METHOD AND SYSTEM FOR TRANSFERRING A SATELLITE FROM AN INTIAL ORBIT INTO A MISSION ORBIT

Abstract

A system and method for transferring a satellite from an initial orbit into a mission orbit. The method includes anchoring to the satellite of an external unit having a tank containing a reserve of propellants. The system includes an autonomous spacecraft having an electric propulsion module and a small internal reserve of propellants, located in a parking orbit close to the initial orbit. The spacecraft with the external unit attached to the satellite is docketed in an initial orbit, to produce a fluidic connection of the propellant tank of the external unit to the propulsion module of the spacecraft. The external unit and satellite is transferred into the mission orbit by the electric propulsion module of the spacecraft supplied with propellants directly from the external unit, thereby releasing the satellite into the mission orbit.

Claims

1-22. (canceled)

23. A method of transferring a satellite from an initial orbit to a mission orbit of the satellite around a celestial body, comprising the steps of: in the initial orbit, docking of a spacecraft to an external propellant supply unit, the spacecraft comprising an electrical propulsion module and a fluidic propellant supply circuit to supply a propellant to the electrical propulsion module, the external propellant supply unit comprises a propellant tank containing a reserve of the propellant and the external propellant supply unit is anchored to the satellite by cooperating anchors carried respectively by the external propellant supply unit and the satellite; wherein the spacecraft and the external propellant supply unit comprise cooperating docking units to reversibly dock with each other, the cooperating docking units comprise cooperating fluidic connectors to provide a fluidic connection between the propellant tank and the fluidic propellant supply circuit to supply the propellant to the electrical propulsion module from the external propellant supply unit during the docketing step; transferring the docketed spacecraft to an assemblage comprising the external propellant supply unit anchored to the satellite to the mission orbit by the electrical propulsion module of the spacecraft supplied with the propellant from the propellant tank of the external propellant supply unit; and releasing the satellite into the mission orbit.

24. The method as claimed in claim 23, wherein the spacecraft is supplied with the propellant directly from the propellant tank of the external propellant supply unit during the transferring step.

25. The method as claimed in claim 23, further comprising a rendezvous step of the spacecraft rendezvousing with the assemblage in the initial orbit.

26. The method as claimed in claim 23, wherein the cooperating anchors are releasable on a command to produce a separation of the external propellant supply unit and the satellite; and wherein the step of releasing the satellite into the mission orbit is accomplished by the command to release the cooperating anchors to separate the external propellant supply unit and the satellite from each other.

27. The method as claimed in claim 23, wherein the spacecraft comprises a propellant reservoir in a fluidic connection with the fluidic propellant supply circuit; and further comprising, between the docking step and the transferring step, a second transferring step of transferring the propellant from the propellant tank of the external propellant supply unit to the propellant reservoir of the spacecraft.

28. The method as claimed in claim 27, wherein the propellant contained in the propellant tank of the external propellant supply unit is of a gaseous type; and wherein the second transferring step transfers the propellant via a fluidic connection between the propellant tank and the propellant reservoir by an equilibration of pressures between the propellant tank and the propellant reservoir.

29. The method as claimed in claim 23, further comprising, after the step of releasing the satellite into the mission orbit, a step of moving the spacecraft to a parking orbit.

30. A method of remotely controlling the spacecraft for performing the steps of the transfer method as claimed in claim 23, the spacecraft comprising the electrical propulsion module, the fluidic propellant supply circuit to supply the propellant to the electrical propulsion module, the cooperating docketing unit to reversibly dock to the external propellant supply unit, and the cooperating fluidic connector to provide the fluidic connection between the external propellant supply unit to the fluidic propellant supply circuit to supply the propellant to the electrical propulsion module from the external propellant unit; and comprising steps of remotely controlling the spacecraft by a control unit; successively determining and transmitting control signals to the spacecraft by the control unit to perform the steps of the transfer method.

