Abstract: In a system for determining the position of a spacecraft based on vector determinations, a direction vector is measured in a body-fixed coordinate system; a reference direction vector is determined within a reference coordinate system based on the path position of the spacecraft and an orbit model; the overall spin vector of the spacecraft is determined within the body-fixed coordinate system; and a reference overall spin vector of the spacecraft is determined within a reference coordinate system by time propagation of known initial values of the overall spin of the spacecraft or by time tracking of a reference model. The position of the spacecraft is determined based on the four vectors.

Abstract: In a system for determining the position of a spacecraft based on vector determinations, a direction vector is measured in a body-fixed coordinate system; a reference direction vector is determined within a reference coordinate system based on the path position of the spacecraft and an orbit model; the overall spin vector of the spacecraft is determined within the body-fixed coordinate system; and a reference overall spin vector of the spacecraft is determined within a reference coordinate system by time propagation of known initial values of the overall spin of the spacecraft or by time tracking of a reference model. The position of the spacecraft is determined based on the four vectors.

Abstract: In a method for determining the position of a spacecraft based on vector determinations, a direction vector is measured in a body-fixed coordinate system; a reference direction vector is determined within a reference coordinate system based on the path position of the spacecraft and an orbit model; the overall spin vector of the spacecraft is determined within the body-fixed coordinate system; and a reference overall spin vector of the spacecraft is determined within a reference coordinate system by time propagation of known initial values of the overall spin of the spacecraft or by time tracking of a reference model. The position of the spacecraft is determined based on the four vectors.

Abstract: The invention describes a method for the fuel-optimized selection of a configuration of thrusters on a spacecraft while resolving a linear optimization problem with an initialization phase for finding a first permissible solution and a subsequent iteration phase, in which proceeding on the permissible solution an iterative optimization of an efficiency criterion takes place. In each iteration step a scaled iteration gradient is formed, and the iteration gradient is multiplied with a limiting factor for a maximum iteration step width, which is formed while taking at least one boundary value condition for a permissible solution into account.

Abstract: In a method for determining the position of a spacecraft based on vector determinations, a direction vector is measured in a body-fixed coordinate system; a reference direction vector is determined within a reference coordinate system based on the path position of the spacecraft and an orbit model; the overall spin vector of the spacecraft is determined within the body-fixed coordinate system; and a reference overall spin vector of the spacecraft is determined within a reference coordinate system by time propagation of known initial values of the overall spin of the spacecraft or by time tracking of a reference model. The position of the spacecraft is determined based on the four vectors.

Abstract: The invention describes a method for the fuel-optimized selection of a configuration of thrusters on a spacecraft while resolving a linear optimization problem with an initialization phase for finding a first permissible solution and a subsequent iteration phase, in which proceeding on the permissible solution an iterative optimization of an efficiency criterion takes place. In each iteration step a scaled iteration gradient is formed, and the iteration gradient is multiplied with a limiting factor for a maximum iteration step width, which is formed while taking at least one boundary value condition for a permissible solution into account.

Abstract: In a process and apparatus for controlling the attitude of a spacecraft, drive signals for driving attitude control units of the spacecraft are generated based on control unit signals related to the three principle axes of a spacecraft oriented coordinate system, using the rule k=B.sup.I e+cu.sub.2. The drive matrix B.sup.I, a definite vector u.sub.2 and the constant c result from the application of a matrix decomposition according to the Singular Value Decomposition (SVD) method to a nozzle matrix B. The latter matrix contains as elements, the momentum and force vectors which can be generated by the attitude control units, and the drive matrix B.sup.I is multiplied directly with the vector e derived from the axis-related control unit signals.

Abstract: A triaxially stabilized earth oriented satellite is provided with an attitude control system. The attitude control system contains a controller, final control elements for generating control torques around each of three axes of a system of coordinates (principle axes x, y, z) fixed relative to the satellite. Additionally, the attitude control system comprises, as measuring transducers, exclusively a direction vector measurement device, like a magnetometer or star sensor, a sun sensor array and an earth sensor. The process of the invention makes use of the apparatus as noted above and includes the steps of:A. determining an estimate for the rotating velocity of the satellite from a direction vector measurement;B. searching for the sun, depending on the size of the field of view of the sensor, by one or two search rotations at right angles to the optical axis;C. adjusting the direction of the sun relative to the optical axis; andD. adjusting the direction vector measurement to a reference direction S.sub.

Abstract: A measurement arrangement is useful for controlling the attitude of a three-axis stabilized satellite equipped with sun-sensors for determining the orientation of the sun (sun vector) with respect to a satellite-fixed coordinate system, as well as with speed gyroscopes for detecting components of the satellite speed of rotation vector .omega.. It is necessary that the measurement range of the sun-sensors cover the round angle in a preselectable plane (for example XY plane) and perpendicularly thereto a limited angular range of maximum .+-..alpha..sub.2max on both sides of the plane. In addition, only an integrating speed gyroscope carrying out measurements in a single measurement axis that encloses with the plane an angle of at least (.pi./2)-.alpha..sub.2max should be provided.

Abstract: A three-axis stabilized, earth-oriented satellite has an attitude control system with regulators, actuators, an earth sensor that carries out measurements along two axes and a sun-sensor arrangement that also carries out measurements along two axes. The field of view of the sun-sensor arrangement covers the round angle on a plane of the satellite-fixed coordinate system. Only the sun-sensor arrangement and the earth sensor act as measurement transducers. A sun and earth acquisition process for such a satellite has the following steps: seeking the sun; setting the sun vector in a first direction of reference; setting the speed of rotation of the satellite around the sun vector at a constant value; setting the sun vector in a second direction of reference, so that by rotating the satellite around the latter the optical axis of the earth sensor sweeps over the earth; and picking up the earth. Special regulating rules are disclosed.

Abstract: A digital modulation process and apparatus for converting an input signal comprising a time variable series of sampling values, each of which sampling values remains constant within a fixed sampling period, into output signals representing a sequence of positive and/or negative pulses having a constant amplitude and a variable pulse width and frequency. The input signal and modulator parameters are supplied to a data processor. In order to reduce computing expenditures and at the same time achieve an increase in precision, by means of the computing system, the switch-on and switch-off points in time t.sub.se, t.sub.sa of the pulses as well as their operational signs are computed in one operation directly for each sampling period .DELTA.t.sub.k.

Abstract: A method and arrangement for aligning a space vehicle with respect to a reference object is disclosed. A measuring direction is obtained by determining the angular orientation direction of the space vehicle with respect to the reference object and subsequently the space vehicle is aligned by means of actuators such that the measuring direction corresponds to a reference direction on the space vehicle. The actuators are driven by control signals which are derived from the measuring direction to cause control torques on the space vehicle. The orientation of the space vehicle and reference object is determined only with respect to a single measuring direction along a main axis of a direction sensor.