Patent Publication Number: US-2023150694-A1

Title: Method for producing a satellite from a generic configuration of antenna elements

Description:
FIELD OF THE INVENTION 
     The present invention belongs to the field of communication or observation satellites carrying parabolic reflector antennas on a face of the satellite. The invention relates in particular to a manufacturing method allowing to facilitate the integration of the antennas on the satellite during the design of the satellite. 
     PRIOR ART 
     A communication or observation satellite is generally placed in orbit around the Earth to carry out a specific mission associated with one or more geographic zones on the surface of the Earth. To carry out its mission, the satellite includes one or more antennas that generally each include a radiofrequency source for emission or for reception, a reflector with specific geometric parameters (diameter, focal distance, profile, etc.), a support structure for the reflector, devices for holding and for releasing the reflector and a device for deploying the reflector. 
     The conventional methods for manufacturing a satellite to carry out a specific mission generally include a relatively tedious step of determining the parameters of the various antennas of the satellite to cover the various geographic zones of the mission of the satellite. This step involves for example determining for each antenna, according to the geographic zone that must be covered by the antenna, the diameter of the reflector of the antenna, its focal distance, the position of the radiofrequency source(s) of the antenna, the dimensions of the support structure of the reflector, the position of the devices for holding and for releasing the reflector, the position of the deployment device, the shape of the profile of the reflector, etc. 
     Technical constraints relative to the structure of the satellite can affect the determination of these parameters. Moreover, modifications made to a specific parameter affect the other parameters, and it is generally necessary to proceed by trial and error with a large number of iterations to obtain a set of parameters suitable and compatible with each other to satisfy the needs of the mission. 
     The number of antennas on board a satellite tends to increase. Thus, the phase of integrating the antennas onto the satellite is particularly long and costly. This integration phase is often part of the critical path of the manufacturing of the satellite. Indeed, for each new satellite to be manufactured, it is necessary to redefine all of the parameters of each antenna according to the mission of the satellite. 
     The document Modern Communication Satellite Antenna Technology (A. Ghulam) describes in detail the various parameters of the antenna system of a satellite that should be defined during the manufacturing of the satellite in order for the latter to be able to carry out a specific mission. This document also describes various software tools for assisting in the determination of these parameters. 
     When the mission of a satellite is associated with several different geographic zones on the surface of the Earth, the deployment device can be controlled (this is called “steering”) to move the reflector in order to aim it at various geographic zones of interest. The patent application US 2002/135532 A1 discloses such a deployment device allowing to adjust the position of a reflector. In such a case, the parameters of the antenna must be defined in such a way that the antenna allow to satisfy the mission for each geographic zone of interest. A compromise must thus be found since the antenna is not therefore optimised for a single specific position of the reflector but for several positions corresponding to the various geographic zones of interest. 
     DISCLOSURE OF THE INVENTION 
     The goal of the present invention is to overcome all or a part of the disadvantages of the prior art, in particular those disclosed above, by proposing a solution that facilitates the phase of integration of the antennas during the manufacturing of a satellite. 
     For this purpose, and according to a first aspect, the present invention proposes a method for manufacturing an artificial satellite intended to be placed in orbit around the Earth to carry out a mission associated with one or more geographic zones on the surface of the Earth. The satellite comprises one or more reflectors on a face of the satellite. Each reflector is connected to said face of the satellite via a deployment device allowing to make the reflector go from a retracted position to a deployed position. When there are several reflectors, the reflectors are stacked above one another facing the face of the satellite when the reflectors are in the retracted position. Each reflector is associated with a radiofrequency source for emission or for reception to form a radio beam carrying a radio signal to or from a geographic zone on the surface of the Earth. 
     The manufacturing method comprises first of all a preliminary step of determining a generic configuration, independently of the mission of the satellite (that is to say independently of the geographic zone(s) associated with the mission of the satellite), by setting at least the following parameters for each reflector:
     diameter of the reflector,   focal distance of the reflector,   offset distance of the reflector when the reflector is in the deployed position,   position of the deployment device on the face of the satellite,   position of the source associated with said reflector on the face of the satellite,   
 in such a way that a beam coming from the source is aimed at the centre of the Earth when the satellite is in orbit and the reflector is in the deployed position.
     The manufacturing method then comprises a specific configuration step, according to the mission of the satellite (that is to say according to the geographic zone(s) associated with the mission of the satellite). The specific configuration step comprises the following substeps for each reflector:
     determining an adjusted deployed position of the reflector obtained by controlling the deployment device, so that a beam coming from the source is aimed at a geographic zone of the mission of the satellite when the satellite is in orbit and the reflector is in the adjusted deployed position,   shaping the surface of the reflector according to said geographic zone and according to the adjusted deployed position thus determined.   

