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In order to guide itself to a successful target intercept, the missile must obtain information about the target. The prelaunch and post launch information too, must be gathered. Before a missile is launched, that is, during the prelaunch phase, the missile needs to know where to go. It knows that it is supposed to go to the target, but it must be told where the target is. The missile is told where the target is by electrical signals entering through the umbilical from the launcher. These signals are head aimed (to point the missile head at the target), English bias (to point the missile at the intercept point), and an estimate of true target Doppler on the simulated Doppler line. [ABSTRACT FROM AUTHOR]

Unlike the traditional ways adopted by most of the researchers, the rotating angle of relative velocity is adopted in this work as the independent variable to analyze the analytical solution of impact angle, capture region (with or without finite line of sight (LOS) turn rate constraint), range to go, time to go, and a defined fuel cost of a traditional pure proportional navigation (PPN) guidance law for both nonmaneuvering moving, and stationary targets. [ABSTRACT FROM PUBLISHER]

This paper considers integrated guidance and control system with input saturation for homing missiles based on the three channels independence design idea. An auxiliary system is first introduced into the integrated guidance and control model for the pitch channel. Then, a smooth function and a Nussbaum type function are employed in the dynamic surface control design to deal with the problem of input saturation constraint. The proposed control scheme guarantees the desired interception performance and the stability of the missile dynamics. The similar control scheme is further applied to the yaw channel and roll channel respectively. Nonlinear 6-DOF simulations are carried out to demonstrate the effect of the approach. [ABSTRACT FROM AUTHOR]

Six-degree-of-freedom ballistic trajectory model can be used to verify design indicators, performance parameters, and correctness so they have been put forward in the missile argumentation and design process. Mathematical simulation is the basis of late semiphysical simulation and field firing testing, and it can greatly reduce the field shooting test time, which shortens the missile test cycle and so reduces costs. [ABSTRACT FROM AUTHOR]

This paper proposes a novel composite integrated guidance and control (IGC) law for missile intercepting against unknown maneuvering target with multiple uncertainties and control constraint. First, by using back-stepping technique, the proposed IGC law design is separated into guidance loop and control loop. The unknown target maneuvers and variations of aerodynamics parameters in guidance and control loop are viewed as uncertainties, which are estimated and compensated by designed model-assisted reduced-order extended state observer (ESO). Second, based on the principle of active disturbance rejection control (ADRC), enhanced feedback linearization (FL) based control law is implemented for the IGC model using the estimates generated by reduced-order ESO. In addition, performance analysis and comparisons between ESO and reduced-order ESO are examined. Nonlinear tracking differentiator is employed to construct the derivative of virtual control command in the control loop. Third, the closed-loop stability for the considered system is established. Finally, the effectiveness of the proposed IGC law in enhanced interception performance such as smooth interception course, improved robustness against multiple uncertainties as well as reduced control consumption during initial phase are demonstrated through simulations. [ABSTRACT FROM AUTHOR]

Novel sliding-mode control laws are proposed in this paper to stabilize a class of uncertain nonlinear systems. Considering that the sign function, which is often used in the sliding-mode control can cause chattering of control input, the saturation function is used to take its place in proposed controllers. What is more, not only the proposed controllers are continuous but also convergence to the origin asymptotically and in finite time can both be guaranteed in theory. For asymptotic convergence, it is only required that uncertainties and disturbances are bounded and the bounds may be unknown by virtue of adaptive laws. The obtained results are then applied to partial integrated guidance and control design for a missile, completing desired terminal impact angle and hit-to-kill interception. Finally, simulations are conducted on the nonlinear longitudinal missile model and results demonstrate the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

Two types of autopilot-guidance schemes for linear quadratic multi-input missile systems are compared: single-and two-loop configurations. In both cases, all the states are fed back into the guidance loop. In the two-loop case, the inner autopilot loop is designed separately from the outer guidance loop, whereas in the single-loop scheme, the guidance command is injected directly to the actuator, without a definite autopilot. It is proven that the two schemes achieve the same performance if and only if the number of guidance commands is identical to the number of available controllers. Thus, using the full-state two-loop scheme may be preferred, because it ensures a stabilized airframe whether the guidance loop is active or not. The theorem and guidance laws are illustrated using a dual-control missile model. [ABSTRACT FROM AUTHOR]

