Proportional Navigation Guidance Law with Variable Coefficient Gravity Compensation

Based upon a discussion on the merits and limitations of proportional navigation（PN）guidance law in which constant gravity compensation is included as a part,a counterpart having varying compensations,which changes with pitching angle and line-of-sight angle,is substituted.Flight trajectory simulation over a submissile model is conducted,resulting in increased impact angle,shorter miss distance,smaller maximum normal overload and narrower terminal angle of attack.

Design and Analysis of Terminal Guidance Law for Kinetic Interceptors based on Pulse Engine

A new terminal guidance law is proposed based on a solid propellant pulse engine and an improved proportional navigation method to address the terminal guidance issue for kinetic interceptors.On this basis,the start-stop curve of the pulse motor during the terminal guidance process is designed,along with its start-up logic.The effectiveness of the proposed guidance strategy is verified through simulation.

Impact time control using biased proportional navigation for missiles with varying velocity

A feasible guidance scheme with impact time constraint is proposed for attacking a stationary target by missiles with time-varying velocity.The main idea is to replace the constant velocity with the future mean velocity;therefore, the existing time-to-go estimation algorithm of the proportional navigation guidance law can be improved to adapt to varying conditions.In order to obtain the prediction of the velocity profile, the velocity differential equation to the downrange is derived, which can be numerically integrated between the current downrange and the target position by the on-board computer.Then, a third-order polynomial is introduced to fit the velocity profile in order to calculate the future mean velocity.At the beginning of each guidance loop, the future mean velocity is predicted and the time-to-go information is updated, based on which a novel biased proportional navigation guidance law is established to achieve the impact time constraint.Finally,numerical simulation results verified the effectiveness of the time-to-go estimation algorithm and the proposed law.

Extended differential geometric guidance law for intercepting maneuvering targets

Without assumptions made on motion states of missile and target, an extended differential geometric guidance law is derived. Through introducing a line of sight rotation coordinate system, the derivation is simplified and has more explicit physical significances. The extended law is theoretically applicable to any engagement scenarios. Then, on basis of the extended law, a modified one is designed without the requirement of target acceleration and an approach is proposed to determining the applied direction of commanded missile acceleration. Qualitative analysis is carried out to study the capture performance and a criterion for capture is given. Simulation results indicate the two laws are effective and make up the deficiency that pure proportional navigation suitable for endoatmospheric interceptions cannot deal with high-speed maneuvering targets. Furthermore, the correctness of the criterion is validated.

Multiconstraint adaptive three-dimensional guidance law using convex optimization

The traditional guidance law only guarantees the accuracy of attacking a target. However, the look angle and acceleration constraints are indispensable in applications. A new adaptive three-dimensional proportional navigation(PN) guidance law is proposed based on convex optimization. Decomposition of the three-dimensional space is carried out to establish threedimensional kinematic engagements. The constraints and the performance index are disposed by using the convex optimization method. PN guidance gains can be obtained by solving the optimization problem. This solution is more rapid and programmatic than the traditional method and provides a foundation for future online guidance methods, which is of great value for engineering applications.

Ideal proportional navigation for exoatmospheric interception

Ideal proportional navigation （IPN） is a natural choice for exoatmospheric interception for its mighty capture capability and ease of implementation. The closed-form solution of two- dimensional ideal proportional navigation was conducted in previous public literature, whereas the practical interception happens in the three-dimensional space. A novel set of relative dynamic equations is adopted in this paper, which is with the advantage of decoupling relative motion in the instantaneous rotation plane of the line of sight from the rotation of this plane. The dimension-reduced IPN is constructed in this instantaneous plane, which functions as a three-dimensional guidance law. The trajectory features of dimension-reduced IPN are explored, and the capture regions of dimension-reduced IPN with limited acceleration against nonmaneuvering and maneuvering targets are analyzed by using phase plane method. It is proved that the capture capability of IPN is much stronger than true proportional navigation （TPN）, no matter the target maneuvers or not. Finally, simulation results indicate that IPN is more effective than TPN in exoatmospheric interception scenarios.

