A spacecraft including a set of thrusters for changing a pose of the spacecraft. At least two thrusters mounted on a gimbaled boom assembly and are coupled together sharing the same gimbal angle. A model predictive controller (MPC) to produce a solution for controlling thrusters of the spacecraft by optimizing a cost function over a receding horizon using a model of dynamics of the spacecraft effecting a pose of the spacecraft and a model of dynamics of momentum exchange devices of the spacecraft effecting an orientation of the spacecraft. A modulator to modulate magnitudes of the thrust of the coupled thrusters determined by the MPC as pulse signals specifying ON and OFF states of each of the coupled thruster, wherein the ON states of the coupled thrusters sharing the same gimbal angle do not intersect in time. A thruster controller to operate the thrusters according to their corresponding pulse signals.

A control system for controlling an operation of a spacecraft. A model predictive controller (MPC) produces a solution for controlling thrusters of the spacecraft. The MPC optimizes a cost function over a finite receding horizon using a model of dynamics of the spacecraft effecting a pose of the spacecraft and a model of dynamics of momentum exchange devices of the spacecraft effecting an orientation of the spacecraft. The optimization is subject to hard and soft constraints on angles of thrusts generated by thrusters. Further, the hard constraints require the angles of thrusts in the solution to fall within a predetermined range defined by the hard constraints. The soft constraints penalize the solution for deviation of the angles of thrusts from nominal angles corresponding to a torque-free thrust passing through the center of the mass of the spacecraft. A thruster controller operates the thrusters according to the solution of the MPC.

A free-falling body verification device for a drag-free spacecraft comprises a spacecraft simulation device (1), used for carrying out free-falling body motion on the ground; an inertial sensor or accelerometer (2), used for measuring the residual disturbance acceleration of the spacecraft simulation device (1); an attitude sensor (3), used for measuring attitude parameters of the spacecraft simulation device (1); a drag-free controller (4), used for processing the residual disturbance acceleration and the attitude parameters so as to obtain a feedback control signal; and a propeller (5), used for generating thrust action applied on the spacecraft simulation device (1) under the control of the feedback control signal, so as to enable the spacecraft simulation device (1) to overcome the residual disturbance of the external environment and maintain the attitude. The space operating environment is simulated by means of the free-falling body motion of the spacecraft on the ground within short time; the inertial sensor or accelerometer (2), the attitude sensor (3), the drag-free controller (4), and the propeller (5) are combined, so that the performance and function test verification for a space drag-free aerospace system is realized in the technical ground environment within short time.

Disclosed is a method for controlling the attitude guidance of a satellite with respect to an orbital reference system including a velocity axis, an orbital axis, and a Nadir axis; the satellite moving in the direction of the velocity axis, the satellite including an optical instrument having an observation axis, a solar generator defining a functional surface having a normal, an attitude control device, and a control unit. The method includes a step (104) of transmitting guidance commands so as to direct the observation axis of the optical instrument towards regions to be imaged or to orient the normal to the functional surface in the direction of the solar radiation. The guidance commands are commands to rotate the satellite about the velocity axis only, the angle of rotation about the orbital axis and Nadir axis within the orbital reference system being kept substantially at zero.

An operation of a spacecraft is controlled using an inner-loop control determining first control inputs for momentum exchange devices to control an orientation of the spacecraft and an outer-loop control determining second control inputs for thrusters of the spacecraft to concurrently control a pose of the spacecraft and a momentum stored by the momentum exchange devices of the spacecraft. The outer-loop control determines the second control inputs using a model of dynamics of the spacecraft including dynamics of the inner-loop control, such that the outer-loop control accounts for effects of actuation of the momentum exchange devices according to the first control inputs determined by the inner-loop control. The thrusters and the momentum exchange devices are controlled according to at least a portion of the first and the second control inputs.

A multi-degree-of-freedom electromagnetic machine includes a first structure, a second structure, and a control. The first structure is configured to rotate about a spin axis and about a tilt axis that is perpendicular to the spin axis, and includes a first spin conductor, a second spin conductor, and a tilt conductor, which together form a general shape of a surface. The second structure is disposed adjacent to the first structure and includes a plurality of magnets. The control is configured to controllably supply alternating current (AC) to the first and second spin conductors and direct current (DC) to the tilt conductor, wherein the first structure continuously rotates about the spin axis in response to the AC being supplied to the first and second spin conductors, and rotates about the tilt axis to a tilt position in response to the DC being supplied to the tilt conductor.

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.

Methods, apparatus, and articles of manufacture to minimize command dynamics of a satellite are disclosed. An example apparatus includes a steering law module to calculate a first set of vectors to maneuver a space vehicle, and calculate a second set of vectors based on projecting the first set of vectors onto a fixed plane. The apparatus further includes an attitude controller to generate an attitude command based on the first and the second sets of vectors to prevent an unplanned rotation by the space vehicle.

The purpose is to easily achieve verification of an integrated attitude angle based on an inertial sensor. An attitude angle calculating device may include an integrated attitude angle calculating module, a reverse-calculated value calculating module, a reference value calculating module, and a determining module. The integrated attitude angle calculating module may calculate an integrated attitude angle using an output of the inertial sensor and a positioning signal. The reverse-calculated value calculating module may reverse calculate, using the integrated attitude angle, a physical quantity obtained based on the positioning signal that is used for calculating the integrated attitude angle. The reference value calculating module may calculate a reference physical quantity corresponding to the reverse-calculated physical quantity, based on an observation value of the positioning signal. The determining module may determine whether to reset the calculation of the integrated attitude angle by using the reverse-calculated physical quantity and the reference physical quantity.

A dipole cancellation system and method may include a plurality of magnetometers for measuring a device magnetic field associated with a plurality of device coils generating a device magnetic field having a primary magnetic dipole moment. A compensating coil carrying a compensating current running a first direction that generates a compensating magnetic field having a compensating magnetic dipole moment. The compensating coil may be positioned and the first current may be selected so that the compensating magnetic dipole moment completely cancels the primary magnetic dipole moment. A method may use the system to stabilize a spacecraft by calculating an estimated torque of the spacecraft, receiving a value for an external magnetic field, receiving a value for a device magnetic field, and calculating and applying a compensating current may be then applied to the compensating coil to cancel the primary magnetic dipole moment, wherein the spacecraft is stabilized.