三维修正比例导引弹道仿真研究

导弹的弹道需用过载是重要的弹道特性指标.而纯比例导引法制导下的弹道需用过载过大,为此通过建立三维空间内的修正比例导引弹道仿真,绘制了理想弹道曲线和各参数变化曲线,仿真结果表明该修正比例导引法制导下的导弹与纯比例导引法制导下的导弹相比,弹道需用过载明显减小.     

基于Monte Carlo方法的战斗部部件质心偏移量仿真分析

针对导弹技术远程化、高命中率的发展趋势,为在设计阶段掌握质心偏移的情况,以尺寸误差为主要分析对象,探讨了质心偏移产生的原因。以某导弹战斗部部件为例,采用极值法与Monte Carlo方法计算了质心偏移量。结果表明,采用极值法计算得到的结果属于小概率事件,而Monte Carlo方法基于大多数机械化零件在公差范围内呈正态概率分布的事实,能够更真实地描述质心偏移的情况。     

摇臂式探测车质心域的计算

为了进行摇臂式探测车的移动性能分析及低重力模拟试验,须确定其质心域.在对摇臂悬架简化的基础上,提出摇臂探测车悬架设计参数的表示方法.将探测车简化为质点系,基于坐标变换法确定各质点的位置矢量,推导出任意运动位置时探测车的质心位置矢量和质心域半径公式.最后以研制的摇臂探测车为例,计算其设计质心、任意运动位置时的质心,并通过优化方法确定该探测车的最大质心域半径.结果表明,摇臂式探测车的质心域为中分面上的平面区域,车体质心的偏置对任意位置时探测车的质心影响较大,而对质心域的变化没有影响.所采用的方法也可用于其他被动铰接式探测车质心域的计算.     

基于MATLAB的某空对地导弹弹道仿真研究

针对某空对地导弹的弹道设计问题,利用MATLAB中的Simulink仿真工具对其进行仿真研究。对某空对地导弹建立了弹道数学模型和仿真模型,得到了导弹运动轨迹、速度变化、弹道倾角等数据曲线。仿真结果表明该仿真系统能够直观地反映出导弹弹道设计中的一些参数问题,为导弹的弹道设计提供理论和数据支持;仿真模型简单,结果直观。     

国外弹道式导弹方位瞄准技术及其发展

弹道式导弹发展趋势向着固体火箭发动机、全机动发射方式、制导技术先进、命中精度高、飞行时间短、体积小、机动性强、反应速度快、突防能力强等方面发展,因此,与之相关导弹的瞄准技术也得到相应发展.导弹的方位瞄准直接影响导弹横向命中精度,得到了各国的重视和发展.本文介绍了国外铁路机动、地下发射井和潜艇发射导弹三种典型的瞄准系统,论述了导弹瞄准技术的发展.     

利用标记分水岭法实现夏克-哈特曼波前传感器质心探测

由于光斑质心探测精度直接影响夏克-哈特曼波前传感器的波前探测精度,本文提出基于标记分水岭法来确定阵列光斑质心探测窗口。首先,对采集到的夏克-哈特曼阵列光斑图像进行平滑并求出其梯度图像;然后,利用大津（OTSU）阈值法在求出的梯度图像上进行目标标记,最后在标记过的梯度图像上进行分水岭分割,确定出每个光斑的探测窗口。由于该方法确定的质心探测窗口是对光斑实际大小进行匹配,故有效地抑制了噪声对质心探测的影响。实验结果表明：利用该方法确定光斑探测窗口所计算的质心精确度和稳定性均比传统的在子透镜窗口中计算光斑质心的方法要高。统计多幅图像计算得到的窗口质心标准差的平均值为0.010 9,比传统法计算出的平均值0.073 4提高了6倍,满足哈特曼波前传感器对光斑质心计算稳定性和精确度的要求。     

基于MATLAB的迫击炮外弹道模型仿真研究

以常规迫击炮外弹道为研究对象,建立了弹丸质心运动的数学模型,运用MATLAB软件对外弹道模型进行仿真计算,得到了外弹道各参数的仿真数据,通过分析仿真结果,为迫击炮外弹道的优化设计提供了合理的依据.     

一种新型智能导引律及其三维可视化仿真

针对空空导弹的导引问题,基于模糊逻辑,提出一种新型智能导引律.将导弹与目标的接近速度以及导弹的视线角速度作为模糊控制器的输入,指令加速度作为输出,并在传统的模糊逻辑控制基础上引入了一个非线性变论域函数,从而实现动态改变模糊变量论域的目的.在此基础上,采用Visual C ＋＋和OpenGL来设计空战的三维可视化仿真平台,并实现了逼真的动画效果.     

