A new steering logic based on a mixed weighted two-norm and least squares optimization solution overcomes a deficiency of being trapped in the momentum saturation singularities. The new steering logic utilizes an additional weighting matrix, W, in conjunction with deterministic dither signals in a weighting matrix V. The new steering logic also provides a simple means of avoiding troublesome, internal elliptic singularities which are commonly encountered by most other pseudoinverse-based steering logic. Various CMG systems, such as a typical pyramid array of four single-gimbal CMGs, two and three parallel single-gimbal CMG configurations, and four parallel double-gimbal CMGs of the International Space Station are described as illustrative embodiments demonstrating the simplicity and effectiveness of the new steering logic.

Singularity escape/avoidance steering logic for control moment gyro systems

Higher Order Sliding Mode (HOSM) control techniques are applied to the Guidance Control (G&C) of interceptor missile in which velocity may be steered by combination of main thrust, aerodynamic lift and lateral on-off divert thrusters, and attitude may be steered by continuous or on-off actuators. Methods include the pointing of the seeker, its associated estimation processes, a guidance law that uses concurrent divert mechanisms, and an attitude autopilot. The insensitivity of the controller to matched disturbances allows the concurrent usage of the divert mechanisms without adverse effect on the accuracy. The controller also allows the de-coupling of the control of roll, pitch and yaw channels, and usage quaternions to represent body attitude and it provides control perfect robustness. While it conceivable to design separately the components of the G&C method, it is widely accepted that designing them in an integrated fashion usually produces a better result.

System and method for guiding and controlling a missile using high order sliding mode control

The present invention, in its various aspects and embodiments, includes a method for planning a mission profile in real time. The method comprises ascertaining a plurality of target information (including a target location, a target velocity, and a target location error) and autonomously determining a pattern from the ascertained target information. In one particular embodiment, the autonomous determination includes projecting along a target axis a distance of the target location error to establish two intersections of the target axis with the target location error; projecting perpendicularly left and right from the intersections to determine a pair of possible start point pairs; selecting the possible start point pair including a closest single start point; selecting the farthest start point of the selected start point pair; identifying an adjusted start point; mirroring the adjusted start point to obtain an adjusted start point pair; and laying out the front-end and back-end traces from the adjusted start point pair.

Autonomous mission profile planning

An interface between an aircraft and a precision-guided missile having automatic target recognition capability receives target image data and target location data from the aircraft, translates the data into formats usable by the missile, and downloads the target data to the missile at any time prior to launch of the missile. The interface includes a mission-planning unit and a terrain and elevation data storage unit enabling the air crew to call up terrain and elevation data on a visual display for planning a missile mission, and to transmit mission parameters to the missile during aircraft flight.

Method and apparatus providing an interface between an aircraft and a precision-guided missile

A ground-based interceptor missile, part of a larger ballistic missile defense system, is provided for intercepting enemy intercontinental ballistic missiles The interceptor employs a method based on numerical partial derivatives for computing required divert velocity corrections to remove predicted ground-based interceptor miss errors for engagement during the midcourse of the enemy missile flight

Method of computing divert velocity for the ground-based interceptor using numerical partial derivatives

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.

Malvin G. Johnson

Papers

Real time missile guidance system