Sliding mode control based guidance law with impact angle constraint

The terminal guidance problem for an unpowered lifting reentry vehicle against a sta- tionary target is considered. In addition to attacking the target with high accuracy, the vehicle is also expected to achieve a desired impact angle. In this paper, a sliding mode control （SMC）-based guidance law is developed to satisfy the terminal angle constraint. Firstly, a specific sliding mode function is designed, and the terminal requirements can be achieved by enforcing both the sliding mode function and its derivative to zero at the end of the flight. Then, a backstepping approach is used to ensure the finite-time reaching phase of the sliding mode and the analytic expression of the control effort can be obtained. The trajectories generated by this method only depend on the initial and terminal conditions of the terminal phase and the instantaneous states of the vehicle. In order to test the performance of the proposed guidance law in practical application, numerical simulations are carried out by taking all the aerodynamic parameters into consideration. The effec- tiveness of the proposed guidance law is verified by the simulation results in various scenarios.     

基于滑模控制的剖面跟踪制导律

针对高超声速飞行器滑翔段的高精度制导问题，考虑复杂多约束条件以及干扰和不确定因素的影响，设计了一种基于全局积分滑模控制的剖面跟踪制导方法。首先，将多重约束转化为阻力加速度-速度（D-V）平面内的再入走廊；然后，以终端精度和总吸热量为性能指标，采用分段函数的形式优化设计出一条标准D-V剖面；再基于简化的动力学模型，推导得到关于阻力加速度微分和速度的二阶非线性模型；最后，基于滑模控制理论，设计全局积分滑模面和指数趋近律，获得控制量幅值大小，并结合侧向方位误差走廊确定控制量符号，从而实现对标准剖面的有效跟踪。采用CAV-H滑翔再入模型进行数值仿真，分析验证了提出的基于滑模控制的剖面跟踪制导律具有较好的跟踪性能和精度。     

总能量控制在RLV自动着陆中的应用

为了解除重复使用运载器(RLV)自动着陆段飞行速度和航迹之间的非线性耦合,设计了基于总能量原理的纵向飞行控制律。该原理的核心思想是通过RLV的阻力板改变阻力来控制总能量,用升降舵调节总能量的分配,通过协调阻力板和升降舵指令,达到合理调节和分配能量的目的,确保控制系统具有良好的非线性解耦控制能力和较强的鲁棒性。仿真结果表明,基于总能量的飞行控制系统,实现了RLV自动着陆段高度和速度的解耦控制,满足不确定条件下RLV自动着陆的要求。     

自适应严格收敛非奇异终端滑模制导律

针对机动目标的末制导拦截问题，设计了一种带终端角度约束的有限时间收敛终端滑模制导律。首先，分析了现有非奇异终端滑模制导律存在的滑模面不能严格有限时间收敛的问题，进而构造了一种新型的非奇异终端滑模面。其次，设计了一种对目标机动上界的自适应估计，提出了一种自适应严格收敛非奇异终端滑模制导律的设计方法。最后，基于Lyapunov稳定性理论，证明了设计的制导律能够使得制导系统在有限时间内收敛到零，并且保证了滑模面在收敛过程中不存在非收敛因子，是严格有限时间收敛的。仿真实验验证了该制导律能够有效地拦截机动目标，同时和与现有的非奇异终端滑模制导律以及基于转换滑模面的非奇异制导律相比，拦截时间更短，终端攻击角度精度更高，导弹机动消耗的能量更少。     

舰载机理想着舰点垂直运动的预估与补偿

理想着舰点的垂直运动是影响着舰精度和安全的一个主要因素,因此必须加强舰载机对理想着舰点垂直运动的同步跟踪能力.为此,提出对理想着舰点垂直运动的位置和速度信号进行预估和补偿的方法,将垂直运动的位置和速度信号经过预估和补偿后分别引入到纵向自动着舰引导系统和飞控系统的垂向速度通道中,使得舰载机可以准确跟踪理想着舰点的垂直运动,以减小甲板运动对着舰的影响.针对不同海况条件,对设计的补偿器和预估器进行仿真验证,并与其他方法进行比较.结果表明本文提出的理想着舰点垂直运动预估与补偿方法可有效地补偿由甲板运动引起的着舰误差,显著提高了着舰的安全性和精确性.     

