• 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:
    The invention relates to a photovoltaic generator (12a, 12b) deployable for satellite (10) stabilized type three axes, said photovoltaic generator (12a, 12b) having a set of planar panels (20, 21) articulated with respect to each other , and an attachment arm (18a, 18b) to the structure of a satellite body (11), said photovoltaic generator (12a, 12b) being able to adopt a first position, called the launch position, in which the planar panels (20, 21) are folded over each other, and a second position, said extended position, in which the planar panels (20, 21) are fully deployed, at least a portion of the planar panels (20) being photovoltaic panels. At least one plane panel (21) consists of a thermal radiator, its radiative face facing away from the face of the photovoltaic panels carrying the photovoltaic sensors, when the photovoltaic generator (12a, 12b) is in the deployed position, this radiative face being said face "shadow", and the opposite face of the radiator panel (21) being said face "sun".
    公开号:FR3014849A1
    申请号:FR1362606
    申请日:2013-12-13
    公开日:2015-06-19
    发明作者:Jerome Lemaire；Yann Miegeville
    申请人:Centre National dEtudes Spatiales CNES；Thales SA；Astrium SAS；
    IPC主号:
    专利说明:

    [0001] The present invention relates to the field of satellite architectures stabilized on 3 axes, including telecommunications satellites intended to be placed in geostationary orbit.
    [0002] Preamble and Prior Art To cope with the ever-increasing volume of telecommunications in the world, telecommunications satellites are becoming increasingly powerful. Their electronic payload (traveling wave tubes in particular) thus releases a thermal power of increasing importance, about 60% of the electric power of the satellite, that is to say commonly of the order of several kW, and the cooling of this payload becomes more and more difficult to achieve. In the usual way, this cooling is achieved by using the so-called north and south faces of the satellite, that is to say the faces substantially perpendicular to the terrestrial equator when the satellite is stationed, as radiators, by placing therein fluid loops. or heat pipes that evacuate the heat generated by the dissipative equipment of the satellite. The dimensions of these faces then become a limit to the cooling capacity of the satellite.
    [0003] Various techniques have been considered to solve this power / cooling problem for high power satellites, but their weight or complexity are limiting factors of their effective use on satellites. The present invention therefore aims to remedy all or part of these disadvantages. SUMMARY OF THE INVENTION The invention aims, for this purpose, in a first aspect, a deployable photovoltaic generator for satellite type stabilized three axes, said photovoltaic generator comprising a set of panels articulated with respect to each other, and an arm attachment to the structure of a satellite body, allowing a free rotation of the photovoltaic generator vis-à-vis the body of the satellite to track the path of the sun during the orbit of the satellite, said photovoltaic generator can adopt a first position, said launch position, wherein the panels are folded over each other, a second position, said fully deployed, in which the panels are fully deployed, at least a portion of the panels being photovoltaic panels having a sensor face photovoltaic cells intended to be oriented towards solar radiation, the photovoltaic generator being such that at least one panel is constituted by a heat radiator, its radiative face being oriented opposite to the face of the photovoltaic panels carrying the photovoltaic sensors, when the photovoltaic generator is in the deployed position, this radiative face being said "shadow" face, and the opposite face of the radiator panel being said face "sun". In this way, the radiator remains oriented in the shade as long as the photovoltaic cells are exposed to the sun, that is to say the vast majority of the time. The radiator is then fully effective. In addition, the deployment of this radiator is done with that of the other panels of the deployable generator, which represents a simplification of implementation. Finally, the fixing devices of the radiator to the satellite structure during the launch phase are then the same as those of the deployable generator, which results in an economy of means. A radiator of maximized rejection capacity is thus obtained, while reducing as much as possible the impact of its implementation on the architecture of the satellite. Deployable radiators implemented on satellites are actually already known, these devices making it possible to increase the radiative surface of the north and south faces. This is most often panels articulated around an axis of attachment to the body of the satellite, or a deployment arm Such deployable radiators constitute additional mechanical complexity to be taken into account in the design of the satellite, they require a dedicated deployment structure and a clean fastening system, which increases the overall complexity of the satellite, and the deployable radiators on other satellites are fixed orientation and are therefore regularly These solutions also require the necessary space in the near-satellite environment, usually between the solar generator and the satellite body, which poses problems for both