机动目标的鲁棒几何法导引律仿真分析

基于机动目标,以空间微分几何和李雅普诺夫稳定性原理为基础,通过仿真分析鲁棒几何法导引律在追击不同机动目标时的弹道。在不同发射角下追击相同目标时,比较比例导引法和鲁棒几何法两种导引律的弹道特性。仿真结果表明鲁棒几何导引法具有以下优点:不需要额外的测量信息,兼具比例导引律的易执行性和微分几何制导律对视线旋转抑制的有效性。该制导律可以拦截大机动目标,且性能优于现有的变系数比例导引算法,同时拦截过程中过载曲线变化更为合理,弹道平直不需要得到目标精确的加速度和速度方位信息,对目标机动具有强的鲁棒性。     

基于扩展滑动模态控制的末制导律设计

对于三维目标拦截问题，提出了一种新的具有鲁棒性的扩展连续滑动模态末制导律。基于变结构控制理论的方法和零化弹目视线角速率的思想，选择一个合适的滑动区域代替传统变结构滑动模态的设计，同时将目标的机动加速度视为已知的有界扰动。设计得到了具有鲁棒性的三维扩展的连续滑动模态末制导律。该方法利用Lyapunov稳定理论严格证明了扩展的连续滑动模态末制导律在滑动区的可达性和渐近稳定性。该方法简单，易于理解，便于工程应用，数字仿真验证了所提出的制导律更适合拦截大机动目标。     

Time-to-go weighted optimal trajectory shaping guidance law

For maneuvering target,the optimal trajectory shaping guidance law which can simultaneously achieve the designed specifications on miss distance and final impact angle was deduced using optimal control theory based on the time-to-go weighted function.Based on the same cost function,the closed-form solutions of the guidance law were derived when the initial displacement of missile,final impact angle,heading error and target maneuver was introduced into the lag-free guidance system.To validate the closed-form solutions,the simulation of the lag-free system was done and the simulation results exactly matched the closed-form solutions and only when the exponent is greater than zero,the final acceleration approaches to zero.     

拦截大机动目标自适应模糊制导律

为了满足拦截高速大机动目标、高精度制导的需要,将制导律设计问题转化为反馈控制问题,基于模糊控制理论,提出了一种解析描述模糊控制规则的新型自适应模糊导引律.该模糊控制器将比例制导律指令及其微分作为模糊控制的输入量,模糊控制规则及模糊推理用解析式表达,易于计算、调整,适合实时在线控制.该导引律能够根据目标加速度和目标速度的变化自适应地改变模糊控制规则,因此具有较强的鲁棒性.对拦截高速大机动目标的大量仿真结果表明,所提出的导引律在脱靶量、拦截时间等指标方面显著优于传统的比例导引律.     

Performance of 3D TPN against true-arbitrarily maneuvering target for exoatmospheric interception

True proportional navigation (TPN) guidance law is widely used for exoatmospheric interception, for its robustness and ease of implementation. The performance of TPN against nonmaneuvering target or the maneuvering target with a specific acceleration had been analyzed before. However, the obtained results are not suitable for the realistic exoatmospheric interception scenario, since the target may maneuver along an arbitrary direction with an arbitrary but upper-bounded acceleration in the three-dimensional (3D) space, which is the so-called “true-arbitrarily maneuvering target” in this paper. With the help of the line-of-sight (LOS) rotation coordinate system, the performance of 3D TPN against the true-arbitrarily maneuvering target is thoroughly analyzed using the Lyapunov-like approach. The upper-bound of the 3D LOS rate is obtained, and so is that of the commanded acceleration of 3D TPN. After that, the capture region of 3D TPN is presented on the initial relative velocity plane. The nonlinear 3D relative kinematics between the interceptor and the target is taken into full account. Finally, the new theoretical findings are demonstrated by numerical simulations.     

HP nonlinear guidance law design based on smooth sliding mode control

To eliminate the perturbation of interceptor detection induced by aerodynamic heating,the head pursuit (HP) guidance law for three-dimensional interception was presented. The guidance law positioned the interceptor ahead of the target on its flight trajectory,and the speed of interceptor was required to be lower than that of the target. On the basis of a novel head pursuit three-dimensional guidance model,a nonlinear guidance law was developed based on smooth sliding mode control theory. At the same time,a special observer was designed to estimate the target acceleration,and a numerical example on maneuvering ballistic target interception verified the effectiveness of the presented guidance law.     