考虑驾驶仪动态特性的固定时间收敛制导律

针对打击机动目标的制导问题，设计了一种同时考虑攻击角度约束、自动驾驶仪动态特性和固定时间收敛的新型制导律。首先，基于非奇异终端滑模控制和固定时间稳定性理论，采用反步递推方法设计制导律。在制导律设计过程中，设计了一种固定时间收敛的非奇异终端滑模面，基于固定时间控制和滑模控制，设计虚拟控制律，构造一种非线性一阶滤波器解决传统反步设计中的"微分膨胀"问题。基于超螺旋算法和固定时间稳定性理论，设计了一种固定时间收敛的滑模干扰观测器，用于估计目标机动等干扰。然后，基于Lyapunov稳定性理论，对制导律的固定时间稳定性进行了证明，并给出了收敛时间的表达式。最后，通过仿真分析，验证了所提制导律的有效性，和现有制导律相比，所提制导律具有较高的制导精度和角度约束精度、较快的系统收敛速度以及较少的能量消耗。     

Accelerated Landweber iteration based control allocation for fault tolerant control of reusable launch vehicle

This paper presents a novel Fault Tolerant Control(FTC) scheme based on accelerated Landweber iteration and redistribution mechanism for a horizontal takeoff horizontal landing reusable launch vehicle(RLV). First, an adaptive law based on fixed-time non-singular fast terminal sliding mode control(NFTSMC), which focuses on the attitude tracking controller design for RLV in the presence of model couplings, parameter uncertainties and external disturbances, is proposed to produce virtual control co...     

基于光学导引的无人机自动着陆控制系统设计

针对目前无人机采用的某些着陆导引方式存在的覆盖面积小、易受干扰等缺点,设计了基于光学导引系统的无人机自动着陆控制系统.建立了某型无人机的纵向和横侧向的线性方程,结合光学导引系统的性能指标提出了光学导引自动着陆系统的总体结构,设计了理想下滑轨迹和自动着陆控制律方案,将光学导引系统和自动着陆控制系统结合在一起,运用经典PI...     

带有双闭环滤波器的有限时间稳定变结构制导律

 在末制导过程中,为了取得较高的命中精度,制导律必须使视线(LOS)角速率较快收敛。为了提高制导律的性能,在假设未知不确定时变项有界的前提下,对带有双闭环滤波器的变结构制导律进行了研究。提出的变结构制导律能够使视线角速率及其一阶导数在有限时间内收敛至滑动模态域内,得到更加精确的等效控制并有效克服干扰的影响。通过引入高阶滑动模态,对有关定理进行了重新证明,简化了证明过程,加强了定理的结论,弱化了定理成立的条件。最后利用数值仿真验证了所研究方法的有效性。     

Three-dimensional guidance law based on adaptive integral sliding mode control

For the terminal guidance problem of missiles intercepting maneuvering targets in the three-dimensional space, the design of guidance laws for non-decoupling three-dimensional engage-ment geometry is studied. Firstly, by introducing a finite time integral sliding mode manifold, a novel guidance law based on the integral sliding mode control is presented with the target acceler-ation as a known bounded external disturbance. Then, an improved adaptive guidance law based on the integral sliding mode control without the information of the upper bound on the target accel-eration is developed, where the upper bound of the target acceleration is estimated online by a designed adaptive law. The both presented guidance laws can make sure that the elevation angular rate of the line-of-sight and the azimuth angular rate of the line-of-sight converge to zero in finite time. In the end, the results of the guidance performance for the proposed guidance laws are pre-sented by numerical simulations. Although the designed guidance laws are developed for the con-stant speed missiles, the simulation results for the time-varying speed missiles are also shown to further confirm the designed guidance laws.     