the solar power and the solar power. storage of the photovoltaic panels and the deployable radiator itself before the launch, and, once in orbit, to manage the occupied volume p ar deployable radiators within a space also used by solar generators or antennas or the propulsion system. It should be noted that the panels optionally also include a position called semi-deployed position, in which the panel or panels of the end of the generator are deployed. According to a particular embodiment, the panel closest to the support arm is a heat radiator panel. This arrangement makes it possible to reduce the length of the cooling fluid circulation tubes between the satellite and the radiator. It also makes it possible to move the photovoltaic sensors away from the ejectors of the attitude control nozzles of the satellite, which reduces the degradation of the sensors by the jets of said nozzles. According to a particular architecture, the deployable generator comprises a matrix assembly of rectangular photovoltaic panels arranged in lines and a radiator panel aligned with the central column of photovoltaic panels. This implementation allows a simple deployment of the deployable generator. This matrix assembly may consist of a single line, consisting for example of 4 to 5 panels, or in several lines, for example 3 lines, each line being constituted for example by 3 to 4 panels According to a particular embodiment, the generator comprises an insulating zone between the radiator panel and the solar panels, so as to reduce heat transfer from the radiator panel to the solar panels. According to a particular embodiment offering a good efficiency of the radiator, at least one radiator panel is traversed by tubes of a fluid loop or a heat pipe, its shadow face being covered with a radiative device, and its sun face being covered with a multilayer insulation. The invention aims, in a second aspect, a satellite of stabilized type on three axes, said satellite comprising at least one deployable array of photovoltaic panels as exposed, said photovoltaic panels being stored in folded position at launch, and then deployed once the satellite is set, the radiator panel being arranged closer to the body of the satellite. In a particular embodiment, the satellite comprises a rotary joint disposed between the body of the satellite and the set of panels, arranged at the support arm of the deployable photovoltaic generator, said rotary joint having its axis of rotation coincides with the axis of rotation of the arm, and comprising means for passing at least one fluid and means for transferring electric current. This arrangement allows the passage of cooling fluids between the satellite and the radiator, as well as the transfer of electrical energy, while ensuring a free rotation of the arm carrying the photovoltaic generator, so as to allow tracking of the position of the Sun during of the satellite's orbit. In a first embodiment, the rotary joint comprises two concentric devices. In this case, adapted to the case of cooling by fluid loop, the flow of fluids is in the central part of the rotary joint. This gives a device allowing a permanent rotational movement of the deployable generator, following the Sun. In a particular embodiment of simple implementation, the means of passage of the electric current consist of a disc-shaped conductive track, traversed by a broom also conductive. Alternatively, the passage of the heat transfer fluid is through a hose winding adapted to the transport of fluid between the body of the satellite and the radiator panel, this hose winding being secured, at one of its ends, to a cooling circuit. heat dissipation from the satellite body, and at the other end to the radiator panel. The hose winding comprises for example two fluid passage tubes, if the satellite is equipped with a fluid loop cooler.
    [0004] This implementation is simpler than a rotary joint adapted to the passage of fluids. Here, the hose winding, for example in helical form, makes it possible to maintain the circulation of the cooling fluids during the rotational movement of the deployable generator, up to a predetermined angular value, for example +180 or + 360 °. Then, a rotational movement in the opposite direction of the deployable generator is made, up to -180 ° or -360 °, to bring the hose winding in the reverse position. In this case, to simplify the drive mechanism of the solar generator, in a particular embodiment (using a hose and no rotating seal for the transport of the fluid) the passage of the current is achieved by at least one flexible strand and not by tracks on a rotating disc. The invention aims, in another aspect, on a method for controlling a photovoltaic satellite generator as described, the method comprising a step in which, during the rotation of the photovoltaic generator around its axis relative to the body of the satellite, the The winding of the hose unfolds progressively but makes it possible to maintain the transport of fluids between the body of the satellite and the photovoltaic generator, and a step in which, at regular intervals, for example every 24 hours at a time when the zone of the earth which the satellite faces is at night time, the photovoltaic generator rotates rapidly in the opposite direction of its direction of sun tracking, so as to bring the winding hose to neutral position, or maximum deployment position in the 'other way.
