Reflector satellite and satellite system comprising such a satellite

Abstract

The invention relates to a reflector satellite (1) comprising at least one curved reflecting surface (11), characterized in that it comprises at least one support structure (10) connected to at least four positioning and orientation systems (12) of the reflector satellite for achieving attitude and orbit control, said positioning and orientation systems comprising at least one propulsion system, said positioning and orientation systems being synchronized and arranged so as to allow orientation of the reflector satellite along at least two axes of rotation. A satellite assembly comprising a source satellite and a reflector satellite is also described. Abstract figure: Figure 1

Description

Title of the invention: Reflector satellite and satellite assembly comprising such a satellite technical field

[0001] This disclosure relates to the field of satellites using radio frequency waves, particularly for telecommunications or radiometry applications, and relates more particularly to a reflector satellite and an assembly consisting of a reflector satellite and a source satellite separate from the reflector satellite. Previous technique

[0002] Satellites using radio frequency waves, such as telecommunication satellites, classically include an antenna comprising a radio frequency transceiver and one or more reflectors which are mounted on movable arms allowing the reflectors to be positioned relative to the transceiver, so as, for example, to ensure a telecommunication link.

[0003] Document EP3595088 entitled "Array fed reflector antenna" describes, for example, an antenna in which the position of the focal point is adjustable by means of a telescopic arm allowing the reflector to be moved relative to an array of sources.

[0004] US5021798, entitled "Antenna with positionable reflector", teaches an antenna equipped with a steerable reflector comprising support arms.

[0005] However, satellites, such as telecommunications satellites, must meet increasingly stringent requirements regarding the gain and directivity of radio frequency radiation emitted to or received from Earth to ensure high throughput and reception accuracy. To meet these performance requirements, one area of ​​development is to increase the diameter of satellite reflectors to dimensions of approximately 50 meters, or even 100 meters in diameter, while maintaining a pointing angle accuracy of approximately 0.1° or better.

[0006] Currently, radio frequency wave reflectors in satellites are generally limited to a maximum size of 20 meters in diameter, due to the complexity that this induces, particularly at the mechanical level, for the launch as well as for the deployment of the satellite. Summary

[0007] This disclosure improves the situation.

[0008] In particular, one purpose of this disclosure is to propose a solution for obtaining a reflector with a diameter exceeding 50 meters while exhibiting satisfactory pointing accuracy.

[0009] Another aim of the invention is to provide a simplified mechanical architecture.

[0010] A reflector satellite is proposed comprising at least one curved reflecting surface, characterized in that it comprises at least one carrier structure connected to at least four positioning and orientation systems of the reflector satellite to achieve attitude and orbit control, said positioning and orientation systems comprising at least one propulsion system, said positioning and orientation systems being synchronized and arranged so as to allow orientation of the reflector satellite along at least two axes of rotation.

[0011] In embodiments, each positioning and orientation system is implemented from an autonomous satellite platform, the satellite platforms being mechanically linked together by the supporting structure and piloted in a synchronized manner.

[0012] In some embodiments, the reflection surface is parabolic.

[0013] In some embodiments, the reflecting surface has a larger diameter at 30 meters.

[0014] In some embodiments, the reflecting surface 11) has a diameter greater than 50 meters.

[0015] In embodiments, the supporting structure includes at least one joint and / or at least one telescopic mast for deploying the positioning and orientation systems and extending a canvas forming said reflective surface.

[0016] In embodiments, the load-bearing structure in the deployed position comprises at least four beams distributed angularly around a central point where the beams meet at one of their ends, their other distal end being connected to one of the positioning and orientation systems, the tensioned fabric being connected to the central point and to the distal ends of the beams.

[0017] In embodiments, each positioning and orientation system includes at least one orientable solar panel, the orientation of the solar panels being controlled so as to compensate at least partially for a disturbing solar torque.

[0018] In embodiments, each positioning and orientation system includes at least one reaction wheel, the reaction wheels being controlled so as to compensate at least partially for attitude disturbances due to positioning maneuvers and the spatial environment.

[0019] In some embodiments, the propulsion systems are controlled so as to modify an orbit of the reflector satellite and so as to compensate at least partially for perturbations due to the space environment.

[0020] According to another object, a satellite set is also described, comprising a reflector satellite according to the preceding description and further comprising at least one other satellite called source satellite comprising at least one radio frequency source, the reflector satellite and said source satellite being distant and slaved in position relative to each other so that said reflecting surface collaborates with said radio frequency source to constitute a transmitting and / or receiving system capable of transmitting and / or receiving radio frequency radiation to and / or from the earth.

