动飞行模式操作控制的集成模型验证提供了参考。

复杂环境下飞控系统PBIT测试优化策略研究

为提前发现多余度电传飞行控制系统潜在故障,三代机、四代机普遍采用了BIT技术,在飞行前检测飞行控制系统各部件的完好性。本文对飞控系统BIT测试内容、设计方法和影响设计门限的多种因素进行了分析,并提出了相应的PBIT测试优化策略。

高超声速边界层转捩研究现状与发展趋势

高超声速飞行器边界层容易经历层流/湍流转捩,层流流动和湍流流动在摩擦阻力、热交换、噪声和掺混等方面有巨大差别,转捩问题已成为制约高超声速技术突破的基础科学问题之一,是当前国际学术研究的热点与难点。本文详细分析了国内外高超声速边界层转捩研究现状,并将其归为三类:已知主要原因的现象与规律、已知部分原因的现象与规律、未知或矛盾的现象。其中已知主要原因的现象与规律包括壁温、马赫数和噪声影响;已知部分原因的现象与规律主要有头部钝度、熵层和攻角影响;未知或矛盾的现象主要有单位雷诺数影响、转捩区长度、转捩区摩阻和热流分布等。同时介绍了高超声速边界层转捩影响因素研究、转捩机理研究、转捩预测方法及模型研究、促进/推迟转捩的控制方法研究、以及一些公开的飞行试验等方面的进展。最后指出,在今后的高超声速边界层转捩研究中,建议把单个影响因素独立出来研究,尽量避免多因素相互干扰;高超声速边界层失稳研究需要特别关注横流失稳、熵层和模态相互作用;转捩预测需考虑三维边界层和来流扰动的影响;转捩控制研究应重点关注高效、低阻、低热的控制方法;转捩飞行试验十分重要,飞行试验和静音风洞发挥的作用会越来越明显。过去60多年的研究经验表明在未来的研究中应该注重多种手段相结合。