    [0005] DESCRIPTION OF THE FIGURES The characteristics and advantages of the invention will be better appreciated thanks to the description which follows, description which sets out the characteristics of the invention through a non-limiting example of application. The description is based on the appended figures which represent: FIG. 1: a schematic representation of a three-axis stabilized satellite orbiting the Earth, FIG. 2: a schematic representation in perspective of a 3-axis stabilized satellite, comprising Figure 3: a schematic representation of a similar satellite, with the deployable generator stored, Figure 4: a representation of the satellite of Figure 2, the deployable generator being in the process of deployment. DETAILED DESCRIPTION OF ONE EMBODIMENT OF THE INVENTION The present invention is intended to be used on a satellite of the type intended to be set to and stabilized on 3 axes. As can be seen in FIGS. 1 and 2, in a purely illustrative and in no way limiting example, it is assumed here that this is a satellite 10 comprising a body 11 of substantially parallelepipedal shape, the centers of the faces of which define X, Y, Z axes. Such a three-axis stabilized satellite can typically be a telecommunications satellite placed in geostationary orbit. The invention, however, is not limited to the geostationary orbit. It is applicable to any orbit and any satellite with a solar generator oriented towards the sun. This is particularly the case when the position of the solar generator (GS 20 by abbreviation) to the sun is provided by a yaw steering law ("yaw steering" in English), possibly combined with a rotation of the GS by a mechanism . When the satellite 10 is placed stationary on its orbit, its faces are oriented so that a North-South axis Y is parallel to the axis of rotation of the Earth, an axis X says East-West is parallel to the axis of displacement of the satellite on its orbit, and a Z axis said Earth-anti-Earth is perpendicular to the two others. A telecommunication satellite in geostationary orbit 10 always has the same face towards the Earth during its orbit, and its North 14 and South faces 15, 30 substantially parallel to the terrestrial equator when the satellite 10 is set, are the least illuminated by the sun S, during the orbit of the satellite around the Earth. In the present example, the satellite 10 considered comprises two photovoltaic generators 12a, 12b, attached to these North 14 and South 15 faces. Each photovoltaic generator 12a, 12b is attached to the body 11 of the satellite 10 by an arm 18a, 18b (commonly called "Yoke" in English). In this way, these photovoltaic generators 12a, 12b remain permanently out of the shadow generated by the satellite 10 itself. A mechanism allows both the orientation of the photovoltaic generator towards the Sun, and the current flow from the photovoltaic generator to the satellite. The North 14 and South 15 faces also usually comprise radiator zones (perpendicular to the North axis). South and not shown in the figure), thus taking advantage of their low exposure to the Sun.
