• 专利附录
  • 专利说明
  • 权利要求
  • 类似技术
  • 同族专利
  • 引用文献
  • 法律状态
  • 优先权
    • 专利摘要:
    Satellite (1) comprising: - at least one optical pickup instrument (3) comprising a main objective having an optical axis (V) and the optical instrument (3) having a field of view; at least one launcher interface system (2), intended to be removably attached to a satellite interface system (2 ') of a launcher of the satellite; a device (7) for connection between the launching interface (2) and the optical instrument (3) extending substantially parallel to the optical axis (V) of the main objective between an upper end (9) and a lower end (10); the launcher interface system (2) is connected to the connecting device (7) by the lower end (10) and the optical axis (V) of the optical instrument (3) is directed from the end (9). ) to the lower end (10) of the connecting device (7), the launcher interface system being outside the field of view of the instrument.
    公开号:FR3041939A1
    申请号:FR1559387
    申请日:2015-10-02
    公开日:2017-04-07
    发明作者:Frederic Faye；Eric Beaufume；Jacques Cottier
    申请人:Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    The invention relates to the field of space vehicles, and more specifically to satellites whose mission involves the presence of optical instruments, such as observation or measurement satellites.
    In order to be dropped into space, the satellite is first mounted and secured to a launcher. The launcher is then propelled into space, then the satellite is separated from the launcher to be dropped on the determined orbit.
    Before being dropped, but also during the dropping phase, the satellite in the launcher is subjected to numerous constraints related to shocks and vibrations, and the securing of the satellite in the launcher and the satellite itself must be able to resist. However, the transmission of stresses to the satellite must be carefully controlled, and more particularly in the case where the satellite carries a fragile optical instrument, that shocks and vibrations can disrupt or damage.
    An optical instrument for space missions is typically formed of at least one dioptric, catadioptric or mirror objective, for focusing rays, for example light rays, to obtain an image in a focal plane equipped with detection systems. The line of sight of the optical instrument, that is to say the direction in which the optical instrument looks, may be confused with the optical axis of the lens of the instrument or may form an angle with Optical tax by means of deflecting mirrors. When the optical instrument is a shooting instrument, that is to say comprising at least one sensor making it possible to form an image of a region, for example a region of the terrestrial ground, the optical instrument defines also a field of view corresponding to the truncated cone extending from the functional surface of the sensor, that is to say the surface of the sensor on which the images are formed to the region of shooting. The optical instrument is typically mounted on a support structure, for example a platform itself mounted on the launcher, the axis of view of the instrument being either perpendicular to the platform or parallel to the platform. More precisely, the objective is carried by the platform, its optical axis being perpendicular to the platform, reflecting mirrors for tilting the line of sight. The support structure may further carry other satellite equipment. The integrity of the instrument and the alignment of its components can be altered by shock and vibration during launch and release, potentially resulting in performance degradation.
    Thus, to ensure both good mechanical strength of the satellite in the launcher and the protection of the optical instrument, it is customary to secure the satellite to the launcher by assembling the support structure to the launcher, via an interface ring satellite of the launcher, so that the line of sight points to the opposite of the interface ring, or in a perpendicular direction. The optical instrument is moved away from the interface ring by the support structure, limiting the transmission of shock and vibration from the interface ring to the optical instrument.
    This arrangement is also the consequence of the production and assembly line of the satellite. Indeed, the support structure and the optical instrument are generally manufactured separately at two separate locations, and then assembled. Thus, the instrument is attached to the support structure, and it is then natural to orient the axis of sight opposite or perpendicular to the support structure. The internal volume of the structure is also used for the accommodation of satellite equipment, including the electronics and the or propellant tanks and thus form a platform of servitude.
    The document FR 2 959 490 describes an example of such a satellite. According to this example, the structure of the satellite comprises an equipment carrier and load-bearing walls, in this case four, forming a service box and fixed rigidly to a launching interface ring, this ring being intended to be fixed on the satellite interface ring of a launcher. According to one embodiment, the satellite comprises a payload box fixed at one end to the load-bearing walls and at another end to a plate, this plate supporting an optical instrument, the opening of which is oriented either opposite to the interface ring on one side. Each of the boxes may include various equipment for the operation of the satellite and the optical instrument. The optical instrument is thus moved away from the interface ring by the service box and the payload box, making it possible to limit the transmission of the stresses from the launcher to the optical instrument.
