战术导弹垂直发射快速转弯复合控制研究

战术导弹垂直发射的关键技术之一是实现快速转弯,研究了垂直发射快速转弯复合控制技术.针对空气舵/扰流片推力矢量复合控制技术,提出了战术导弹垂直发射初段程序转弯方案;建立了导弹的空气舵/扰流片推力矢量复合控制数学模型;通过设计线性组合的舵复合控制策略,将系统由多输入-单输出,转变为单入-单出系统,推导了复合控制弹体的传递函数,设计了复合控制回路,并利用频域分析方法进行了回路分析,纵向通道频域特性分析表明,系统具有良好的鲁棒性和动态品质;最后通过弹道的六自由度仿真,结果验证了所设计的复合控制策略的可行性.     

球型收敛段矢量喷管建模与仿真研究

研究球型收敛段矢量喷管的优化设计问题。由于球型收敛段矢量喷管对矢量特性的要求较高,但是过分强调喷管的矢量会引起推力的下降,由于存在推力与矢量角折中选择方面的难题。为了解决以上问题,通过数值仿真的方法,借助于计算流体动力学软件,分别对不同喉部宽高比和不同矢量偏转角的球型收敛段矢量喷管进行了数值仿真,并分析了喷管内外流场的特性,又研究了压比和来流马赫数对喷管矢量特性的影响规律。仿真结果表明,喉部宽高比和矢量偏转角等结构上的变化对喷管推力等内性能造成的损失很小,说明球型收敛段矢量喷管是一种高效的矢量喷管。     

扰流片诱导的可压缩非定常流场数值模拟

振动扰流片非定常分离流动直接与通过外部激发对剪切流场进行控制的问题有关.例如,可用扰流片产生的非定常分离涡改变飞行器大攻角下背风面产生的分离、旋涡破裂、以及各种非定常现象,以改变飞行器的飞行特性和操纵性能.达到高升力和机动飞行的目的.因此,对于扰流片非定常分离涡的机理研究引起了人们广泛的兴趣.国内这方     

高空飞艇飞行控制系统的线性模型跟随设计

研究飞艇高空定点优化控制问题,针对飞艇在高空时空气舵面操纵不足、舵效控制力低,为解决上述问题,提出了采用气动力舵面和偏转螺旋桨复合控制的操纵模式对高空飞艇的姿态进行控制的方法.首先给出了采用气动力和螺旋桨拉力矢量组合控制的高空飞艇横侧向数学模型,把气动力舵面和偏转螺旋桨复合控制看作一个双输入单输出系统,采用线性模型跟随方法设计了姿态控制系统,使控制系统的性能尽可能地跟随给定的理想参考模型系统的性能.通过仿真表明偏转螺旋桨操纵控制可以有效弥补空气舵面效率的不足,所设计控制系统具有良好的动态品质,并验证了方案的可行性和方法的有效性.     

绝缘子在偏心冲洗时有效冲洗面积的研究

带电水冲洗在绝缘子的清洗作业中得到了大量的运用,因此进行带电水冲洗的相关研究就显得十分必要。论文通过理论分析计算出颗粒脱附所需的临界速率,利用fluent对绝缘子在小水、中水、大水的冲洗方式下的偏心冲洗进行三维仿真计算,根据仿真所得到的绝缘子表面的剪切应力云图,对比绝缘子在不同压力、不同的水平偏转角以及不同冲洗方式下有效冲洗区域的大小,同时对比了在相同压力下有效面积与流量的比值,以此来比较水的利用率。结果表明:利用小水和中水的冲洗方式对绝缘子壁面进行冲洗时,随着喷嘴水平偏转角度的增大,有效冲洗面积先增大然后减小;且随着压力的增大,有效冲洗面积的随偏转角度变化的转折点更靠近小的水平偏转角度;而大水冲洗方式下有效冲洗面随着水平偏转角度的增大而持续减少。     

城轨交通用直线感应电机模糊PI矢量控制

本文将模糊控制的快速性与PI调节的静态无差相结合, 提出一种基于矢量控制的模糊逻辑与PI结合的复合型控制器, 同时设计了推力补偿器对动态纵向边端效应引起的推力的衰减进行了补偿. 通过与PI调节的矢量控制系统的数值仿真结果的对比, 验证了该复合模糊PI调节器对矢量控制下直线感应电机(LIM)动态性能的改善.     

推力矢量可倾转四旋翼自抗扰飞行控制方法

针对常规四旋翼难以实现位置和姿态独立控制问题, 研究了一种具有全向推力矢量的可倾转四旋翼飞行 器系统. 为克服系统的大范围不确定性、强耦合性及外部风扰影响, 设计了基于自抗扰控制(ADRC)技术的飞行控 制器. 通过建立风扰下的系统动力学模型, 分析阵风对旋翼气动力的影响. 接着将系统解耦为六通道单回路结构并 分别设计自抗扰控制器, 引入扩张状态观测器估计系统的内外扰动, 利用非线性状态误差反馈律输出扰动补偿控 制. 在此基础上, 通过变量代换线性化控制分配矩阵, 将控制器输出直接映射到旋翼转速和倾转角. 仿真结果表明, 所设计的自抗扰飞行控制器具有良好的位置和姿态独立控制能力, 能够有效地估计和补偿紊流风扰动, 同时对系统 的部分动力失效故障有较强的鲁棒性.     

