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This paper proposes a novel approach for guidance law design to satisfy the impact-time constraints for a certain class of homing missiles. The proposed guidance law provides proper lateral acceleration commands that make the impact time error converge to zero by the time of impact. This scheme can be applied to any existing guidance law for which a formula of predicted time to go is available. Convergence of time-to-go errors is supported by Lyapunov stability. The optimal guidance law and the impact angle control guidance law are extended by the proposed method for impact-time-control guidance and impact-time-and-angle-control guidance, respectively. The performance of the extended guidance laws is demonstrated by numerical simulation. [ABSTRACT FROM AUTHOR]

The proposed robust nonlinear disturbance observer based adaptive guidance method considers the nonlinear coupled missile dynamics and rapidly maneuvering target. It also includes unavailable information such as high-order line-of-sight rates and target acceleration as time-varying parametric uncertainties and nonparametric disturbances. An integrated guidance and control model, a robust nonlinear disturbance observer, a nonlinear missile jerk observer, and a robust nonlinear guidance law are proposed, the validity of which is demonstrated by stability analysis and simulations. [ABSTRACT FROM AUTHOR]

In air combat, an active countermeasure against an attacking missile homing into a Target aircraft entails the launch of a defending missile. The Target is protected by the Defender, which aims to intercept the Attacker before the latter reaches the Target aircraft. A differential game is presented where the Target–Defender team strives to maximize the terminal separation between the Target and the Attacker at the time instant where the Attacker is intercepted by the Defender, whereas the Attacker strives to minimize the said separation. This paper discusses the case where the Defender is endowed with a positive capture radius. Optimal strategies for the three agents are derived and simulation examples illustrate the effectiveness of the proposed approach. [ABSTRACT FROM PUBLISHER]

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]

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]

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]

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]

A new precision guidance law is presented for three-dimensional intercepts against a moving target. In contrast to previously published guidance laws, it does not require knowledge of the range to the target. This makes it appropriate for use on platforms which have an imaging device, such as a video camera, as a primary sensor. We prove that with idealized dynamic model, the guidance law results in zero miss distance, and a formula is given for impact angle error which tends to zero as does target speed, making this method particularly suitable against slow moving targets. Computer simulations are used to test the law with a more realistic model, with a video camera and optical-flow algorithm providing target information. It is shown to perform well compared with another law from the literature, despite requiring less information. [ABSTRACT FROM AUTHOR]

A novel missile guidance law that is dependent on the conditional probability density function of the estimated states is presented. The guidance law is derived by analyzing an interception scenario in the framework of an linear quadratic Gaussian (LQG) terminal control problem with bounded acceleration command. The nonlinear saturation function is represented by the equivalent random input describing function. Since for the investigated problem the certainty equivalence property is not valid, the resulting controller depends on the measurement noise level and on the saturation limit. In comparison to the classical optimal guidance law (OGL), the maximal value of the effective navigation gain is achieved during the engagement instead of near the terminal time. Thus, the saturation limit is reached earlier so as to have enough time to reduce the guidance errors. Using Monte-Carlo simulations, the superiority of the new guidance law over the classical OGL is shown. This validates the new approach of designing an estimation statistics dependent guidance law by using a random input describing function to approximate the missile's acceleration saturation. [ABSTRACT FROM AUTHOR]

A novel stochastic optimal control guidance law is presented for a missile with bounded acceleration. The nonlinear optimal guidance law (NOGL) is obtained by numerically solving the stochastic optimization problem. Since the certainty equivalence principle is not valid in the investigated problem, the resulting NOGL depends on the conditional probability density function of the estimated states. It is shown that the NOGL is also nonlinear in the estimated zero effort miss distance, and that the probability density function of the miss distance is non-Gaussian. The dependence of the new guidance law on the acceleration limit is investigated and it is shown that only for an extremely large acceleration limit does the proposed guidance law degenerate to the classical optimal linear one. [ABSTRACT FROM AUTHOR]

