Satellite system comprising two satellites attached to each other and method for launching them into orbit

Abstract

A satellite system includes a so-called carrier satellite and a so-called piggyback satellite, each one having an Earth face. The piggyback satellite is attached to the carrier satellite by fastening elements that can be released on command. The piggyback satellite includes propulsion elements suitable for maintaining same in orbit, and the carrier satellite includes propulsion elements for performing a change of orbit of the satellite system including the carrier satellite and the piggyback satellite. The piggyback satellite is attached to the Earth face of the carrier satellite in such a way that the Earth face of the piggyback satellite is essentially perpendicular to the Earth face of the carrier satellite.

Claims

1. A satellite system, comprising: a first satellite, called carrier satellite; and a second satellite, called piggyback satellite, said piggyback satellite being releasably fixed to said carrier satellite, said piggyback satellite being releasable, on command, from said carrier satellite to provoke separation of said carrier satellite and from said piggyback satellite, each of said carrier and piggyback satellites comprising an Earth face, wherein, said piggyback satellite comprises: a first propulsion system with a first propellant reserve suitable for keeping said piggyback satellite in orbit, and said carrier satellite comprises: a second propulsion system with a second propellant reserve configured to perform a change of orbit of said carrier satellite and of said piggyback satellite fixed to said carrier satellite, and said piggyback satellite is fixed onto the Earth face of the carrier satellite, such that the Earth face of said piggyback satellite is substantially at right angles to the Earth face of said carrier satellite.

2. The satellite system as claimed in claim 1, wherein the carrier satellite and the piggyback satellite are arranged in a same launch vehicle.

3. The satellite system as claimed in claim 1, wherein the carrier satellite and the piggyback satellite do not share any data bus.

4. The satellite system as claimed in claim 1, wherein the piggyback satellite remains passive while fixed to the carrier satellite, and becomes active when separated from the carrier satellite.

5. The satellite system as claimed in claim 1, wherein the carrier satellite comprises a substantially cylindrical rigid support structure which defines a longitudinal axis of said carrier satellite extending between the Earth face and an opposite anti-Earth face of said carrier satellite, and said piggyback satellite is fixed onto said support structure.

6. The satellite system as claimed in claim 1, further comprising heating lines located in said piggyback satellite.

7. The satellite system as claimed in claim 1, comprising a plurality of piggyback satellites stacked one on top of the other on the Earth face of the carrier satellite, the Earth face of each of said piggyback satellites being substantially at right angles to said Earth face of said carrier satellite.

8. A method for stationing on a mission orbit at least one of the satellites of a set of satellites comprising a first satellite, called carrier satellite, and a second satellite, called piggyback satellite, each of said satellites comprising an Earth face, said method comprising steps of: a) forming a satellite system as claimed in claim 1, b) placing said satellite system in a launch vehicle suitable for transferring the satellite system from the Earth's surface to an initial orbit, c) injecting said satellite system into said initial orbit by said launch vehicle, d) transferring said satellite system, by the propulsion system of said carrier satellite, into or in proximity to the mission orbit of said piggyback satellite, and e) separating said piggyback satellite from said carrier satellite.

9. The method as claimed in claim 8, wherein step a) comprises: forming a satellite system comprising a plurality of piggyback satellites stacked one on top of the other on the Earth face of the carrier satellite, the Earth face of each of said piggyback satellites being substantially at right angles to said Earth face of said carrier satellite, and and repeating steps d) and e) for each of said piggyback satellites.

10. The method as claimed in claim 8, further comprising, after step e), a step f) of transferring said carrier satellite into a carrier satellite orbit.