Strapdown Homing Guidance System Design for Some Ammunition

The guidance is one of the main development di-rections of conventional ordnance.Strapdown homingis one of the main measures for ammunition guidancebecause of the li mitation of mass,volume,cost andlaunch overload.Such guided ammunition couldn’tdirectly use the conventional proportional guidancelawasthe strapdownseeker could only providethein-formation of line-of-sight angles in body coordinates,which could be used in attitude pursuit guidance sys-tem[1].It is well known that the attitude pur…     

前向拦截自适应滑模制导律

为拦截高速飞行目标,基于前向拦截的思想,给出了一种自适应滑模制导律．通过弹道的不断调整,置拦截导弹于具有较快飞行速度的目标前方预测弹道上,使得二者的飞行方向满足一种特定的几何关系．该制导律的推导考虑到了拦截导弹与目标的自动驾驶仪动态及其模型误差,且由于采用了自适应滑模的设计方法,不需要知道目标的加速度界和模型误差界．采...     

指向预测命中点的最短时间制导

拦截时间最短为指标函数的最优制导策略要求导弹沿命中点视线方向飞行,但制导指令形成方式并未得到有效解决。根据最短时间制导策略原理,以相对预测命中点的制导误差渐进收敛为条件设计了新的制导算法,解决了非线性拦截系统最短时间拦截制导策略的指令形成问题。针对末制导系统信息来源和拦截系统参数变化特点,提出了简化算法。新的制导算法能够保持最短时间拦截策略的最优性,适合大离轴角发射情况下对高机动目标的拦截。仿真结果表明新制导算法确实具有拦截时间短,脱靶量小的优点。     

Three-dimensional suboptimal guidance law for fly-over and shoot-down smart ammunition via virtual target

An optimal burst height is required for the fly-over and shoot-down smart ammunition with an EFP warhead at the instant of explosion which brings a special requirement to the miss distance of the terminal guidance law.In this paper,aguidance law based on the virtual target scheme is proposed.First,the practical pursuit-evasion issue between the ammunition and the target with specific miss distance is transformed into a virtual pursuit-evasion problem with zero miss distance.Secondly,a complete three-dimensional pursuit-evasion kinematics model is established without any simplifications.And then,a suboptimal guidance law is designed based on theθ-D method which has constraints of the elevation and azimuth angular velocity of the virtual line of sight(LOS).Finally,in order to verify the performance of the proposed guidance law,three test cases are conducted.Numerical results show that under the proposed terminal guidance law,the smart ammunition not only can fly above the target with an optimal burst height but also have a smaller normal acceleration on the terminal trajectory.     

Differential geometric guidance command with finite time convergence using extended state observer

For improving the performance of differential geometric guidance command(DGGC), a new formation of this guidance law is proposed, which can guarantee the finite time convergence(FTC) of the line of sight(LOS) rate to zero or its neighborhood against maneuvering targets in three-dimensional(3D) space. The extended state observer(ESO) is employed to estimate the target acceleration, which makes the new DGGC more applicable to practical interception scenarios. Finally, the effectiveness of this newly proposed guidance command is demonstrated by the numerical simulation results.     

空基反助推段导弹制导技术综述

为研究空基反助推段导弹的发展现状和拦截制导面临的关键技术与挑战,对空基反助推段导弹相关的文献进行了分析与总结.空基反助推段导弹由载机发射,对处于助推段飞行的导弹实施拦截,是导弹防御体系的重要组成部分.首先概述空基反助推段系统、导弹和制导技术的发展现状,介绍了典型系统方案的作战流程、拦截弹的结构与性能以及国内外助推段拦截制导技术的研究进展.其次分析了空基助推段拦截制导面临的诸多挑战,体现在拦截弹能力有限、可用时间短、目标机动性强、拦截条件复杂和不确定性强等方面.然后提出了空基助推段拦截制导需要解决的关键技术,包括动力系统配置、发射诸元快速解算、制导指令在线修正、中末制导交班方案和拦截机动目标的高精度制导律等技术.最后总结未来空基助推段拦截制导技术的发展方向为:在线制导、高精度多模复合制导、智能化和协同制导.研究表明:空基反助推段导弹制导技术尚在发展阶段,还存在一些关键问题亟需研究解决,同时,自主规划、人工智能等理论可为关键技术的突破提供参考.     

Algebraic solution of differential geometric guidance command and time delay control

According to the three-dimensional geometry of the engagement,the explicit algebraic expression of differential geometric guidance command(DGGC)is proposed.Compared with the existing solutions,the algebraic solution is much simpler and better for the further research of the characteristics of DGGC.Time delay control(TDC)is a useful method to tackle the uncertainty problem of a control system.Based on TDC,taking the target maneuvering acceleration as a disturbance,the estimation algorithm of the target maneuvering acceleration is presented,which can be introduced in DGGC to improve its performance.Then,the augmented DGGC(ADGGC)is obtained.The numerical simulation of intercepting a high maneuvering target is conducted to demonstrate the effectiveness of ADGGC.     