[0021] In embodiments, said source satellite comprises a source array and a reflector called a sub-reflector, such that the assembly consisting of the reflecting surface, the source array and the sub-reflector constitutes an antenna capable of emitting or receiving radio frequency radiation to or from Earth.

[0022] In embodiments, said source satellite comprises a source array and first and second sub-reflectors, wherein these two sub-reflectors are positioned so that a beam emitted by the source array is reflected by the first sub-reflector to the second sub-reflector, and then reflected by the second sub-reflector to the reflecting surface of the source satellite, so that the assembly consisting of the reflecting surface, the source array, the first sub-reflector and the second sub-reflector constitutes an antenna capable of emitting or receiving radio frequency radiation to or from Earth.

[0023] In some embodiments, said source satellite is master and said reflector satellite is slave, said source satellite controlling each of the positioning and orientation systems of the reflector satellite, said source satellite and the reflector satellite constituting an active antenna establishing an onboard and / or ground control loop for adjusting the positions and orientation of the source satellite relative to the reflector satellite or pointing the active antenna.

[0024] In some embodiments, said radio frequency source generates a beam of parallel electromagnetic waves reflected by the reflecting surface of the reflecting satellite into an expanded beam

[0025] Thus, a telecommunications link can, for example, be established by two separate satellites, one carrying a radio frequency source and the other acting as a reflector with a diameter exceeding 50 meters. These two satellites are not held in place by a mechanical link but by a relative position control system that continuously meets the pointing accuracy requirements. Furthermore, the reflector according to the invention offers increased pointing robustness.

[0026] The structure of each satellite is clearly simplified compared to a single satellite comprising a reflector to be deployed after the launch of the satellite, whether in terms of sizing, hold at launch or deployment kinematics.

[0027] Advantageously, the power requirement of the source satellite is also reduced thanks to the significant gain of the large diameter reflector.

[0028] The absence of a physical link between the so-called source satellite and the reflector satellite also provides greater flexibility for a satellite array applicable to different missions. It is indeed possible to replace the source satellite, for example with a source satellite of more recent design or with different functionalities, by simple means during the reflector satellite's lifetime. Brief description of the drawings

[0029] Other features, details and advantages will become apparent from reading the detailed description below and from analyzing the accompanying drawings, in which: Fig. 1

[0030] [Fig.l] schematically represents a reflecting satellite according to one embodiment. Fig. 2

[0031] [Fig.2] schematically represents a satellite array according to a mode of rea lisation. Fig. 3

[0032] [Fig.3] schematically shows the variation in induced solar radiation pressure on a reflecting system depending on its position in its orbit. Fig. 4

[0033] [Fig.4] schematically shows the correction of a longitudinal drift of the satellite reflector relative to a source satellite by an orientation correction of the reflector satellite. Fig. 5

[0034] [Fig.5] schematically shows a positioning and orientation system for a reflector satellite. Fig. 6a

[0035] [Fig. 6a] schematically shows a reflector satellite in storage configuration in a launcher. Fig. 6b

[0036] [Fig. 6b] schematically shows a reflector satellite during deployment in one embodiment. Fig. 6c

[0037] [Fig. 6c] schematically shows a reflector satellite being deployed in a method of implementation. Fig. 7a

[0038] [Fig.7a] schematically shows the deployment of an example of an arm according to the invention. Fig. 7b

[0039] [Fig.7b] schematically shows an example of web deployment in relation to an arm according to the invention.

[0040] [Fig.7c]

[0041] [Fig.7c] schematically shows another example of web deployment by ratio to one arm according to the invention. Fig. 8a

[0042] [Fig.8a] schematically shows the deployment of a web forming the reflective surface. Fig. 8b

[0043] [Fig.8b] schematically shows the deployment of a web forming the reflection surface on an X-shaped structure. Fig. 9

[0044] [Fig.9] schematically shows an example of double offset. Fig. 10

[0045] [Fig. 10] schematically shows an example of solar sailing. Fig. II

[0046] [Fig. 11] shows an example of a launcher carrying a source satellite and a reflector satellite according to the invention. Fig. 12

[0047] [Fig. 12] shows the lower loading part of the launcher of [Fig. 11], carrying the reflector satellite. Fig. 13