31. A spacecraft to perform the transfer method as claimed in claim 23, comprising: the electrical propulsion module; the fluidic propellant supply circuit to supply the propellant to the electrical propulsion module; a propellant reservoir in a fluidic connection with the fluidic propellant supply circuit; the cooperating docking unit to reversibly dock the spacecraft to the external propellant supply unit, and comprising the cooperating fluidic connector to provide the fluidic connection between the fluidic propellant supply circuit and the external propellant supply unit; and wherein the fluidic propellant supply circuit is configured to supply the propellant directly to the electrical propulsion module from the external propellant supply unit, without passing through the propellant reservoir of the spacecraft, and to establish a fluidic connection between the propellant reservoir and the external propellant supply unit.

32. The spacecraft as claimed in claim 31, wherein the fluidic propellant supply circuit comprises a sealing member configured to move between a closure position, preventing circulation of a fluid in the fluidic propellant supply circuit between the propellant reservoir and the electrical propulsion module, and an opening position enabling the circulation of the fluid; and an actuator to actuate the sealing member between the closure position and the opening position.

33. The spacecraft as claimed in claim 31, wherein the fluidic propellant supply circuit comprises a sealing member enabling a passage of a fluid to the propellant reservoir and preventing the passage of the fluid from the propellant reservoir.

34. The spacecraft as claimed in claim 31, further comprising a propulsion system to rendezvous with the external propellant supply unit located in the initial orbit.

35. The spacecraft as claimed in claim 31, further comprising an electrical connector to establish an electrical connection between the spacecraft and the external propellant supply unit.

36. An external propellant supply unit for implementing the transfer method as claimed in claim 23, comprising: the propellant tank; the cooperating anchor to the external propellant supply unit to the satellite, situated in a first zone of the external propellant supply unit; and the cooperating docking unit configured to cooperate with the cooperating docking unit of the spacecraft to reversibly dock with each other, the spacecraft situated in a second zone of the external propellant supply unit and comprising the cooperating fluidic connector to provide the fluidic connection between the propellant tank and the spacecraft.

37. The external propellant supply unit as claimed in claim 36, wherein the cooperating anchor is releasable on a command to produce a separation of the external propellant supply unit and the satellite.

38. The external propellant supply unit as claimed in claim 36, further comprising an electrical connector to establish an electrical connection between the external propellant supply unit and the spacecraft.

39. The external propellant unit as claimed in claim 37, further comprising an electrical connector to establish an electrical connection between the external propellant supply unit and the spacecraft; and wherein the release of the cooperating anchor is slaved to the electrical connector such that the release of the cooperating anchor is controllable from the spacecraft.

40. The external propellant supply unit as claimed in claim 36, wherein the cooperating anchor is configured to simultaneously anchor the external propellant supply unit to a plurality of satellites.

41. An assemblage, comprising: the external propellant supply unit as claimed in claim 36, the propellant tank of the external propellant supply unit contains the propellant for an electrical propulsion; the satellite comprising the cooperating anchor configured to cooperate with the cooperating anchor of the external propellant supply unit; and wherein the external propellant supply unit is secured, in a reversible manner, to the satellite by a cooperation of the cooperating anchor of the satellite and the cooperating anchor of the external propellant supply unit.

42. The assemblage as claimed in claim 41, further comprising a plurality of satellites; wherein the cooperating anchor of the external propellant supply unit is configured to simultaneously anchor the external propellant supply unit to said plurality of satellites, each satellite comprises a cooperating anchor configured to cooperate with the cooperating anchor of the external propellant supply unit, and the external propellant supply unit is secured, in the reversible manner, to each satellite by the cooperation of the cooperating anchor of said each satellite and the cooperating anchor of the external propellant supply unit.

43. A system to transfer of a satellite from an initial orbit to a mission orbit around a celestial body, comprising the spacecraft as claimed in claim 31 and an external propellant supply unit for implementing the transfer method, the external propellant supply unit comprising: the propellant tank; the cooperating anchor to the external propellant supply unit to the satellite, situated in a first zone of the external propellant supply unit; and the cooperating docking unit configured to cooperate with the cooperating docking unit of the spacecraft to reversibly dock with each other, the spacecraft situated in a second zone of the external propellant supply unit and comprising the cooperating fluidic connector to provide the fluidic connection between the propellant tank and the spacecraft.

