An impact angle constraint integral sliding mode guidance law for maneuvering targets interception

An integral sliding mode guidance law(ISMGL)combined with the advantages of the integral sliding mode control(SMC)method is designed to address maneuvering target interception problems with impact angle constraints.The relative motion equation of the missile and the target considering the impact angle constraint is established in the longitudinal plane,and an integral sliding mode surface is constructed.The proposed guidance law resolves the existence of a steady-state error problem in the traditional SMC.Such a guidance law ensures that the missile hits the target with an ideal impact angle in finite time and the missile is kept highly robust throughout the interception process.By adopting the dynamic surface control method,the ISMGL is designed considering the impact angle constraints and the autopilot dynamic characteristics.According to the Lyapunov stability theorem,all states of the closed-loop system are finally proven to be uniformly bounded.Simulation results are compared with the general sliding mode guidance law and the trajectory shaping guidance law,and the findings verify the effectiveness and superiority of the ISMGL.

Adaptive weighting impact angle optimal guidance law considering seeker's FOV angle constraints

In this paper, a new adaptive optimal guidance law with impact angle and seeker’s field-of-view(FOV) angle constraints is proposed. To this end, the generalized optimal guidance law is derived first. A changeable impact angle weighting(IAW) coefficient is introduced and used to modify the guidance law to make it adaptive for all guidance constraints. After integrating the closed-form solution of the guidance command with linearized engagement kinematics, the analytic predictive models of impact angle and FOV angle are built, and the available range of IAW corresponding to constraints is certain. Next, a calculation scheme is presented to acquire the real-time value of IAW during the entire guidance process. When applying the proposed guidance law, the IAW will keep small to avoid a trajectory climbing up to limit FOV angle at an initial time but will increase with the closing target to improve impact position and angle accuracy, thereby ensuring that the guidance law can juggle orders of guidance accuracy and constraints control.

A trajectory shaping guidance law with field-of-view angle constraint and terminal limits

In this paper, a trajectory shaping guidance law,which considers constraints of field-of-view(FOV) angle, impact angle, and terminal lateral acceleration, is proposed for a constant speed missile against a stationary target. First, to decouple constraints of the FOV angle and the terminal lateral acceleration, the third-order polynomial with respect to the line-ofsight(LOS) angle is introduced. Based on an analysis of the relationship between the looking angle and the guidance coefficient,the boundary of the coefficient that satisfies the FOV constraint is obtained. The terminal guidance law coefficient is used to guarantee the convergence of the terminal conditions. Furthermore, the proposed law can be implemented under bearingsonly information, as the guidance command does not involve the relative range and the LOS angle rate. Finally, numerical simulations are performed based on a kinematic vehicle model to verify the effectiveness of the guidance law. Overall, the work offers an easily implementable guidance law with closed-form guidance gains, which is suitable for engineering applications.

Trajectory shaping guidance law based on virtual angle with terminal constraints

A trajectory shaping guidance law based on virtua angle （TSGLBVA） is proposed for a re-entry vehicle with the constraints of terminal impact angles and their time derivatives. In the view of differential properties of the maneuvering trajectory, a virtual angle and a virtual radius are defined. Also, the shaping trajectory of the vehicle is established by the polynomials of the virtual angle. Then, four optimized parameters are selected according to the theorem of parameters transformation presented in this paper. Finally, a convergent variant of the Nelder-Mead algorithm is adopted to obtain the reference trajectory, and a trajectory feedback tracking guidance law is designed. The simulation results demonstrate that the TSGLBVA ensures the re-entry vehicle to impact a target precisely from a specified direction with smal terminal load factor command, as well as to obtain a maximum or constrained terminal velocity according to various requirements.

Three-dimensional cooperative guidance law for multiple missiles with impact angle constraint

This paper proposes a cooperative guidance law for attacking a ground target with the impact angle constraint based on the motion camouflage strategy in the line-of-sight(LOS)frame.A dynamic model with the impact angle constraint is established according to the relative motion between multiple missiles and the target.The process of cooperative guidance law design is divided into two stages.Firstly,based on the undirected graph theory,a new finite-time consensus protocol on the LOS direction is derived to guarantee relative distances reach consensus.And the value of acceleration command is positive,which is beneficial for engineering realization.Secondly,the acceleration command on the normal direction of the LOS is designed based on motion camouflage and finite-time convergence,which can ensure the missiles reach the target with the desired angle and satisfy the motion camouflage state.The finitetime stability analysis is proved by the Lyapunov theory.Numerical simulations for stationary and maneuver targets have demonstrated the effectiveness of the cooperative guidance law proposed.

