The invention relates to small satellites capable to fly in formation (10), in particular nano- or picosatellites with a mass of 10 kg or less, for LEO applications, comprising a housing (12) and at least one plug-in board (14) arranged in the housing (12) with a predetermined functionality and a propulsion system (16) for generating a directed pulse in the direction of the flight trajectory T.sub.k.

It is proposed that the small satellite (10) comprises an independent and autonomously working collision avoidance system (18), which is capable of adapting a trajectory correction T.sub.kk of the trajectory T.sub.k by the propulsion system (16), when a collision with an object (30) is expected.

In a further independent aspect, the invention relates to a formation (100) composed of several small satellites capable to fly in formation (10), wherein a relative position and flight trajectory T.sub.k of each small satellite (10) is modifiable via the independently and autonomously working collision avoidance system (18).

The invention concerns a method for estimating the angular velocity (and, preferably, also the attitude) of a space platform (for example, a satellite, a space vehicle, or a space station) using only the information provided by one or more optical sensors, such as one or more star trackers, one or more colour and/or black and white cameras or video cameras, one of more infrared sensors, etc.

A dual solution candidate searcher receives an input of information about a constraint coefficient matrix and a cost vector, determines a dual problem of a linear programming problem being a primal problem and all active sets representing combinations of active formulas in constraints of the dual problem, finds, for each of the active sets, a feasible dual solution candidate meeting constraints, and stores the dual solution candidate into a storage in a manner associated with a corresponding one of the active sets. An optimal solution calculation device receives an input of a constraint vector as, selects an optimal one of the active sets as an optimal active set based on an inner product of the constraint vector and the dual solution candidate stored in the storage, and finds and outputs a basic feasible solution corresponding to the selected active set as an optimal solution.

The present technology relates to an artificial satellite and a control method thereof that enable to ensure quality of a captured image while suppressing battery consumption. An artificial satellite includes: an imaging device configured to perform imaging of a predetermined region on the ground; and a management unit configured to change accuracy of attitude control in accordance with a remaining battery amount at an instructed imaging time, and configured to change an imaging condition in accordance with accuracy of the attitude control. The present technology can be applied to, for example, an artificial satellite or the like that performs satellite remote sensing by formation flight.

A satellite in orbit around a planetary body includes a bus and a drag flap coupled to the bus. The drag flap is used to increase the drag torque applied to the satellite. The bus may house sensors and actuators, such as a star tracker, a gyroscope, a reaction wheel, and a global position system (GPS) receiver to monitor the attitude of the satellite in response to the applied drag torque. The measurements from the sensors and actuators may be used to determine the drag torque applied to the satellite. An estimate of the atmospheric density may be then be determined based on the drag torque. Compared to conventional approaches, the satellite and methods described herein estimates the atmospheric density at comparable, if not better, resolution and bandwidth. The atmospheric density estimates may also be acquired in real-time using a cheaper, lighter, and smaller satellite.

Provided is a single-axis pointing pure magnetic control algorithm for a spacecraft based on geometrical analysis to realize single-axis pointing control of the spacecraft through the pure magnetic control algorithm in which a magnetic torque is only output by a magnetorquer to interact with a geomagnetic field to generate a control torque. The algorithm uses a spatial geometry method to obtain an optimally controlled magnetic torque direction, thereby designing a PD controller. The problem that the traditional magnetic control method is low in efficiency and even cannot be controlled is overcome. The algorithm is simple and easy, can be used in the attitude control field of spacecrafts, and achieves the pointing control in point-to-sun of a solar array and point-to-ground of antennae.

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.

A method of controlling the attitude of a spacecraft in spinning around itself with a non-zero total angular momentum H.sub.TOT. The spacecraft includes a set of inertia flywheels configured to form an internal angular momentum H.sub.ACT. The axis of the total angular momentum H.sub.TOT is aligned with a principal axis of inertia of the spacecraft, in the course of which the inertia flywheels are controlled to form an internal angular momentum H.sub.ACT. The following expression, in which J is the inertia matrix of the spacecraft:
H.sub.actJ.sup.1(H.sub.tot.Math.J.sup.1H.sub.tot)
is negative if the principal axis of inertia targeted is the axis of maximum inertia of the spacecraft and is positive if the principal axis inertia targeted is the axis of minimum inertia of the spacecraft.

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.

Provided is a method of comparing satellite attitude control performances, the method including generating, by a controller, a satellite task execution command, receiving, by an input and output (I/O) unit, a result of a simulation performed on a satellite attitude control by a satellite attitude control simulator based on the satellite task execution command, and transmitting, by the I/O unit, the satellite task execution command to a satellite based on the result of the simulation.