发动机推力矢量测试系统的设计

介绍了关于发动机推力矢量的测试系统,阐述了利用传感器进行发动机推力矢量的测试。利用设计的压电式测力平台和其相应的测试软件以及测试系统,用反求法求得空间力矢量大小、推力矢量的推力偏移量、推力偏斜角。通过试验结果表明:此方案是可行的。     

推力矢量飞行器的自抗扰控制设计及控制分配

推力矢量飞行器往往需要在大功角等具有大不确定性和强非线性的区域高质量地完成飞行动作,因此,如何应对大范围不确定性是推力矢量飞行器控制设计的关键问题.另一方面,推力矢量飞行器包含多种控制输入并且不同控制输入具有不同物理特性.因此,控制输入分配也是推力矢量飞行器控制设计的关键问题.为了对付大范围的不确定性,本文引入虚拟控制量的概念,采用自抗扰控制技术实现对飞行过程中的总扰动的实时估计和补偿.进一步,考虑控制输入的物理约束条件,提出了保证虚拟控制量达到设计值并使得发动机能耗最小的控制输入分配方案.通过建立对应的优化问题,严格分析其最优解的性质并提出了有限步求解最优控制分配输入量的算法.在仿真环境下,提出的控制算法有效实现了推力矢量飞行器大功角区域的机动动作,并能应对大范围的气动参数不确定性.     

基于传递函数的高速互连线串扰分析

针对互连线系统在高速信号激励下产生的串扰问题,导出高速互连线的时域分布参数系统模型,建立从激励电压到攻击线末端和受害线两端电压相应的传递函数,从系统的角度研究串扰噪声问题。仿真结果表明了该方法的有效性,与数值方法相比更具有一般性。     

基于CFD的绝缘子水冲洗临界水平偏转角研究

利用流体仿真软件fluent对绝缘子冲洗过程中存在的非轴心射流工况进行仿真模拟,分析了水射流以不同水平偏转角对绝缘子进行冲洗时的压力大小以及作用范围;通过Matlab对相同射流压力以及喷管直径在不同水平偏转角时所得到的打击力进行曲线拟合,从而找到在该压力下的临界水平偏转角,对不同压力下临界水平偏转角的大小进行了对比,并利用Matlab的cftool对不同压力和不同喷口直径下的临界水平偏转角度进行了多项式拟合,根据该多项式可以计算出压力在1MPa~4MPa,喷口直径在3mm~8mm任一工况下的临界偏转角度。结果表明:临界偏转角度随着入口压力的增大而减小,其变化率随喷口直径的增大而增大;当入口压力大小相同时,临界偏转角度随喷口直径的增大而减小,为绝缘子的实际清洗工作提供了一定的理论参考。     

Transition Flight Modeling of a Fixed-Wing VTOL UAV

This paper describes a complete six-degree-of-freedom nonlinear mathematical model of a tilt rotor unmanned aerial vehicle (UAV). The model is specifically tailored for the design of a hover to forward flight and forward flight to hover transition control system. In that respect, the model includes the aerodynamic effect of propeller-induced airstream which is a function of cruise speed, tilt angle and angle of attack. The cross-section area and output velocity of the propeller-induced airstream are calculated with momentum theory. The projected area on the UAV body that is affected by the propeller-induced airstream is specified and 2D aerodynamic analyses are performed for the airfoil profile of this region. Lookup-tables are generated and implemented in the nonlinear mathematical model. In addition, aerodynamic coefficients of the airframe are calculated by using CFD method and these data are embedded into the nonlinear model as a lookup-table form. In the transition flight regime, both aerodynamic and thrust forces act on the UAV body and the superimposed dynamics become very complex. Hence, it is important to define a method for hover-to-cruise and cruise-to-hover transitions. To this end, both transition scenarios are designed and a state-schedule is developed for flight velocity, angle of attack, and thrust levels of each of the thrust-propellers. This transition state schedule is used as a feedforward state for the flight control system. We present the simulation results of the transition control system and show the successful transition of TURAC in experiment.     

导弹异形卷弧翼安装选型

为在空中发射武器外形设计中引入异形卷弧翼,将其作为导弹主升力面,研究其相对于弹身的安装选型问题．模型设计采用1对异形卷弧翼,并将其沿弹体纵向平面对称布置．定义异形卷弧翼相对于弹身的安装位置角和安装偏角参数,选取2组计算模型,基于计算流体力学（Computational Fluid Dynamics,CFD）仿真计算评估这2个角度的变化对异形卷弧翼-弹身组合体纵向超音速气动参数的影响,得到组合体升力因数、阻力因数以及升阻比随2个角度参数变化而变化的规律．结果表明,在设计任务中,当安装位置角等于120°,安装偏角等于-10°时,组合体的升力因数和升阻比达到最大值．该方法可以改善导弹的升力特性,提高导弹的升阻比,使导弹获得更好的飞行性能．     

Identification of Aerodynamic Coefficients of Longitudinal Movement and Error Estimates for Onboard Measurements of Supercritical Angles of Attack

We propose an identification method for the aerodynamic coefficients of the lift force, drag force, and pitch moment in the supercritical range of angles of attack. The investigated models take into account the hysteresis with respect to the angle of attack, which is the main specific feature of aerodynamic coefficients in the specified range. Also, we consider the problem to find and correct errors of onboard measurements in the channels of the angle of attack and the air speed and the computational model for the forces and moments created by engines with a deflected thrust vector. We provide processing examples for the flight test data obtained during maneuvers on supercritical angles of attack; they confirm the efficiency of the proposed methods.     