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 timevarying 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.

Extended optimal guidance law with impact angle and acceleration constriants

The extended optima straints of miss distance and Schwartz inequality. To reduce guidance law with terminal conmpact angle is derived by the terminal acceleration and eliminate gravity disturbance absolutely, the object function, which designs the weight of control command to be the power function of time-to-go＇s reciprocal, is given. And the gravity is considered when building the state equation. Based on the parsing express of the guidance command change with varying time and adjoint system analysis method, the command characteristics and the non-dimensional miss distance of the guidance law are analyzed, a design principle of guidance order coefficients is discussed. Finally, based on the requirement of engineering, the method to calculate the guidance condition and maximal required acceleration of the guidance law is given. The simulation demonstrates that not only the guidance law can satisfy the terminal position and impact angle constraints, but also the terminal acceleration can be converged toward zero, which will support a good situation for the terminal angle of attacking control.

Three-dimensional impact angle constrained distributed cooperative guidance law for anti-ship missiles

This paper investigates the problem of distributed cooperative guidance law design for multiple anti-ship missiles in the three-dimensional(3-D)space hitting simultaneously the same target with considering the desired terminal impact angle constraint.To address this issue,the problem formulation including 3-D nonlinear mathematical model description,and communication topology are built firstly.Then the consensus variable is constructed using the available information and can reach consensus under the proposed acceleration command along the line-of-sight(LOS)which satisfies the impact time constraint.However,the normal accelerations are designed to guarantee the convergence of the LOS angular rate.Furthermore,consider the terminal impact angle constraints,a nonsingular terminal sliding mode(NTSM)control is introduced,and a finite time convergent control law of normal acceleration is proposed.The convergence of the proposed guidance law is proved by using the second Lyapunov stability method,and numerical simulations are also conducted to verify its effectiveness.The results indicate that the proposed cooperative guidance law can regulate the impact time error and impact angle error in finite time if the connecting time of the communication topology is longer than the required convergent time.

Robust Trajectory and Communication Design for Angle-Constrained Multi-UAV Communications in the Presence of Jammers

This paper studies a multi-unmanned aerial vehicle(UAV)enabled wireless communication system,where multiple UAVs are employed to communicate with a group of ground terminals(GTs)in the presence of potential jammers.We aim to maximize the throughput overall GTs by jointly optimizing the UAVs’trajectory,the GTs’scheduling and power allocation.Unlike most prior studies,we consider the UAVs’turning and climbing angle constraints,the UAVs’three-dimensional(3D)trajectory constraints,minimum UAV-to-UAV(U2U)distance constraint,and the GTs’transmit power requirements.However,the formulated problem is a mixed-integer non-convex problem and is intractable to work it out with conventional optimization methods.To tackle this difficulty,we propose an efficient robust iterative algorithm to decompose the original problem be three sub-problems and acquire the suboptimal solution via utilizing the block coordinate descent(BCD)method,successive convex approximation(SCA)technique,and S-procedure.Extensive simulation results show that our proposed robust iterative algorithm offers a substantial gain in the system performance compared with the benchmark algorithms.

Impact angle constrained fuzzy adaptive fault tolerant IGC method for Ski-to-Turn missiles with unsteady aerodynamics and multiple disturbances

An impact angle constrained fuzzy adaptive fault tolerant integrated guidance and control method for Ski-to-Turn(STT)missiles subject to unsteady aerodynamics and multiple disturbances is proposed.Unsteady aerodynamics appears when flight vehicles are in a transonic state or confronted with unstable airflow.Meanwhile,actuator failures and multisource model uncertainties are introduced.However,the boundaries of these multisource uncertainties are assumed unknown.The target is assumed to execute high maneuver movement which is unknown to the missile.Furthermore,impact angle constraint puts forward higher requirements for the interception accuracy of the integrated guidance and control(IGC)method.The impact angle constraint and the precise interception are established as the object of the IGC method.Then,the boundaries of the lumped disturbances are estimated,and several fuzzy logic systems are introduced to compensate the unknown nonlinearities and uncertainties.Next,a series of adaptive laws are developed so that the undesirable effects arising from unsteady aerodynamics,actuator failures and unknown uncertainties could be suppressed.Consequently,an impact angle constrained fuzzy adaptive fault tolerant IGC method with three loops is constructed and a perfect hit-to-kill interception with specified impact angle can be implemented.Eventually,the numerical simulations are conducted to verify the effectiveness and superiority of the proposed method.