     The specific configuration step is implemented without modifying the parameters set during the preliminary step of determining the generic configuration. 
     In other words, the method for manufacturing a satellite according to the invention involves reusing a generic configuration that was defined independently of a specific mission, and defining a limited number of specific parameters according to the specific mission of the satellite. The generic configuration is determined in such a way that each antenna is aimed by default at the centre of the Earth. The specific configuration is then made possible by controlling (steering) the deployment device, that is to say by a modification of the aiming of the antenna once the reflector is deployed. The generic configuration can be reused for each new satellite to be manufactured. Only the specific configuration must be redefined for a new satellite to be manufactured. 
     For the generic configuration, the “deployed position” of a reflector corresponds to a “default deployed position” defined in such a way that the reflector is aimed at the centre of the Earth. For the specific configuration, the “adjusted deployed position” corresponds to a deployed position different than the default deployed position. When the reflector is in the adjusted deployed position, the reflector is aimed at a specific geographic zone on the surface of the Earth to carry out a specific mission. 
     In specific embodiments, the invention can further include one or more of the following features, taken alone or according to all the technically possible combinations. 
     In specific embodiments, each reflector is associated with holding and releasing devices comprising a portion detachable on command and allowing to hold the reflector in the retracted position during a phase of launch of the satellite and release the reflector into the deployed position once the satellite is in orbit. The preliminary step of determining a generic configuration also comprises setting, for each reflector, the position of the holding and releasing devices on the face of the satellite. 
     In specific embodiments, the face includes at least two reflectors. 
     In specific embodiments, the reflectors share the same holding and releasing devices. The method according to the invention is indeed particularly well adapted in such a case. 
     In specific embodiments, the reflectors all have a diameter between 220 and 270 centimetres. 
     In specific embodiments, the specific configuration is optimised by iterating several times the substeps of determining an adjusted deployed position and shaping the surface of the reflector. 
     According to a second aspect, the present invention proposes a face for manufacturing a satellite according to a manufacturing method according to any one of the above embodiments. The face includes one or more reflectors, each reflector being connected to said face of the satellite via a deployment device allowing to make the reflector go from a retracted position to a deployed position. The reflectors are stacked above one another facing the face of the satellite when the reflectors are in the retracted position. For each reflector, the face also includes a radiofrequency source for emission or for reception for forming a beam carrying a radio signal to or from the Earth. Each reflector, each source and each deployment device are arranged so that a beam coming from a source is aimed at the centre of the Earth when the satellite is in orbit and the reflector is in a default deployed position. The deployment device is further configured to move each reflector into an adjusted deployed position in which the reflector is aimed at a specific geographic zone corresponding to a mission of the satellite. 
     Such a face meets the requirements of a generic configuration of the manufacturing method according to the invention. To finalise the phase of integration of the antennas during the manufacturing of a satellite intended to carry out a specific mission associated with one or more geographic zone(s) on the surface of the Earth, all that remains to do is therefore to determine for each reflector an adjusted deployed position, and shape the surface of the reflector, in such a way that a beam coming from the source associated with said reflector covers a specific geographic zone of the mission of the satellite. 
     In specific embodiments, the invention can further include one or more of the following features, taken alone or according to all the technically possible combinations. 
     In specific embodiments, the face includes between one and three reflectors, each reflector satisfying one of the following sets of parameters (each set of parameters include a diameter of the reflector, a focal distance of the reflector, and an offset distance measured between the vertex of a paraboloid in which the reflector is inscribed and the tip of the reflector closest to said vertex):
     diameter between 220 and 270 centimetres, focal distance between 380 and 420 centimetres, offset distance between 65 and 105 centimetres,   diameter between 220 and 270 centimetres, focal distance between 358 and 398 centimetres, offset distance between 90 and 130 centimetres,   diameter between 220 and 270 centimetres, focal distance between 375 and 415 centimetres, offset distance between 90 and 130 centimetres.   

     In specific embodiments, the face includes at least two reflectors. 
     In specific embodiments, the face includes several reflectors. The face further includes holding and releasing devices comprising a portion detachable on command and allowing to hold the reflectors in the retracted position during a phase of launch of the satellite, and release the reflectors into the deployed position once the satellite is in orbit. Advantageously, the various reflectors share the same holding and releasing devices. 