[0048] [Fig. 13] shows the upper loading part of the launcher of [Fig. 11], carrying the source satellite. Fig. 14

[0049] [Fig. 14] shows an example of synchronization by control loop. Description of embodiments

[0050] Reference is now made to [Fig. 1], which schematically represents an example of a reflecting satellite 1 according to the invention. The reflecting satellite 1 comprises a curved reflecting surface 11 mounted on a supporting structure 10, the reflecting surface being adapted to reflect electromagnetic radiation, and in particular radio frequency waves. The supporting structure also comprises a curved support interface. In this respect, the reflecting surface 11 may have, for example, a parabolic shape and be formed of a metal mesh having a mesh size adapted to the application frequency. Alternatively, the surface of Reflection can also be achieved using a flexible metallized fabric. The reflective surface can also be made of carbon or another surface that reflects radio frequencies.

[0051] With reference to [Fig. 2], the reflector satellite 1 is adapted to be positioned relative to another satellite, called the source satellite 2, for the transmission or reception of radio frequency waves. The source satellite comprises, for example, a radio frequency wave source 20. The reflector satellite 1 and the source satellite 2 are separated from each other and are not mechanically linked. In fact, the reflector satellite 1 and the source satellite 2 are servo-controlled in orientation and position relative to each other, so that the reflecting surface 11 of the reflector satellite 1 works with the radio frequency source 20 to constitute a transmitting and / or reflecting system capable of transmitting and / or receiving radio frequency radiation to and / or from Earth.

[0052] An example of double offset emission is illustrated in [Fig. 9]. The source 20 consists of an array of horns generating a beam of radiofrequency waves reflected by a sub-reflector 22 of the source satellite and then by the reflector 11 of the reflecting satellite. The adjustment of the distance between the source and the larger diameter reflector is advantageously not subject to mechanical constraints.

[0053] Thus, for example, it is possible to scan a wide swath of the Earth. An area equivalent to a quarter of the Earth will, for example, be covered with a 1.5m diameter antenna using a 50m diameter reflector.

[0054] A reflector satellite according to the invention is used, for example, for telecommunications applications where the distribution of digital data can reach speeds of several Tbit / s. The reflector can, for example, be used for telecommunications in the S, L, C, X, Ku, and Ka bands. The Ku band is generally defined from 12 GHz to 18 GHz. The Ka band is generally defined from 26 to 40 GHz. The frequencies used are, for example, between 1 GHz and 40 GHz.

[0055] A reflector satellite according to the invention can also be used for radar detection or for scientific applications using microwaves (radiometry, for example), thanks in particular to its high antenna gain. The large diameter and pointing accuracy of the reflector allow for increased performance and greater flexibility in its uses.

[0056] The reflector according to the invention can also advantageously be used for very high throughput applications, also designated as VHTS (Very high throughput applications).

[0057] The fact that the reflector satellite 1 is not mechanically attached to the source satellite implies that the spatial perturbations exerted on the reflector satellite must be taken into account, and that attitude and orbit control of this satellite must be implemented. A reflector satellite is used to compensate for these spatial perturbations. In particular, the reflector satellite is typically subject to perturbations in its orbit due to the varying gravity exerted by the Sun and the Moon, as well as irregularities in Earth's gravity.

[0058] Furthermore, since the reflecting surface of the reflecting satellite is always oriented towards the Earth, the solar radiation pressure exerted on the reflecting surface varies according to the position of the reflecting satellite in its orbit, causing an eccentricity in this orbit that must be compensated for. In [Fig. 3], the different positions of the reflecting satellite 1 in its geostationary orbit around the Earth T are schematically represented, and the relative importance of the solar radiation pressure exerted by the sun S on the reflecting satellite as a function of its position in its orbit and therefore its orientation relative to the sun is schematically represented by arrows of different sizes.

[0059] Figure 10 illustrates examples of solar panel orientations for positioning and orientation systems designed to compensate for solar pressures exerted by solar radiation on the reflecting satellite, and in particular on its solar panels, also referred to as solar generators. For example, one pair of positioning and orientation systems has a first orientation of its solar panels, while another pair has a second orientation distinct from the first. By way of exception, the reflecting satellite may comprise two solar panels per positioning and orientation system. Alternatively, positioning and orientation systems could also be considered, each with one solar panel.