Based on the discrete form of the target-missile relative motion equations in plane, a discrete sliding-mode guidance (DSMG) law is proposed. All previous missile seeker's measurements are used in the design of the DSMG law to estimate the target acceleration such that noises in the seeker's measurements are effectively being smoothened. It is proved that the proposed DSMG law is finite time convergent. Quasi sliding-mode bands of the DSMG law are discussed, and the formula for calculating the terminal miss distances of the missile under the DSMG law are presented. Simulation results from a space interception process verify the effectiveness of the proposed method. [ABSTRACT FROM AUTHOR]

According to the characteristics of the target ship in motion in the net recovery process of UAV, this paper put forward the guidance law based on angle of sight by taking missile guidance law of proportional guidance law for reference, and introduced the backstepping guidance law to improve the adaptability of guidance law. Since the angle of sight guidance law is to make the UAV flight path angle and angle of sight be proportional to the rate of change, flight path angle can be tracked by controlling the angle of sight. That taking the guidance law into use can reduce the sensitivity of UAV motion on the ship parameters change, thereby to obtain more stable trajectory. The simulation shows the feasibility of this guidance law, and has strong robustness. [ABSTRACT FROM AUTHOR]

The disturbance rejection rate ( DRR) of seeker has the effect on the output accuracy of seeker and the guidance performance. An online inhibition method based on disturbance observer is proposed for the problems of the guidance precision of missile caused by DRR. The errors of seeker models, the output frequency of detector and the effect of measurement noise on DRR inhibition are presented. The result shows that the spring torque and the damping torque are the main factors of the DRR generation, and the missile guidance precision would be reduced. The online inhibition method based on disturbance observer can inhibit DRR effectively; but the errors of seeker models, the output frequency of the detector and the measurement noise reduce the inhibition effects to some extent. [ABSTRACT FROM AUTHOR]

This paper investigates the development and implementation of a new homing-missile guidance system for intercepting highly maneuverable targets using arbitrary features of the homogeneous high-order sliding-mode controllers and observers. The concept of this guidance system involves artificially increasing the orders of controllers and observers to enhance accuracy and robustness of future interceptors. Two advanced guidance laws, that is, augmented proportional navigation law and direct collision or hit-to-kill law, are considered to construct acceleration-based and velocity-based attractive sliding manifolds, respectively; and high-order robust exact differentiators are used to compute high-order time derivatives of these manifolds. Real-time estimation of target maneuver from measurements using these differentiators is also considered. The resulting guidance laws are implemented in computer simulation using nonlinear interceptor dynamics and engagement kinematics. The results prove that the guidance system achieves high-level performance and successfully meets direct collision guidance requirements. In addition, this system shows good adaptability and strong robustness to target evasive maneuvers, parameter uncertainties, and measurement noises. [ABSTRACT FROM AUTHOR]

The implementation of missile guidance laws on collision course trajectories with bearings-only measurements may result in bad range observability. This in turn may result in poor engagement performance for those guidance algorithms that exploit range information in their steering law. Maneuvering away from the collision course trajectory improves range observability. To propose a sizing criterion for this maneuver, this work presents a new guidance strategy that exploits the information from the error covariance matrix of the homing loop integrated Kalman filter in the framework of a pursuit–evasion game. The new strategy has been tested in several scenarios against other guidance laws, such as the classical formulation of the pursuit–evasion game and proportional navigation. Numerical simulations on a set of Monte Carlo samples show that the proposed solution is able to improve engagement performance in terms of both miss distance and range observability. [ABSTRACT FROM AUTHOR]

The article explores a passive guidance law against surface targets only with the line-of-sight (LOS) information applied to angle control in vertical plane. Topics discussed include the schematic view of engagement geometry in the vertical plane, the true passive guidance law with only LOS information and the results of the comparative simulation between the proposed method and the mathematical model describing the dynamics of a flight vehicle around its pitch axis.

