Aeronautics and astronautics – Spacecraft – With payload accommodation
Reexamination Certificate
2000-02-14
2002-04-30
Carone, Michael J. (Department: 3641)
Aeronautics and astronautics
Spacecraft
With payload accommodation
C244S172200
Reexamination Certificate
active
06378810
ABSTRACT:
BACKGROUND OF THE INVENTION
The present invention relates to satellites for placing in non-equatorial low orbit (LEO) and for carrying a payload which is generally a telecommunications payload. Such a satellite is articularly, but non-exclusively, suitable for use as a telecommunications satellite belonging to a constellation of satellites that are distributed in longitude in such a manner as to be directly in line of sight of one another.
The electrical power required for operating the payload of a satellite, e.g. a telecommunications satellite, is supplied by a solar generator comprising panels whose orientation relative to the platform constituted of the satellite need be modified as the satellite travels round its orbit so as to ensure that the panels present the greatest possible apparent area to sunlight.
For a satellite on a low orbit which is inclined relative to the equator, that means that the solar generator need to be given angular motion about a plurality of axes so as to keep the panels properly oriented and so as to obtain high efficiency. Such motion must not interfere with the field of view of other units of the satellite and in particular its antennas and transducers for links with the ground and for links between the satellites of a constellation that are on orbits that are at the same altitude, but that are mutually offset.
Furthermore, the solar generator must be apt to be taken from a launch configuration in which it has a compact configuration to an operating configuration in which it is spread.
SUMMARY OF THE INVENTION
It is an object of the present invention to provide a satellite whose solar generator is organized and mounted on the platform of the satellite in such a manner as to enable it to perform movements necessary for achieving high energy efficiency, and to do so while avoiding interfering with radio or optical links and/or with detection elements.
To this end, there is provided a satellite having a platform that is elongate in a determined direction (generally designed to constitute its roll axis) and that has a longitudinally-extending face that carries a deployable arm with the solar generator being mounted at the end thereof by means that enable the generator to be rotated through a range of 360° about a first axis which is orthogonal to the direction of elongation of the platform and that enable it to rotate through a limited angular range about a second axis that is orthogonal to the first axis and parallel to the plane of the generator. This limited angular movement can be referred to as “bias”.
Typically, the means for rotating the generator are designed to rotate it about the first axis (the pitch axis) at the orbital period. In contrast, the bias movement is seasonal, between two extreme positions corresponding to orbits for local times close to 6:00 AM and 6:00 PM.
The hinged arm is advantageously deployable from a storage condition in which the folded or rolled-up solar generator can contact the side surfaces of the platform, to a deployed state in which it holds the solar cells away from the platform.
Over one orbit or one day, the generator must deliver at least as much electrical energy as is consumed by the satellite. Unfortunately, the amount of power delivered depends both on the surface area of the generator and on the inclination of the generator relative to the solar flux. Ideal bias motion would enable the solar cells to be oriented under all circumstances so as to be orthogonal to the light flux. However various considerations often put a limit on bias.
In a particular embodiment of the invention, where the solar generator comprise two wings disposed symmetrically relative to the arm, it is possible to reduce the consequences of bias limitation for the wing which extends towards the half-space which is on the Earth side of the satellite (towards nadir). For this purpose, the wings are provided with independent bias motors. The motors are controlled symmetrically responsive to time at the orbital period so long as there is no risk that the bias which would be optimum would result in interfering with the fields of view of link or detection elements. Beyond that limit position, the bias motors are controlled:
to give the wing on the side of the half-space away from the Earth motion that ensures optimum orientation therefor at all times, and
to limit the amplitude of the motion imparted to the wing on the side of the half-space placed towards the Earth.
Maximum bias therefore may be larger away from the platform than towards the platform, since in the away direction there is no risk of interference with the fields of view of transducers or of antennas.
The cross-section of the platform is advantageously approximately trapezoidal. The platform may have a rigid wall or panel occupying its small base and carrying the most service elements. The arm is advantageously arranged to enable the wings of the solar generator to be maintained in a location and in an orientation such that in a storage configuration they are maintained applied against the side faces of the platform for launch. The length of the arm, and thus the height at which the main axis of rotation (pitch axis) of the wings of the solar generator is located, constitutes a compromise between eliminating interference with the members carried by the platform and accomodating the increased offset of the center of gravity with increasing arm length, and also the increasing difficulties in organizing the storage configuration.
The invention also provides a method of bringing a satellite of the above-defined kind from a transfer orbit to a final low earth orbit,comprising the steps of:
deploying the solar generator while leaving the arm in storage condition, so as to keep down disturbing torque;
orienting a main thruster member so as to take account both of the position of the center of gravity of the satellite while in this condition and of the required direction for upward thrust so as to generate spiral motion that raises the altitude of the orbit;
once the altitude reached is compatible with the capacity for controlling disturbing torques acting on the satellite (generally in the range 550 km and 600 km), deploying the arm; and
continuing upward movement by operating the main thruster until the satellite has reached its station, with the satellite always remaining pointed towards the Earth.
During the first portion of upwards movement, while the solar generator is deployed but the arm is not deployed, it is possible to drive the solar cells about the pitch axis and about the orthogonal axis, or to leave them stationary.
The main thruster may be a nozzle capable of taking up two orientations relative to the platform of the satellite in order to take account of the variation in the position of the center of gravity depending on whether the arm is or is not deployed. In both cases, the angular position of the nozzle is selected in optimum manner to take disturbances into account.
The above features and others will appear better on reading the following description of a particular embodiment of the invention, given by way of non-limiting example. The description refers to the accompanying drawings.
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Kelly W.D., Hill C.D.: “Solar panel tracking in Low Earth Orbit with Earth-oriented spacecraft”—Advances in the Astronautical Sciences, vol. 76, No. 1, Aug. 19-22, 1991, pp. 743-762, XP002115699—San Diego, CA, U.S.A., Abstract, Fig. 1A, col. 4, line 33-line 48.
Patent Abstracts of Japan—vol. 013, No. 567 (M-908), Dec. 15, 1989 & JP 01 237296 A (Mitsubishi Electric Corp.), Sep. 21, 1989, Abstract.
Pham Philippe
Polle Bernard
Carone Michael J.
Larson & Taylor PLC
Matra Marconi Space France
Sukman Gabriel S
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