[0060] As shown in [Fig.2] or [Fig.9], the source satellite 2 and the reflector satellite 1 are configured to form a transmitter and / or receiver assembly. This transmitter and / or receiver assembly is, for example, in geostationary orbit or in low Earth orbit (LEO / MEO), with the constraint that the reflector satellite must be at a slightly different altitude from that of the source satellite, for example between 50 meters and 150 meters, in order to ensure the relative positioning of these two satellites and to establish the radio frequency link with the Earth.

[0061] With reference to [Fig.4], this difference in altitude which induces a drift in longitude Tx of the reflecting satellite relative to the source satellite, can be compensated by a correction of the orientation of the reflecting surface (along the axis represented by the arrow Ry on the figure) to maintain the pointing direction.

[0062] In view of these various disturbances, and in order to control the orientation of the reflector satellite 1, it comprises a set of positioning and orientation systems 12 for the reflector satellite 1, connected to the support structure 10, each comprising at least one propulsion system. The positioning and The orientation systems 12 enable attitude and orbit control of the reflector satellite. More specifically, the reflector satellite 1 comprises at least four positioning and orientation systems 12 arranged and synchronized to allow the reflector satellite to be oriented along at least two axes of rotation, in order to implement this attitude and orbit control. In one embodiment, the support structure 10 has a cross shape, which is described in more detail below, and each positioning and orientation system 12 is mounted on one end of the cross. The axes of rotation along which the reflector satellite can be oriented correspond in this case to the two arms of the cross.

[0063] With reference to [Fig.5], an example of a positioning and orientation system 12 has been schematically represented. Such a system may include a chassis 129, mechanically assembled to the supporting structure, to which the various components of the positioning and orientation system are connected.

[0064] Each propulsion system of a positioning and orientation system comprises, for example, one or two thrusters, such as electric thrusters 120. The thrust axes of the thrusters are preferably aligned with the center of mass of the positioning and orientation system. The propulsion systems are used, in particular, in orbit control of the reflector satellite, that is, to compensate for perturbations involving a change in the reflector satellite's orbit.

[0065] Furthermore, each positioning and orientation system may include a set of sensors and actuators involved in attitude and orbit control. For example, each positioning and orientation system may include at least one star tracker 121, enabling the determination of the reflector satellite's attitude relative to one or more specified stars. Each positioning and orientation system may also include an optical and / or radio frequency transmitter and / or receiver 122, adapted to establish a link with a corresponding transmitter / receiver on the source satellite so as to determine the position and relative orientation of the reflector satellite with respect to the source satellite.

[0066] Each positioning and orientation system may further include one or more solar generators 123, providing electrical power to the thruster(s). In some embodiments, each positioning and orientation system includes one or more steerable solar generators 123, the orientation of which can be controlled to at least partially compensate for a solar disturbance torque caused by solar radiation pressure, by implementing conventional solar navigation techniques. For example, each positioning and orientation system 12 may include two solar generators, each mounted on a mast extending perpendicularly to two opposite walls of the frame. 129, the two solar generators being orientable in rotation around said mast.

[0067] Positioning and orientation systems may also include one or more magnetic actuators (not shown), i.e. coils creating a magnetic field on which the Earth's magnetic field acts, creating a torque to control the attitude of the reflecting satellite.

[0068] In addition, each positioning and orientation system may further include at least one reaction wheel 124 controlled so as to compensate at least partially for disturbances due to positioning maneuvers and / or to correct errors in position or relative orientation of the reflector satellite with respect to the source satellite.

[0069] The propulsion system of each positioning and orientation system can also be used to reduce the saturation of the reaction wheels when too much angular momentum is accumulated there.

[0070] In some embodiments, each positioning and orientation system 12 of the reflector satellite can be implemented using an autonomous satellite platform, the satellite platforms being mechanically linked to each other by the support structure and piloted synchronously. This allows the direct use of a functional satellite, for example, a microsatellite or nanosatellite, to implement each positioning and orientation system 12, resulting in design and manufacturing savings. For example, each positioning and orientation system can be obtained from a satellite designed for use in the OneWeb constellation.

[0071] In certain embodiments, and in particular in the case where the positioning and orientation systems are carried out from autonomous satellite platforms, they may also include a payload (not shown), comprising for example one or more observation instruments or additional antennas.

[0072] As indicated above, the positioning and orientation systems 12 are controlled in a synchronized manner to achieve attitude and orbit control of the reflector satellite, so that the reflector satellite is correctly positioned and oriented with respect to the source satellite despite the absence of a physical or mechanical link between the source satellite and the reflector satellite, and with respect to the Earth in order, for example, to ensure radio frequency communication between the source satellite and the Earth.