ILS-a safe bet for your future landings

An argument is presented in favor of the ILS (instrument landing system), the principal components of which are the localizer, which provides lateral guidance to the pilot, and the glide slope, which gives the vertical guidance. A pilot flying these signals precisely will end up over the hard surface runway with the ILS receiving antenna, wherever it is mounted on the aircraft, being about 20 ft above the touchdown point. The safety record and evolution of the ILS are examined, and a pilot's view of ILS (that of the author, an engineer who has worked with ILS since 1957 and a pilot who has used ILS safely for over 30 years) is given. System availability is discussed, and comments concerning future use are offered     

Saturated super-twisting sliding mode missile guidance

This paper concentrates on developing a missile terminal guidance law against a highly maneuvering target whose maneuvering acceleration is very close to that of the missile or even exceeds the missile normal acceleration in a finite period of time. A new saturated super-twisting algorithm is proposed and applied to the design of missile guidance law. The proposed algorithm has the advantages of simple structure, easy parameter tuning rules and a full utilization of the limit control input. The designed saturated super-twisting sliding mode guidance law is then employed in a missile guidance system. Simulation and its superior performance against strong maneuvering targets is demonstrated.     

Autonomous navigation and guidance for landing on asteroids

This paper presents an autonomous landing navigation and guidance scheme for future asteroid sample return mission. An autonomous navigation scheme based on feature tracking technology is brought out firstly; secondly, desired descent landing trajectories with the initial and terminal constraints are planned in order to achieve arrive-at-time landing on an asteroid; then, two guidance control laws, based on error phase analysis method and PD plus PWPF method respectively, are designed to track reference descent trajectory; finally, the validity of the proposed scheme is confirmed by computer simulation.     

An adaptive fast fixed-time guidance law with an impact angle constraint for intercepting maneuvering targets

Sliding mode guidance laws based on a conventional terminal sliding mode guarantees only finite-time convergence, which verifies that the settling time is required to be estimated by selecting appropriate initial launched conditions. However, rapid convergence to a desired impact angle within a uniform bounded finite time is important in most practical guidance applications. A uniformly finite-time/fixed-time convergent guidance law means that the convergence (settling) time is predefined independently on initial conditions, that is, a closed-loop convergence time can be estimated a priori by guidance parameters. In this paper, a novel adaptive fast fixed-time sliding mode guidance law to intercept maneuver targets at a desired impact angle from any initial heading angle, with no problems of singularity and chattering, is designed. The proposed guidance law achieves system stabilization within bounded settling time independent on initial conditions and achieves more rapid convergence than those of fixed-time stable control methods by accelerating the convergence rate when the system is close to the origin. The achieved acceleration-magnitude constraints are rigorously enforced, and the chattering-free property is guaranteed by adaptive switching gains. Extensive numerical simulations are presented to validate the efficiency and superiority of the proposed guidance law for different initial engagement geometries and impact angles.     

Hybridizing contemporary glide slopes to provide vertical guidancefor GPS approaches

Vertical guidance for an instrument approach to landing during conditions of reduced visibility is a crucial element with respect to safety of flight. It is noteworthy that this vertical component-to be most useful and safe-must desirably provide the pilot with no more than several feet of uncertainty. Unfortunately, with GPS the vertical portion of the position information supplied the pilot by GPS signals is the least precise because of geometries involved. Augmentation for enhancement of accuracy is quite important and the assertion is that it is necessary for totally safe vertical guidance. Evidence from approximately 60 years of experience with electronic landing systems serving the public is that there has been no aircraft accident due to a defective vertical guidance signal. Visibilities as low as 600 feet horizontally can exist at certain airports and landings can still be accomplished. These landings, while not common, are being accomplished flawlessly by contemporary aircraft and equipment. Many aspects of the contemporary UHF glide slope have been studied during its half-century of use. This paper reports on advances, some of them quite recent, that make it realistic to claim that a glide slope can be sited for Category I operation at any runway that meets the physical requirements for this type of operation. For the UHF glide slope, siting imperfections are accommodated by using one of five available types of contemporary glide slope systems or derivatives thereof. This paper reviews how accommodations are achieved in practical cases. Results from earlier tests are identified which show hybridizing of UHF glide slopes and GPS can be used to provide good approach guidance to aircraft making fully automatic approaches to touchdown     