44. A method of positioning a satellite in a mission orbit around a celestial body, comprising the steps of: forming the assemblage as claimed in claim 41 by attaching the external propellant supply unit to the satellite and introducing the propellant for electrical propulsion into the propellant tank of the external propellant supply unit; launching the assemblage into space, and releasing the assemblage into the initial orbit around the celestial body; and performing the steps of the transfer method to transfer the assemblage from the initial orbit to the mission orbit.

Description

[0113] The characteristics and advantages of the invention will appear more clearly in light of the following example embodiments and modes of implementation, provided merely as an illustration and in no way limiting the invention, supported by FIGS. 1 to 4, in which:

[0114] FIG. 1 represents schematically an external unit according to the invention, attached on the one hand to a satellite and on the other hand to a launch vehicle, in launch configuration;

[0115] FIG. 2 represents schematically an external unit according to the invention, attached on the one hand to a satellite and on the other hand to a spacecraft according to the invention;

[0116] FIG. 3 illustrates schematically the fluidic connection between an external unit and a spacecraft according to the invention, as well as different configurations for circulation of fluid within said spacecraft; and

[0117] FIG. 4 illustrates schematically the different steps of a method of positioning a satellite in its mission orbit, according to a particular embodiment of the invention.

[0118] In the following description, for greater convenience the invention shall be described with reference to a launch configuration of a satellite from the Earth's surface, for a stationing in a geostationary mission orbit (GEO) realized in two steps, involving a first step of injection of the satellite into an initial low Earth orbit (LEO). However, the invention is in no way limited to such a configuration, and it likewise applies in similar fashion to all other combinations of launch surface, initial orbit and mission orbit.

[0119] A system according to the invention for the positioning of a satellite 10 in its mission orbit involves the following different components: an external unit 20, a spacecraft 30, which shall be called hereinafter the transfer vehicle, and a control unit of the spacecraft 50.

[0120] In the following description the case will be taken of the stationing of a single satellite, it being understood that this teaching is directly transposable to the stationing of a plurality of satellites launched at the same time by the same launch vehicle, and all anchored to the same external unit. The external unit thus comprises for this purpose anchoring means for reversible anchoring to each of said plurality of satellites, and means of releasing each of these satellites. These means of releasing can be actuatable both jointly and separately, so that by choice the satellites can be released either at the same time into the same orbit, or successively into different orbits.

[0121] Schematically, the external unit 20 is designed to be secured to the satellite 10 on the Earth's surface, in order to form an assemblage intended to be injected into the LEO orbit. As for the transfer vehicle 30, this is designed to take over this assemblage in the LEO orbit, in order to carry out its transfer to the GEO orbit. The transfer vehicle, after having released the external unit 20 into a graveyard orbit, if applicable, is intended to return to station in a so-called parking orbit, in proximity to the LEO orbit, for the purpose of taking over a new assemblage comprised of a new satellite and a new external unit, to be transferred in turn to the GEO orbit (or another mission orbit).

[0122] The satellite 10, shown schematically in FIG. 1, is classical in itself. In the area of its surface 11 normally intended to form an interface with a launch vehicle, this satellite 10 is provided with anchoring means for anchoring to the external unit 20. The external unit 20 comprises for this purpose cooperating anchoring means. These anchoring means, not shown in FIG. 1, can be of any type customarily used in the field of space travel. In particular, they can be pyrotechnical devices, magnetic or electromagnetic coupling means, or mechanical rings, such as the so-called cup and cone system, pyrotechnical bolts or mechanisms of clamp-band type currently used in the area of satellite/launcher interfaces. Such anchoring systems are well known to the skilled person.

[0123] The anchoring means for anchoring the external unit 20 to the satellite 10 are configured in particular to realize, besides a mechanical connection, an electrical and communication connection between these two elements.

[0124] The external unit 20 itself comprises means of reversible fixation, not visible in FIG. 1, to a launch vehicle 61. This launch vehicle 61 is classical in itself and adapted to realize, from Earth, the injection of the assemblage formed by the satellite 10 and the external unit 20 into the LEO orbit.