[0073] In this regard, the control of the positioning and orientation systems can be carried out according to formation flight techniques of said systems, either by being coordinated with each other with respect to a master positioning and orientation system, or by being coordinated with respect to the source satellite.

[0074] To achieve this, each positioning and orientation system 12 includes before at least one controller 125. In embodiments, the source satellite 2 may include a master controller 21 communicating independently with each slave controller 125 of each positioning and orientation system 2 to acquire data acquired by the sensors of each positioning and orientation system and to actuate the propulsion system and / or actuators of each positioning and orientation system to control the attitude and orbit of the reflector satellite.

[0075] Alternatively, one of the positioning and orientation systems may include a master controller 125 in communication with a source satellite controller, and the master controller may control the slave controllers associated with the other positioning and orientation systems.

[0076] Figure 14 illustrates the implementation of a three-step control loop which can be executed consecutively or independently of each other.

[0077] In some embodiments, in a first step, the positioning and orientation systems 12 of the reflector satellite 1 are configured to implement a first control loop for the attitude of the reflector satellite. This control loop is based on attitude measurements taken by the stellar sensors present in each positioning and orientation system. Radio frequency links 30 are established, for example, between the positioning and orientation systems to allow them to operate in a synchronized manner.

[0078] In a second step, the positioning and orientation systems can also be configured to implement a second control loop, based on the position and relative attitude of the reflector satellite with respect to the source satellite, obtained from the optical and / or radio frequency link established between each positioning and orientation system and the source satellite. For example, an optical link 32 is established between the source satellite and optical positioning points 35, also referred to as pinballs, arranged on the reflector. A radio frequency control link 31 is established between the source satellite and one of the positioning and orientation systems.

[0079] In a third step, in some embodiments, the satellite array formed by the source satellite 2 and the reflector satellite 1 can constitute an active antenna establishing an onboard and / or ground control loop, depending on the accuracy of the antenna's pointing, to adjust the position and orientation of the source satellite relative to the reflector satellite. For example, the active antenna can emit a test beam that is detected by one or more ground terminals. The ground terminal(s) can measure a pointing error value of the test beam and transmit this information to the satellite array to correct this pointing error by adjusting the position and / or orientation of the source satellite and / or reflector satellite or a correction of the beam pointing by the active antenna. A dedicated radio frequency radiation 33 is, for example, pointed towards the earth, while a pointing correction 34 is, for example, emitted back towards the source satellite.

[0080] The reflector satellite remains mechanically independent of the source satellite, while precisely controlling its position and orientation to achieve satisfactory pointing accuracy, despite the various disturbances mentioned above. This allows for a considerable simplification of the structure of both the source and reflector satellites, and also eliminates the constraints associated with launching a satellite comprising both a radio frequency source and a deployable reflector.

[0081] In particular, the reflecting surface can be large, and have a diameter greater than 30 meters, or even greater than 50 meters.

[0082] In some embodiments, the load-bearing structure 10 can be a structure fabricated in space. Reference may be made, for example, to French patent FR3070049 entitled "Method for fabricating a large-dimension structure in space," which describes a lattice made from a winding of wire or metal strip and a surface formed by tiles mounted on the lattice.

[0083] Alternatively, and as schematically represented in [Fig.6a], the support structure 10 can be a deployable structure manufactured on the ground and designed to allow its launch into the fairing of a launcher in a storage configuration.

[0084] For example, the supporting structure 10 may have a star shape comprising a set of beams 101 distributed angularly around a central point 102 at which the beams meet at one of their ends; the other end of each beam carrying a positioning and orientation system 12 for the reflecting satellite. In some embodiments, the supporting structure comprises an even number of beams that are aligned in pairs. As mentioned above, in one embodiment, the reflecting satellite comprises four positioning and orientation systems 12, and in this case, the supporting structure has a cross shape comprising four beams 101, each carrying a positioning and orientation system.Alternatively, the supporting structure can also comprise six or eight beams, and in this case the reflecting satellite comprises six or eight positioning and orientation systems respectively, each mounted at the end of a respective beam.