11. The method as claimed in claim 10, wherein the mission orbit of the carrier satellite is a geostationary orbit.

12. The method as claimed in claim 8, wherein the mission orbit of the piggyback satellite is a geostationary orbit.

13. A satellite system, comprising: a first self-propelled satellite configured for changing an orbit of the satellite system, the first self-propelled satellite being called a carrier satellite; and a second self-propelled satellite, called a piggyback satellite, said piggyback satellite being releasably fixed to said carrier satellite, said piggyback satellite being releasable, on command, from said carrier satellite to provoke separation of said carrier satellite and from said piggyback satellite, wherein each of said carrier and piggyback satellites comprises an Earth face, and wherein said piggyback satellite is fixed onto the Earth face of the carrier satellite, such that the Earth face of said piggyback satellite is substantially at right angles to the Earth face of said carrier satellite.

Description

BRIEF DESCRIPTION OF THE FIGURES

(1) The invention will now be described more specifically in the context of preferred embodiments, which are in no way limiting, represented in FIGS. 1 to 3, in which:

(2) FIG. 1 schematically represents a satellite system according to a particular embodiment of the invention, comprising a single piggyback satellite;

(3) FIG. 2 schematically represents a satellite system according to a different particular embodiment of the invention, comprising two piggyback satellites;

(4) FIG. 3 is a block diagram illustrating the steps of a method for stationing the satellites of the satellite system of FIG. 1 or of FIG. 2; and

(5) FIG. 4 is a schematic illustration of the carrier satellite and the piggyback satellite being arranged in a same launch vehicle.

DETAILED DESCRIPTION OF AN EMBODIMENT OF THE INVENTION

(6) A satellite system according to a particular embodiment of the invention is schematically represented in FIG. 1.

(7) This system comprises a first satellite 10, called carrier satellite, and a second satellite 20, called piggyback satellite.

(8) Each of these satellites is equipped with payloads, with sufficient power capacity to operate during its mission, with a propulsion system 102, 29 with a propellant reserve that are sufficient for the orbit keeping maneuvers, aiming to correct the position of the satellite on its mission orbit, and, if appropriate, for orbit-change maneuvers.

(9) All of these elements are well known to those skilled in the art and will not be described in detail in the present description.

(10) More particularly, the carrier satellite 10 comprises a substantially cylindrical rigid structure 11, called support structure, which is represented by dotted lines in FIG. 1, because it is arranged inside an outer jacket 101 of the carrier satellite, and which defines a longitudinal axis 12 of the carrier satellite 10, also called Earth axis. In the mission orbital position of the carrier satellite 10, this Earth axis 12 joins the center of mass of the satellite and the center of the Earth, the direction toward the Earth being indicated as 13 in FIG. 1.

(11) In a manner that is conventional in itself, the carrier satellite 10 comprises solar energy collection means, in the form of generally deployable solar panels, not represented in FIG. 1. The carrier satellite 10 further comprises communication means using deployable reflector antennas 14, represented schematically in FIG. 1 in the folded-down launch configuration. The support structure 11 supports and protects the propulsion means, the electrical power supply, control and navigation, telemetry and communication and other such systems with which the satellite is equipped, as well as the electric wiring. All of these elements, arranged in the outer jacket 101 of the satellite, are not represented in FIG. 1.

(12) The carrier satellite 10 comprises an Earth face 15 and an opposite anti-Earth face 16, which are globally substantially at right angles to the Earth axis 12.

(13) The carrier satellite 10 is intended to be fixed, for its launch, to a launch vehicle 50, schematically represented in FIG. 4, via its anti Earth face 16. More specifically, the carrier satellite 10 is intended to be arranged under the fairing of the launch vehicle, the anti-Earth face 16 then being positioned facing the Earth's surface, in the normal launch configuration.

(14) On its Earth face 15, the carrier satellite 10 bears sensors, deployable telemetry, control and other such antennas, one of which is represented by way of example in FIG. 1, identified by the reference 17, at the end of a mast 18 which can be dedicated to it or consist of an extension of a radiator of the satellite.

(15) The propulsion means of the carrier satellite 10, comprising the propulsion system in itself, such as nozzles, valves, tanks, etc., and the propellant reserve, are configured in such a way as to be suitable for ensuring at least an orbit transfer of the system comprising the carrier satellite 10 and the piggyback satellite 20 from an initial orbit into which this system is injected by a launch vehicle, to a mission orbit of at least one of the carrier 10 and/or piggyback 20 satellites, and for keeping the carrier satellite 10 alone in mission orbit.