[0125] In FIG. 1, the satellite 10, the external unit 20 and the launch vehicle 61 are shown attached to each other, in a configuration for launching into space.

[0126] The external unit 20 comprises a propellant tank 21, adapted to receive a volume 22 of propellant in the form of a gas, such as xenon.

[0127] It moreover comprises docking means for reversible docking to the transfer vehicle 30, able to cooperate with cooperating docking means carried by the latter, and preferably arranged in the area of a surface of the external unit 20 opposite the surface where the anchoring means for anchoring to the satellite 10 are arranged.

[0128] FIG. 2 shows schematically the satellite 10 attached to the external unit 20, itself attached to the transfer vehicle 30, in a so-called transfer configuration.

[0129] The cooperating docking means of the external unit 20 and the transfer vehicle 30 can be of any type known in themselves to the skilled person, enabling a reversible docking. These docking means are for example of the mechanical, magnetic or electromagnetic type. They comprise a first means 23 disposed on the external unit 20 and a second means 33 disposed on the spacecraft 30, able to cooperate and ensure the docking in reversible manner, and illustrated schematically in FIG. 2.

[0130] The transfer vehicle 30 comprises means of rendezvous with the external unit 20 which are classical in themselves. The external unit 20 comprises cooperating passive means of rendezvous, also classical in themselves, and comprising for example a series of light-emitting diodes 24, whose signal is used by the spacecraft 30 to realize the approach and coupling maneuvers.

[0131] The transfer vehicle 30 constitutes an autonomous space vehicle. In particular, it comprises an electrical propulsion module 37, a fluidic propellant supply circuit 39 for supplying propellant to this propulsion module 37, means of communication, especially with a remote control unit 50, especially on the ground, means of maneuvering, etc. It can also comprise a propellant reservoir 35.

[0132] Besides the means of electromagnetic connection, the docking means of the transfer vehicle 30 and of the external unit 20 likewise comprise means of electrical and communication connection. In particular, the external unit 20 is configured so as to constitute an interface between the spacecraft 30 and the satellite 10, to allow the transmission of signals from one to the other. Moreover, the anchoring means for anchoring the external unit 20 to the satellite 10 are slaved to the means of electrical connection between the external unit 20 and the transfer vehicle 30, so that the releasing of the satellite 10 can be controlled from the transfer vehicle 30.

[0133] The docking means for docking the transfer vehicle 30 and the external unit 20 also comprise means of fluidic connection to each other, more precisely between the tank 21 of the external unit 20 and the fluidic propellant supply circuit 39 of the propulsion module 37 of the transfer vehicle 30.

[0134] These means of fluidic connection, as well as the fluidic circuit 39 of the transfer vehicle 30, are illustrated schematically in FIG. 3. This figure also illustrates the different configurations of fluid circulation from the external unit 20, and within the transfer vehicle 30.

[0135] The propellant tank 21 of the external unit 20 is dimensioned to contain a volume of propellant 22 adapted to the supplying of the electrical propulsion module 37 of the spacecraft 30 to carry out the following phases: transfer of the system composed of the transfer vehicle 30, the external unit 20 and the satellite 10 from the injection orbit LEO to the mission orbit GEO of the satellite; transfer of the system composed of the transfer vehicle 30 and the external unit 20 from this mission orbit GEO to a graveyard orbit for the external unit 20; return of the transfer vehicle 30 to a parking orbit close to the initial orbit; maintaining of the transfer vehicle 30 in the parking orbit; transfer of the transfer vehicle 30 to the initial orbit, to take over a new satellite whose stationing has to be carried out.

[0136] The propellant reservoir 35 of the transfer vehicle 30 for its part is dimensioned to contain a volume of propellant 36 adapted to the supplying of the electrical propulsion module 37 of the transfer vehicle 30 for the performance of the following phases: return of the transfer vehicle 30, from the graveyard orbit of the external unit 20, to a parking orbit near the initial orbit; maintaining of the transfer vehicle 30 in the parking orbit; transfer of the transfer vehicle 30 to the initial orbit, for taking over a new satellite whose stationing has to be carried out.

