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.

When calculating a gimbal angle trajectory that satisfies boundary conditions set by an attitude boundary condition setter 2131 of the ground station 21, a gimbal angle trajectory calculator 2132 calculates the gimbal angle trajectory that minimizes a period of an acceleration interval within a range that satisfies driving restrictions of a gimbal, based on a gimbal angle .sub.0i of a start time and a gimbal angle .sub.ci of a fixed interval of an attitude change. Also, the gimbal angle trajectory is calculated that minimizes a period of a deceleration interval within a range that satisfies the driving restrictions of the gimbal, based on the gimbal angle .sub.ci of the fixed interval and a gimbal angle .sub.fi of a completion time of the attitude change. The obtained .sub.0i, .sub.ci, .sub.fi and an attitude change period are transmitted to the artificial satellite as gimbal angle trajectory parameters, and the control moment gyros are controlled based on the gimbal angle trajectory parameters.

A spacecraft (10) that includes a body (1); a rocket engine (2) installed in the body; an aerodynamic element (5) which is installed in the body and on which aerodynamic force acts; a measurement quantity acquiring system (7) configured to acquire at least one measurement quantity of the spacecraft; and a control device (8) configured to calculate an operation quantity to operate at least one of a gimbal angle of the rocket engine and an aerodynamic characteristic of the aerodynamic element. The control device (8) is configured to calculate the operation quantity according to the measurement quantity by a non-linear optimal control using a stable manifold method in the attitude change such that the attitude angle of the spacecraft (10) changes to the target attitude angle.

A method for capturing and deorbiting space debris includes: deploying a space debris capturing device in planetary orbit; receiving an initial target set including a first database of space debris targets that are within range of the space debris capturing device; performing a first algorithm to convert the initial target set to an accessible target set including a second database of space debris targets that are within range of the space debris capturing device; performing a second algorithm to convert the accessible target set to a final target set including a third database of space debris targets to be captured by the space debris capturing device capturing the first space debris target via a capture mechanism of the space debris capturing device; jettisoning the capture mechanism and the first captured space debris target into a decaying orbit; one of the remaining space debris targets of the third database; and positioning the space debris capturing device and the final captured space debris target into a decaying orbit.

A method, apparatus and system for controlling an attitude of a spacecraft, the spacecraft including an attitude control system operatively associated with a ground-based spacecraft control system. According to an exemplary embodiment, the spacecraft attitude control system uses a B-spline interpolator for commanding the spacecraft and a Kalman filtering process is used to estimate B-spline interpolator coefficients. The methods and systems disclosed herein can be implemented in, for example, executable machine code and/or integrated circuit hardware.

An on-orbit servicing spacecraft includes an engagement system to engage a space vehicle or object to be serviced or tugged, so as to form a space system, and an electronic reaction control system to cause the spacecraft to rotate about roll, yaw, and pitch axes to control attitude and displacement along given trajectories to cause the spacecraft to carry out given maneuvers. The electronic reaction control system includes (i) a sensory system to directly sense physical quantities or allow physical quantities to be indirectly computed based on sensed physical quantities, including one or more of position, attitude, angular rates, available fuel, geometrical features, and on-board systems state, (ii) attitude control thrusters mounted so as to allow their positions and orientations to be adjustable, and (iii) an attitude control computer in communication with the sensory system and the attitude control thrusters and programmed to receive data from the sensory system and to control, based on the received data, positions, orientations, and operating states of the attitude control thrusters so as to control attitude and position of the spacecraft. The attitude control computer is programmed to cause the spacecraft to carry out a given mission including an engagement step, in which the engagement system and the attitude control thrusters are controlled by the attitude control computer to engage a space vehicle or object to be serviced or tugged, and one or more operating steps, in each of which the attitude control thrusters are controlled by the attitude control computer to meet one or more requirements established for the operating step.

A reaction wheel detects an angular momentum. A satellite controller selects a target thruster based on the detected angular momentum. A power supply apparatus changes adjustment electric power for the target thruster. A flow rate adjustment apparatus supplies a propellant to the target thruster at a flow rate corresponding to the adjustment electric power. This changes a thrust of the target thruster. When a discharge current of the target thruster does not become a target current, the power supply apparatus further changes the adjustment electric power for the target thruster.

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.

Described herein are systems and methods for attitude determination using infrared Earth horizon sensors (EHSs) with Gaussian response characteristics. Attitude information is acquired by detecting Earth's infrared electromagnetic radiation and, subsequently, determining the region obscured by Earth in the sensors' fields of view to compute a nadir vector estimation in the spacecraft's body frame. The method can be applied when two sensors, each with known and distinct pointing directions, detect the horizon, which is defined as having their fields of view partially obscured by Earth. The method can be implemented compactly to provide high-accuracy attitude within small spacecraft, such as CubeSat-based satellites.

When the number of CMGs is represented by n (n is an integer of 4 or more), (n3) gimbal angles out of n gimbal angles corresponding to the n CMGs are set as free parameters, and an algebraic equation representing a relationship among three gimbal angles out of the n gimbal angles, the free parameters, and an angular momentum of all the CMGs is used to solve the algebraic equation while changing the free parameters within set ranges, to thereby obtain solutions of the gimbal angles of the plurality of CMGs required for achieving a given angular momentum.