[0085] In the case where the reflector satellite is designed to be deployed from a launcher, the support structure 10 includes, for example, a set of joints 103 and / or telescopic masts for deploying the positioning and orientation systems 12, allowing for a folded configuration sufficiently compact to be placed in the fairing of a launcher, and a deployed configuration in which the positioning and orientation systems are located at the ends of the deployed beams of the support structure. In this case as well, the support structure 10 may further include an extension device 104 for a fabric forming the reflective surface 11, so that the fabric, once deployed, extends from the central point 102 of the support structure to the distal ends of the beams.

[0086] According to a non-limiting example shown in [Fig. 6b], each beam 101 of the supporting structure 10 extending from the central point 102 can be formed of a foldable articulated arm, where at each joint 103 a spring exerts a force tending to deploy the joint. An arm can comprise two or more segments folded against each other. A retaining cable 105, connecting, for example, the central point and two opposing positioning and orientation systems, can be unwound in a controlled manner to regulate the deployment speed, a cable reserve R being stored in the positioning and orientation systems for this purpose. Advantageously, the mechanical stresses applied to the structure to be deployed, and in particular to the arms articulated with each other, are thus reduced.

[0087] According to a non-limiting example shown in [Fig. 6c], the nozzles 36 and 37 of the positioning and orientation systems are actuated to generate two opposing deployment forces 38 and 39. Advantageously, the deployment structure is greatly simplified.

[0088] The arms are for example articulated and / or telescopic arms.

[0089] Alternatively, a retaining cable can be used in conjunction with the nozzles 36 and 37 of the positioning and orientation systems.

[0090] It is also possible to consider the assembly of a reflector in space, as proposed in patent FR3083216 entitled "structure assembled and adjusted in space" which teaches a lattice structure supporting several tiles adjustable in position and able to form the reflector.

[0091] Figure 7a shows an example of articulated and telescopic arms. The arms extend by aligning their portions and by extending their telescopic portions. The extension is controlled, for example, by a set of pulleys 50 and retaining cables 49. A support 45 is connected, for example, by a pivot joint 46 to a telescopic arm 47. The assembly is symmetrical with respect to a central pivot joint 48 connecting two telescopic arms 47.

[0092] Figure 7b shows an example of a canvas held on a folded arm. The segments of the arm 51 are connected to each other, for example, by a deployment axis 50 comprising a spring tending to spread the arm segments apart towards their alignment position. The mesh 52, or meshing in English, is held, for example, by pins 53 fixed to the arm 51 and secured to the mesh 52 by rivets. The pins thus allow certain points of the mesh to be positioned relative to the arm to determine its position relative to the extended arm when the mesh is taut. This allows the mesh, forming the reflective surface, to be given a curved shape.

[0093] Figure 7c shows another example of a web support arm. The segments of the The arms are connected to each other, for example, by a deployment axis including a spring that tends to move the arm segments towards their alignment position. On each face of the arm, a cable 54 is provided, for example, passing through guides 55. This cable 54 can thus be moved relative to the guides 55 to pull a fabric and extend it above the arm. In this example, the arm segments can also be telescopic.

[0094] The extension device 104 of the fabric forming a reflective surface may in particular include a set of pulleys arranged in each positioning and orientation system 12, and cables extending between the central point and each pulley, to which the fabric forming the reflective surface is assembled.

[0095] Each beam 101 of the supporting structure can also be in the form of a telescopic arm to the end of which is attached a deployment cable for the fabric forming the reflective surface, so that the deployment of the telescopic arm also allows the deployment of the fabric.

[0096] The supporting structure is also configured to give the reflective surface, once deployed, a curved shape, for example, a parabolic profile. For example, the beams of the supporting structure may include deployable guides that allow this shape to be given to the reflective surface once it is stretched.

[0097] In some embodiments, whether the support structure is manufactured on the ground, launched and deployed in space, or manufactured in space, the support structure may have a cross or star shape as described above, or another shape. For example, the support structure 10 may have a polygonal shape with as many vertices as there are positioning and orientation systems. In the case where the reflecting satellite includes, for example, four positioning and orientation systems, the support structure may have a square shape.

[0098] As shown in [Fig. 8a], the support structure can be deployed prior to the deployment of the fabric forming the reflective surface. The deployed support structure includes, at the ends of the arms, wheels 58 for controlling cables 59 passing through guide elements 60 relative to the support structure. The cable guides 60 are, for example, attached to the arms. As can be seen in the top view diagram 63a, the unstretched fabric is, for example, initially placed at center of the structure. The canvas 62 is, for example, held in the center by pins 61 attached to the center. The canvas 62 is also fixed to other pins 61 arranged around the periphery of the canvas and opposite one of the arms. The pins can, for example, be driven by cables 59, set in motion by control wheels 58.