[0137] In an initial configuration, illustrated in (a) in FIG. 3, the propellant tank 21 of the external unit 20 is filled with propellant, and the propellant reservoir 35 of the transfer vehicle is substantially empty. This configuration corresponds to a phase of rendezvous, in the initial orbit, of the transfer vehicle 30 and the external unit 20.

[0138] The external unit 20 comprises means 25 of fluidic connection of the propellant tank 21 with the transfer vehicle 30. The transfer vehicle 30 comprises cooperating means 38 of fluidic connection of the external unit 20 to the fluidic propellant supply circuit 39 for supplying propellant to the electrical propulsion module 37.

[0139] This fluidic circuit 39 comprises a so-called supply branch 40, connecting the means of fluidic connection 38 and the electrical propulsion module 37. In FIG. 3, this branch forms a right angle.

[0140] The fluidic circuit 39 furthermore comprises a diversion branch 41, extending between two points of the supply branch 40. The propellant reservoir 35 is connected to this diversion branch 41 by a channel 42.

[0141] On the fluidic circuit 39 there are mounted two antireturn flaps which enable the circulation of fluid from the means of fluidic connection 38 to the reservoir 35, yet prevent any circulation of fluid in the opposite direction. A first antireturn flap 43 is mounted on the supply branch 40, upstream of the diversion branch 41 in the direction of circulation of the fluid coming from the fluidic connection means 38. A second antireturn flap 44 is mounted on the diversion branch 41, upstream of the branching point of the channel 42, in the direction of circulation of the fluid coming from the means of fluidic connection 38.

[0142] The fluidic circuit 39 also comprises a valve 45, actuatable to open and close, especially a pyro-valve, mounted on the diversion branch 41, so as to be able to seal the latter in the position of closure, downstream from the branching point of the channel 42, in the direction of circulation of the fluid coming from the reservoir 35 and going to the electrical propulsion module 37. The transfer vehicle comprises means to control the opening and the closure of the valve 45.

[0143] In the initial configuration illustrated in (a) of FIG. 3, the valve 45 is in the closed position.

[0144] During the docking of the transfer vehicle 30 to the external unit 20, as illustrated in (b) of FIG. 3, the fluidic connection is established between the means of fluidic connection 25 of the external unit 20 and the means of fluidic connection 38 of the transfer vehicle 30. The tank 21 and the reservoir 35 are then in fluidic connection with each other. By a spontaneous equilibration of their internal pressures, and owing to the substantial difference in internal volumes of the tank 21 on the one hand and the reservoir 35 on the other, a transfer of propellant then occurs from the tank 21 to the reservoir 35, as indicated at 26 in the figure, resulting in a filling of the reservoir by a volume of propellant 36 substantially corresponding to its maximum capacity. The pressure after the transfer remains elevated in the tank 21. Such a configuration of the external unit 20 and the transfer vehicle 30 advantageously makes it possible to dispense with the use of costly, heavy and complicated equipment, such as pumps, to realize the transfer of propellant from the tank 21 of the external unit 20 to the reservoir 35 of the transfer vehicle 30.

[0145] In particular embodiments of the invention, the spacecraft 30 furthermore comprises a cold gas propulsion system, and a specific reservoir different from the reservoir 35 used by the electrical propulsion module 37, known as the main reservoir (elements not shown in the figures). This specific reservoir is advantageously arranged in the spacecraft 30 so that it can be supplied before the main reservoir 35, in order to maximize the pressure therein and thereby maximize the specific impulse of the cold gas propulsion system, which proves to be advantageous as compared to the use of a joint reservoir when the cold gas propulsion is implemented primarily during the rendezvous phase.

[0146] During the phase of transfer from the LEO orbit to the GEO orbit, the electrical propulsion module 37 of the transfer vehicle 30 draws from the propellant reserves available to it. The valve 45 being always closed, and the antireturn flap 44 being operating, its supplying is accomplished directly and solely from the tank 21 of the external unit 20, via the supply branch 40, as indicated at 27, in (c) of FIG. 3. The quantity of propellant contained in the tank 21 allows it perform the transfer to the GEO orbit and then from there to the graveyard orbit of the external unit 20.