有限视场下的攻击时间和角度多约束制导律

针对精确打击任务存在导引头视场受限这一过程约束的制导问题,提出了一种考虑视场限制以及攻击角度和时间多种约束的制导律.首先,在以比例导引项实现精确命中的基础上,利用最优控制理论,采用弹目距离加权控制能耗形式的指标函数,设计了攻击角度控制项,并推导了相应的剩余飞行时间解析预测表达式.其次,基于最优误差动力学方法,设计了攻击时间相关的控制项,保证了攻击时间误差的有限时间收敛.同时,针对导引头视场有限的问题,分析了各控制项对前置角收敛性的影响,利用特定的限制函数确保前置角在整个制导过程中始终处于有效视场内,并给出了前置角的有界性及收敛性证明.数学仿真结果验证了所提出制导律对多约束制导问题的有效性.

改进RRT算法的全向机器人路径规划

针对快速扩展随机树(RRT)算法在进行全局路径规划时生成路径时间长,长度长,在采样时缺乏目标导引性。所以改进了一种目标偏向采样策略,通过引入由初始点,终点,障碍物和随机点构成的虚拟力场,促使随机采样导向目标点。对于障碍物远近的影响,设定障碍物影响的阈值,引入了自适应系数λ,在扩展步长方面,加入了角度约束,根据偏向目标点角度,选择不同的扩展步长。引入冗余节点剔除策略,缩短路径长度,最后加入贪心策略,减少不必要采样节点的个数。首先,在MATLAB2018a仿真进行实验对比,其次在实际环境中进行实验。实验结果显示,在路径搜索时间,拐点数量以及路径距离等均有明显的提高。

针对机动目标的三维领从协同制导律

针对多飞行器协同拦截机动目标的实际需求,提出一种三维领从时空协同制导律。在视线方向上,基于齐次度理论和固定时间稳定理论设计一种领从协同制导律,确保从飞行器与领飞行器的相对距离和相对速度在固定时间内趋于一致,并且给出一致性误差收敛时间的上界,实现多飞行器的同时刻拦截。在视线法向上为实现以特定的碰撞角度拦截目标,设计一种固定时间收敛的非奇异终端滑模制导律,实现攻击角度的固定时间收敛,并避免了制导指令奇异引起的控制饱和问题。考虑到目标的机动对制导系统产生的不利影响,引入固定时间扰动观测器对目标加速度进行前馈补偿。运用李雅普诺夫稳定性理论来证明系统的固定时间稳定特性。通过仿真实验验证所设计制导律可以有效控制多飞行器在领从框架下同时打击机动目标,并以期望的终端角度攻击目标。

拦截机动目标的固定时间事件触发协同制导律

针对无向通信拓扑结构下的多枚导弹考虑角度约束同时攻击机动目标的问题,提出一种事件触发控制下的固定时间收敛分布式多弹协同制导律.首先,建立导弹-目标相对运动学模型,在视线坐标系下解耦为沿视线和垂直视线两个方向.其次,在视线切向考虑时间约束,提出一种基于事件触发机制的固定时间收敛协同制导律,保证攻击时间的一致性;在视线法向上考虑空间角度约束,提出一种基于固定时间收敛理论的分布式空间协同制导律,使视线角和视线角速率快速收敛到期望值,以满足脱靶量和终端角度约束;同时,针对目标机动引起的系统扰动,采用固定时间收敛的观测器进行在线估计与补偿.利用Lyapunov理论,证明了所提出协同制导律的稳定性.最后,通过数学仿真,验证了所提出的协同制导律的有效性.

基于RSS_GN RRT算法的狭长空间路径规划

针对RRT算法在狭长空间中存在的收敛速度慢及规划路径不平滑的问题,提出了一种RSS_GN RRT算法。为了提升算法的收敛速度,提出了引导节点导向策略与分区域采样策略,极大地减少了算法对无效区域的搜索;其次,算法引入了采样角度约束策略来提高规划路径的质量,并采用父节点拓展选择的方法有效解决了由角度约束引起的迭代次数增加的问题。此外,算法可根据感知信息进行地图地动态重构并规划避障路径,提高了算法在低速动态环境中的适应性。仿真结果显示,在狭长通道环境中,RSS_GN RRT算法在规划路径的耗时上比RRT、Goal_bias RRT、角度约束下的RRT、Informed RRT*及DR-RRT算法分别减少了77.3%,51.9%,84.7%,98.8%和60.3%。在迭代次数上,相比于上述算法,分别减少了95.9%,92%,98.3%,98.3%和89.5%。路径的平均曲率也分别降低了94.1%,93.2%,88.7%,91%和92.9%。仿真结果证明了RSS_GN RRT算法在提升规划速度和改善路径质量方面具有显著优势。同时,本文采用了阿克曼模型的小车实测了算法的局部避障能力。经测试,小车可对行驶途中出现的障碍物进行合理避让。