A linear quadratic based missile guidance law is presented that imposes a predetermined terminal body angle. Solutions are provided for arbitrary order autopilot dynamics, with explicit analytical expressions given for first-order and zero-lag dynamics. Performance is demonstrated through simulation in a planar interception scenario. [ABSTRACT FROM AUTHOR]

This paper presents an interactive multiple-model-based guidance law estimation filter. An antiship missile's threat index, such as the remaining time of flight and impact angle, can be estimated by using the guidance law estimation filter. The filter bank of the filter was composed of models based on proportional navigation and impact angle control guidance laws. In the guidance law estimation filter, state variables, rather than fixed values of navigation constant or impact angle, are used for estimation. To that end, the modified interactive multiple model, which is a modified version of the conventional interactive multiple model method is presented in this paper. The modified interactive multiple model method minimizes the degradation of estimation performance by systematic approach for different state vectors. The effectiveness of the modified interactive multiple model method and the performance of the guidance law estimation filter are examined through the numerical simulation of various engagement scenarios. [ABSTRACT FROM AUTHOR]

A multi-criteria decision making (MCDM) method is proposed for the design parameter optimization of missile guidance and control system with the considerations of system performance, cost and other factors. The optimal parameters of the guidance system can be easily obtained by transforming the parameter optimization problem into a MCDM problem. Proportional navigation guidance system and the technique for order preference by similarity to ideal solution (TOPSIS) are used to show the effectiveness of the proposed method. In the selection of decision-making criteria, both system performance and cost are considered. Also, the influence of different weights for the criteria on decision making is studied. The results show that the method is able to select the optimal design alternative in multi-criteria sense. [ABSTRACT FROM AUTHOR]

The extended weighted and object functions are proposed based on the time-to-go exponential function. The optimal control theory is used to deduce a family of extended optimal trajectory shaping guidance laws for the constant maneuvering target. According to Schwartz inequality, the analytical solution of the guidance law acceleration command is derived by introducing the initial displacement, initial heading error, target maneuver and final impact angle into the lag-free guidance system. The analysis shows that the final acceleration command approaches to zero when the exponent of the time-to-go exponential function is greater than zero. The non-dimensional position and angle miss distance of guidance system with first order lag are studied using the non-dimensional method and the adjoint method. The results show that the position and angle miss distance induced by the heading error and final impact angle approach to zero when the missile terminal guidance time is about 15 times of the system lag time constant. And also, the position and angle miss-distance are smaller when the signs of initial heading error angle and final impact angle are opposite. [ABSTRACT FROM AUTHOR]

The signal received by the missile seeker is processed in the monopulse receiver to determine the direction (bearing and elevation) of the target. The sum and difference voltages generated by the signal are electronically manipulated to obtain the boresight error (angle) to the missile seeker. This angle is processed to provide appropriate guidance to the missile. In the presence of multipath in sea environment, this signal is a combination of the direct target return and a return seemingly emanating from the target image beneath the sea surface. This concept of multipath is utilized in terrain bounce jamming to jam the missile seeker. This paper focuses on the study of the effect of terrain bounce jamming on missile guidance based on analysing the boresight error to the missile seeker. [ABSTRACT FROM AUTHOR]

The paper presents an overview of the different errors created in an interception endgame by the measurement noise and the need of using an estimator in the guidance loop. Approaches for reducing the effects of the estimation error are described and some directions for further investigations are pointed out. [ABSTRACT FROM AUTHOR]

A novel formulation of sliding mode control (SMC) based proportional navigation (PN) guidance law is presented. Unlike conventional SMC-based guidance laws, the law presented here does not need any knowledge of bounds of target acceleration. The target acceleration is estimated using the so-called inertial delay control (IDC). Closed-loop stability for the guidance loop is established. Simulations are carried out by considering highly-maneuvering targets and constant as well as varying missile velocity and the results are presented to demonstrate the effectiveness of the proposed formulation. [ABSTRACT FROM AUTHOR]