    [0006] The satellite 10 furthermore comprises a set of transmitting antennas 13 (see FIG. 2), attached mainly to the East 16 and West 17 faces, and thus always facing the Earth during the orbit of the satellite 10. This architecture general of the satellite 10 is well known to those skilled in the art, and as such comes out of the scope of the invention. It is not described further here. While the satellite still keeps the same face facing the Earth, the photovoltaic generators 12a, 12b remain, as much as possible, always oriented towards the Sun S. The photovoltaic generators 12a, 12b thus rotate about the North axis. South compared to the satellite in 24 hours to follow the Sun. These photovoltaic generators 12a, 12b comprise, in the present example of implementation, deployable assemblies formed of an articulated assembly of substantially planar and stackable solar panels 20. This configuration is illustrated in particular in FIG. 2. These solar panels 20 are here rectangular panels, the unit area of which is generally close to or less than the surface of a North 14 or South 15 lateral face of the satellite 10. At launch, the panels Solar cells 20 forming the photovoltaic generators 12a, 12b are folded over each other, so as to form a compact assembly attached to a wall of the body of the satellite 10 (see Figure 3). After the satellite 10 has been placed in geostationary orbit, the solar panels 20 are deployed, according to a kinematics known per se, to offer their maximum surface area to solar radiation. The mechanism and the mode of deployment of the solar panels, between their folded configuration and their deployed configuration, are assumed to be known per se. The solar panels 20, once deployed, form in orbit a substantially flat rectangular paving structure, comprising for example three columns (along the North-South axis Y) and five rows (along the East-West axis X). The assembly of solar panels 20 could naturally, alternatively, include any number of columns and rows, depending on the needs related to the mission of the satellite 10. Each solar panel 20 has a covered surface of photovoltaic cells, this face being intended to be turned towards solar radiation. In the embodiment given here by way of non-limiting example and in particular illustrated in FIG. 2, each photovoltaic generator 12a, 12b comprises, between the assembly of the solar panels 20 itself and the body 11 of the satellite 10, a radiator panel 21 In the present exemplary embodiment, the radiator panel 21 is connected by only one of its sides to the solar panel array assembly 20. More precisely, in the configuration illustrated in FIG. 2, the radiator panel 21 is connected to a assembly of three columns of five rows of solar panels 20 by its edge farthest from the body 11 of the satellite 10. This radiator panel 21 is aligned with the central column of solar panels 20. It comprises deployment means similar to those of the panels solar cells, and outside the scope of the present invention. The radiator panel 21 is attached to the body 11 of the satellite via the arm 18. The radiator panel 21 also comprises means for transmitting the mechanical forces experienced by the photovoltaic generator 12 on the arm 18, so that these are taken up by the structure of the satellite 10. In the present embodiment, the radiator panel 21 is connected by its outer edge to the assembly of solar panels 20. In fact it is connected here to the inner edge of the solar panel disposed in the central column and the row closest to the satellite body.
    [0007] It may also, alternatively, be connected to this assembly of solar panels by another mechanical support arm, aligned with the arm 18 (configuration not shown in the figures). As seen in Figure 3, when the satellite 10 is in the launch position, the panels 20, 21 forming a photovoltaic generator 12a, 12b are folded over each other and are placed in front of a face (eg North or South) of the satellite. The radiator panel 21 is then stacked with the other panels. The configuration of the satellite 10 in the launching position is unchanged compared to a conventional satellite configuration (devoid of photovoltaic generators as exposed), with however thicker panel stacks here of the thickness of the radiator panel 21. launch configuration, each support arm 18a, 18b of a generator 12a, 12b is retracted inside the body 11 of the satellite 10.
    [0008] FIG. 4 then illustrates a configuration of a photovoltaic generator in the course of deployment, when the satellite 10 is put into service once it has been placed in geostationary orbit. During this deployment, each arm 18a, 18b comes out of its housing in the body of the satellite. Then the panels 20, 21 are deployed in a conventional kinematic to adapt their maximum surface position to solar radiation. In the present example, the first panel 21 is the radiator panel, each subsequent row of panels carrying three solar panels 20. In the present embodiment, in no way limiting, each radiator panel 21 comprises a path of one or more circulation tubes. of fluid. They may be heat pipe tubes and / or fluid loop. This or these tubes perform the heat exchange between the fluid flowing inside said tubes, and the outside environment, in a direction not exposed to solar radiation.
