• 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 spacecraft comprising: - a body (4) having a + Y face (18) and a Y face (16) opposite the + Y face (18); a radiator (22, 24) carried by at least one of the + Y face (18) and the Y face (16); said radiator (22, 24) having an outer face (30); characterized in that it comprises an assembly (34) having at least one absorption face (40) adapted to absorb a portion of the incident solar radiation during the spring and autumn equinoxes, said assembly (34) being adapted to transmitting said absorbed heat to the radiator (22, 24); said absorption face (40) being inclined with respect to the outer face (30) of the radiator (22, 24) at an angle substantially between 20 ° and 26 °.
    公开号:FR3030457A1
    申请号:FR1462565
    申请日:2014-12-17
    公开日:2016-06-24
    发明作者:Fabrice Mena；Patrick Coutal
    申请人:Airbus Defence and Space SAS；
    IPC主号:
    专利说明:

    [0001] The present invention relates to a spacecraft and in particular to a geostationary satellite.
    [0002] Due to the earth's revolution around the sun, the different faces of a geostationary satellite do not receive the same amount of solar radiation during the seasons. This results in cyclic temperature variations on the + Y face and the Y-face. Thus, during the winter and summer equinoxes (EQ), the - Y and + Y faces have lower temperatures than during the winter (WS) and summer (SS) solstices as shown in the figure. 1. In addition, the temperature of the satellite faces also fluctuates over time. This temperature is approximately equal to 20 ° C at the beginning of life of the satellite and 70 ° C at the end of life of the satellite. These temperature variations over the seasons and over time are reproduced on the ground in a vacuum atmosphere during satellite validation tests. These validation tests are long and constraining to achieve. To facilitate ground validation tests, it has been envisaged to decrease temperature variations over seasons and over time. To this end, the + Y and - Y sides of the satellite were warmed at the time of the equinox by electric heaters. Nevertheless, the heaters are limited in efficiency and require over-sizeing the satellite's electrical power system. This oversizing significantly increases the cost of the satellite. Furthermore, to regulate the temperatures of the earth observation satellite faces, louvers (in English "louvers") were fixed vis-à-vis radiators mounted on the outer faces of the satellites. These louvers obscure more or less the outer face of the radiator according to the cooling requirements of the satellites. The louvers are pivotally mounted about an axis disposed along a longitudinal edge of the louvers. In the closed position, such louvers obscure the entire outer face of the radiator. These shutters have the sole function of preventing solar radiation from reaching the radiator. In addition, these louvers require a complex mechanism to rotate the louvers. This mechanism must have good behavior in a variable thermal environment.
    [0003] The object of the present invention is to provide a low cost system that would limit the temperature variations between the equinoxes and the solstices.
    [0004] For this purpose, the subject of the invention is a machine comprising: a body having a + Y face and a Y face opposite the + Y face; a radiator carried by at least one face among the face + Y and the face -Y; said radiator having an outer face; characterized in that it comprises an assembly having at least one absorption surface adapted to absorb a portion of the incident solar radiation (Ic) during the spring and autumn equinoxes, said assembly being adapted to transmit said absorbed heat to the radiator ; said absorption face being inclined with respect to the outer face of the radiator at an angle substantially between 20 ° and 26 °.
    [0005] Advantageously, the assembly is adapted to absorb the solar radiation at the equinox and to transfer, by radiation or by fluid transfer (heat pipe), the heat generated by these solar radiation to said radiator to heat it. The increase of the radiator temperature generated by the assembly during the equinox makes it possible to reduce the temperature variations of this face of the vehicle between the seasons. Advantageously, the decrease in temperature variations of the faces of the spacecraft makes it possible to reduce the duration of the qualification tests in thermal vacuum. This would result in a significant gain in terms of the cost and design time of a spacecraft.
    [0006] Advantageously, this assembly is passive. They do not need electrical power to heat the radiator. In addition, this assembly can be performed at a low cost. It has a low mass and a small footprint especially during the launch phase of the satellite.