基于高斯分布的星像点精确模拟及质心计算

为提高星敏感器仿真系统的精度,提出了基于高斯规律的模拟星像点的灰度扩散方法和高斯质心提取算法.按照二维高斯分布规律置灰度值来模拟星像点像素,对称中心是映射坐标而不是取整的中心像素坐标,以便准确模拟实际星像点散焦及像差导致的灰度扩散.高斯质心提取亚像素定位过程包括像素粗定位和偏差精定位两个步骤,基于高斯规律建立了一个分段函数实现偏差精定位.在星像点噪声为N(0, 1.22)的仿真条件下,整个偏差区间[-0.5,0.5)pixel内高斯质心定位标准差为0.007 pixel,远小于灰度重心法的0.041 pixel和加权灰度重心法的0.026 pixel,而3种方法对模拟星图的处理结果一致.仿真实验表明:模拟星像点高斯灰度扩散法是合理准确的,而高斯质心提取算法简单精确,精度高于传统灰度重心法.     

刚性机械零件质心测量平台的研究与设计

在卫星、导弹、精密机械等复杂设备的研制生产过程中,产品质心位置是一个十分重要的指标,因此通过试验测量产品的质心是必不可少的环节。如何快速准确地测量产品的质心是人们一直在研究的课题。目前,质心的测量方法大致可归纳为试探法、垂线法、配平法、多支点称重法和综合求解法等类型。本文采用多质点称重型质心测量方案,以小型机械零件为测量对象研究设计了一种效率高、精度大、测量便捷的质心测量平台,对于卫星等精密机械零件的质心测量,能很好的满足其精密性、完备性的要求。     

采用伪谱法的再入飞行器最优反馈制导方法

将伪谱最优反馈控制理论应用于再入飞行器制导研究,使用伪谱法进行在线轨道重构,实时反馈更新当前轨道控制量迎角和倾斜角,达到实时最优反馈制导的目的,并采用无量纲化、弹性约束和自适应反馈更新等策略保证算法的实时性。再入飞行仿真表明,轨道重构可以满足实时性要求,阵风干扰下飞行器能达到所要求的终端约束条件,并且制导指令不会出现增加控制难度的抖动现象。     

On the design of fuzzified trajectory shaping guidance law

Midcourse guidance is commonly designed to save as much energy as possible so that the missile’s final speed can be maximized while entering the homing stage. For this purpose, a competitive guidance design should be able to generate an admissible flight trajectory as to bring the interceptor to a superior altitude for a favorable target engagement. In this paper, a new adaptive trajectory shaping guidance scheme based on the adaptive fuzzy inference system, which is capable of generating a variety of trajectories for efficient target interception, is presented. The guidance law is developed with the aim of saving the interceptor’s energy conservation while improving performance robustness. Applications of the presented approach have included a variety of mission oriented guidance, such as cruise missile guidance, anti-ballistic missile guidance, etc.     

采用伪谱法的再入飞行器最优反馈制导方法

将伪谱最优反馈控制理论应用于再入飞行器制导研究,使用伪谱法进行在线轨道重构,实时反馈更新当前轨道控制量迎角和倾斜角,达到实时最优反馈制导的目的,并采用无量纲化、弹性约束和自适应反馈更新等策略保证算法的实时性。再入飞行仿真表明,轨道重构可以满足实时性要求,阵风干扰下飞行器能达到所要求的终端约束条件,并且制导指令不会出现增加控制难度的抖动现象。     

滑翔导弹末段多约束智能弹道规划

滑翔导弹末段飞行时空复杂度高、不确定性强、约束多,给弹道规划与制导算法带来了较大的建模和求解难度。针对这一问题,同时增大末段机动范围并提高弹道规划效率,本文提出一种利用连续型深度置信神经网络（Convolutional Deep Brief Networks,CDBN）预测机动能力、设计经由点状态实现末段多约束智能弹道规划的方法。过程中采用CDBN对机动能力进行在线预测,快速判定经由点状态的可行性,并且通过经由点状态智能设计,实现前后段能量的优化分配,扩大弹道机动包络;通过设计三角函数型弹目视线角实现末段弹道摆动机动,推导机动弹道最优末制导律对视线角进行跟踪,并调节机动频率以满足速度约束。仿真结果表明,CDBN相对BP网络具有更高的机动能力预测精度;本文所提智能弹道规划方法在满足末端速度约束的前提下,可以实现弹道摆动机动并大幅增加飞行包络。弹道规划能够在0.5 s内完成,满足工程应用的快速性要求。     