    [0009] Each radiator panel 21 has a thermally insulating face, oriented towards the solar radiation, and therefore coplanar with the face of the solar panels 20 carrying the photovoltaic cells. This insulating face is for example, but not limited to, covered by multilayer insulation MLI (Multi-Layer Insulation), of a type known per se. This insulating face may also include a layer of any other thermally insulating material. The purpose of this isolation is to minimize radiation inputs on the "Sun" side. The face opposite to the thermally insulating face, that is to say the face oriented back to the sun, is, in the present embodiment, covered with a radiative device. This radiative device is here, but not limited to, white paint. It may also be OSR (Optical Solar Reflector) or any other means known per se for this purpose. As mentioned above, the photovoltaic generator 12a, 12b remains almost permanently oriented towards the sun, while the body of the satellite remains facing the earth. As a result, the photovoltaic generator is rotated around an axis of rotation represented by the arm 18a, 18b with respect to the body of the satellite 10. The transport of heat transfer fluid between the body of the satellite 10 and the radiator panel 21 of the photovoltaic generator 12a, 12b must therefore take into account this rotational movement. In a first implementation, the photovoltaic generator 12a, 12b comprises a rotary joint disposed between the structure of the satellite 10 and the panel assembly 20, 21, for example at the support arm of the deployable photovoltaic generator, said seal rotating having its axis of rotation coincides with the axis of rotation of the arm. This rotary joint comprises means for passing at least one fluid and means for transferring electric current. In one embodiment, the rotary joint comprises in the central portion one or two concentric fluid flow devices (depending on whether the device uses a heat pipe or a fluid loop). Such rotating joints for the passage of a fluid are already known per se and, as such, are within the scope of the present invention. For such examples of rotary joints used in the space, reference may be made, for example, to AIAA Paper 86- 1298 Heizer, Goo, Rhodes et al, "A steerable radiator for spacecraft application", 4th seal thermophysics and heat transfer conference. The electrical current passage means are, for example, consist of a disc-shaped conductive track, traversed by a broom 5 also conductive. The path of the broom on the track ensures the passage of the current regardless of the respective angular position of these means. In a second implementation, the photovoltaic generator 12a, 12b comprises a hose winding (coil without hub) adapted to the passage of fluid between the body of the satellite and the radiator panel 21. This hose winding is secured to one of its ends, the heat removal circuit of the body of the satellite, and at its other end, the radiator panel 21. Here again, the hose winding may comprise one or two fluid passage tubes, depending on the satellite is equipped with a heat pipe cooler or fluid loop .. Such a winding "pigtail", has at rest a small length, turns of tube coming into contact with each other, and has in position deployed an arbitrarily large length, determined by the number of turns and their diameter. Such devices are already known per se and come out, as such, from the scope of the present invention. In this way, when the photovoltaic generator is rotated about its axis relative to the body of the satellite, the winding of flexible deploys gradually but allows the transport of fluids (and possibly electricity) between the body of the satellite and the generator. At regular intervals, for example every 24 hours at a time when the area of the earth to which the satellite faces is at night, the photovoltaic generator rotates rapidly (for example in a few tens of minutes) in the opposite direction its direction of sun tracking, so as to bring the hose winding to the neutral position, or to the maximum deployment position in the other direction. It is thus possible to provide a maximum travel of the generator around its axis of rotation of + - 190 °, and to control every 24 hours, when the generator is less than 10 ° of the maximum unwinding of the hose winding, a fast rotation to the opposite position. The electrical current passage means are, for example, consist of a disk-shaped conductive track, traversed by a broom also conductive. The path of the broom on the track ensures the passage of the current regardless of the respective angular position of these means. Alternatively, the flow of current can also be through flexible strands in the manner of the winding of the coolant pipe, to aim to simplify the need for the rotation mechanism.
    [0010] In the implementation that has been described here, only one or both deployable generators comprise a radiator panel 21. In a variant, one of the photovoltaic generators 12a, 12b comprises two aligned and connected radiator panels 21 arranged according to the central panel column 20 of the photovoltaic generator 12a, 12b. It is clear that, in another variant embodiment, any number of panels forming the deployable generator 12a, 12b may consist of radiator panels 21, according to the specifications of the mission of the satellite 10, and, in particular, according to the thermal power that must be dissipated.
    [0011] Similarly, only one of the two deployable photovoltaic generators may comprise a radiator panel 21. In this case, the other photovoltaic generator is modified in size, so as to ensure the balancing couples of the satellite.
    [0012] In another variant, at least one radiator panel 21 is attached to the rear face (opposite to the photovoltaic sensors) of one of the solar panels 20 of the photovoltaic generator 12a, 12b.