    FIG. 1 schematically illustrates such a satellite 100 according to the state of the art, in an exploded view. The satellite 100 according to the state of the art comprises a launcher interface ring 101 intended to be secured to a satellite interface ring of a launcher, a support plate 102 fixed to the launcher interface ring, an optical instrument occupying a volume 103 represented by a cylinder in broken lines mounted on the support 102, a support structure 104 fixed to the support plate 102 and a structure 105 intended to support the optical instrument and possibly the electronics associated with this instrument . According to this design, the optical instrument 103 comprises an axis of aiming parallel to the axis 107 of the interface ring 101 and oriented opposite the ring 101, so that the opening 108 of the the instrument 103 is directed away from the ring 101. As already shown, in a variant, the aiming axis 106 may be perpendicular to the axis 107 of the ring 101, so that the opening 108 is on one side. These are the only two possible arrangements for a satellite according to this design.
    A disadvantage of this design is that it limits the performance in terms of resolution including the optical instrument.
    Indeed, the performance of the optical instrument is, in general, related to its diameter, that is to say the diameter of the lens: the greater the latter, the better the performance in terms of resolution and radiometric sensitivity. This is particularly the case when the optical instrument is a telescope, and more specifically a Korsch type telescope commonly used in the space field for its compactness, in which the diameter of the mirror or mirrors and the length of the focal length are linked. Thus, if the telescope's performance needs to be increased, it is necessary to increase its diameter, and its length, which implies an increase in the dimensions of the satellite.
    However, in the launcher, the space available for the satellite is limited in width and height by the volume available under the cover. In the case of a double launch, the double launch structure, for example the VESPA structure on the VEGA launcher, comprises a lower compartment in which the dimensions of the passenger are particularly constrained.
    According to the design of the state of the art, the length of the satellite is already partly occupied by the support structure, so that the length of the instrument and therefore the diameter of the instrument are limited by the diameter of the headdress or double launch structure.
    There is therefore a need for a new satellite design overcoming the aforementioned drawbacks.
    An object of the present invention is then to provide a new satellite comprising an optical instrument with improved performance while ensuring the protection of the instrument. For this purpose, according to a first aspect, the invention relates to a satellite comprising: at least one optical pickup instrument comprising a main objective having an optical axis and the optical instrument having a field of view; at least one launcher interface system, adapted to be releasably secured to a satellite launcher satellite interface system; a connecting device between the launching interface and the optical instrument extending substantially parallel to the optical axis of the main objective between an upper end and a lower end.
    The launcher interface system is then connected to the connecting device by the lower end and the optical axis of the optical instrument is directed from the upper end to the lower end of the connecting device, the interface system launcher being outside the field of view of the instrument.
    According to one embodiment, the connecting device comprises a cylindrical side wall with an axis parallel to the line of sight, and comprises an inner surface facing the line of sight, defining an interior space in which extends at least part of the main objective of the optical instrument. The main objective is for example a telescope comprising at least one primary mirror, the optical axis of the primary mirror being the optical axis of the objective, and the primary mirror extending, preferably completely, in the interior space .
    The lateral wall may be cylindrical with a circular or cylindrical guide curve, or a cylindrical directional square or rectangular curve, so that the connecting device comprises four walls.
    According to one embodiment, the satellite further comprises at least one secondary equipment attached to the connecting device. The secondary equipment comprises for example a propellant tank for propulsion or a gas tank for electric propulsion.
    According to one embodiment, the launcher interface system is an interface ring, the line of sight of the optical instrument passing through the interface ring.
    For example, the optical instrument has a diameter greater than 50 cm, and which is for example 100 cm.
    According to one embodiment, the upper end of the connecting device comprises an auxiliary interface system intended to cooperate with another satellite to form a stack. Other advantages and features will emerge in the light of the description of particular embodiments of the invention accompanied by the figures in which:
    Figure 1 is a schematic representation of an exploded view of a satellite according to the state of the art.
    Figure 2 is a schematic representation of an exploded view of a satellite according to the invention.
    Figure 3 is a schematic representation of a sectional view of an example of a satellite according to the invention.
    Figure 4 is a three-dimensional view of an exemplary embodiment of a satellite.
    FIG. 5 is an exploded view of the satellite of FIG. 4.
    Figure 6 is a bottom view of the satellite of Figures 4 and 5.
    Figure 7 is a schematic representation of a sectional view of a stack of two satellites according to the invention.
    FIG. 8 is a schematic representation of a cap of a VEGA launcher for a double launch, comprising two compartments, one of which is a lower VESPA compartment in which the satellite of FIGS. 4 to 6 is placed.
    Figure 1 has already been described in the introduction.
    FIGS. 2 and 3 show schematically a first embodiment of a satellite 1 according to the invention, comprising a launcher interface system 2, designed to be removably attached to a system 2 satellite interface of a satellite launcher shown in broken lines in Figure 3. The interface system 2 is, in common and as will be the case in the following description, an interface ring, The diameter of the interface ring 2 is generally chosen from the standard dimensions in the spatial domain, which are: 937 mm, 1194 mm and 1666 mm. The satellite interface system 2 'is then of annular shape and of complementary dimensions. The two rings 2, 2 'are assembled using a locking mechanism, not shown here, which is for example in the form of a tightening belt, also called strap, which is secured to one of the two rings, and advantageously of the launching interface ring 2 of the satellite 1.