Abstract: In directed infrared countermeasures (DIRCM), the purpose of jamming toward missiles is making missiles miss the target (aircraft of our forces) in the field of view. Since the DIRCM system directly emits the pulsing flashes of infrared (IR) energy to missiles, it is more effective than present flare method emitting IR source to omni-direction. In this paper, we implemented a reticle seeker simulation tool using MATLAB-SIMULINK, in order to analyze jamming effect of spin-scan and con-scan reticle missile seeker used widely in the world, though it was developed early. Because the jammer signal has influence on the missile guidance system using its variable frequency, it is very important technique among military defense systems protecting our forces from missiles of enemy. Simulation results show that jamming effect is greatly influenced according to frequency, phase and intensity of jammer signal. Especially, jammer frequency has the largest influence on jamming effect. Through our reticle seeker simulation tool, we can confirm that jamming effect toward missiles is significantly increased when jammer frequency is similar to reticle frequency. Finally, we evaluated jamming effect according to jammer frequencies, by using correlation coefficient as an evaluation criterion of jamming performance in two reticle missile seekers. [Copyright &y& Elsevier]

The present paper focuses mainly on the semi-strapdown homing guided system design, which includes control system design of the two-degree-of-freedom semi-strapdown homing seeker and estimation of the guidance information. According to Euler dynamics theory and momentum moment theorem, the two-degree-of-freedom nonlinear dynamic model of the semi-strapdown homing seeker is built in common coordinate frames. Then the control-oriented state-space model of the semi-strapdown homing seeker system is proposed innovatively, and a nonlinear controller is designed by integrating the feedback linearization and linear lead controller. Next, the estimation and construction method of the guidance information based on the unscented Kalman filter is presented for the semi-strapdown homing guided missile. Finally, a three-dimensional interception flight simulation is made using the six-degree-of-freedom axial-symmetry nonlinear aerodynamic model of the semi-strapdown homing guided missile. Simulation results show that guidance information of the semi-strapdown homing guided missile can be estimated via the unscented Kalman filter, and can be used in shaping seeker control command. The seeker nonlinear controller based on feedback linearization can improve the disturbance rejection effect of the semi-strapdown homing seeker effectively with good target tracking ability. [ABSTRACT FROM PUBLISHER]

This article considers the problem of optimal guidance laws for cooperative attack of multiple missiles based on the optimal control theory. New guidance laws are presented such that multiple missiles attack a single target simultaneously. Simulation results show the effectiveness of the proposed algorithms. [ABSTRACT FROM PUBLISHER]

In this paper, sliding-mode-control-based guidance laws to intercept stationary, constant-velocity, and maneuvering targets at a desired impact angle are proposed. The desired impact angle, which is defined in terms of a desired line-of-sight angle, is achieved in finite time by selecting the missile's lateral acceleration to enforce terminal sliding mode on a switching surface designed using nonlinear engagement dynamics. The conditions for capturability are also presented. In addition, by considering a three-degree-of-freedom linear-interceptor dynamic model and by following the procedure used to design a dynamic sliding-mode controller, the interceptor autopilot is designed as a simple static controller to track the lateral acceleration generated by the guidance law. Numerical simulation results are presented to validate the proposed guidance laws and the autopilot design for different initial engagement geometries and impact angles. [ABSTRACT FROM AUTHOR]

A new derivation of an optimal missile guidance law is presented. The guidance law derived is an extension of previously published versions of explicit guidance in that gravity compensation is incorporated in an optimal fashion. Performance of the extended guidance law is demonstrated via simulation comparisons with generalized vector explicit guidance and Kappa guidance, both of which have appeared previously in the literature. The new law is then incorporated into a hybrid guidance scheme that allows the specification and satisfaction of a desired time-of-arrival constraint. Performance of this capability is also demonstrated via simulation examples. Finally, an approach to coordinating desired times of arrival among a group of cooperating missiles to achieve simultaneous target intercepts or missile rendezvous is presented. [ABSTRACT FROM AUTHOR]

Abstract: A novel nonlinear missile guidance law against maneuvering targets is designed based on the principles of partial stability. It is demonstrated that in a real approach which is adopted with actual situations, each state of the guidance system must have a special behavior and asymptotic stability or exponential stability of all states is not realistic. Thus, a new guidance law is developed based on the partial stability theorem in such a way that the behaviors of states in the closed-loop system are in conformity with a real guidance scenario that leads to collision. The performance of the proposed guidance law in terms of interception time and control effort is compared with the sliding mode guidance law by means of numerical simulations. [Copyright &y& Elsevier]

