Missile Guidance Law Design Research Articles

Missile Guidance Law Design Based on Free-Time Convergent Error Dynamics

Missile Guidance Law Design Based on Free-Time Convergent Error Dynamics

Design of Missile Guidance Law Using Takagi-Sugeno-Kang (TSK) Elliptic Type-2 Fuzzy Brain Imitated Neural Networks

The exploiting of missile guidance law is one of the most important issue of defense systems. However, it is a complex nonlinear guidance control problem. In recent years, several intelligent algorithms have been employed for the missile guidance law design. This paper aims to provide a more effective intelligent neural network, and then apply it to missile guidance control. An elliptic type-2 function (ET2F) combined with a Takagi-Sugeno-Kang (TSK) fuzzy system (FS) based on a wavelet function and a brain imitated neural network (BINN) is proposed to produce a new Takagi-Sugeno-Kang elliptic type-2 fuzzy brain imitated neural network (TSKET2FBINN). The proposed TSKET2FBINN is then incorporated into the missile guidance law. The proposed missile guidance control system includes the TSKET2FBINN controller and a robust compensation controller. A Lyapunov function is used to generate parameter adjustment adaptive laws so as to guarantee system stability with fast tracking performance of missile guidance control system. By presenting three missile control scenarios, the effectiveness of the proposed method is then demonstrated. The miss-distance (MD) of the proposed method is calculated and compared to other recent guidance methods to demonstrate its superior performance.

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.

A New Missile Guidance Law Design based on Interception Point Strategy and Fuzzy Logic against High Maneuvering Targets

A New Missile Guidance Law Design based on Interception Point Strategy and Fuzzy Logic against High Maneuvering Targets

Robust terminal guidance law design for missiles against maneuvering targets

This paper describes an application of a new terminal guidance law for missiles against maneuvering targets in three-dimensional (3D) environment during the terminal phase. This novel approach is suitable for practical implementation of the guidance law under the circumstance of a guided missile against a maneuvering target. The design procedure is divided into two steps. First, a feedback linearization control scheme is designed to approximately achieve desired tracking of the uncertain missile–target system. Next, a combined optimal robust control scheme is employed to minimize the worst-case effect arising from external disturbances (target's maneuvers) in improving the tracking performance. Simulation examples are presented to illustrate engagement performance of the proposed approach.

Adaptive-sliding-mode guidance law design for missiles with thrust vector control and divert control system

This approach addresses an adaptive-sliding-mode guidance law for missiles equipped with thrust vector control and divert control system, for the task of intercepting of a theatre ballistic missile with manoeuvring acceleration. The aim of the present study is to achieve a bounded target interception under the three degrees of freedom (DOF) control and the zero-effort-miss phase to improve the intercepting accuracy of the missile by narrowing the distance between the intercepting missile and the target missile to within effective range for triggering the missile’s explosion. Considering the external disturbances and the variation in the missile’s mass, the 3 DOF adaptive-sliding-mode guidance law is designed using adaptive control and sliding-mode control to minimise the distance between the centre of the intercepting missile and that of the target missile asymptotically. Also, the overall system stability is verified using the Lyapunov stability theory. Furthermore, successful simulation results are conducted to validate the effectiveness of the proposed adaptive-sliding-mode guidance law.

Missile Guidance Law Design Using Adaptive Cerebellar Model Articulation Controller

An adaptive cerebellar model articulation controller (CMAC) is proposed for command to line-of-sight (CLOS) missile guidance law design. In this design, the three-dimensional (3-D) CLOS guidance problem is formulated as a tracking problem of a time-varying nonlinear system. The adaptive CMAC control system is comprised of a CMAC and a compensation controller. The CMAC control is used to imitate a feedback linearization control law and the compensation controller is utilized to compensate the difference between the feedback linearization control law and the CMAC control. The online adaptive law is derived based on the Lyapunov stability theorem to learn the weights of receptive-field basis functions in CMAC control. In addition, in order to relax the requirement of approximation error bound, an estimation law is derived to estimate the error bound. Then the adaptive CMAC control system is designed to achieve satisfactory tracking performance. Simulation results for different engagement scenarios illustrate the validity of the proposed adaptive CMAC-based guidance law.

Integrated Guidance and Control of Missiles with θ-D Method

Optimal formulations allow designers to bring together concerns about guidance laws and autopilot responses under one unified framework. They lead to a natural integration of these different functions. By modifying the cost functions different responses, saturations (control related), miss distance (guidance related) etc., which are of primary concern to a missile system designer can be easily studied. Missile guidance law design and autopilot are formulated into a single unified state space. The cost function is chosen to reflect both guidance and control concerns. The ultimate control input is the missile fin deflections. A new suboptimal control method, called the θ-D method, gets an approximate closed-form solution to this nonlinear guidance problem based on approximations to the Hamilton-Jacobi-Bellman equation. A nonlinear six-degree-of-fireedom missile simulation is used to demonstrate the potential of this new integrated guidance and control approach.