Analytical impact time and angle guidance via time-varying sliding mode technique

To concretely provide a feasible solution for homing missiles with the precise impact time and angle, this paper develops a novel guidance law, based on the nonlinear engagement dynamics. The guidance law is firstly designed with the prior assumption of a stationary target, followed by the practical extension to a moving target scenario. The time-varying sliding mode (TVSM) technique is applied to fulfill the terminal constraints, in which a specific TVSM surface is constructed with two unknown coefficients. One is tuned to meet the impact time requirement and the other one is targeted with a global sliding mode, so that the impact angle constraint as well as the zero miss distance can be satisfied. Because the proposed law possesses three guidance gain as design parameters, the intercept trajectory can be shaped according to the operational conditions and missile׳s capability.To improve the tolerance of initial heading errors and broaden the application, a new frame of reference is also introduced. Furthermore, the analytical solutions of the flight trajectory, heading angle and acceleration command can be totally expressed for the prediction and offline parameter selection by solving a first-order linear differential equation. Numerical simulation results for various scenarios validate the effectiveness of the proposed guidance law and demonstrate the accuracy of the analytic solutions.     

弹体质量、质心及质偏心测试误差补偿

文中介绍了弹丸质量、质心以及质偏心的三点支承测量方法,论述了各个参数的测量原理,对各参数产生误差的因素以及各个误差的大小进行了定量分析,并提出了相应的误差补偿方式。实验证明,使用此方法可以大大提高测量精度。     

Fuzzy logic sliding mode control for command guidance law design

Recently, the combination of sliding mode and fuzzy logic techniques has emerged as a promising methodology for dealing with nonlinear, uncertain, dynamical systems. In this paper, a sliding mode control algorithm combined with a fuzzy control scheme is developed for the trajectory control of a command guidance system. The acceleration command input is mathematically derived. The proposed controller is used to compensate for the influence of unmodeled dynamics and to alleviate chattering. Simulation results show that the proposed controller gives good system performance in the face of system parameters variation and external disturbances. In addition, they show the effectiveness of the proposed missile guidance law against different engagement scenarios where the results demonstrate better performance over the conventional sliding mode control.     

Non-singular terminal dynamic surface control based integrated guidance and control design and simulation

In this paper, a novel cascade type design model is transformed from the simulation model, which has a broader scope of application, for integrated guidance and control (IGC). A novel non-singular terminal dynamic surface control based IGC method is proposed. It can guarantee the missile with multiple disturbances fast hits the target with high accuracy, while considering the terminal impact angular constraint commendably. And the stability of the closed-loop system is strictly proved. The essence of integrated guidance and control design philosophy is reached that establishing a direct relation between guidance and attitude equations by “intermediate states” and then designing an IGC law for the obtained integrated cascade design model. Finally, a series of simulations and comparisons with a 6-DOF nonlinear missile that includes all aerodynamic effects are demonstrated to illustrate the effectiveness and advantage of the proposed IGC method.     

弹道导弹中段突防弹道设计与验证

针对中段多脉冲弹道规划及评价问题,提出了一种新的弹道设计及评价方法。借鉴轨迹规划中节点搜索及扩展的思想,将多脉冲弹道的脉冲点火点视为节点,建立了多脉冲机动变轨模型。基于该模型,对突防弹道进行了设计及优化。运用"博弈论"的思想,在考虑敌方防御系统的探测机理、性能及部署基础上将敌方预测弹道与实际弹道的差值作为评价函数,以尽量降低敌方预测弹道的精度。最后,设计了导弹突防效能地面验证平台以验证所设计弹道的突防效率与弹道评价方法的有效性。选取了一个例子进行分析及验证,结果表明:相比最省能量弹道,设计的导弹防御系统对突防弹道的预测误差提高了5到15倍,达到了100~300km,飞行时间缩短了8.53%,而且规划时间不到140s。实验显示本文提出的弹道设计方法,能够在很短时间内规划出一条突防概率很高的弹道。     

Game optimal receding horizon guidance laws and its equivalence to receding horizon guidance laws

In this paper, a game optimal receding horizon guidance law (GRHG) is proposed, which does not use information of the time-to-go and target maneuvers. It is shown that by adjusting design parameters appropriately, the proposed GRHG is identical to the existing receding horizon guidance law (RHG), which can intercept the target by keeping the relative vertical separation less than the given value, within which the warhead of the missile is detonated, after the appropriately selected time in the presence of arbitrary target maneuvers and initial relative vertical separation rates between the target and missile. Through a simulation study, the performance of the GRHG is illustrated and compared with that of the existing optimal guidance law (OGL).