     In specific embodiments, the reflectors are concentric when the reflectors are in the retracted position. This means that the reflectors are “substantially” concentric, that is to say for example that the distance projected onto the face of the satellite between the centres of two reflectors is less than 10% of the smallest of the diameters of the reflectors, or even less than 5% of the smallest of the diameters of the reflectors. 
     In specific embodiments, the sources and/or the deployment devices are aligned according to a transverse axis of the face. This means that these elements are “substantially” aligned, that is to say for example that the projected distance between two elements according to an axis orthogonal to said transverse axis is less than 20 cm, or even less than 10 cm. 
     According to a third aspect, the present invention relates to a satellite including a face according to any one of the above embodiments. 
    
    
     
       PRESENTATION OF THE DRAWINGS 
       The invention will be better understood upon reading the following description, given as an example that is in no way limiting, and made in reference to  FIGS.  1  to  10    which show: 
         FIG.  1    a representation of a face of a satellite having several reflectors stacked in the retracted position, 
         FIG.  2    a diagram of a satellite with a reflector in the retracted position on a face of the satellite, 
         FIG.  3    a diagram of a side view of the satellite illustrated in  FIG.  2    when the reflector is in the retracted position, 
         FIG.  4    a diagram of the satellite illustrated in  FIG.  3    when the reflector is in the deployed position, 
         FIG.  5    a diagram of the main geometric parameters for a reflector, 
         FIG.  6    a diagram of a satellite with two reflectors stacked in the retracted position on a face of the satellite, 
         FIG.  7    a diagram of the mains steps of a method for manufacturing a satellite according to the invention, 
         FIG.  8    a diagram of a satellite in orbit around the Earth with a reflector in the deployed position according to a generic configuration, 
         FIG.  9    a diagram of a satellite in orbit around the Earth with a reflector in the deployed position according to a specific configuration, 
         FIG.  10    a diagram illustrating how the devices for holding and for releasing the reflectors of a satellite can interfere with a beam coming from a radiofrequency source of the satellite, 
         FIG.  11    a diagram of a face of a satellite with three reflectors in the deployed position when different holding and releasing devices are used for the various reflectors, 
         FIG.  12    a diagram of three reflectors stacked in the retracted position sharing the same holding and releasing devices, 
         FIG.  13    a representation of a holding and releasing device intended to be associated with three different reflectors, 
         FIG.  14   a diagram of a face of a satellite with three reflectors in the deployed position when the same holding and releasing devices are used for the various reflectors. 
     
    
    
     In these drawings, references identical from one drawing to another designate identical or analogous elements. For reasons of clarity, the elements shown are not necessarily on the same scale, unless otherwise mentioned. 
     DETAILED DESCRIPTION OF AN EMBODIMENT OF THE INVENTION 
     In the rest of the description, the case is considered of the manufacturing of a telecommunications satellite intended to be placed in geostationary orbit around the Earth. Geostationary orbit means a circular orbit around the Earth characterised by an orbital inclination of zero (the orbit is thus located in the equatorial plane) and an orbital period equal to the period of rotation of the Earth. 
     The satellite is placed in orbit to carry out a specific mission associated with one or more geographic zone on the surface of the Earth. The satellite includes for this purpose one or more antennas with a parabolic reflector. Each reflector is intended to cover a geographic zone of the mission of the satellite to ensure exchanges of messages by radio communication between a station located in said geographic zone on the surface of the Earth and the satellite. 
       FIG.  1    shows an example of a face  11  of such a satellite  10  including three reflectors  20 . Each reflector is fastened to a support structure  40  connected to the face  11  of the satellite via a deployment device  50 . In the example considered and illustrated in  FIG.  1   , the support structure  40  includes several rectilinear metal elements assembled together and arranged in the shape of a rectangle, as well as an additional rectilinear metal element acting as an arm  41  for linking with the deployment device  50 . Nothing, however, prevents, in alternatives, the elements forming the support structure  40  from being curved, from being assembled in a shape other than a rectangle, and/or from being manufactured from a non-metal material (for example from carbon fibres). 
     The deployment device  50  (ADTM for Antenna Deployment and Trimming Mechanism) allows to make the reflector  20  go from a retracted position to a deployed position. The retracted position is the position taken up by a reflector  20  during the phase of launch of the satellite (that is to say during the take-off and the flight of the launcher of the satellite, and during the release of the satellite). In the retracted position, each reflector  20  is positioned facing the face  11  of the satellite. When there are several reflectors on the same face  11  of the satellite, as is the case in the example illustrated in  FIG.  1   , the reflectors are stacked above one another facing the face  11  of the satellite when they are in the retracted position. When the reflectors  20  are in the retracted position, the bulk of the satellite  10  is reduced, which is advantageous to place the satellite in a nose cone of a launcher. Moreover, when the reflectors  20  are in the retracted position, they better resist the mechanical stresses caused by the vibrations during the launch phase. 