    The satellite 1 comprises an optical instrument 3, shown in FIG. 2 in the form of a cylinder in broken lines representing the space occupied by the optical instrument 3. The objective of the optical instrument 3 is for example a telescope and comprises a primary mirror Mi, also called an input mirror having an optical axis which corresponds here to the V axis of sight of the instrument. The optical instrument 3 is here part of the satellite payload, that is, the main equipment of the satellite 1 mission.
    The primary IVL mirror of the satellite 1 is fixed on a support plate 6. The support plate 6 is in the form of a plate comprising an upper surface 6a and a lower surface 6b, these two surfaces 6a, 6b being substantially perpendicular to the sighting axis V of the instrument 3.
    The adjectives "upper" and "lower" and their variant are used here for the sake of clarity with reference to the natural orientation of the figures, and correspond to the position of the satellite in the launcher when it is in the launch position.
    More specifically, according to the example shown in the figures, the rear of the mirror IVL of the instrument 3 is in contact with the lower surface 6b of the support plate 6.
    The satellite 1 further comprises a device 7 connecting the optical instrument 3 and the starter interface ring 2. According to the embodiment presented here, but in a nonlimiting manner, the connecting device 7 forms the main body 7, that is to say a carrier structure of the satellite 1, on which, as will be seen later, in more than the optical instrument 3, secondary equipment can be attached. More specifically, in the following, the equipment designated as secondary are all equipment other than the optical instrument 3, and include for example the satellite control electronics, but also equipment ensuring the proper functioning of the instrument 3 optical.
    According to the embodiment presented here, the main body 7 has at least one side wall 8 extending substantially parallel to the optical instrument sighting axis V between a first end 9 said upper and a second end 10 said lower .
    Alternatively, the connecting device 7 may be one or more bars or rods connecting the launching interface ring 2 to the optical instrument 3. The satellite 1 can then include an additional structure on which the secondary equipment can be fixed.
    For reasons of simplification, in the embodiment that follows, the connecting device 7 will be called the main body of the satellite 1. The upper end 9 of the side wall 8 is fixed on the support plate 6, and more precisely on the lower surface 6b of the support plate 6. For example, the upper end 9 is in contact on its entire surface with the lower surface 6b of the plate 6. Linear connection means, that is to say, extending continuously on the all of the surface of the upper end 9, or near-point ensure the attachment of the main body 7 on the plate 6. The launching interface ring 2 of the satellite 1 is connected to the main body 7 by the lower end 10 , that is to say the interface ring 2 is disposed relative to the main body 7, following the axis V of sight, the side of the lower end 10 and the connection between the ring 2 d interface and the main body 7 bears on the lower end.
    Thus, for example, the lower end of the side wall 8 bears directly on the interface ring 2, and the side wall 8 is attached to the interface ring 2. In other words, at least one surface portion of the lower end of the side wall 8 is in contact with at least one upper surface portion of the interface ring 2.
    In another example, the lower end of the main body 7 is not in direct abutment on the interface ring 2, but a vibration damping system is placed between the upper face of the ring 2, interface and the lower end.
    Thus, by passing the connection between the main body 7 and the interface ring 2 by the lower end, the V-axis of sight of the optical instrument is substantially parallel to the axis A of the ring. 2 interface. In addition, the rear of the primary mirror M1 being in contact with the plate 6, which is fixed to the upper end 9 of the main body 7, the V-axis of sight is directed towards the interface ring 2.
    In general, according to the invention, the optical axis of the objective here coincides with the optical axis V of the optical instrument 3, is directed from the upper end 9 to the lower end, and the attachment of the objective is remote from the interface ring 2 to protect it from shocks and vibrations, which are at least partly absorbed by the main body 7. When the objective is a telescope with an entrance mirror it is then removed from the interface ring 2, protecting the input mirror IVh.
    In addition, according to the invention, the interface device 2 is outside the field of view of the optical instrument 3. In other words, the interface device 2 does not block some of the rays of the field of view of the optical instrument 3, for optimal resolution of the shooting region. Indeed, according to the example, the interface device 2 in the form of a ring defines a closed contour, with a free space in the middle, through which passes the field of view of the instrument 3.
    In what follows, the term "longitudinal" and its variants designates the direction parallel to the axis A of the interface ring 2 and to the sighting axis V; the adjective "transverse" and its variants designates the directions perpendicular to the longitudinal direction.