带有双曲正切加权函数的落角约束最优制导律

为降低末制导律对初始状态误差的敏感度、提高导弹的末端抗干扰能力,针对带有落角约束的末制导问题,考虑基于双曲正切函数的一类加权函数,提出了一种基于间接Gauss伪谱法的最优末制导律. 首先,基于目标位置和期望落角建立了落角坐标系,并在该坐标系中建立了导引运动关系方程,得到了带有落角约束的末制导模型;然后,根据极小值原理推导出了用于求解最优制导律的两点边值问题,运用Gauss伪谱法进行离散,把两点边值微分方程转换为一系列代数方程;最后,通过显式求解代数方程快速得到了最优控制律,该方法避免了求解黎卡提微分方程,不需要进行繁琐的积分运算,计算量小. 所提制导律在推导过程中不依赖于加权函数的具体形式,可非常方便地处理复杂加权函数. 仿真结果表明:通过设计不同形式的加权函数,可灵活改变导弹运动轨迹及制导指令的分布,以实现不同的制导目标;所提方法能有效降低制导律对初始状态误差的敏感度,而且还可以提高导弹的末端抗干扰能力,在很大程度上提高了制导律的设计灵活性.     

仅有角测量系统鲁棒跟踪研究

传统的仅有角测量跟踪器不能跟踪机动目标。为此作者在「２」中提出了跟踪机 仅有角测量系统。在文「２」基础上本文提出了新的变结构鲁跟踪算法，解决了令有角测量系统鲁棒跟踪机动目标问题，该系统由机动目标检测，噪声污染模型以及鲁棒跟踪器构成。最后应用本结果提出了仅有角测量系统鲁棒制导规律，给出了纺真计算结果。     

Effect of Autopilot Structure on Proportional Navigation Guidance

The dynamic characteristics of acceleration autopilot and attitude autopilot are discu.ssed in detail. Also, a comparison study was made between these two different types of control schemes for guidance loop. By means of simulation, it is concluded that the guidance accuracy is mainly determined by the slowest subsystem among different system dynamics. For air-to-ground missiles, with limited terminal guidance time, the control scheme of acceleration autopilot combined with proportional navigation guidance （PNG） law is the better choice.     

战术导弹的PID型比例最佳制导

本文提出了一种PID型导引方法。这种导引方法以广义视线角误差(视线角速度积分,比例,微分的函数)绝对值极小为准则,可以实现准平行接近法导引,它对弹体的过载要求低,可以攻击高机动目标。线性制导方式即为PID型比例导引,它是一种次最佳制导策略。PID型比例导引使制导系统的稳定性,动态品质和准确度得到协调提高,对制导参数,导弹自动驾驶仪参数摄动的鲁棒性高。本文的PID型比例导引克服了[3]的PID型比例导引结构上存在的缺陷,仿真结果证明了上述结论。     

拦截导弹自适应动态滑模制导

对于平面拦截问题,提出了一种具有强鲁棒自适应性的动态滑模制导律。将目标机动视为一类具有有界扰动的不确定因素,以视线法向上的相对运动速度及其导数构筑滑模切换面,并基于Lyapunov稳定性理论进行制导律的设计,由于采用了动态滑模的思想,有效降低了抖振的影响。理论分析与数字仿真表明该制导律具有优良的弹道特性,可对连续高机动目标进行有效拦截,同时控制结构相对简单,易于工程实现。     

Optimal Guidance Law to Maximize Terminal Velocity for Missiles with Impact Angle Constraint

In this paper, an optimal guidance law for missiles with impact angle and miss distance constraints is proposed to achieve the maximal terminal velocity. The normal acceleration command that includes the time-varying coefficients is introduced to satisfy the desired impact angle as well as zero miss distance according to the geometric relation and relative motion parameters between missile and target. The problem is formulated as an optimal control problem by defining the angle of velocity error and flight-path angle as state variables and maximizing a performance index of the terminal velocity. The analytical form of the proposed guidance law is obtained as the solution of the optimal control problem combining optimal control theory and numerical value computation method. Nonlinear simulations of various situations demonstrate the performance and feasibility of the proposed optimal guidance law.     

升力式再入飞行器全程三维自主制导方法

为增强升力式再入飞行器任务的灵活性,减少射前准备,实现快速发射和精确打击,研究一种升力式再入飞行器全程三维自主制导方法.首先,再入段制导提出通过再入走廊上下边界内插得到阻力加速度剖面,倾侧角采用两次反转的设计方法,通过调节阻力加速度剖面的内插值系数和倾侧角的反转点来满足约束和精度的要求;然后,下压段仅依靠飞行器当前信息和目标点位置采用比例导引的形式来设计导引律,并通过对导引系数的实时更新实现以一定的弹道倾角对目标的精度打击;最后通过对再入段和下压段衔接点处控制指令的平滑过渡,得到全程三维自主制导方法.蒙特卡洛打靶仿真结果表明,该制导方法能够导引飞行器在满足再入约束的情况下以规定的弹道倾角对目标实施精确打击,其中打击偏差小于10 m,弹道倾角偏差小于1.3°.