     Holding and releasing devices  60  (HRM for “Hold and Release Mechanism”) allow to hold the reflectors  20  in the retracted position during the phase of launch of the satellite then to release the reflectors  20  in order for them to be deployed once the satellite  10  is in orbit. 
     In the deployed position, a reflector is moved away from the face  11  of the satellite via the deployment device  50  and the linking arm  41  in such a way that the reflector is positioned facing the Earth. The separation of the reflector from the face  11  of the satellite can be defined by an offset distance of the reflector when the reflector is in the deployed position. This offset distance will be defined later in a more precise manner in reference to  FIG.  5   . 
     Each reflector  20  is associated with at least one radiofrequency source  30  for emission or for reception. A radiofrequency source allows to form a beam carrying a radio signal to or from the geographic zone on the surface of the Earth covered by the reflector  20 . Radio signal means an electromagnetic wave, the frequency of which is lower than 300 GHz. A radiofrequency source  30  of the satellite  10  is for example adapted to emit a radio signal on a downstream link of the Ku layer (range of microwave frequencies ranging from 10.70 to 12.75 GHz) and/or to receive a radio signal on an upstream link of the Ku layer (range of microwave frequencies ranging from 12.75 to 18 GHz). According to other examples, a radiofrequency source  30  can also be adapted to emit on another frequency band for example such as one of the bands L, S, C, X, Ka, Q or V. 
     In the example considered, each reflector  20  is associated with a single radiofrequency source  30  for example in the form of a corrugated horn (the inside of the horn has annular machinings to improve the radiation diagram of the antenna). Each source forms a radio beam. It should be noted, however, that the invention could also be applied to an antenna of the SFPB type (acronym for Single Feed Per Beam) for which the reflector is associated with several sources, each source being associated with a radio beam, or to an antenna of the MFPB type (acronym for Multiple Feeds Per Beam) for which the reflector is associated with a network of several sources allowing to form different beams, each beam being formed by several sources of the network. In the case of an SFPB or MFPB antenna, the single radiofrequency source  30  shown in  FIG.  1    would be replaced by a set of several radiofrequency sources. 
       FIG.  2    is a simplified diagram of a face  11  of a satellite  10  with a reflector  20  in the retracted position. The source  30  associated with the reflector  20 , the support structure  40 , the linking arm  41 , the deployment device  50  and the devices  60  for holding and releasing the reflector  20  are also shown in  FIG.  2   .  FIG.  2    also highlights linking points  42 , for example in the form of threaded inserts, allowing to fasten the reflector  20  to the support structure  40 , as well as a point  43  for linking the arm  41  with the rectangular part of the support structure  40 . 
     The deployment device  50  allows to make the reflector  20  go from the retracted position to the deployed position once the satellite  10  is in orbit and the devices  60  for holding and releasing the reflector have released the reflector. For this purpose, the deployment device  50  allows to apply a movement of rotation to the assembly formed by the support structure  40  (including the linking arm  41 ) and the reflector  20  about an axis according to the direction y passing through the deployment device  50  (pitch) and about an axis according to the direction x passing through the deployment device  50  (roll). The pitch allows in particular to move the reflector  20  away from the face  11  of the satellite in order for the reflector  20  to be positioned facing the Earth. The pitch and roll movements also allow to orient the reflector  20  to aim at a specific geographic zone on the surface of the Earth. In the example considered, when the satellite is in orbit around the Earth, the direction z is aimed at the centre of the Earth, the direction x is aimed east and the direction y is aimed south. The face  11  shown in  FIG.  2    thus corresponds to the western face of the satellite. 