Design of Closed Loop Optimal Guidance Law Using Neural Networks

It is generally impossible to obtain the analytic optimal guidance law for complex nonlinear guidance systems of homing missiles,and the open loop optimal guidance law is often obtained by numerical methods,which can not be used directly in practice.The neural networks are trained off line using the optimal trajectory of the missile produced by the numerical open loop optimal guidance law,and then,the converged neural networks are used on line as the feedback optimal guidance law in real time.The research shows that different selections of the neural networks inputs,such as the system state variables or the rate of LOS(line of sight),may have great effect on the performances of the guidance systems for homing missiles.The robustness for several guidance laws is investigated by simulations,and the modular neural networks architectures are used to increase the approximating and generalizing abilities in the large state space.Some useful conclusions are obtained by simulation results.

基于粒子群优化的组合导引律模型

针对导弹打击机动目标的问题,提出了一种基于组合导引律的粒子群优化算法以应对战场环境中的机动目标打击问题。首先对三维弹目运动关系进行分析,设计了一种包含角度和时间偏置项的比例导引律,旨在有效抑制终端角度变化,同时确保打击时间的精确控制。为实现最优导引律参数的求解,研究构建了以打击角、打击时间、法向加速度和脱靶量等约束条件为基础的代价函数。采用标准粒子群算法进行自适应寻优,通过迭代计算确定使代价函数最小化的最优导引律参数,进而得到满足约束条件的最优弹道。为验证算法的有效性,设置了机动目标场景,并设计了与原始比例导引律的对比仿真实验以及多种打击时间测试实验。实验结果表明,相较于传统比例导引律,基于粒子群优化的组合导引律算法在应对机动目标时能够生成满足打击时间要求及其他约束条件的弹道,并在不同打击时间设定下均展现出良好的打击效果,具有广阔的应用前景。

一种基于迁移学习的多任务制导算法

针对典型的飞行器制导任务,利用深度学习算法可以有效地拟合导弹飞行状态与制导指令之间的函数关系。然而当制导任务发生变化时,其二者之间的映射关系也会随之改变,从而导致在当前环境下预训练好的模型无法直接作用于新环境,重新训练制导模型需要大量的弹道数据和巨额的时间成本。为解决上述问题,基于迁移学习的思想引入域对抗神经网络,提出基于迁移学习的多任务制导算法。以1个含有大量标签数据的源域任务辅助2个含有极少量标签数据的目标域任务进行迁移学习,从而克服预训练与在线控制之间的环境差异。使用特征提取器和域判别器提取出对任务环境不敏感的关键特征,使神经网络学习到各个任务所共享的底层信息;为提高预测精度,分别设计针对不同任务的偏置加速度预测器。数值仿真结果表明:基于迁移学习的多任务制导算法实现了导弹在不同任务中的加速度指令预测。

攻击机动目标的多导弹分布式协同制导律

在分布式通信和增广比例导引的基础上,基于网络同步原理设计了一种多导弹拦截机动目标的分布式协同制导律.该分布式协同制导律由导弹本地导引律和分布式协调策略两部分组成:本地导引律由增广比例导引得到;基于网络同步原理设计了导弹之间的分布式协调策略,实现多弹对机动目标的同时协同攻击.该分布式协同制导律只需要视距内相邻导弹之间传输状态信息,具有通信量小、可扩展性好等特点.仿真结果表明多弹网络系统采用该制导律能够有效地协同攻击机动目标.

比例导引律研究现状及其发展

在研究和分析国内外大量关于比例导引律的研究成果基础上,系统地论述了纯比例导引、理想比例导引、扩展P ID型比例导引,变结构比例导引和修正比例导引等改进型比例导引的研究现状。分析了它们的原理、性能及特点,讨论了在理论与实践中的优缺点,对其发展进行了展望,为导弹制导与控制及相关问题研究提供了参考。

激光末制导炮弹比例导引律性能研究

通过分析激光末制导炮弹比例导引制导律几何关系和原理,研究了激光末制导炮弹的工作原理,建立了含非线性进动的激光陀螺式导引头动力学模型和线性化弹体模型的末制导炮弹比例导引回路模型。根据制导控制系统是否利用导引头非线性进动信号作为制导信号,建立了两种制导控制系统数学模型,采用不同的制导时间以及弹体参数,进行了末制导炮弹比例导引回路仿真,对比分析了两种制导控制系统的性能。最后,针对非线性系统易引起弹体姿态较大摆动的问题,提出了在制导时间有限时的基于衰减系数的激光末制导炮弹改进制导方案。