Design of Missile Guidance Law via Variable Structure Control

The missile guidance law utilizing variable structure control is proposed. The acceleration command input is determined considering the target acceleration as an uncertainty. The proposed guidance law uses only the information for the target acceleration bound; therefore, the precise measuring of target acceleration during the maneuver is not required, and the robustness to the target maneuver is achieved. It is also shown that the proposed guidance law can be classie ed into the augmented true proportional navigation or the augmented realistic true proportional navigation guidance law. Numerical simulations show that the proposed guidance law yields better performance compared to existing guidance laws. ROPORTIONAL navigation guidance (PNG) was e rst developed during the 1950s, and during the 1970s and 1980s various PNG laws such as pure proportional navigation (PPN), true proportional navigation (TPN), the optimal guidance law (OGL), generalized TPN (GTPN), and realistic TPN (RTPN) have been developed. 1;2 Lots of studies have been performed to obtain analytical solutions as well as to analyze the capture regions of the guidance laws. With the development of accurate avionic sensors, augmented proportional navigation (APN) and the predictive guidance law (PGL) utilizing the information about target acceleration were also proposed. By analyzing the various guidance laws, the characteristics of the guidance laws, capture regions, and pursuing performance were compared. 1;2

Three‐dimensional nonlinear H∞ guidance law

AbstractThis paper proposes a novel three‐dimensional missile guidance law design based on nonlinear H∞ control. The complete nonlinear kinematics of pursuit–evasion motion is considered in the three‐dimensional spherical co‐ordinates system; neither linearization nor small angle assumption is made here. The nonlinear H∞ guidance law is expressed in a simple form by solving the associated Hamilton–Jacobi partial differential inequality analytically. Unlike adaptive guidance laws, the implement of the proposed robust H∞ guidance law does not require the information of target acceleration, while ensuring acceptable interceptive performance for arbitrary target with finite acceleration. The resulting pursuit–evasion trajectories for both the H∞‐guided missile and the worst‐case target are determined in closed form, and the performance robustness against variations in target acceleration, in engagement condition, and in control loop gain, is verified by numerical simulations. Copyright © 2001 John Wiley & Sons, Ltd.

Observability analysis for target maneuver estimation via bearing-only and bearing-rate-only measurements

The implementation of modern guidance laws requires the design of tracking filters that provide information about the target maneuver. Because of the nonlinearity of the tracking problem, the stability of extended Kalman filters in tracking applications involving bearing-( rate-) only observations is highly dependent on the intercept geometry. This is due mainly to the lack of observability in the absence of range measurements. In this paper, the observability of target maneuvers via bearing-( rate-) only measurements is discussed. Intercept scenarios that result in the loss of observability are identified. They play an important role for both the initialization of tracking filters and the design of missile guidance laws based on bearing-( rate-) only measurements.

Improved guidance law design based on the mixed-strategy concept

The paper summarizes a new approach for missile guidance law design in a noise-corrupted environment. This approach is based on the concept of mixed strategies known from game theory. Two examples demonstrate that this innovative design approach leads to an improved missile performance compared to all guidance laws previously discussed in the open literature. For the sake of simplicity, a two-dimensional version of the terminal phase of a missile-aircraft encounter is adopted. It is assumed that the game takes place in the vicinity of the collision course (see Fig. 1). In this section, the missile and the aircraft will be called pursuer and evader, respectively. Throughout the duration of the game the pursuer measures the relative range R, the closing velocity Vc, and the LOS angle Ji, relative to a reference line. It is assumed that the range and velocity measurements are exact and that the angle is corrupted by noise. The velocity and range information is processed to give an accurate estimate of the time-to-go fgo, while the range and angle information is processed and yields a noisy measurement Z of the evader's relative position perpendicular to the reference line. The input to the estimation process is Z and it is given by:

Analytical solution of optimal trajectory-shaping guidance

This paper presents a combined midcourse and terminal guidance law design for missiles to achieve range enhancement with excellent intercept performance. We derive analytic solutions of a closed-loop, nonlinear optimal guidance law for three-dimensional flight for both the midcourse and terminal phases. This combined guidance law can quickly modify the missile trajectory during midcourse guidance when the target direction changes. Zero heading error is achieved at handover from the midcourse to the terminal phase. The guidance algorithm is in a feedback form, either in inertial coordinates or in seeker coordinates. It is sufficiently simple for onboard implementation and has been applied successfully for on-line operation.