    According to an exemplary embodiment, the side wall 8 is cylindrical in shape, of circular or polygonal section, about the V axis of sight. For example, in order to form substantially planar surfaces as will be seen later, the section of the side wall 8 may advantageously be of square section. Therefore, the side wall 8 separates an inner space 11 of the body 7 of the external environment. More specifically, the side wall 8 has an inner surface 12 facing the sighting axis V, and an outer surface 13 facing away from the sighting axis V. The inner space 11 is then delimited by the inner surface 12, and between the two ends 9, 10 of the lateral wall 8, the upper end 9 being closed by the plate 6, the upper end being open to leave the R-rays enter the optical instrument 3 and reach the primary IVL mirror that extends into the inner 11 space. Thus, the lower end of the lateral wall 8 is in abutment, directly or indirectly, on the interface ring 2 so that the optical axis V of the optical instrument 3 passes through the ring 2. interface.
    Thus, only the main body 7 of the satellite is in contact with the interface ring 2, so that the stresses transmitted to the satellite 1 by the launcher necessarily pass through the main body 7, which absorbs at least some of these constraints. , and protects the optical instrument 3.
    Secondary equipment, that is to say other than the optical instrument 3, can be mounted on the main body 7 and the plate 6. In particular, secondary equipment can be mounted on the outer surface 13 of the side wall 8 that is to say, they are in direct contact with the outer surface 13. The cylindrical side wall 8 may be centered, but not necessarily, on the viewing axis V, so that the optical instrument 3 is centered in the interior space 11. The optical instrument 3 can also be decentered in the inner space 11, so as to clear an area for fixing, in direct support, on the inner surface 12 of the secondary equipment, and in particular the electronic equipment related to the operation of the 3 optical instrument. The objective of the optical instrument 3 is, for example, a telescope, of the Korsch type, comprising the primary M-ι mirror and a secondary M2 mirror. The primary IVL mirror has a hole 14 at its center. The two mirrors IVL and M2 are arranged facing each other, so that a ray R entering the instrument 3 along the V-axis of view is first reflected by the primary mirror IVL on the secondary mirror M2 to be reflected again by the secondary mirror M2 to the mirror IVL where it passes through the bore 14. The piercing 14 of the primary mirror IVL coincides with a piercing 15 of the support plate 6 to let the ray R through the plate 6 to a detection system of the optical instrument 3, mounted for example outside the main body 7. The detection system comprises in particular an external mirror M3 and at least one sensor 16, mounted on the outer surface 6a of the plate 6. The external mirror M3 is placed opposite the bore 15 of the plate 6, so as to reflect the beam R towards the functional surface of the sensor 16 mounted on the upper surface 6a of the support plate 6.
    The plate 6 extends transversely beyond the transverse wall 8, that is to say that it has a transverse dimension greater than the transverse dimension of the lateral wall 8, which makes it possible to increase the focal length of the instrument 3 without increasing its length. Indeed, the sensor 16 is placed on a peripheral edge of the upper surface 6a of the plate, so that the larger the transverse dimension of the plate 6, the greater the distance between the sensor 16 and the external mirror M3 may be important . Advantageously, the surface of the sensor 16 oriented opposite its functional surface, that is to say towards space when the satellite is in orbit, may be covered with a radiative material for dissipating the heat generated at the within the satellite. Thus, the distance of the sensor 16 from the external mirror M3, and therefore with respect to the optical instrument 3, also allows better dissipation of heat.
    The two mirrors IVL and M2 of the optical instrument 3 are placed in the space 11 inside the main body 7, so that the side wall 8 forms a protection for the optical instrument 3.
    The lateral wall 8 of the main body 7 then advantageously forms a protective device for the optical instrument 3. For example, as already mentioned, the side wall 8 may have a beam barrier function that is not parallel to the V-axis of sight.
    The arrangement of the mirrors IVL and M2 of the instrument 3 makes it possible to move the mirror Mi away from the interface ring 2, and thus to protect it from the stresses transmitted from the launcher via the launcher interface ring 2.
    Numerous variants of the satellite 1 are possible, for example in the shape of the main body 7, in the type of optical instrument 3, in the dimensions and in the additional functions that the main body 7 can provide.
    With reference to FIGS. 4 to 6, an embodiment of the satellite 1 according to the invention will now be described, in which the optical instrument 3 is a Korsch type telescope as shown above. The same references will be used to denote elements or components identical or similar to those presented with reference to Figures 2 and 3.
    In FIG. 4, the plate 6 is represented in transparency, revealing the primary mirror Mi and sensors 16, the external mirror of the detection system being omitted.
    According to this embodiment, the side wall 8 of the main body 7 is of rectangular or square section, formed of four substantially flat walls 17 disposed substantially at 90 °. Each of the four walls 17 then forms a substantially flat inner surface on the inner surface 12 and a substantially flat outer surface on the outer surface 13 of the lateral wall 8, making it possible to attach a so-called secondary equipment, ie to say other than the 3 otic instrument, participating in the proper functioning of the satellite and the smooth running of its mission.