       FIG.  3    is a diagram of a side view of the satellite  10  illustrated in  FIG.  2    when the reflector  20  is in the retracted position. In the example illustrated in  FIG.  3   , each device  60  for holding and releasing the reflector  20  includes a base  61  and a portion  62  detachable on command. The base  61  of the device  60  for holding and releasing the reflector  20  is fastened to the face  11  of the satellite  10 . The detachable portion  62  connects the support structure  40  of the reflector  20  to the base  61  of the holding and releasing device  60  via a pyrotechnic element  63 . When the pyrotechnic element  63  is actuated on command, the detachable portion  62  is disconnected from the base  61  of the holding and releasing device  60 . To deploy the reflector  20 , the pyrotechnic elements  63  of the various devices  60  for holding and releasing the reflector  20  are actuated at the same time to release the reflector  20 . Other mechanisms are possible to detach the detachable portion  62 . For example, non-explosive mechanisms could be used to control the separation of the detachable portion  62  from the base  61  of the holding and releasing device  60 . The choice of a specific type of devices  60  for holding and releasing the reflector  20  merely corresponds to an alternative of the invention. 
       FIG.  4    is a diagram of a side view of the satellite  10  illustrated in  FIG.  2    when the reflector  20  is in the deployed position. To be thus positioned, the deployment device  50  applied to the arm  41  a movement of rotation about an axis according to the direction y and passing through the deployment device  50 . The arm  41  thus drove the support structure  40  and the reflector  20  into a deployed position, separated from the face  11  of the satellite, in order for the reflector  20  to be positioned facing the Earth. The bases  61  of the devices  60  for holding and releasing the reflector  20 , which remain fastened onto the face  11  of the satellite  10  after deployment of the reflector  20 , are also shown in  FIG.  4   . 
     For reasons of simplification,  FIGS.  2  to  4    only show a single reflector  20  associated with the face  11  of the satellite. It should be noted, however, that the face  11  of the satellite can include several reflectors  20 . The reflectors  20  are thus stacked above each other facing the face  11  of the satellite when they are in the retracted position, and they are deployed next to each other to face the Earth once the satellite is in orbit. It should also be noted that the face of the satellite opposite to the face  11  described in reference to  FIGS.  2  to  4    can also include one or more reflectors. Two opposite faces of the satellite can thus be totally or partly symmetrical. The two faces can for example be totally symmetrical if they include the same number of reflectors  20 , the same reflectors, and the same elements positioned identically for each reflector (source  30 , deployment device  50 , holding and releasing devices  60 ). The two faces can be partly symmetrical for example if one face includes less reflectors than the other or if different radiofrequency sources  30  are used to operate in different frequency bands (the positions of the sources  30  can nevertheless remain symmetrical). 
       FIG.  5    is a diagram of the main geometric parameters for a reflector  20 . As illustrated in  FIG.  5   , the reflective surface of the reflector  20  is inscribed in a paraboloid  21  of revolution that has as a vertex a point noted as V (vertex of the paraboloid). The focus of the paraboloid is noted as S. The source  30  associated with the reflector  20  is placed at this focal point S. The focal distance F of the reflector  20  thus corresponds to the distance between the vertex V of the paraboloid  21  and the focal point S of the paraboloid  21 . The axis passing through the points S and V corresponds to the focal axis of the paraboloid  21 . The diameter of the reflector  20  is noted as D. As illustrated in  FIG.  5   , in the present application “diameter of the reflector 20” means the diameter of the projection of the reflector in a plane orthogonal to the focal axis and passing through the vertex V. The separation of the reflector  20  from the face  11  of the satellite can be defined by an offset distance of the reflector when the reflector  20  is in the deployed position. According to a first example, the offset distance can be defined as the distance Δ1 between the vertex V of the paraboloid  21  and the point on the reflective surface of the reflector  20  that is closest to the vertex V (this distance is generally called “offset to the edge” or “clearance”). According to a second example, the offset distance can be defined as the distance Δ2 between the vertex V of the paraboloid  21  and the centre of the reflective surface of the reflector  20  (this distance is generally called “offset”). According to a third example, the offset distance can be defined as the distance Δ3 between the face  11  of the satellite and the point on the reflective surface of the reflector  20  that is closest to said face  11 . It should be noted that from the moment that the position of the source  30  with respect to the face  11  is set, the distances Δ1 , Δ2 and Δ3 all three allow to define the same deployed position of the reflector  20 , that is to say with the same separation of the reflector  20  from the face  11  of the satellite. The manner of defining the offset distance of the reflector when the reflector is in the deployed position is not very important. Ways other than those illustrated in  FIG.  5    are also possible for defining an offset distance of the reflector when the reflector is in the deployed position (for example the distance between the face  11  and the centre of the reflective surface of the reflector  20 ). 
     In order for the reflector  20  to be aimed at the centre of the Earth in the deployed position, the direction taken by a beam coming from the source  30  and reflected by the reflector  20  should be parallel to the axis z. This direction (aiming direction) is also parallel to the focal axis of the paraboloid  21  (that is to say the axis passing through the source S and the vertex V). 