[0147] When the external unit 20 is released by the transfer vehicle 30, as illustrated in (d) of FIG. 3, the means of fluidic connection 25 and 38 are disconnected from each other. The reservoir 35 contains a substantially maximum volume of propellant 36. In order to ensure the supply of propellant to the electrical propulsion module 37, the valve 45 is actuated to open, so as to allow the circulation of fluid from the reservoir 35 toward the electrical propulsion module 37, as indicated at 46 in FIG. 3. The antireturn flap 43 prevents any return of fluid to the means of fluidic connection 38.

[0148] The electrical propulsion module 37 of the transfer vehicle 30 comprises a plurality of thrusters, chosen to provide a strong thrust. For example, these thrusters have a power of 20 kW each.

[0149] The propulsion module 37 is preferably configured such that these strong-thrust thrusters are located in the area of one surface of the transfer vehicle 30, distinct from the surface containing the docking means for docking to the external unit 20, for example opposite to this surface.

[0150] These thrusters can have variable thrust and specific impulse, enabling an operation with strong thrust/weak specific impulse or weak thrust/strong specific impulse, such as the thrusters using the Elwing technology.

[0151] In variants of the invention, the thrusters with strong thrust and weak specific impulse are associated with thrusters of weak thrust and strong specific impulse, for the implementing of operations of movement of the transfer vehicle 30 after separation from the satellite 10. These supplemental thrusters can be disposed in the same zone of the transfer vehicle 30 as the thrusters of strong thrust 47, or preferably in a distinct zone.

[0152] The system according to the invention furthermore comprises a remote control unit 50, configured to drive the different phases carried out by the transfer vehicle 30. For this purpose, the control unit 50 and the transfer vehicle 30 each comprise conventional remote communication means.

[0153] The control unit 50 is adapted to determine control signals which are sent to the transfer vehicle 30. These control signals are determined, for example, as a function of measurement signals received from the transfer vehicle 30, which are determined by different sensors (gyroscope, star tracker, etc.) of the latter.

[0154] The control unit 50 comprises for example at least one processor and at least one electronic memory in which a computer program product is stored, in the form of a set of program code instructions to be executed in order to carry out the different steps of a method of control of the transfer vehicle 30. In one variant, the control unit 50 likewise comprises a programmable logic circuit or circuits of FPGA, PLD, etc., type and/or specialized integrated circuits (ASICs) adapted to carry out some or all of said steps of the method of control. In other words, the control unit 50 comprises a set of means configured in software (specific computer program product) and/or hardware (FPGA, PLD, ASIC, etc.) mode to carry out the different steps of a method of control as described below.

[0155] The different steps of a method of positioning the satellite 10 in its mission orbit GEO 62, from the Earth's surface 60, by means of the system as described above, are illustrated in FIG. 4.

[0156] In a first step of the method, implemented at the Earth's surface 60, the satellite 10, comprising means of propulsion dimensioned solely to be able to ensure its maintaining in position in its mission orbit, such as a geostationary orbit, is attached to the external unit 20, by their cooperating anchoring means. This anchoring realizes a mechanical, electrical and communication connection between these two elements. The propellant tank 21 is filled with a volume 22 of gaseous propellant, such as xenon, and attached to the launch vehicle 61, shown in exploded view in FIG. 4.

[0157] In a subsequent step, the assemblage comprising the satellite 10 and the external unit 20 is launched from the Earth's surface 60 and injected, by the launch vehicle 61, into the LEO orbit 63, as illustrated at 72 in FIG. 4.

[0158] The transfer vehicle 30 has been previously placed in orbit, in conventional manner, in an Earth orbit called a parking orbit 64, near the LEO orbit 63. Its orbital maintenance in this orbit is ensured by its propulsion module 37, and a small quantity of propellant contained in its propellant reservoir 35.