    The square or rectangular shape of the section of the side wall 8 optimizes the size of the transverse dimensions in the launcher taking into account the secondary equipment mounted on the outer wall 13. However other polygonal shapes can be used, including hexagonal or octagonal.
    Each wall 17 is attached to the interface ring 2 by the lower end. More specifically, two disjointed surface portions of the lower end of each wall 17 are in direct contact with an upper surface of the interface ring 2, forming almost two points of contact or contact areas. The connection between each wall 17 and the interface ring 2 is then provided for example by a quasi-point connection, screw type, at each point or contact area. Alternatively, each wall 17 may have only one point or one contact area with the interface ring 2.
    When the section of the lateral wall 8 is circular, the diameter advantageously corresponds to that of the interface ring 2. Linear type connecting means, such as gluing, stapling or welding, can then be continuously positioned over the entire surface of the lower end and the upper surface of the ring 2. interface, improving the mechanical strength.
    Preferably, no other element of the satellite 1 is in contact with the interface ring 2, so that all the constraints are transmitted from the launcher to the main body 7.
    Thanks to the substantially flat facets formed by the walls 17, the assembly of secondary equipment is easy. In particular, in the embodiment presented here, in order to have the highest possible resolution, the telescope occupies most, if not all, of the interior space 11 of the main body 7, that is to say that the primary mirror Mi has a maximum diameter. The secondary equipment is then preferably fixed on the outer surface 13 of the walls 17, that is to say that they are in direct contact with the outer surface 13 of the walls 17. The flatness of the walls 17 is particularly suitable for mounting electronic equipment, but not exclusively.
    Thus, on the outer surface 13 of the main body 7 is mounted a propulsion system 18. The use of electric propulsion is advantageous because the volume of take-away propellants is much lower than that of a conventional chemical propulsion. The gas reservoir, generally xenon can thus be placed easily outside on the outer surface 13 of the walls 17, maintaining a size in the acceptable transverse directions in the launcher, which allows to leave the space 11 inside available for the optical instrument 3, and the primary mirror M1 may extend especially its diameter in the inner space 11. In the context of a short-term mission, conventional chemical propulsion can however be used, the necessary propellant volume being low and the reservoirs can be accommodated on the outer surface of the main body. Other secondary equipment 19 may also be attached to the outer surface 13 of the walls 17, such as batteries, control boxes, or sensors.
    The satellite 1 may further comprise retractable solar panels 21 fixed to the outer surface 13 of the main body 7 by means of pivoting arms 22.
    Actuator means, such as CMG 23 (acronym for Control Momentum Gyroscope) can also be mounted on the outer surface 13 of the walls 17.
    Thus, the support plate 6 and the main body 7 together support all the equipment of the satellite 1. This results in a great modularity, the secondary equipment can be arranged on the outer surface 13 of the walls 17 independently of the optical instrument 3 . The path of the stresses transmitted by the interface ring 2 necessarily passes through the lateral wall 8 of the main body 7, protecting the optical instrument 3. The length of the optical instrument 3, that is to say its dimension along its viewing axis V, can then be increased, for a total length of the satellite 1 lower compared to the state of the art. When the optical instrument 3 is a telescope as described above, increasing the length of the optical instrument makes it possible in particular to increase the distance between the primary mirror Mi and the secondary mirror M2, and therefore to increase the diameter of the mirrors Mi and M2 while respecting the dimensional requirements of optical principles. By increasing the diameter of the mirrors Mi and M2, until the primary mirror Mi fills the inner space 11, the resolution of the telescope is increased. By way of comparison, whereas the diameter of a primary mirror of an optical instrument of a satellite according to the state of the art can reach a diameter of the order of 40 to 50 cm (centimeters) in a VESPA coping volume for VEGA, the primary mirror Mi of the satellite according to the present invention can go beyond, until reaching the double, or reach a diameter up to 1 m, in a configuration where the launching interface is of 1194 mm and still in the same volume of VESPA cap for VEGA.
    In other words, thanks in particular to the design of the satellite 1 in which the stresses pass through the lateral wall 8 of the main body 7, and by orienting the optical instrument 3 so that its axis V of aiming is directed towards the lower end in connection with the interface ring 2, the total length of the optical instrument 3, and therefore the satellite, can be reduced to maintain performance at least equivalent to that of the state of the art. However, as presented in the introduction, the length of the satellite 1 is the most critical dimension in terms of space in the launcher.
    The satellite 1 according to the invention is then particularly adapted to take the place in the compartment of the smallest size in the case of a dual launch system, usually the lower compartment as for example in a VESPA structure for a launcher VEGA. FIG. 7 thus schematically illustrates the volume of the cap 24 of a VEGA launcher comprising a VESPA structure. Two independent compartments are formed: a lower compartment and an upper compartment 26, the adjectives "lower" and "upper" being used here with reference to the natural orientation of FIG. 7, which corresponds to the orientation of a launcher placed on the ground for the assembly of satellites. Each compartment 25, 26 is intended to receive a satellite attached to a satellite interface ring.