Three-dimensional Large Landing Angle Guidance Based on Two-dimensional Guidance Laws

 Both the design process and form of the three-dimensional (3D) suboptimal guidance law (3DSGL) are very complex. Therefore, we propose the use of two-dimensional (2D) guidance laws to meet the guidance requirements of 3D space. By analyzing the relationship between the flight-path angle and its projections on OXY and OXZ planes, we obtain the ideal design requirements of the guidance laws. Based on the requirements, we design a 2D suboptimal guidance law used in the horizontal plane; combining the 2D vertical suboptimal guidance law, we create a whole ballistic simulation of six degree-of-freedom. The results are compared with those using other three guidance modes in the case of large windage of the initial azimuth angle. When the proportional navigation guidance (PNG) law is used in the horizontal planes, the landing angle will obviously decrease. With the proposed guidance mode, the large landing angle can be realized and meet the guidance precision requirements. Moreover, the required overload can decrease to meet the control requirement. The effects of the proposed guidance mode are close to that of 3DSGL despite its very simple form.     

基于蛇行变轨的反舰导弹末制导律研究

提出了一种超声速反舰导弹针对海面机动目标的主动蛇行变轨导引设计方案。首先建立了反舰导弹、拦截弹和舰艇在航向平面的运动几何关系模型,然后基于Terminal滑模控制的思想进行了制导律设计,采用约束视线角和视线角速度的方法使反舰导弹在打击舰艇的过程中实现蛇行变轨,以能够规避敌方拦截导弹。仿真结果表明,该制导律具有期望的性能要求。     

GA-Based Model Predictive Control of Semi-Active Landing Gear

Semi-active landing gear can provide good performance of both landing impact and taxi situation, and has the ability for adapting to various ground conditions and operational conditions. A kind of Nonlinear Model Predictive Control algorithm （NMPC） for semi-active landing gears is developed in this paper. The NMPC algorithm uses Genetic Algorithm （GA） as the optimization technique and chooses damping performance of landing gear at touch down to be the optimization object. The valve＇s rate and magnitude limitations are also considered in the controller＇s design. A simulation model is built for the semi-active landing gear＇s damping process at touchdown. Drop tests are carried out on an experimental passive landing gear systerm to validate the parameters of the simulation model. The result of numerical simulation shows that the isolation of impact load at touchdown can be significantly improved compared to other control algorithms. The strongly nonlinear dynamics of semi-active landing gear coupled with control valve＇s rate and magnitude limitations are handled well with the proposed controller.     

Three-dimensional time-varying sliding mode guidance law against maneuvering targets with terminal angle constraint

This paper deals with the problem of intercepting maneuvering targets with terminal angle constraints for missiles subjected to three-dimensional non-decoupling engagement geometry.To achieve the finite-time interception and satisfactory overload characteristics, a time varying sliding mode control methodology is developed based on a time base generator function. The main feature of the proposed guidance law guarantees the Line-of-Sight(LOS) angles to converge to small neighborhoods of the desir...     

A fault-tolerant neural aided controller for aircraft auto-landing

This paper presents a neural-aided controller that enhances the fault tolerant capabilities of a high performance fighter aircraft during the landing phase when subjected to severe winds and failures such as stuck control surfaces. The controller architecture uses a neural controller aiding an existing conventional controller. The neural controller uses a feedback error learning mechanism and employs a dynamic Radial Basis Function neural network called Extended Minimal Resource Allocating Network (EMRAN), which uses only on-line learning and does not need a priori training. The conventional controller is designed using a classical design approach to achieve the desired autonomous landing profile with tight touchdown dispersions called herein as the pillbox. This design is carried out for no failure conditions but with the aircraft being subjected to winds. The failure scenarios considered in this study are: (i) Single faults of either aileron or elevator stuck at certain deflections, and (ii) double fault cases where both the aileron and elevator are stuck at different deflections. Simulation studies indicate that the designed conventional controller has only a limited failure handling ability. However, neural controller augmentation considerably improves the ability to handle large faults and meet the strict touchdown dispersion requirements, thus enlarging the fault-tolerance envelope.