Modern 4th generation air-to-air missiles are quite capable of dealing with today's battlefield needs. Advanced aerodynamics, highly efficient warheads and smart target acquisition systems combine to yield higher missile lethality than ever. However, in order to intercept highly maneuverable targets, such as future unmanned combat air vehicles (UCAV), or to achieve higher tracking precision for missiles equipped with smaller warheads, further improvement in the missile guidance system is still needed. A new concept is presented here for deriving improved differential-game-based guidance laws that make use of information about the target orientation, which is acquired via an imaging seeker. The underlying idea is that of using measurements of the target attitude as a leading indicator of target acceleration. Knowledge of target attitude reduces the reachable set of target acceleration, facilitating the computation of an improved estimate of the zero-effort miss (ZEM) distance. In consequence, missile guidance accuracy is significantly improved. The new concept is applied in a horizontal interception scenario, where it is assumed that the target maneuver direction, constituting a partial attitude information, can be extracted via processing target images, acquired by an imaging sensor. The derivation results in a new guidance law that explicitly exploits the direction of the target acceleration. The performance of the new guidance law is studied via a computer simulation, which demonstrates its superiority over existing state-of-the-art differential-game-based guidance laws. It is demonstrated that a significant decrease in the miss distance can be expected via the use of partial target orientation information. [ABSTRACT FROM AUTHOR]

Since the proportional navigation guidance law was first introduced, many of the researchers had proposed different methodologies to investigate the corresponding performances of all the existing guidance laws. Even though a unified approach was proposed a few years ago, other authors found that under the proposed framework, all the existing guidance laws, namely ideal proportional navigation (IPN), true proportional navigation (TPN), and pure proportional navigation (PPN), were indeed special cases of the mentioned general guidance law. However, the results were restricted to two-dimensional space. In this paper, the author not only extends the results to three-dimensional space, but also to general LPN (GIPN), general TPN (GTPN), and PPN. Unlike conventional researchers, a modified polar coordinate (MPC) is adopted. It is shown that with the property of this MPC, for the line of sight (LOS) based guidance laws (GIPN and GTPN) the number of differential equations required to fully describe the relative dynamics can be reduced from six to three, however, for the missile's velocity-based guidance law, i.e., PPN, five differential equations are required. All the terms of differential equations involve only products and additions of variables. For all the mentioned guidance laws in this paper, only two transformed variables are required to describe the capture region, while the third variable is required to provide the condition of finite turn rate. [ABSTRACT FROM AUTHOR]

Addressed here is the precision missile guidance problem where the successful intercept criterion has been defined in terms of both minimizing the miss distance and controlling the missile body attitude with respect to the target at the terminal point. We show that the H[sup&infin]; control theory when suitably modified provides an effective framework for the precision missile guidance problem. Existence of feedback controllers (guidance laws) is investigated for the case of finite horizon and non-zero initial conditions. Both state feedback and output feedback implementations are explored. [ABSTRACT FROM AUTHOR]

Presents information on a study which proposed a terminal guidance law for impact angle control for active homing engagement scenarios. Ways to increase terminal effectiveness of tactical missile systems; Results; Conclusions.

Proposes a composite guidance scheme based on the command to line-of-sight (CLOS)+infrared terminal homing (IRTH) for a short-range surface-to-air missile system. Reason for the proposal of the scheme; Analysis of launch boundaries for a successful guidance handover; Development of practical guidance laws for the IRTH phase in the presence of inherent missile heading errors.

Motivated by spectral separation, guidance and control systems of interceptors are traditionally designed separately. However, interception of highly maneuverable targets and increased interception accuracy requirements lead to intense maneuvering during the last phase of the engagement, thus violating the spectral separation assumption of the conventional two-loop design. This has boosted the interest in integrated guidance-control designs that demonstrate superior interception accuracies by accounting for the coupling between the kinematics and dynamics of the problem. One of the approaches used for the integrated design is the sliding mode control (SMC) methodology. The results presented so far assumed full state information. We examine the effect of estimation on the performance of the integrated sliding mode controller design. In particular, a sliding mode observer/differentiator (SMOD) is used to estimate the target states. Through an intensive numerical study it is shown that the integrated guidance-control approach has a clear advantage over a conventional separated two-loop design. This advantage is attributed to the robustness of the integrated sliding mode controller design that better deals with the estimation errors and measurements noises. [ABSTRACT FROM AUTHOR]