    The lower compartment is of limited size, especially in the direction of the length of the satellite placed therein. Thus, the satellite 1 according to the invention, whose length is reduced while maintaining the expected performance, is particularly suitable for being placed in the lower compartment.
    The satellite 1 may be particularly adapted to be stacked with another satellite of the same design or different design. For this purpose, according to another embodiment, the connecting device 7 is in the form of a central cylinder of longitudinal main axis, for example merged with the axis A of the interface ring 2. The satellite 1 can always comprise walls 17, which are fixed on the central cylinder 7. The support plate 6 is fixed to the central cylinder 7. For example, the optical instrument 3 and the support plate 6 are housed inside the central cylinder 7. The detection system can be mounted outside the central cylinder 7, on the surface of the upper end 9. The central cylinder 7 can emerge beyond the walls 17 on either side in the longitudinal direction, so that the launching interface ring 2 can be fixed by the lower end to the central cylinder 7, and the the upper end 9 of the central cylinder 7 is available to mount an auxiliary interface system 27 for cooperating with a complementary interface system of another satellite. As for the launcher interface ring 2, the auxiliary interface system 27 may be an interface ring, and will thus be designated in what follows. The auxiliary interface ring 27 has a bottom surface attached to the central cylinder 7.
    In order to facilitate the stacking of two satellites 1 of design according to the invention, the auxiliary interface ring 27 of a first satellite 1 is intended to cooperate with the launching interface ring 2 of the second satellite 1.
    Thus, the two satellites 1 according to the invention can be superimposed in the following manner.
    A first satellite 1 is attached to a satellite interface ring of a launcher 28 by its launcher interface ring 2. The second satellite 1 is placed on the first satellite 1 so that their axes V of sight are coincident, or at least parallel. The upper end 9 of the first satellite 1 is vis-à-vis the lower end of the second satellite 1, and the auxiliary interface ring 27 of the first satellite is matched with the ring 2 of launcher interface of the second satellite 1. The two rings 2, 27 provide the link between the two satellites 1.
    Optionally, the detection system mounted on the surface of the upper end 9 of the central cylinder 7 extends beyond the walls 17 in the longitudinal direction. In this case, the second satellite 1 of the stack comprises a space for the detection system of the first satellite 1 to be housed there when the two satellites 1 are stacked.
    In such a stack, all the stresses transmitted by the satellite interface ring 2 'of the launcher to the first satellite 1 pass through the lateral wall 8 of the main body 7 of the first satellite 1 and are transmitted to the lateral wall 8 of the main body 7 of the second satellite 1, again protecting the optical instrument 3 of the second satellite 1. Thus, the mechanical path through which the stresses pass is located in the side wall 8 of the main body 7 of the two satellites 1.
    It is the same when more than two satellites 1 according to the invention are stacked as well.
    The design satellite 1 according to the invention therefore makes it possible to present a compact structure while guaranteeing a performance, particularly in terms of resolution, of the optical instrument 3 at least equivalent to the state of the art.
    The compactness of the satellite 1 allows it, in addition to reducing the congestion performance equal to the state of the art, reduce its inertia and thus facilitate control of the attitude of the satellite to reduce energy consumption .
    In addition, the length of the satellite being decreased, the surface exposed to the speed vector is decreased, thereby decreasing the drag, and thus again facilitating the control of the attitude to reduce energy consumption.
    When the satellite 1 is mounted in the launcher, the launcher interface ring 2 secured to the satellite interface ring 2 ', the sighting axis V is oriented downwards in the direction of gravity. Thus, since the environment in the launcher is not free from particles such as dust, by orienting the optical instrument with its axis of aim downward, the mirror IV is protected from particulate contamination which would degrade the particles. performance of the optical instrument 3.
    权利要求:
    Claims (13)
    [1" id="c-fr-0001]
    A satellite (1) comprising: at least one optical pickup instrument (3) comprising a main lens having an optical axis (V) and the optical instrument (3) having a field of view; at least one launcher interface system (2), to be releasably secured to a satellite launcher satellite interface system (2 '); a device (7) for connection between the launching interface (2) and the optical instrument (3) extending substantially parallel to the optical axis (V) of the main objective between an upper end (9) and a lower end (10); the satellite (1) being characterized in that the launcher interface system (2) is connected to the connecting device (7) by the lower end (10) and in that the optical axis (V) of the optical instrument (3) is directed from the upper end (9) to the lower end (10) of the connecting device (7), the launcher interface system being outside the field of view of the instrument .
    [2" id="c-fr-0002]
    2. Satellite (1) according to claim 1, wherein the connecting device (7) comprises a cylindrical side wall (8) with an axis parallel to the sighting axis (V), and comprises a surface (12). ) facing towards the axis (V) of sight, defining an interior space (11) in which extends at least a portion of the main objective of the optical instrument (3).