(16) The piggyback satellite 20, which can be of the same class as the carrier satellite 10 or of a different class, comprises a structure 21, which defines its longitudinal axis, or Earth axis, 22. The direction of gaze toward the Earth of the piggyback satellite 20 in its mission orbit is indicated as 23 in FIG. 1.

(17) The piggyback satellite 20 comprises, just like the carrier satellite 10, an Earth face 25 and an opposite anti-Earth face 26, both globally substantially at right angles to the Earth axis 22 of the satellite.

(18) The propulsion means of the piggyback satellite 20, comprising the propulsion system in itself and the propellant reserve, are configured in such a way as to be suitable for ensuring at least its mission orbit-keeping. Preferentially, they are ill-suited to ensuring an orbit transfer of the piggyback satellite 20 from the initial orbit to its mission orbit, so that the weight and the volume occupied by the propulsion means are limited.

(19) As illustrated in FIG. 1, the piggyback satellite 20 is fixed onto the Earth face 15 of the carrier satellite 10, directly above the support structure 11 of the carrier satellite 10. It is fixed there in such a way that its Earth axis 22 is substantially at right angles to the Earth axis 12 of the carrier satellite 10, that is to say that the Earth face 25 of the piggyback satellite 20 is globally substantially at right angles to the Earth face 15 of the carrier satellite 10, more specifically globally forms, with the latter, an angle of less than 10 degrees. The piggyback satellite 20 is thus fixed onto the carrier satellite 10 by a so-called lateral first face 24, distinct from the Earth 25 and anti-Earth 26 faces.

(20) Such an arrangement of the piggyback satellite 20 on the carrier satellite 10 proves very advantageous in terms of optimization of the volume under the fairing of the launch vehicle. In effect, it makes it possible to stack, on the carrier satellite 10, a piggyback satellite 20 that has a significant volume, and notably an Earth face 25, which bears communication equipment, of large surface area. In such an arrangement, an increase in the volume of the piggyback satellite 20 is in fact advantageously little constrained by the dimensions of the fairing of the launch vehicle.

(21) Furthermore, such an arrangement of the piggyback satellite 20 on the carrier satellite 10 advantageously makes it possible to easily accommodate, on the Earth face 15 of the carrier satellite 10, sensors and/or antennas 17 that require, particularly in the phase of orbit transfer of the system formed by the carrier satellite 10 and the piggyback satellite 20, from the initial orbit to the mission orbit, a wide field of view 19 including the Earth axis 12 of the carrier satellite 10. As illustrated in FIG. 1, obtaining such a result does not require the implementation of a mast 18 of very great length bearing the sensor or the antenna 17.

(22) The piggyback satellite 20 is fixed onto the carrier satellite 10 by attachment means that can be released on command and without generating debris, for example by a series of pyrotechnic bolts.

(23) The interface between the carrier satellite 10 and the piggyback satellite 20 is, furthermore, reduced to the minimum, and implemented in such a way that no data bus links the two satellites, or their avionics. The piggyback satellite 20 is advantageously configured to remain passive as long as it is fixed to the carrier satellite 10. The system comprises means for slaving the starting up of the piggyback satellite 20 to the releasing of the attachment means, so that the piggyback satellite 20 is automatically started up upon its separation from the carrier satellite 10. These slaving means are notably arranged in the piggyback satellite 20. For its part, the command to release the attachment means is preferentially implemented from the carrier satellite 10.

(24) As long as the piggyback satellite 20 is attached to the carrier satellite 10, its avionics remaining off, the system is preferably configured in such a way that the carrier satellite 10 is able to heat up the piggyback satellite 20, via an extension to the piggyback satellite of its heating harness (not visible in FIG. 1). This heating harness is preferentially controlled and powered by means provided for this purpose and located in the carrier satellite 10.