     The position of a reflector  20  with respect to the face  11  when the reflector  20  is in the deployed position is thus completely determined when the position of the source  30 , the focal distance F and the offset distance are set. After also having set the diameter of the reflector  20  and the position of the deployment device  50 , it becomes possible to define the dimensions of the support structure  40  and of the linking arm  41 . 
     It should be noted than in the example illustrated in  FIG.  5   , the plane containing the beam coming from the source  30 , the beam reflected by the reflector  20  and the focal axis of the paraboloid  21  is orthogonal to the face  11  of the satellite. This plane is not, however, necessarily orthogonal to the face  11  of the satellite. 
       FIG.  6    is a diagram of a face  11  of a satellite  10  with two reflectors  20 - 1 ,  20 - 2  stacked in the retracted position. The elements associated with the first reflector  20 - 1  are shown with solid lines: support structure  40 - 1 , linking arm  41 - 1 , holding and releasing devices  60 - 1 , deployment device  50 - 1 , radiofrequency source  30 - 1 . The elements associated with the second reflector  20 - 2  are shown with dotted lines: support structure  40 - 2 , linking arm  41 - 2 , holding and releasing devices  60 - 2 , deployment device  50 - 2 , radiofrequency source  30 - 2 . In the example illustrated in  FIG.  6   , the first reflector  20 - 1  is stacked above the second reflector  20 - 2  when the reflectors  20 - 1 ,  20 - 2  are in the retracted position. In the example illustrated in  FIG.  6   , the first reflector  20 - 1  has a greater diameter than the second reflector  20 - 2 . Nothing prevents, however, all the reflectors from having the same diameter. 
       FIG.  7    schematically illustrates the main steps of the method  100  for manufacturing a satellite  10  according to the invention. 
     The manufacturing method  100  comprises first of all a preliminary step of determining  110  a generic configuration, independently of the mission of the satellite, that is to say independently of the geographic zone(s) associated with the mission of the satellite, by setting at least the following parameters for each reflector  20  associated with a face  11  of the satellite  10 :
     diameter of the reflector  20 ,   focal distance of the reflector  20 ,   offset distance of the reflector  20  when the reflector  20  is in the deployed position,   position of the deployment device  50  on the face  11  of the satellite  10 ,   position of the source  30  associated with said reflector  20  on the face  11  of the satellite  10 .   
 The various aforementioned parameters are determined in such a way that a beam coming from the source  30  is aimed at the centre of the Earth when the reflector  20  is in the deployed position.
     The manufacturing method then comprises a specific configuration step  120 , according to the mission of the satellite, that is to say according to the geographic zone(s) associated with the mission of the satellite. The specific configuration step  120  comprises the following substeps for each reflector  20 :
     determining  121  an adjusted deployed position of the reflector  20  obtained by controlling the deployment device  50 , so that a beam coming from the source  30  is aimed at a specific geographic zone of the mission of the satellite,   shaping  123  the surface of the reflector  20  according to the geographic zone to be covered and according to the adjusted deployed position thus determined.   

     It is important to note that the specific configuration step is implemented without modifying the parameters set during the preliminary step of determining  110  the generic configuration. 
     In other words, the method  100  for manufacturing a satellite  10  according to the invention involves reusing a generic configuration that was defined independently of a specific mission, and defining a limited number of specific parameters according to the specific mission of the satellite. The generic configuration is determined in such a way that each antenna is aimed by default at the centre of the Earth. The specific configuration is then made possible by controlling the deployment device, that is to say by a modification of the aiming of the antenna once the reflector is deployed. The generic configuration can be reused for each new satellite to be manufactured, and for each face of the satellite that must include reflector antennas. Only the specific configuration must be redefined for a new satellite to be manufactured and/or for each face of the satellite. 
     The manufacturing method  100  according to the invention is made possible by the use of a deployment device  50  that has a steering margin in terms of pitch and roll sufficient to be able to orient each reflector  20 , starting from the default deployed position, to be aimed at any geographic zone on the surface of the Earth that is visible by the satellite  10  once the satellite is in orbit. 
       FIG.  8    schematically illustrates a satellite  10  in orbit around the Earth  70  with a reflector  20  in the deployed position according to a generic configuration. As illustrated in  FIG.  8   , the generic configuration is determined so that a beam  31  coming from the radiofrequency source  30  associated with the reflector  20  is aimed at the centre  71  of the Earth  70  when the reflector  20  is in the default deployed position. As illustrated in  FIG.  8   , the beam  31  is reflected by the reflector  20  to be directed towards the Earth  70 . 
