Abstract: A three-axis attitude control for a low-orbiting satellite. A sensor which measures in two axes is a magnetometer, and a satellite is positioned on an inclined orbit at an inclination angle of approximately 25 degrees to approx. 90 degrees. The satellite has been provided with an overall spin by spin wheels, and actuators generate torque on all three axes of the satellite. Accordingly, this is applicable to three-axis attitude control of a low-orbiting satellite that includes a sensor which measures in two axes and that includes one or several spin wheels.

Abstract: A method for operating a space vehicle equipped with at least one device cooperating with a celestial body, wherein rotation of the space vehicle takes place around at least one axis of the space vehicle. The rotation of the space vehicle is slowed when the device cooperating with the celestial body is oriented substantially toward the celestial body. The device can be a solar panel intended to face the sun and produce electrical energy.

Abstract: In a process for identifying an incorrectly measuring sensor which is part of a sensor arrangement on a spacecraft, external directional vectors are measured, relative to external objects, such as the earth, the sun or selected stars, or to external field vectors, such as the earth's magnetic field. Based on actual time and space coordinates of the spacecraft, respective actual external directional vectors are calculated with respect to an inertial system of coordinates, and the angles between these vectors are determined. These angles are compared with corresponding angles which exist between the directional vectors measured by a function of the sensor arrangement relative to a spacecraft-fixed system of coordinates. Based on this comparison a sensor which may measure incorrectly can be identified reliably.

Abstract: A low earth orbiting satellite control arrangement according to the invention has an earth sensor for measuring roll and pitch of the satellite relative to an earth oriented moving reference system, a spin wheel mounted along an axis perpendicular to the orbital plane of the satellite for measuring spin of the satellite, and two magnet coils for generating control. A first of the two magnet coils is aligned substantially in parallel to a roll axis of the satellite, and the second is aligned substantially parallel to the yaw axis. Alternative control algorithms are disclosed for controlling attitude, nutation and spin of the satellite based on information from the earth sensor, spin wheel and an observer which is also mounted on the satellite.

Abstract: A method and apparatus for steering a low-earth-orbit communication satellite requiring sun-orientation for solar-power accumulation is disclosed. Momentum-bias both maintains nadir pointing and adds yaw-steering moments for sun-trackig attitude control, simultaneously. The method has two principal steps: 1) open-loop momentum decoupling, for correcting the calculated steering torque, and closed-loop attitude compensation, to correct for perturbations about the calculated attitude in accordance with one of two control law definitions. This combines the advantage of stable gyroscopic attitude control with those of open-loop yaw steering.

Abstract: The invention relates to a method for determining the angular momentum vector of a satellite, located in an external magnetic field, relative to a satellite-integral orthogonal system of coordinates. The satellite has a torque generating system to generate pulse-shaped torques around all three axes. A magnetometer measuring on three axes is used, whose measured values are fed to a computer. In the latter, at successive points in time and using the current measured values, an observer equation is integrated, with new estimated values being determined for the angular momentum vector. Using these as well as the current measured values, an inequality is continuously checked; when it is not fulfilled, the torque generating system operates briefly, so that finally the fulfillment of the inequality is forced and identity of the last estimated value that has been determined with the rotary pulse vector to be determined is achieved.

Abstract: A reacquisition process is used in the event that an earth orientation of the yaw axis Z of an earth satellite is lost, in that the angular velocity .omega. of a flywheel, the rotational axis of which is oriented in the direction of a pitch axis Y, falls below a predetermined value as a result of an insufficient energy supply. After restoration of the energy supply, the pitch attitude of the earth satellite and, accordingly, the earth orientation of the yaw axis Z, are to be restored. For this purpose, constantly alternating increases and decreases in speed are commanded for the flywheel in a determined manner in correlation with the zero passages of the pitch deviation, whereby the satellite is subjected to a constant reversal of its direction of rotation. This occurs until the angular momentum of the satellite falls below a predetermined value during a zero passage of the pitch deviation.