[0099] Thus, the web 62 can be moved relative to each arm, with the peripheral pawls 61 each advancing along one arm. For example, two pawls are pulled by cables along two adjacent arms to extend one quarter of the web, as shown in 63b. Then, two more pawls are pulled along the next two adjacent arms to extend another quarter of the web, as shown in 63c. In 63d, three-quarters of the web has been stretched by two more pawls along the next two adjacent arms. Finally, by setting in motion the cables pulling the pawls for the remaining quarter, the web is fully extended, as shown in 63e.

[0100] Alternatively, the canvas can be stretched simultaneously with the deployment of the supporting structure.

[0101] As shown in [Fig. 8b], the arms, as shown in Figures 6b, 6c, 7a, 7b, or 7c, can be integrated into the deployable structure. The arms 75 are connected one after the other to form several long axes. The structure, for example, has an X shape comprising four long axes. The deployable canvas 77 is, for example, stored in the center of the structure.

[0102] Returning to [Fig. 2], the source satellite 2 may comprise a radio frequency source 20 that generates a beam of parallel radio frequency waves. This radio frequency source may be an array of horns generating a multi-wave beam. Alternatively, several sources or a single source generating radio frequency radiation may be considered.

[0103] The source can either be oriented directly towards the reflecting surface 11 of the reflecting satellite 1 to be reflected back towards the area pointed on Earth, or the reflecting satellite can also include one or more reflectors, called sub-reflector(s) 22, which are intercalated between the source and the reflecting surface of the reflecting satellite so that the assembly consisting of the reflecting surface, the radio frequency source and the sub-reflector(s) constitutes an antenna capable of emitting or receiving radio frequency radiation towards or from Earth.

[0104] For example, as shown in [Fig.9], an antenna may have a so-called double offset configuration, comprising the radio frequency source of the source satellite, the reflecting surface of the reflecting satellite and a sub-reflector carried by the source satellite and positioned so as to reflect radiation emitted by the source towards the reflecting satellite, so that the latter reflects it towards the Earth, and vice versa.

[0105] With reference to [Fig.2], the antenna may also have a so-called configuration with Triple eccentricity or triple offset in which the source satellite 2 comprises two reflectors 22, called respectively the first and second sub-reflectors, which are positioned such that a beam emitted by the source 20 is reflected by the first sub-reflector to the second, then reflected by the second sub-reflector to the reflecting surface 11 of the reflecting satellite, which in turn reflects the beam back to Earth. Conversely, radiation or a beam can be emitted from Earth to the source satellite.

[0106] In these cases, the sub-reflector(s) carried by the source satellite can be deployed to adopt their operational configuration after launch. The parallel electromagnetic beam generated by the source, or the array of sources, is thus broadened at each reflection, whether by an intermediate sub-reflector or by the reflecting surface of the reflecting satellite. Therefore, double or triple offset configurations make it possible to significantly extend the ground coverage of the antenna while maintaining a small size for the source.

[0107] Figure 11 shows an example of a launch vehicle carrying the source satellite and the Reflector satellite. The launcher comprises, for example, a first stage 71 surmounted by a second stage 72, itself surmounted by an upper stage 73 and then by a fairing including a space 74 for the payload. The first stage includes the launcher's nozzles 75 in its lower part. The upper stage includes an engine, such as a Vinci engine, operating on liquid oxygen and liquid hydrogen, in its lower part. The first and second stages operate, for example, on solid propellant. The payload includes an upper and a lower position for receiving the elements constituting the payload.

[0108] The lower loading section of the launcher, as shown in [Fig. 12], carries, for example, the reflector satellite in its folded position 77. The reflector satellite in the folded position 77 is, for example, placed under the separation frame of a dual launch system of the Ariane launcher, also known as SYLDA. The reflector satellite will, for example, be injected into geostationary orbit.

[0109] The upper loading section of the launcher, as shown in [Fig. 13], carries, for example, the source satellite 78 in its folded position. The source satellite 78 in its folded position is, for example, placed above the separation frame of the SYLDA. The source satellite will, for example, be injected into an orbit close to geostationary orbit. Thus, the source satellite and the reflector satellite can be launched in a dual launch, in the same way as two independent satellites.