[0159] After the injection of the satellite 10 and the external unit 20 into the LEO orbit 63, the method of positioning the satellite provides for the controlling of the transfer vehicle 30, by the remote control unit 50 situated for example on the Earth's surface 60, to carry out the following successive phases.

[0160] In a first phase, indicated at 73 in FIG. 4, the transfer vehicle performs a change of orbit, moving into the LEO orbit 63. Here it carries out rendezvous operations with the external unit 20, in a manner conventional per se. For this purpose, the external unit 20 moves into the adequate orbital position.

[0161] In a second phase 74, still in the LEO orbit 63, the spacecraft 30 docks with the external unit 20, by the cooperating docking means carried respectively by the spacecraft 30 and the external unit 20. This docking produces, between these two elements, a fluidic, pneumatic, electrical and communication connection. In particular, it produces a fluidic connection between the propellant tank 21 of the external unit 20 and the fluidic propellant supply circuit 39 for supplying propellant to the electrical propulsion module 37 of the transfer vehicle 30.

[0162] After realizing this fluidic connection, a transfer of propellant is carried out by spontaneous equilibration of pressures between the tank 21 of the external unit 20 and the reservoir 35 of the transfer vehicle 30. Once the transfer is finished, the following phase 75 is carried out.

[0163] This phase 75 consists in a transfer of the system formed by the transfer vehicle 30, the external unit 20 and the satellite 10 from the LEO orbit 63 to the mission orbit GEO 62. This transfer is realized by propulsion by means of the propulsion module 37 of the transfer vehicle 30, supplied with propellant directly from the tank 21 of the external unit 20. It is rapid, by virtue of the advantageous characteristics of the electrical propulsion module 37. The time of exposure of the satellite 10 to the radiation of the Van Allen belts is consequently reduced.

[0164] The transfer vehicle 30 then controls the releasing of the anchoring means anchoring the external unit 20 to the satellite 10, for the injection of the latter into the GEO orbit 62.

[0165] In one variant of the invention, it is the assemblage composed of the external unit 20 and the satellite 10, joined to each other, which is injected into the GEO orbit 62 by the transfer vehicle 30, which then continues its mission alone.

[0166] During all these phases, the satellite 10 has remained passive, in particular its means of telecommunication remaining non-deployed. However, its control from the control unit 50 has been made possible via the transfer vehicle 30, and the external unit 20 forming an interface for the transmission of signals between the latter and the satellite 10. The satellite 10 in particular can be configured such that the releasing of its anchoring means for anchoring to the external unit 20 automatically triggers its placement in service.

[0167] In certain cases, it may be advantageous during one or more of these phases to deploy the solar panels of the satellite, in order to boost the power available for the transfer vehicle 30. Then the satellite is no longer passive. In particular embodiments of the invention, the spacecraft 30, the external unit 20 and the satellite 10 are provided for this purpose with means of transmission of power between the satellite 10 and the spacecraft 30.

[0168] From the GEO orbit 62, the next phase 77 may consist in a transfer of the transfer vehicle 30, still docked to the external unit 20, to a graveyard orbit 65. In this graveyard orbit 65, the transfer vehicle 30 is controlled to release the docking means for docking to the external unit 20, as indicated at 78 in FIG. 4, so as to release the external unit 20 there. The graveyard orbit 65 may be situated above or below the GEO orbit 62. When the graveyard orbit 65 is located below the geostationary arc, this operation does not require additional Delta V.

[0169] During all of these phases of transfer from the LEO orbit 63, the supplying of propellant to the electrical propulsion module 37 of the transfer vehicle 30 has been done directly from the propellant tank 21 of the external unit 20. After or at the moment of separation of the external unit 20 and the transfer vehicle 30, the valve 45 of the latter is opened, so as to establish a fluidic connection between the propellant reservoir 35 and the electrical propulsion module 37 of the transfer vehicle 30. The latter then becomes entirely autonomous, and begins to use its own reserve of propellant, formed just after the implementing of the phase 74 of docking to the external unit 20.