    [3" id="c-fr-0003]
    The satellite (1) according to claim 2, wherein the main objective is a telescope comprising at least one primary mirror (IVh), the optical axis of the primary mirror (IVh) being the optical axis of the objective, the primary mirror (IVh) extending into the inner space (11).
    [4" id="c-fr-0004]
    4. Satellite (1) according to claim 2 or claim 3, wherein the side wall (8) is cylindrical circular guide curve.
    [5" id="c-fr-0005]
    The satellite (1) according to claim 2 or claim 3, wherein the side wall (8) is cylindrical with a polygonal guide curve.
    [6" id="c-fr-0006]
    6. Satellite (1) according to claim 5, wherein the wall (8) side is cylindrical square or rectangular guide curve, so that the device (7) connecting comprises four walls (17).
    [7" id="c-fr-0007]
    7. Satellite (1) according to any one of the preceding claims, further comprising at least one equipment (18, 19, 20, 21, 22) secondary attached to the device (7) connecting.
    [8" id="c-fr-0008]
    8. Satellite (1) according to claim 7, wherein the equipment (18, 19, 20, 21, 22) secondary comprises a propellant tank for propulsion.
    [9" id="c-fr-0009]
    9. Satellite (1) according to claim 7, wherein the equipment (18, 19, 20, 21, 22) secondary comprises a tank (18) of gas for electric propulsion.
    [10" id="c-fr-0010]
    10. Satellite (1) according to any one of the preceding claims, wherein the system (2) launcher interface is an interface ring, the axis of view (V) of the instrument (3) through optical the interface ring.
    [11" id="c-fr-0011]
    11. Satellite (1) according to any one of the preceding claims wherein the optical instrument (3) has a diameter greater than 50 cm.
    [12" id="c-fr-0012]
    The satellite (1) of claim 11, wherein the diameter of the optical instrument (3) is 100 cm.
    [13" id="c-fr-0013]
    The satellite (1) according to any one of the preceding claims, wherein the upper end (9) of the connecting device (7) comprises an auxiliary interface system (24) for cooperating with another satellite to form a stack.
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    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    FR2959490A1|2010-04-28|2011-11-04|Astrium Sas|SATELLITE HAVING A SIMPLIFIED, ALLEGE AND ECONOMIC STRUCTURE AND ITS IMPLEMENTING METHOD|
    WO2012116366A2|2011-02-25|2012-08-30|Utah State University Research Foundation|Multiple petal deployable telescope|
    EP2662287A1|2012-05-11|2013-11-13|The Boeing Company|Multiple space vehicle launch system|
    CN104648693A|2014-12-23|2015-05-27|中国科学院长春光学精密机械与物理研究所|Satellite structure for platform and load integration|
    US5169094A|1990-02-26|1992-12-08|Aerospatiale Societe Nationale Industrielle|Geostationary earth observation satellite incorporating liquid propellant apogee maneuver system and hollow antennas|
    FR2667299B1|1990-10-02|1994-01-14|Aerospatiale Ste Nationale Indle|GEOSTATIONARY-TYPE SATELLITE WITH HANDLING SYSTEM HAVING LIQUID ERGOLS AND HOLLOW ANTENNAS.|
    US5337980A|1992-09-21|1994-08-16|General Electric Co.|Spacecraft-to-launch-vehicle transition|
    JPH07294279A|1994-04-21|1995-11-10|Nec Corp|Image pickup position detector|
    JP2003167196A|2001-12-04|2003-06-13|Nikon Corp|Cata-dioptric system|
    FR2869292B1|2004-04-23|2006-07-07|Cnes Epic|SATELLITE, METHOD AND FLEET OF SATELLITES FOR OBSERVING A CELESTED BODY|
    JP3922719B2|2004-06-03|2007-05-30|独立行政法人宇宙航空研究開発機構|Spin imparting method, spin imparting mechanism, spacecraft, and artificial satellite separating spring assembly|
    US7631839B1|2004-08-20|2009-12-15|Lockheed Martin Corporation|Enhanced multiple instrument distributed aperture sensor|
    US7240879B1|2005-05-06|2007-07-10|United States of America as represented by the Administration of the National Aeronautics and Space Administration|Method and associated apparatus for capturing, servicing and de-orbiting earth satellites using robotics|
    US7780119B2|2006-12-11|2010-08-24|Michael David Johnson|Modular spacecraft|