[0110] The reflector satellite and the source satellite can also be injected during two separate launches.

Claims

Demands

1. Reflector satellite (1) comprising at least one curved reflecting surface (11) adapted to reflect electromagnetic radiation, characterized in that it comprises at least one carrier structure (10) of said curved reflecting surface, the carrier structure connecting together at least four positioning and orientation systems (12) of the reflector satellite (1) to achieve attitude and orbit control, said positioning and orientation systems each comprising at least one propulsion system (120), said positioning and orientation systems being synchronized and arranged so as to permit orientation of the reflector satellite along at least two axes of rotation.

2. Reflector satellite (1) according to claim 1, wherein the curved reflecting surface is formed of a grid, a cloth or a material reflecting radio frequency waves.

3. Reflector satellite (1) according to claim 1 or 2, wherein each positioning and orientation system (12) is made from an autonomous satellite platform, the satellite platforms being mechanically linked together by the carrier structure (10) and piloted in a synchronized manner with each other.

4. Reflecting satellite (1) according to any one of the preceding claims, wherein the reflecting surface (11) is parabolic.

5. Reflector satellite (1) according to any one of the preceding claims, wherein the reflecting surface (11) has a diameter greater than 30 meters.

6. Reflector satellite (1) according to any one of the preceding claims, wherein the reflecting surface (11) has a diameter greater than 50 meters.

7. Reflector satellite (1) according to any one of the preceding claims, wherein the support structure (10) comprises at least one joint (10) and / or at least one telescopic mast for deploying the positioning and orientation systems (12) and for extending (104) a fabric forming said reflecting surface (11).

8. A reflector satellite (1) according to claim 6, wherein the support structure (10) in the deployed position comprises at least four beams (101) distributed angularly around a central point (102) where the beams meet at one of their ends, their other end distal being connected to one of the positioning and orientation systems (12), the tensioned web being connected to the central point and the distal ends of the beams.

9. Reflector satellite (1) according to any one of the preceding claims, wherein each positioning and orientation system (12) comprises at least one steerable solar panel (123), the orientation of the solar panels being controlled so as to at least partially compensate for a disturbing solar torque.

10. Reflector satellite (1) according to any one of the preceding claims, wherein each positioning and orientation system (12) comprises at least one reaction wheel (124), the reaction wheels being controlled so as to compensate at least partially for attitude disturbances due to positioning maneuvers and the space environment.

11. Reflector satellite (1) according to any one of claims 1 to 9, wherein the propulsion systems (120) are controlled so as to modify an orbit of the reflector satellite and so as to compensate at least partially for perturbations due to the space environment.

12. Satellite assembly, comprising a reflector satellite (1) according to any one of the preceding claims and further comprising at least one other satellite called a source satellite (2) comprising at least one radio frequency source (20), the reflector satellite and said source satellite being distant and servo-controlled in position relative to each other so that said reflector surface (11) collaborates with said radio frequency source (20) to constitute a transmitting and / or receiving system capable of transmitting and / or receiving radio frequency radiation to and / or from the earth.

13. Satellite assembly according to claim 12, wherein said source satellite (2) comprises a source array and a reflector called sub-reflector (22) such that the assembly consisting of the reflecting surface, the source array and the sub-reflector constitutes an antenna capable of transmitting or receiving radio frequency radiation to or from Earth.

14. Satellite assembly according to claim 12, wherein said source satellite comprises a source array and first and second sub-reflectors (22), wherein these two sub-reflectors are positioned such that a beam emitted by the source array is reflected by the first sub-reflector towards the second sub-reflector, then reflected by the second sub-reflector towards the reflecting surface of the source satellite, so that the assembly consisting of the reflecting surface, the source array, the first sub-reflector and the second sub-reflector constitutes an antenna capable of emitting or receiving radio frequency radiation to or from Earth.

15. Satellite assembly according to any one of claims 12 or 13, wherein said source satellite (2) is master and said reflector satellite (1) is slave, said source satellite driving each of the positioning and orientation systems (12) of the reflector satellite, said source satellite (2) and the reflector satellite (1) constituting an active antenna establishing an onboard and / or ground control loop for adjusting the positions and orientation of the source satellite relative to the reflector satellite or pointing of the active antenna.

16. Satellite assembly according to any one of claims 12 to 14, wherein said radio frequency source (20) generates a beam of parallel electromagnetic waves reflected by the reflecting surface (11) of the reflecting satellite into an expanded beam.