[0170] In a later phase 79, the transfer vehicle 30 is then taken back to its parking orbit 64. This orbital transfer is preferably performed with a low thrust, so as not to consume much propellant. The transfer vehicle 30 is maintained in this parking orbit 64 until its next mission.

[0171] In variants of the invention, less advantageous than those described above in terms of complexity of the transfer vehicle 30 equipment, after the phase 76 of releasing the satellite 10 in the mission orbit 62 the transfer vehicle 30 is brought directly to the parking orbit 64, still docked to the external unit 20. In its next mission, it docks to a new assemblage formed by a satellite to be stationed attached to an external supply unit carrying its own reserve of propellant. After the positioning of this new satellite in its mission orbit, the transfer vehicle 30 returns to the graveyard orbit 65, to release there the external unit 20, empty of propellant, from the previous mission. Thus, the transfer vehicle is always docked to an external unit, which supplies it with propellant for all its maneuvers. In such a context, the transfer vehicle 30 need no longer have its own propellant reservoir 35.

[0172] As an example, the dimensioning of the system according to the invention and the orbital transfer parameters can be as follows.

[0173] The Delta V for the transfer from the LEO orbit 63 to the GEO orbit 62 is 4549 m/s, and the Delta V for the transfer from the GEO orbit 62 to the graveyard orbit 65 is 11 m/s.

[0174] The dry weights are 3000 kg for the transfer vehicle 30, 2500 kg for the satellite 10 and 100 kg for the external supply unit 20. The transfer vehicle 30 uses 5 thrusters, each of 20 kW, providing a total thrust of 7.2 N in Hall effect technology, with a specific impulse of 2000 s, and 5 thrusters of 20 kW with a total thrust of 3.2 N in ionic grid technology, with a specific impulse of 4500 s. To provide redundancy in the event of a failure, and compensate for engine wear, a total of 10 thrusters for each technology has been taken into account in the weight balance sheet.

[0175] The propellant used is xenon. The useful volume of the propellant tank 21 of the external unit 20 is 821 l. The initial pressure in this propellant tank filled with propellant is 250 bar.

[0176] The useful volume of the propellant reservoir 35 of the transfer vehicle 30 is 174 l. The initial pressure in this reservoir, prior to docking with the external unit 20, is 10 bar.

[0177] After docking and pressure equilibration, the pressure in each of the tank 21 and the reservoir 35 is 208 bar.

[0178] The specific impulse during the phase of transfer of the satellite from the LEO orbit 63 to the GEO orbit 62 is 2000 s, and the specific impulse during the phases of transfer from the GEO orbit 62 to the graveyard orbit 65 and the return to the parking orbit 64 is 4500 s.

[0179] The weight of propellant used to carry out the transfer of the satellite 10 to the GEO orbit 62, and then the transfer of the transfer vehicle 30 and the external unit 20 to the graveyard orbit 65, is 1542 kg.

[0180] The residual weight of propellant in the external unit 20 after its release in the graveyard orbit 65 is 78 kg.

[0181] The weight of propellant used for the return of the transfer vehicle 30 to its parking orbit 64 is 327 kg.

[0182] The residual weight of propellant in the reservoir 35 of the transfer vehicle 30 after its return to the parking orbit is 17 kg.

[0183] By comparison, for a classical launching solution, consisting in launching the satellite into a GTO transfer orbit, for the same dry weight as in this example of implementation of the method according to the invention, the weight launched in the GTO orbit would be 3950 kg with a chemical propulsion system (specific impulse of 320 s) and 2700 kg with an electrical propulsion system (specific impulse of 1750 s).

[0184] The gain provided by the present invention differs in the two instances. In the first case, it substantially translates into a gain in the capacity of the launch vehicle, since the latter needs to provide 150 GJ for the case with transfer according to the invention, compared to 215 GJ for a classical chemical transfer. In the second case, there is a significant gain in the duration of the transfer (around 6 months for the classical launch solution with electrical transfer, compared to around 2 months according to the present invention), and consequently in the operating costs and the financial costs for satellite amortization, as well as the radiation dose to which the satellite is exposed (this reduction represents between 30 and 80% of the absorbed dose, depending on the case).