Design of maximum thrust plug nozzles with variable inlet geometry

An optimization analysis is presented for axisymmetric plug nozzles with varible inlet geometry. The analysis is based on the governing gas dynamic relations for a rotational flow of a frozen or equilibrium gas mixture. The problem is formulated to maximize the axial thrust produced by the plug nozzle for a general isoperimetric constraint, such as constant nozzle length or constant nozzle surface area. The effects of base pressure and ambient pressure are included in the thrust expression to be maximized. The governing gas dynamic equations and the differential and integral constraints that the solution must satisfy are incorporated into the formulation by means of Lagrange multiples. The formalism of the calculus of variations is applied to the resulting functional to be maximized. The results of the optimization analysis are a set of partial differential equations for determining the Lagrange multipliers in the region of interest and a set of equations for determining the necessary boundary conditions for the solution. The complete set of equations for the gas dynamic properties and the Lagrange multipliers are system of first order, quasi-linear, non-homogeneous partial differential equations of the hyperbolic type, which can be treated by the method of charac- teristics. The characteristic and compatibility equations for the system are presented. A numerical solution procedure is presented to determine wether or not a given plug nozzle geometry is an optimal solution. An iteration technique is developed which systematically adjusts the plug nozzle geometry until the optimal solution is obtained. Selected parametric studies are presented. These studies illustrate the effect of the specific heat ratio, the design pressure ratio and the base pressure model on the thrust peformance and nozzle geometry of optimal, fixed length, plug nozzles.     

Electrohydraulic thrust vector control of twin rocket engines with position feedback via angular transducers

The thrust vector control (TVC) of rocket engines is used when the aerodynamic surfaces are inadequate to control vehicles or when a greater agility may be required of a missile. Traditionally, in all spacecraft to date actuators used to gimbal the engine have been hydraulic. The subject of this paper is the TVC with gimballed nozzle assembly controlled by an electrohydraulic servosystem, where two linear hydraulic servoactuators gimbal the engine. Each servoactuator is controlled by an electrohydraulic servovalve. The thrust vector direction is a result of the motion of both servoactuators. In this paper the TVC system is treated as a robotic system that allows developing the procedure of solving an inverse kinematics problem as well as the control of the robotic system in the output space instead of in the space of internal dynamics. The position feedback is provided by measuring the direction of the thrust vector, instead of measuring the displacements of the servoactuators. A linear model of the servosystem has been developed and simulated. The proposed control concept has experimentally been validated in the TVC test bench.     

短特型喷管三维型面气动优化设计

以某型冲压发动机的喷管为研究对象,为进一步提高发动机的推力,采用双圆弧法建立了短特型喷管的物理模型.以发动机推力为优化目标,利用自动优化平台,集成了参数化建模和喷管流场数值模拟,并运用ASA和SQP组合优化策略对喷管的内型面进行气动优化,优化后,发动机轴向推力提高了1.16%,从而改善了冲压发动机的性能.对优化前后喷管的扩张损失进行了比较和验证,证明优化结果是可信的.优化结果表明:通过对喷管型面进行组合优化可以提高发动机的推力,这种方法十分有效,能够快速获得最优解,可用于喷管的优化设计.     

A four-level technique for estimation of tactical missile aerodynamic parameters

A four-level technique is described for estimation of aerodynamic moment coefficient Mα, normal acceleration force coefficient Zα, and velocityV. These parameters are associated with a simplified model of pitch plane dynamics for a tactical missile. Contrary to past studies where aerodynamic parameters were estimated from measurements which are very difficult to obtain (e.g., angle of attack), it is shown herein how estimates of the above parameters can be realized from measurements which are more easily obtained. Specifically, parameters are estimated from measurements of rate, normal acceleration, and gimbal angle deflection.     

基于表速控制的无人机高度改变的控制律仿真设计

介绍了一种使无人机以一定的表速实现高度改变的控制算法的设计和仿真验证。无人机在高度改变过程中,纵向采用表速控制的算法,通过控制表速偏差使无人机以一定的表速实现高度的改变。基于总能量的概念,创新性地引入能量角,依据速度差值和能量角补偿来实现高度的改变,同时控制发动机推力,并进行速度保护,使无人机以保持恒定速度的方式爬升或下降至目标高度,保证了因推力变化引起的能量改变,最大限度地转化为势能,提高了经济性及稳定性。该算法通过与自动油门算法相结合,可实现对无人机的纵向高度层改变的自动控制,具有控制精度高、控制平稳等优点。