This article introduces an adaptive guidance law for intercepting a manoeuvring target using information from an imaging seeker. The relative missile-target kinematics is represented in the line-of-sight (LOS) frame, and the image information - instead of the relative range - is introduced to formulate the interception problem as an adaptive control problem with an unknown control coefficient. Considering the target's acceleration as a time-varying bounded external disturbance, an estimator with an adaptive property is derived. It is proven that the proposed guidance law can guarantee the LOS angular rates global uniform ultimate boundedness during the terminal course. The effectiveness of the proposed guidance law has been demonstrated in simulations for two kinds of typical manoeuvring targets. [ABSTRACT FROM AUTHOR]

A three-body interception scenario is considered, where an aircraft launches a defending missile as a counterweapon against an incoming, attacking missile. Line-of-sight guidance is investigated as a prospective strategy for the defender missile, with the aircraft being the moving launch platform. Kinematic results on line-of-sight guidance with a moving/maneuvering platform are derived, showing defender-attacker interception for all attack geometries and attacker maneuvers. The speed and maximum terminal acceleration requirements are shown to be considerably lower when compared with a proportional navigation guided defender missile. It is also shown that the attacker pays maximum penalty for an evasive maneuver from the defender if the defender uses line-of-sight guidance. A cooperative guidance law is proposed for the aircraft to maximize the attacker-to-defender lateral acceleration ratio. Simulations are carried out for various attack geometries, with and without the cooperative strategy, showing better relative control effort performance for the former. [ABSTRACT FROM AUTHOR]

To date many studies of missile and aircraft differential games seem to have tried to find a unique solution, however, the authors of this study found there are many solutions of comparable values to the game. The authors called them "a family of local solutions." This paper shows a method to find the family for a missile and an aircraft differential game, which minimizes and maximizes the essential payoff of the problem, the miss distance, without employing any linearized approximation. For this purpose, the authors introduced the concept of energy maneuverability. The sets of optimal terminal surfaces of both vehicles that maximize their velocities are obtained first. Next, minimax solutions of the vehicles that start from these terminal surfaces, and minimize and maximize the miss distance, respectively, are calculated; then a local minimax family in this game is obtained. The proposed method can solve minimax problems under several vehicles' acceleration constraints. The results may be able to be incorporated into an actual missile guidance system as a knowledge base, and then the performance will be significantly improved. [ABSTRACT FROM AUTHOR]

A new guidance concept is proposed. Implementing it enables intercepting a maneuvering target while also imposing, from early on in the engagement, a predefined angle relative to the target's velocity vector. The scenario of interest is aerial interception between a missile and a maneuvering target. The guidance concept is applicable in all aerial interception geometries: namely, head-on, tail-chase, and the novel head-pursuit. Analytical conditions for existence of these different engagement geometries are derived. The guidance concept is implemented using the sliding-mode approach. The common assumption of flight along an initial collision triangle is not taken, and thus the guidance law is applicable for both midcourse and endgame guidance. The application in the different engagement geometries is studied via simulation. It is shown that the head-on scenario allows the smallest range of intercept angles. It also places the most severe maneuverability requirements on the interceptor. Thus, in some cases, tail-chase or head-pursuit engagements should be considered instead. The choice between the two is dependent on the adversary's speed ratio; for tail-chase, the interceptor must have a speed advantage over its target, while for head-pursuit, it must have a speed disadvantage. [ABSTRACT FROM AUTHOR]

The purpose of this paper is to develop an optimal guidance law that takes into account both the missile acceleration limit and a maximum value allowed for the guidance gain. The maximum value of the guidance gain is dictated mainly by the possible coupling between the missile body motions and the measurement of the inertial target line-of-sight angle rate. The tools for this work are the linear quadratic stochastic Gaussian optimal control theory coupled with the use of random-input describing functions. Although the use of the random-input describing function introduces an approximation in the work, it has been shown in the past that the use of random-input describing functions in miss distance calculations results in good approximations for the predicted miss distance. The paper includes the derivation of the optimal guidance law that accounts for both the missile acceleration limit and the constraint on the guidance gain and a simple one-dimensional Monte Carlo simulation to examine the performance of the guidance law. [ABSTRACT FROM AUTHOR]