    US7686255B2|2007-08-28|2010-03-30|Raytheon Company|Space vehicle having a payload-centric configuration|
    JP2009190661A|2008-02-18|2009-08-27|Mitsubishi Electric Corp|Protective cover for optical apparatus|
    KR101130119B1|2010-04-30|2012-03-28|한국과학기술원|Optical Structure for Aerospace Engineering|
    FR2959460B1|2010-05-03|2012-12-28|Entpr Weber|TRAILERS IN PARTICULAR FOR TRANSPORTING WOOD|
    EP3361561A1|2010-12-15|2018-08-15|Planet Labs Inc.|Integrated antenna system for imaging microsatellites|
    EP2489593A1|2011-02-21|2012-08-22|European Space Agency|Earth observation satellite, satellite system, and launching system for launching satellites|
    FR2974348B1|2011-04-21|2014-01-24|Thales Sa|DEVICE FOR PROTECTING AN OPTICAL INSTRUMENT OF A SATELLITE|
    FR2974787B1|2011-05-05|2014-11-28|Thales Sa|DEVICE FOR PROTECTING A MULTIFACEAL OPTICAL INSTRUMENT|
    DE102012100287B4|2012-01-13|2018-12-13|Grammer Ag|vehicle seat|
    US20140027577A1|2012-07-26|2014-01-30|Alliant Techsystems Inc.|Imaging satellite system and method|
    US9027889B2|2013-02-28|2015-05-12|The Boeing Comapny|Modular core structure for dual-manifest spacecraft launch|
    CN104309824A|2014-09-12|2015-01-28|上海卫星工程研究所|Satellite configuration of embedded compound eye camera load|
    FR3041940B1|2015-10-02|2018-07-13|Airbus Defence And Space Sas|CYLINDRICALLY MAIN BODY SATELLITE, STACK COMPRISING SUCH SATELLITE AND LAUNCH SET FOR SUCH A SATELLITE|
    US10670853B2|2016-01-22|2020-06-02|Hera Systems, Inc.|Imaging system with an optical path and telescope shape optimized for nanosatellites|
    FR3047813B1|2016-02-16|2019-11-08|Airbus Defence And Space Sas|METHOD FOR CONTROLLING SATELLITE ATTITUDE GUIDANCE, SATELLITE, SATELLITE PLURALITIES AND ASSOCIATED COMPUTER PROGRAM|FR3041940B1|2015-10-02|2018-07-13|Airbus Defence And Space Sas|CYLINDRICALLY MAIN BODY SATELLITE, STACK COMPRISING SUCH SATELLITE AND LAUNCH SET FOR SUCH A SATELLITE|
    IL257491A|2018-02-12|2021-02-28|Israel Aerospace Ind Ltd|Deployable space vehicle|
    CN109188649B|2018-09-19|2021-07-02|珠海达理宇航科技有限公司|Protective device for polygonal barrel and space telescope lens|
    法律状态:
    2016-10-25| PLFP| Fee payment|Year of fee payment: 2 |
    2017-04-07| PLSC| Publication of the preliminary search report|Effective date: 20170407 |
    2017-10-27| PLFP| Fee payment|Year of fee payment: 3 |
    2018-10-26| PLFP| Fee payment|Year of fee payment: 4 |
    2019-10-25| PLFP| Fee payment|Year of fee payment: 5 |
    2020-10-26| PLFP| Fee payment|Year of fee payment: 6 |
    2021-09-27| PLFP| Fee payment|Year of fee payment: 7 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1559387A|FR3041939B1|2015-10-02|2015-10-02|SATELLITE COMPRISING OPTICAL OPTICAL INSTRUMENT|FR1559387A| FR3041939B1|2015-10-02|2015-10-02|SATELLITE COMPRISING OPTICAL OPTICAL INSTRUMENT|
    KR1020187012096A| KR101900226B1|2015-10-02|2016-09-29|Satellite containing optical photographic equipment|
    CN201910183771.1A| CN109795718A|2015-10-02|2016-09-29|Satellite including optical photography instrument|
    EP16785247.4A| EP3174794B2|2015-10-02|2016-09-29|Satellite comprising an optical photography instrument|
    CA3038877A| CA3038877C|2015-10-02|2016-09-29|Satellite comprising an optical photography instrument|
    ES16785247T| ES2656841T5|2015-10-02|2016-09-29|Satellite comprising an optical imaging instrument|
    PCT/FR2016/052476| WO2017055750A1|2015-10-02|2016-09-29|Satellite comprising an optical photography instrument|
    JP2018517138A| JP6401422B1|2015-10-02|2016-09-29|Satellite with imaging optics|
    US15/764,667| US10442557B2|2015-10-02|2016-09-29|Satellite comprising an optical photography instrument|
    CN201680057432.3A| CN108137170B|2015-10-02|2016-09-29|Satellite including optical photography instrument|
    PL16785247T| PL3174794T5|2015-10-02|2016-09-29|Satellite comprising an optical photography instrument|
    IL25846218A| IL258462A|2015-10-02|2018-03-29|Satellite comprising an optical photography instrument|
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