A process to design an attitude control system (ACS) controller in each of a plurality of joined entities includes identifying a worst case configuration as a design-to configuration as one or more configurations in a given set S of configurations required for a spacecraft. For the design-to configuration, the process includes deriving one or more system equations in a functional form of equations to determine intermediate design parameters that represent effective proportional and derivative gains of the combined controller, Kp and Kd, respectively. The process also includes determining the design parameters of the ACS controller, namely, gains Kq and K and stiffness and damping coefficients, Ks and Cd respectively of all the interfaces between each of the plurality of joined entities, from the intermediate design parameters Kp and Kd. The process further includes programming the ACS controller with selected values of the design parameters for matrices Kq and K and selecting springs with stiffness Ks and dampers with damping coefficient Cd for all interfaces between each of the plurality of joined entities. The process includes iterating the computer-implemented process after incrementing a convergence requirement parameter threshold when the control performance is not acceptable and until the system achieves acceptable performance, and programming the ACS controller for each of the plurality of joined entities.

A spacecraft for removing space debris is disclosed. The spacecraft includes a satellite bus, a shield member foldable on an outer side face of the satellite bus and disposed facing towards space debris to reduce a movement speed of the space debris, and a support member configured to support the shield member with respect to the satellite bus, in which the shield member includes a central panel configured to overlap one face of the satellite bus, a plurality of first panels connected to peripheral sides of the central panel and radially extended, and a plurality of second panels located between the first panels.

A menu-type design method based on optimized information fusion applied to a GEO satellite control system is provided, which includes: configuring four long-life inertial attitude sensor gyroscopes for a long-life GEO satellite control system; configuring sensors capable of measuring three-axis attitude according to a menu-type design requirement on hardware, where the long-life inertial attitude sensor gyroscopes and the sensors capable of measuring three-axis attitude are combined to form three types of Kalman filters; autonomously sorting, by the satellite-borne computer application software, the Kalman filters; and in a case where an FDIR module detects a fault, autonomously generating, by the FDIR module, an alarm corresponding to the fault, and autonomously performing, by a currently selected Kalman filter, reduced-order filtering, and in a case where the fault is not eliminated within a set time period, issuing, by the FDIR module, a macro instruction sequence to perform autonomous reorganization.

A method of controlling a satellite and a computer-readable recording medium are provided. The method is for controlling a satellite moving along an orbit having an inclination angle from the equatorial plane to capture due-north images. The method includes: determining a position of the satellite; calculating a roll angle and a pitch angle of the satellite for pointing a line-of-sight vector of the satellite to a first ground surface being a photographing point; determining a compensation angle by considering effects of the inclination angle and rotation of the Earth so as to capture images in the due north direction of the photographing point; calculating a yaw angle based on the compensation angle; and rotating the satellite according to the calculated roll angle, pitch angle, and yaw angle.

In an example embodiment, an attitude, orbit, and de-orbit control system (AODCS) for a satellite is provided. In an example embodiment, the AODCS system comprises one or more selectively retractable booms. The one or more selectively retractable booms are collectively configured to provide a selectively adjustable drag during de-orbiting of a satellite over a predefined de-orbiting time.

A system and method of providing an affordable navigation, guidance and control system for arbitrary nano/micro launch vehicles by integrating commercial grade sensors with advanced estimation algorithms in a manner that provides sufficient accuracy of the resulting vehicle state estimates to inject nano/micro satellites into low earth orbits. The system and method uses commercial grade sensors and an advanced sensor-fusion estimator software that estimates and removes the estimated measurement errors and filters noise produced by the commercial grade sensors, resulting in estimated states with suitable accuracy. The filtered data are sent to a guidance and control system where actuator commands are formulated based on the filtered data. A simulated launch and flight of the launch vehicle is performed using the filtered data to validate that the GNC system and launch vehicle are ready for launch.

The present invention relates to an aerospace vehicle navigation and control system comprising a terrestrial illumination matching module for determining spacecraft position and attitude. The method permits aerospace vehicle position and attitude determinations using terrestrial lights using an Earth-pointing camera without the need of a dedicated sensor to track stars, the sun, or the horizon. Thus, a module for making such determinations can easily and inexpensively be made onboard an aerospace vehicle if an Earth-pointing sensor, such as a camera, is present.

Systems and methods are described for a satellite control system that exhibits improved stability and increased efficiency by implementing a non-linear dynamic inversion inner-loop control algorithm coupled with an eigen vector outer-loop control algorithm. Thus, the attitude determination and control system (ADACS) may operate using commands to rotate directly about an eigen vector. Additionally, the outer-loop control system includes a feed-forward control element to enhance pointing accuracy when tracking moving targets.

A satellite attitude data fusion system and method is disclosed, applicable to the earth satellite environment to estimate attitude data of the satellite. When the satellite attitude data fusion system of the present invention is used to perform the satellite attitude data fusion method, the first step is to perform a body rates quaternion attitude data processing operation. Then, the next step is to perform an attitude/rates data fusion processing operation, wherein an attitude data fusion algorithm module receives the first IAE result data from the first EKF, and the second JAE result data from the second EKF, and performs an attitude/rates data fusion algorithm in a subsystem level to evaluate an attitude estimation JAE performance.

Energy efficient satellite maneuvering is described herein. One disclosed example method includes maneuvering a satellite that is in an orbit around a space body so that a principle sensitive axis of the satellite is oriented to an orbit frame plane to reduce gravity gradient torques acting upon the satellite. The orbit frame plane is based on an orbit frame vector.