In this paper, a strength Pareto evolutionary algorithm based approach is proposed for designing a multistage fuzzy-based guidance law which consists of three fuzzy controllers. Each of these controllers is activated in a region of the interception. The distribution of the membership functions and the rules are obtained by solving a nonlinear constrained multiobjective optimization problem where final time, energy consumption, and miss distance are treated as competing objectives. A hierarchical clustering technique is implemented to provide the decision maker with a representative and manageable Pareto-optimal set without destroying the characteristics of the trade-off front. Moreover, a fuzzy-based mechanism is employed to extract the best compromise solution over the trade-off curve. The simulation results show that the proposed design technique was able to generate a missile guidance law which has better performance than the classical proportional guidance law. [ABSTRACT FROM AUTHOR]

A method for the co-ordinated rendezvous of a team of unmanned air vehicles (UAVs) to a formation through the use of sliding mode control is presented. The focus of this work is on the scenario where three 'follower' UAVs must rendezvous in a desired formation about a 'Leader' UAV in a finite time. An assumption is made that each UAV only receives knowledge of the bounds of the manoeuvres that the other UAVs are performing. By not requiring manoeuvre information to be constantly available, a level of robustness to possible communication outages between UAVs is obtained. The proposed method also guarantees that the formation will be achieved in a finite time through the use of a concept of sliding mode terminal guidance (SMTG). SMTG takes advantage of the finite-time reaching phase of sliding mode to ensure that any desired constraint can be fulfilled in a finite time. [ABSTRACT FROM AUTHOR]

Abstract: Along with the development of the optoelectronic interference technology, invisible technology and anti-radiation missile technology, missile guidance with single band or single mode cannot adapt for modern warfare. So the technology of combined guidance is proposed. In the field of combined guidance, MMW/IR combined guidance is more significant. The critical component of the MMW/IR semi-physical simulation will be discussed in this paper, and the selection principle of the substrate and coating materials used for dichroic beam combiner (DBC) is analyzed: MMW/IR beam combiner. By researching the method of the beam combiner of 8–18GHz MW and 3–5μm IR, the theory and simulation are discussed in detail in this paper. And the result is given. At the same time, the distortion effect on the MMW signal is analyzed theoretically and the method is presented. Substrate material is selected. The main problems and study factors include: MMW transmission, insertion phase delay, size of substrate material, and coating. At last, the substrate material is determined, and the thickness of the substrate material is optimized. [Copyright &y& Elsevier]

In this study, it is aimed at comparing the performance of the guidance methods called Linear Homing Guidance (LHG), Parabolic Homing Guidance (PHG), and Proportional Navigation Guidance (PNG) on a missile with two relatively rotating parts against a surface target. In this extent, first the dynamic model of the missile is constructed. Next, the guidance laws LHG, PHG, and PNG are formulated. Modelling the target kinematics as well, the entire guidance and control system is built by integrating all the models mentioned above, and the relevant computer simulations are carried out. Consequently, the simulation results are discussed and the study is evaluated. [ABSTRACT FROM AUTHOR]

The article proposes the cooperative proportional navigation (CPN) homing guidance law for many-to-one engagements. The formulation of the homing problem of multiple missiles against a single target is explained A discussion on the preliminary concepts is presented. The article also characterizes the homing guidance law for the case of two missile.

The impact time control guidance (ITCG) method, which has been proposed recently, can be applied successfully to a salvo attack of multiple missiles. Compared to the proportional navigation guidance law, this guidance method makes additional manoeuvres to synchronize the impact times. However, such manoeuvres do not consider the manoeuvrability and the seeker's field-of-view (FOV) of a missile and may cause the target to move out of the missile seeker's FOV; maintaining the seeker lock-on condition during the engagement is critical for missile guidance. To solve this problem, two methods are presented in this article: one is based on the calculation of minimumandmaximumflight times considering the missile's manoeuvring limit and the seeker's FOV limit to check the available impact time. The other is based on guidance law switching logic that keeps the target look angle of the seeker constant. These methods can provide the boundary limit of the impact time of the salvo attack and prevent the lock-on failure because of the seeker's FOV limit of the missile during the homing phase when the ITCGis used. This method was applied to the case of a time critical salvo-attack of multiple missiles, which have manoeuvring limit and the seeker's FOV limit, and desired results were obtained. [ABSTRACT FROM AUTHOR]