    权利要求:
    Claims (13)
    [0001]
    REVENDICATIONS1. A photovoltaic generator (12a, 12b) deployable for a satellite (10) of stabilized three-axis type, said photovoltaic generator (12a, 12b) comprising a set of panels (20, 21) articulated with respect to each other, and an arm of fastener (18a, 18b) to the structure of a satellite body (11), said photovoltaic generator (12a, 12b) being able to adopt a first position, called launch position, in which the panels (20, 21) are folded over each other, and a second position, said extended position, wherein the panels (20, 21) are fully deployed, at least a portion of the panels being solar panels (20) having a photovoltaic sensor face for to be oriented towards the solar radiation, characterized in that at least one panel (21) consists of a thermal radiator, its radiative face being oriented opposite to the face of the photovoltaic panels carrying the sensor s photovoltaic, when the photovoltaic generator (12a, 12b) is in the deployed position, this radiative face being said "shadow" face, and the opposite face of the radiator panel (21) being said face "sun".
    [0002]
    2. Photovoltaic generator (12a, 12b) according to claim 1, characterized in that the panel closest to the support arm (18) is a panel (21) forming a heat sink.
    [0003]
    3. Photovoltaic generator (12a, 12b) according to claim 2, characterized in that it comprises a matrix array of rectangular solar panels (20) arranged in lines, and a radiator panel (21) aligned with the central column of panels PV.
    [0004]
    4. Photovoltaic generator (12a, 12b) according to any one of claims 1 to 3, characterized in that it comprises an insulating zone between the radiator panel (21) and the solar panels (20).
    [0005]
    5. Photovoltaic generator (12a, 12b) according to any one of claims 1 to 4, characterized in that at least one radiator panel (21) is traversed by tubes of a fluid loop or a heat pipe, its shadow face being covered with a radiative device, and its sun face is covered with a multilayer insulator.
    [0006]
    6. Satellite (10), stabilized type on three axes, characterized in that said satellite comprises at least one expandable photovoltaic generator (12a, 12b) according to any one of claims 1 to 5, said photovoltaic panels being stored in position folded during launch, and then deployed once the satellite set, the radiator panel (21) being disposed closer to the body (11) of the satellite (10).
    [0007]
    7. Satellite according to claim 6, characterized in that it comprises a rotating joint disposed between the body (11) of the satellite (10) and the set of panels (20, 21), disposed at the arm (18). for supporting the deployable photovoltaic generator (12a, 12b), said rotary joint having its axis of rotation coincides with the axis of rotation of the arm, and comprising means for passing at least one fluid and means for transferring electric current .
    [0008]
    8. Satellite (10) according to claim 7, characterized in that, in the case where the photovoltaic generator comprises a radiator panel (21) using a fluid loop, the rotary joint comprises in the central part two concentric fluid flow devices, oriented along the axis of rotation of the rotary joint.
    [0009]
    9. Satellite (10) according to claim 7 or 8, characterized in that the means of passage of the electric current consist of a disc-shaped conductive track, traversed by a broom also conductive.
    [0010]
    10. Satellite (10) according to claim 6, characterized in that the satellite comprises a hose winding adapted to the transport of fluid between the body of the satellite and the radiator panel (21), the hose winding being secured to one of its ends, to a heat removal circuit of the body of the satellite, and at its other end, to the radiator panel (21).
    [0011]
    11. Satellite (10) according to claim 10, characterized in that the means for passing the electric current are constituted by at least one flexible strand.
    [0012]
    12. Satellite (10) according to claim 10 or 11, characterized in that the hose winding comprises two fluid passage tubes, if the satellite is equipped with a fluid loop cooler.
    [0013]
    13. A method of controlling a photovoltaic satellite generator according to claim 12, characterized in that it comprises a step in which, during the rotation of the photovoltaic generator (12a, 12b) about its axis relative to the body (11 ) of the satellite (10), the hose winding unfolds progressively but makes it possible to maintain the transport of fluids between the body of the satellite (10) and the photovoltaic generator (12a, 12b), and a step in which, at regular intervals , for example every 24 hours at a time when the area of the earth to which the satellite faces is at night, the photovoltaic generator (12a, 12b) rotates rapidly in the opposite direction of its direction of tracking of the sun, so as to return the hose winding to the neutral position, or to the maximum deployment position in the other direction.