     In the generic configuration, the default deployed position of a reflector  20  is therefore defined by the pitch angle and the roll angle that the deployment device  50  must apply to place the reflector  20  in a position such that a beam  31  coming from the radiofrequency source  30  associated with the reflector  20  is aimed at the centre  71  of the Earth  70 . 
       FIG.  9    highlights the step of controlling  122  the deployment device  50  of the reflector  20  to place the reflector  20  in an adjusted deployed position. During this step, an additional movement of pitch and/or roll is applied by the deployment device  50  to the linking arm  41  to place the reflector  20  in a position for which a beam  31  coming from the radiofrequency source  30  and reflected by the reflector  20  covers a specific geographic zone  72  belonging to the mission of the satellite. 
     The adjusted deployed position of a reflector  20  determined during the specific configuration step  120  is thus defined by the pitch angle and the roll angle that the deployment device  50  must apply to place the reflector  20  in a position such that a beam  31  coming from the radiofrequency source  30  associated with the reflector  20  is aimed at a specific geographic zone  72  after reflection on the reflector  20 . 
     The substep of shaping  123  the reflector  20  during the specific configuration step  120  allows to optimise the contour of the zone covered by the beam  31  to best cover the geographic zone  72 . As illustrated in  FIG.  7   , the substeps of determining  121  an adjusted deployed position and of shaping  123  the surface of the reflector can be iterated several times during the phase of specific configuration to optimise the performance of the antenna. 
     It should be noted that the substep of shaping  123  is not absolutely indispensable to the invention (in particular if the invention is applied to antennas of the SFPB or MFPB type). However, this substep of shaping  123  is particularly advantageous since it allows to compensate for the fact that the antenna was initially designed to be aimed by default at the centre  71  of the Earth  72 , and not at a specific geographic zone  72 . 
     The number of parameters to be defined for the specific configuration is very limited compared to the conventional methods for which the set of all the parameters relative to all the elements present on the face of the satellite must be defined. Reusing the generic configuration thus allows to facilitate the phase of integration of the antennas during the manufacturing of the satellite. 
     The diameter and the focal distance of each reflector  20  are set by the generic configuration. It is advantageous to provide a generic configuration with reflectors having a relatively large diameter. There will thus be more flexibility in the choice of the missions that could be carried out by a satellite manufactured from this generic configuration. A reflector with a large diameter can indeed be used to cover geographic zones for which a reflector with a smaller diameter would have sufficed. In preferred embodiments, the reflectors  20  have a diameter between 220 and 270 centimetres. 
     Setting the position of the radiofrequency sources  30  on the face  11  in the generic configuration allows to standardise the position of the various interfaces with these sources  30  in the satellite (waveguides, devices for fastening the sources, etc.). 
     Setting the position of the deployment devices  50  on the face  11  in the generic configuration is also advantageous since the study of the technical constraints relative to the position of these devices on the satellite will only need to be done once (and no longer every time a new satellite is manufactured). 
     In specific embodiments, the positions of the various devices  60  for holding and releasing the reflectors  20  are also set in the generic configuration. 
     The position of the devices  60  for holding and releasing the reflectors  20  can have an impact on the performance of the radiofrequency sources  30 . As illustrated in  FIG.  10   , once the reflectors  20  are deployed, the bases  61  of the holding and releasing devices  60  fastened onto the face  11  of the satellite  10  can be located at least partly inside a radio beam  31  coming from a radiofrequency source  30  and cause interference. This is why it is generally necessary, during the design of the satellite, to carry out a study of the effect of these holding and releasing devices  60  on a radio signal emitted by a source  30 , and take the necessary measures to compensate for this effect in the radio transmission or reception chain. Here again, by setting the position of the various elements in the generic configuration, this study will only need to be done once, and no longer every time a new satellite is designed. 
     It is advantageous to reduce the number of the holding and releasing devices  60  to limit their effect on a radio signal emitted by a radiofrequency source  30 . In the example illustrated in  FIG.  10   , there are four holding and releasing devices  60  for each reflector  20 . Since there are three reflectors in the example considered, this makes a total of twelve holding and releasing devices  60  on the face  11  of the satellite. 
       FIG.  11    schematically shows a face  11  of a satellite  10  with three reflectors  20 - 1 ,  20 - 2 ,  20 - 3  in the deployed position when different holding and releasing devices are used for the various reflectors. As illustrated in  FIG.  11   , after deployment there remains on the face  11  the bases  61 - 1  of four holding and releasing devices associated with the first reflector  20 - 1 , the bases  61 - 2  of four other holding and releasing devices associated with the second reflector  20 - 2 , and the bases  61 - 3  of four more other holding and releasing devices associated with the third reflector  20 - 3 . 