带过重力补偿的比例导引制导律参数设计与辨识

研究以增大导弹落角为目标的过重力补偿比例导引律及其工程应用。将重力补偿信号作为加速度补偿指令引入线性比例导引回路,从而建立了仿真数学模型,分析了重力补偿系数、比例导引系数等参数对制导系统性能的影响,并根据遥测数据给出实际系统的过重力补偿系数、比例导引系数的辨识结果从而验证了理论设计结果,得到过重力补偿比例导引律的一般设计方法。

组合末制导律及其仿真研究

传统的比例导引具有很多缺点。第一,在攻击大机动目标时的控制精度差;第二,对导弹的需用过载要求大;第三,增益高的比例导引中任何噪声都会使它的性能大大降低。为了克服上述缺点,提出了一种简单有效的组合末制导规律,即比例加追踪法的组合末制导规律,它继承了纯比例导引所具有的简单、易于实现的优点,同时还可以降低导弹的需用过载,提高导弹的打击精度。经过大量的数字仿真,所提出的组合末制导律,导弹的平均过载和峰值以及终端脱靶量均小于比例导引。

制导规律研究现状及展望

系统地论述了古典和现代的一些制导规律的研究现状 ,分析了它们的优点和不足、适应范围和局限性 ,并讨论了它们的发展前景。其中着重论述了比例导引、最优导引、微分对策和近年来热门研究的神经网络控制。可供从事导弹制导的科研人员对现有制导规律有一个全面的认识 ,以便于更深入地研究。

一种新型的组合末制导律研究

本文提出了一种简单有效的组合末制导律 ,即比例加追踪法的组合制导律 ,它具有纯比例导引的简单、易于实现的优点 ,同时可以降低导弹的需用过载。大量的数字仿真结果表明 ,应用本文提出的组合末制导律 ,导弹的平均过载和峰值过载均小于比例导引 。

角速度零位误差对比例导引制导精度的影响

主要研究了角速度零位误差对比例导引制导精度的影响,首先建立了比例导引制导回路的简化模型,并得到了简化模型的解析解。根据解析解定义了导弹无量纲位置误差和无量纲法向加速度,并分析了比例导引系数对以上两个无量纲量的影响。详细研究了更复杂的比例导引制导回路模型,给出了在比例导引制导回路各参数取不同值的情况下,导弹无量纲法向加速度随时间的变化曲线。最后运用仿真手段研究了在比例导引制导回路各参数取不同值的情况下,导弹无量纲脱靶量随无量纲末制导时间的变化情况,仿真结果表明:角速度零位误差引起的无量纲脱靶量随着无量纲末制导时间的增加而收敛。

高超声速飞行器自适应BTT末制导律

针对高超声速飞行器末段打击问题,提出了具有自适应能力的比例导引律,实现了高精度定点定向打击.通过纵向进入条件的分析,设计了一种提高比例导引进入时间的收敛策略,从而使飞行器有足够的时间进行横纵向调整,保证了落角落点约束的满足.由于制导参数直接影响落点和落角的精度,而飞行器动力学约束会导致无法理想响应制导指令,因此设计了一种闭环非线性自适应律,通过在线选择制导参数确保飞行器以高精度命中目标.仿真结果表明了该末制导方法的有效性和准确性.

几种增大空地导弹落角的制导方式比较

为提高对加固掩体、地下指挥控制中心等硬目标的毁伤效果,新型空地导弹需要使用较大的落角攻击目标。文中首先对两种可实现大落角攻击的制导方式——过重补比例导引和Zarchan弹道成型制导律进行了仿真分析,并提出了"Zarchan弹道+比例导引"组合制导方式,仿真分析表明,该组合制导方式可实现接近90°的末端落角,比较适合于防区外发射空地导弹和高空投放的制导炸弹。

侵彻制导武器终端多约束最优制导律

针对侵彻型制导武器终端多约束制导问题,建立包含一阶弹体动力学的多约束制导模型,采用最优控制方法推导终端多约束制导问题的通解。根据终端约束与罚函数的关系,求解了实际工程中关心的3种最优制导律,并结合当前主流制导控制体制,在小角假设下,将其表述为便于工程实现的形式,即包含弹体动力学的最优比例导引制导律、弹道成型最优制导律和终端多约束最优制导律。理论分析和典型末制导条件仿真表明,后2种制导律可为侵彻制导武器终端位置与角度双约束以及位置、角度与加速度多约束制导问题提供理论基础和工程应用参考。