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    同族专利:
    公开号 | 公开日
    US20160311561A1|2016-10-27|
    CA2969037A1|2015-06-18|
    FR3014849B1|2018-06-15|
    US10207824B2|2019-02-19|
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    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    US4947825A|1989-09-11|1990-08-14|Rockwell International Corporation|Solar concentrator - radiator assembly|
    US5386953A|1991-11-08|1995-02-07|Calling Communications Corporation|Spacecraft designs for satellite communication system|
    EP0945343A1|1998-03-26|1999-09-29|Fokker Space B.V.|System to link the kinematics of neighbouring panels in a panel assembly|
    US6102339A|1998-04-17|2000-08-15|Turbosat Technology, Inc.|Sun-synchronous sun ray blocking device for use in a spacecraft having a directionally controlled main body|
    US6075200A|1999-06-30|2000-06-13|Entech, Inc.|Stretched Fresnel lens solar concentrator for space power|
    US6284966B1|2000-03-06|2001-09-04|The Aerospace Corporation|Power sphere nanosatellite|
    US5806800A|1995-12-22|1998-09-15|Caplin; Glenn N.|Dual function deployable radiator cover|
    US6378809B1|1997-10-10|2002-04-30|Space Systems|AFT deployable thermal radiators for spacecraft|
    FR2823182B1|2001-04-05|2004-06-04|Cit Alcatel|DEPLOYABLE RADIATOR FOR SPACE ENGINE|
    US6854510B2|2001-04-24|2005-02-15|Space Systems/Loral, Inc.|Spacecraft radiator system and method using cross-coupled deployable thermal radiators|
    US20080164012A1|2006-12-11|2008-07-10|Cavanaugh David B|Radiative cooling system|
    WO2010111364A1|2009-03-24|2010-09-30|Lockheed Martin Corporation|Spacecraft heat dissipation system|FR3031969B1|2015-01-27|2017-01-27|Airbus Defence & Space Sas|ARTIFICIAL SATELLITE AND METHOD FOR FILLING A PROPULSIVE GAS TANK OF SAID ARTIFICIAL SATELLITE|
    FR3041608B1|2015-09-25|2018-04-13|Thales Sa|DEPLOYABLE ASSEMBLY|
    US10618677B1|2017-05-03|2020-04-14|Space Systems/Loral, Llc|Articulating sunshield|
    CN110341990B|2019-07-05|2020-09-04|中国人民解放军国防科技大学|Space debris racemization capturing device|
    RU200445U1|2020-07-03|2020-10-26|Федеральное государственное бюджетное образовательное учреждение высшего образования "Рязанский государственный радиотехнический университет имени В.Ф. Уткина"|BINARY SPACE WITH A RECONFIGURABLE ANTENNA COMBINED WITH A SOLAR BATTERY DEPLOYABLE MULTIVECTOR MATRIX ROCKET ENGINES|
    法律状态:
    2015-12-23| PLFP| Fee payment|Year of fee payment: 3 |
    2016-12-29| PLFP| Fee payment|Year of fee payment: 4 |
    2017-12-29| PLFP| Fee payment|Year of fee payment: 5 |
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    优先权:
    申请号 | 申请日 | 专利标题
    FR1362606A|FR3014849B1|2013-12-13|2013-12-13|DEPLOYABLE RADIATOR FOR SATELLITE STABILIZED THREE AXES|FR1362606A| FR3014849B1|2013-12-13|2013-12-13|DEPLOYABLE RADIATOR FOR SATELLITE STABILIZED THREE AXES|
    CA2969037A| CA2969037A1|2013-12-13|2014-12-15|Radiator deployable for a satellite stabilized on three axes|
    US15/103,834| US10207824B2|2013-12-13|2014-12-15|Radiator deployable for a satellite stabilized on three axes|
    EP14812483.7A| EP3079988B1|2013-12-13|2014-12-15|Radiator deployable for a satellite stabilized on three axes|
    PCT/EP2014/077746| WO2015086851A1|2013-12-13|2014-12-15|Radiator deployable for a satellite stabilized on three axes|
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