     In specific embodiments, and as illustrated in  FIG.  12   , the reflectors  20  share the same holding and releasing devices  60 .  FIG.  12    shows three reflectors  20 - 1 ,  20 - 2 ,  20 - 3  stacked in the retracted position on a face  11  of the satellite  10 . The three reflectors share the same holding and releasing devices  60 . Thus, only four holding and releasing devices  60  are necessary for the three reflectors  20 - 1 ,  20 - 2 ,  20 - 3 . For this purpose, each holding and releasing device  60  includes for example three detachable parts  62 - 1 ,  62 - 2 ,  62 - 3  and three pyrotechnic elements  63 - 1 ,  63 - 2 ,  63 - 3 . To deploy the reflectors, the first pyrotechnic element  63 - 1  of each holding and releasing device  60  is activated, which allows to detach the first detachable part  62 - 1  of each holding and releasing device  60  to release the first reflector  20 - 1 . Then, the second pyrotechnic element  63 - 2  of each holding and releasing device  60  is activated, which allows to detach the second detachable part  62 - 2  of each holding and releasing device  60  to release the second reflector  20 - 2 . Finally, the third pyrotechnic element  63 - 3  of each holding and releasing device  60  is activated, which allows to detach the third detachable part  62 - 3  of each holding and releasing device  60  to release the third reflector  20 - 3 . With such arrangements, only four holding and releasing devices  60  are necessary for all three reflectors  20 - 1 ,  20 - 2 ,  20 - 3 . 
       FIG.  13    shows an example of a holding and releasing device  60  intended to be associated with three different reflectors. As illustrated in  FIG.  13   , each detachable part  62 - 1 ,  62 - 2 ,  62 - 3  respectively includes a fastening element  64 - 1 ,  64 - 2 ,  64 - 3  to fasten the detachable part to the associated reflector. In the example illustrated in  FIG.  13   , a fastening element  64 - 1 ,  64 - 2 ,  64 - 3  includes a metal surface with holes to receive for example bolts or rivets. The base  61  of the holding and releasing device  60  is intended to be fastened onto the face  11  of the satellite. 
       FIG.  14    schematically shows the satellite face  11  described in reference to  FIG.  12    when the three reflectors  20 - 1 ,  20 - 2 ,  20 - 3  are in the deployed position. After deployment, there remains on the face  11  only the bases  61  of four holding and releasing devices associated with the three reflectors  20 - 1 ,  20 - 2  and  20 - 3 . 
     It should be noted that in the example illustrated in  FIGS.  12  and  14   , the various reflectors  20 - 1 ,  20 - 2 ,  20 - 3  have substantially the same diameter. Nothing prevents, however, reflectors having different diameters from being able to still share the same holding and releasing devices. In such a case, the support structures associated with the various reflectors have for example different dimensions to be able to be attached to the fastening elements  64 - 1 ,  64 - 2 ,  64 - 3  of the holding and releasing devices  60  shared among the various reflectors. 
     The method  100  according to the invention is particularly well adapted to obtain a configuration in which the reflectors share the same holding and releasing devices, and/or in which the reflectors are concentric when they are in the retracted position (as illustrated in  FIG.  6   ), and/or in which the sources are aligned (as illustrated in  FIGS.  1 ,  6 ,  11  and  14   ), and/or in which the deployment devices are aligned. 
     The above description clearly illustrates that, via its various features and their advantages, the present invention achieves the goals set by proposing a method that facilitates the phase of integration of the antennas during the design of a satellite. 
     In general, it should be noted that the embodiments and implementations considered above have been described as non-limiting examples, and that other alternatives are consequently possible. 
     In particular, the choice of the number of reflectors  20  arranged on a face  11  of the satellite, the choice of a specific shape for the support structure  40  of a reflector  20 , the choice of the number and/or of the type of devices  60  for holding and releasing a reflector  20 , the choice of a specific type of radiofrequency source  30  or of deployment devices  50  are merely alternatives of the invention. 
     The invention has been described while considering the manufacturing of a telecommunication satellite intended to be placed in geostationary orbit around the Earth. Nothing excludes, however, according to other examples, applying the present invention to the manufacturing of another type of satellite, optionally intended to be placed in non-geostationary orbit, or even in orbit around a celestial body other than the Earth.