Missile

About: Missile is a research topic. Over the lifetime, 12829 publications have been published within this topic receiving 94307 citations. The topic is also known as: guided missile & missiles.

Sliding mode control based impact angle control guidance considering the seeker׳s field-of-view constraint

Abstract: The problem of impact angle control guidance for a field-of-view constrained missile against non-maneuvering or maneuvering targets is solved by using the sliding mode control theory. The existing impact angle control guidance laws with field-of-view constraint are only applicable against stationary targets and most of them suffer abrupt-jumping of guidance command due to the application of additional guidance mode switching logic. In this paper, the field-of-view constraint is handled without using any additional switching logic. In particular, a novel time-varying sliding surface is first designed to achieve zero miss distance and zero impact angle error without violating the field-of-view constraint during the sliding mode phase. Then a control integral barrier Lyapunov function is used to design the reaching law so that the sliding mode can be reached within finite time and the field-of-view constraint is not violated during the reaching phase as well. A nonlinear extended state observer is constructed to estimate the disturbance caused by unknown target maneuver, and the undesirable chattering is alleviated effectively by using the estimation as a compensation item in the guidance law. The performance of the proposed guidance law is illustrated with simulations.

Real time missile guidance system

Abstract: A weak Hamiltonian finite element method is used for iterative computation of missile guidance acceleration commands for maximizing a missile's terminal velocity while satisfying control authority limits and terminal attitude constraints. The guidance acceleration commands include commands for controlling the angle of attack (α) and the bank angle (φ) of the missile. The angle of attack (α) and bank angle (φ) are related to a set of virtual control variables selected to avoid convergence problems when the angle of attack is approximately zero. The preferred control variables are β 2 and β 3 such that β 2 =cosφtanα and β 3 =sinφtanα. Iterative convergence is facilitated when control inequality constraint parameters are reached by adjusting iterative solutions between iterations toward satisfaction of the constraints. An approximation to an optimal trajectory is calculated at each guidance cycle during missile flight using data which are revised during each guidance cycle. The revised data include current position data for the target and the current position for the missile. The revised data are taken from the most reliable source currently available, such as on-board target-seeking radar when the target-seeking radar is locked onto the target, uplink data from ground or airborne tracking radar when an uplink is operational, or inertial guidance data. Extracted from the optimal trajectory is an optimal acceleration command for optimally controlling the angle of attack and bank angle of the missile.

Finite-Time Cooperative Guidance Strategy for Impact Angle and Time Control

Abstract: In this article, two cooperative guidance schemes with impact angle and time constraints are proposed for multiple missiles cooperatively intercepting a maneuvering target in the presence or absence of a leader missile. First, cooperative interception with impact angle constraint is formulated based on the relative motion equation between the interceptors. Subsequently, the design of the guidance law is divided into two stages. In the first stage, based on finite-time consensus theory and the super-twisting control algorithm, the cooperative guidance law in the direction of the line of sight (LOS) is derived to control the impact time so that all missiles can simultaneously intercept the target in finite time. Then, based on finite-time sliding mode control and the super-twisting control algorithm, a finite-time convergence cooperative guidance law in the normal direction to the LOS is developed to control the intercept angle and the LOS angle rate so that each missile can accurately hit the target at a desired angle. Finally, by controlling the consensus error of flight time in the presence of a leader missile, a finite-time cooperative guidance method in the “leader–follower” scenario is derived for simultaneous arrival at the target. The proposed methods can effectively suppress chattering and ensure fast convergence in finite time. Simulation experiments are conducted to verify the effectiveness of the proposed cooperative guidance schemes.

Design of robust, time-varying controllers for missile autopilots

Abstract: The design of a longitudinal gain-scheduled autopilot for a tail-controlled missile using mu -synthesis control design techniques is discussed. The goal is to design a single controller to track commanded acceleration maneuvers with a steady state error of 0.5%, a time constant of less than 0.2 s, a step response overshoot of less than 10% and a maximum tail-deflection rate of 25 degrees/s/g. The controller is to provide robust performance over a range of +or-20 degrees angle-of-attack and missile speeds between Mach 2 and 4. The controller must avoid saturating the tail-deflection actuator rate capabilities and destabilizing unmodeled high frequency flexible body modes of the missile. >

Adaptive Techniques for Multiple Actuator Blending

Abstract: Advanced missiles employ multiple actuators to enhance maneuverability and to improve the intercept probability against highly maneuverable targets. Actuators employed in such missiles include aerodynamic control surfaces and reaction jets. While the usage of aerodynamic surfaces are not generally constrained, reaction jet usage has to be minimized due to the limited amount of fuel available on-board. A blending logic is employed to optimally allocate the actuators in response to commands from the autopilot. This paper discusses the development of a fuel conservative actuator blending logic that provides relatively invariant actuator performance over widely varying flight conditions. The invariant performance is obtained using the model reference adaptive control technique. Multiple adaptation strategies are employed to ensure rapid convergence and stable behavior. The performance of the model reference adaptive actuator blending strategy is illustrated using a realistic missile model.