视场角限制下的攻击时间和角度三维矢量制导律设计

为了提高导弹精确打击目标的能力,控制攻击时间和攻击角度三维制导问题在实际应用中具有重要的意义。针对这一问题,本文基于三维矢量制导模型提出了一种视场角约束下的攻击时间和攻击角度控制律。首先,通过将平面矢量制导律扩展至三维空间,提出了一种三维矢量攻击角度约束制导律;其次,在上述制导指令的拦截分量中引入剩余时间偏置项,设计了一种在视场角约束下的三维矢量制导律,并进行了稳定性分析。保证导弹能在视场角约束的条件下,以期望的攻击时间和角度击中目标,并且误差均小于0.01;最后,通过数值模拟验证了所设计制导律的正确性和有效性。

基于落角和时间约束的巡飞弹末端协同打击策略

为解决集群巡飞弹末制导阶段协同定向打击问题,提出了一种具有落角约束的巡飞弹末端协同打击策略。基于小偏差方法,推导了一种具有落角约束的偏置比例制导律,以解决纵向平面定向打击问题。结合所提制导方案,推导了综合考虑比例控制项和角度控制项剩余时间估计方法。在侧向平面内,设计了一种具有时变系数且不存在除零奇异问题的时间控制偏置项,以解决时间控制问题。此外,以剩余时间为协调变量,利用双层协同架构来解决多枚巡飞弹协同定向打击问题。仿真结果验证了该协同制导策略良好的落角控制和协同打击性能。

高速飞行器编队通信保持控制与协同制导技术研究综述

在高速武器平台网络化的背景下,对于高速滑翔飞行器(以下简称“高速飞行器”)的研究不再局限于多约束下的制导方法,而是亟须破解多飞行器协同制导控制技术难题。通过总结部分军事大国研制高速飞行器的发展历程,分析了目前多飞行器协同作战的发展现状与趋势,进而引出研究多飞行器协同制导控制技术的必要性。随后分别对编队通信保持控制技术、带飞行时间约束的协同制导控制技术以及带飞行角度约束的协同制导控制技术的研究现状进行了梳理,为多高速飞行器协同作战探索解决思路,最后对多飞行器协同制导控制技术的未来发展方向进行了展望。

带落角与舵面饱和约束的三维制导控制一体化设计

为使制导控制系统在实际应用过程中适用所面对的多种约束条件,同时避免制导与控制系统设计过程中反复迭代带来的巨大工作量,提出了一种带有落角与舵面饱和约束的三维制导控制系统一体化设计方法。根据导弹六自由度非线性模型以及三维弹目相对运动模型建立全状态耦合的制导控制一体化模型,基于动态面控制器原理,采用4级滑模面设计,构建了能够满足落角与舵面饱和约束的控制器,并给出了控制系统闭环稳定性证明。通过仿真实验有效验证了所提方法的正确性、有效性及鲁棒性。

系统收敛时间和攻击角度控制的滑模导引律

针对打击机动目标带攻击角度约束的导引律设计问题,采用滑模控制理论、任意收敛时间控制方法和干扰观测器,设计了一种攻击角度和收敛时间控制的导引律。根据弹目相对运动关系,将目标机动和建模误差等视为干扰,构建了考虑攻击角度约束的制导系统状态方程。分析了任意收敛时间控制方法在干扰情况下的性能,采用滑模控制和干扰观测器对任意收敛时间控制方法进行改进,提升鲁棒性。将该方法用于滑模面和趋近律设计中,采用干扰观测器对系统扰动进行估计,基于滑模控制理论,结合制导控制系统状态方程,推导出了收敛时间和攻击角度可控的滑模导引律。相比有限时间收敛导引律,该导引律的收敛时间和攻击角度可直接设定,无需根据制导参数计算。设置不同的目标机动方式以及不同的攻击角度和收敛时间,进行了多场景的数值仿真。结果表明,所提导引律能够有效实现攻击角度和收敛时间控制,并精确命中目标,同时相比现有导引律其制导性能更佳。