(25) The system can further comprise a so-called trickle charge mechanism making it possible to avoid the discharging of the batteries of the piggyback satellite 20 as long as the latter is fixed to the carrier satellite 10.

(26) A different embodiment of the satellite system according to the invention is schematically represented in FIG. 2.

(27) This system comprises a carrier satellite 10 and a piggyback satellite 20, similar to those described above with reference to FIG. 1, and arranged similarly relative to one another.

(28) It further comprises a second piggyback satellite 20′, which can be identical to or different from the first piggyback satellite 20, but which does take on the features thereof described above, notably concerning the propulsion means.

(29) This second piggyback satellite 20′ comprises an Earth axis 22′, an Earth face 25′ and an opposite anti-Earth face 26′. It is stacked on the first piggyback satellite 20, in such a way that its Earth axis 22′ is substantially at right angles to the Earth axis 12′ of the carrier satellite 10, and substantially parallel to the Earth axis 22 of the first piggyback satellite 20. In the particular embodiment represented in FIG. 2, the second piggyback satellite 20′ is fixed, by a so-called lateral face 24′ distinct from the Earth face 25′ and the anti-Earth face 26′, to a lateral face 27 of the first piggyback satellite 20 opposite the first lateral face 24.

(30) The stacking of the two piggyback satellites 20, 20′ on the carrier satellite 10 is preferably done along the longitudinal axis 12 thereof.

(31) In the particular embodiment represented in FIG. 2, the two piggyback satellites 20, 20′ are for example arranged head-to-tail, that is to say that their respective Earth faces 25, 25′ are arranged opposite one another. It is understood that other respective orientations of the Earth faces of the piggyback satellites can be envisaged.

(32) The particular arrangement of the piggyback satellites relative to the carrier satellite according to the invention makes it possible to stack thereon a plurality of piggyback satellites one on top of the other, by occupying the volume available under the fairing of the launch vehicle optimally, and by having the weight of the piggyback satellites supported by the support structure 11 of the carrier satellite 10, which is intrinsically strong enough for this purpose. This stacking thus advantageously does not require any specific modification of the structure of the carrier satellite 10.

(33) The second piggyback satellite 20′ is fixed onto the first piggyback satellite 20 by attachment means, which can be released on command for the separation of the piggyback satellites from one another. The release command for these attachment means can be implemented from the carrier satellite 10, or from the first piggyback satellite 20, depending on the sequence of separation of all of these satellites that is envisaged.

(34) The different steps of a nonlimiting exemplary method for stationing, on their mission orbit, satellites of the satellite systems described above with reference to FIGS. 1 and 2, are schematically illustrated in FIG. 3.

(35) In a first step 31, the piggyback satellite 20 is fixed onto the carrier satellite 10, by attachment means, such that the respective Earth faces 15, 25 of these satellites are arranged globally substantially at right angles to one another. These operations are performed on the Earth's surface.

(36) In a second step 32, the duly formed satellite system is assembled under the fairing of a launch vehicle, in a manner that is conventional in itself.

(37) The next step 33 consists of the launch from the Earth's surface and the injection of the satellite system, still with the satellites fixed to one another, into the initial orbit, for example a geostationary transfer orbit.

(38) The satellite system is then transferred, in a step 34, by means of the propulsion system of the carrier satellite 10, into the mission orbit of the piggyback satellite 20, for example the geostationary orbit.

(39) The next step 35 is the separation of the piggyback satellite 20 and of the carrier satellite 10, by the releasing of the attachment means triggered by the carrier satellite 10.

(40) When the system comprises a plurality of piggyback satellites 20, 20′, all of these piggyback satellites can be detached from the carrier satellite 10 in this step. Otherwise, the steps 34 and 35 can be reiterated for each piggyback satellite, as indicated at 36 in FIG. 3.

(41) Finally, in a final step 37, each satellite is brought, by its own propulsion system, into its operational position.

(42) Each of these steps is performed in a manner that is conventional in itself.

(43) The exemplary method according to the invention schematically described above is in no way limiting on the invention, and any variant falls equally within the scope of the invention.