    [0007] According to particular embodiments, the spacecraft comprises one or more of the following characteristics: The assembly comprises at least one lamella having a flat main face disposed in line with the radiator outer face, this flat face being the front face. absorption, the absorbed heat being transmitted to the radiator by radiation. Advantageously, as said at least lamella is inclined at an angle corresponding to the angle formed between the outer face of the radiator and the direction taken by the solar radiation during the solstice, said at least lamella is not heated during solstice. - The machine comprises a support arm carrying said assembly, said support arm extending substantially perpendicularly to the outer face of the radiator, and at least one own drive motor to drive the support arm in rotation along a longitudinal axis of the support arm, the speed of rotation of the support arm being substantially equal to the speed of rotation of the sun around the spacecraft. Advantageously, since the support arm is rotated at a speed corresponding to the speed of rotation of the sun around the spacecraft, during the equinox, the absorption face of each lamella is at right solar radiation All day long. Advantageously, the absorption faces of the lamellae are thus illuminated in the same way throughout a day. They can thus transmit to the radiator the same amount of heat during a day.
    [0008] The assembly comprises at least two lamellae fixed one above the other, and separated from one another by a free space. - The machine comprises at least one solar panel, and wherein said support arm supports said at least one solar panel. Since the solar panels are generally rotated around an axis perpendicular to the outer face of the radiator at a speed equal to the speed of rotation of the sun around the spacecraft, the use of this support arm to carry the assembly of heating slats makes it possible not to add an additional rotation mechanism. This implementation is very advantageous from an economic point of view. The lamellae comprise a polyimide film and rods adapted to stiffen said polyimide film. Said at least absorption face has a coating having a high solar absorptivity, for example a coating of black color. - The outer face of said radiator has a non-specular coating. - Each slat has a thickness less than one millimeter. The assembly comprises at least two cones each having an absorption face and at least one thermally connected heat pipe with absorption faces and with the radiator, the heat absorbed by the absorption face being transmitted by said heat pipe. radiator. The invention will be better understood on reading the description which will follow, given solely by way of example and made with reference to the drawings, in which: FIG. 1 are two curves representative of the variations of temperature of the faces + Y and Y of a spacecraft according to the state of the art during a year; FIG. 2 is a perspective view of the spacecraft according to a first embodiment of the invention, in geostationary orbit; - Figure 3 is a sectional view of an assembly of heating slats; FIG. 4 is a schematic perspective view illustrating the mode of operation of the first embodiment of the invention during the summer solstice; FIG. 5 is a schematic perspective view illustrating the mode of operation of the first embodiment of the invention, during the equinox; FIG. 6 is a perspective view of the spacecraft according to a second embodiment of the invention, in geostationary orbit; and FIG. 7 is a schematic perspective view illustrating a part of the second embodiment of the invention.
    [0009] The present invention is defined with respect to an orthogonal coordinate system R (X, Y, Z) shown in Figures 2 to 5. The direction of the X, Y and Z vectors is defined as the positive direction. The opposite direction is defined as being a negative sense.
    [0010] In the different figures, the same references designate identical or similar elements. With reference to FIG. 2, a spacecraft 2 according to the first embodiment of the invention is in the form of a parallelepipedal box 4. This box 4 always has the same face directed towards the Earth, this face being called the Earth face 6. The face opposite and parallel to the Earth face 6 is itself called the anti-Earth face 8. The face - X, also called East face 10, and the face + X, also called West face 12, are opposite faces, parallel to each other and perpendicular to the direction of movement of the spacecraft 2. Communication antennas 14 are generally fixed on the faces -X 10 and + X 12. The face - Y, also called North face 16, and the face + Y, also called South face 18, are two other faces of the body. They are opposite, parallel to each other and perpendicular to the North-South axis of the Earth. The spacecraft 2 comprises a first radiator 22 and a second main radiator 24 for cooling electronic equipment contained in the body. These electronic equipment not shown in the figures are thermally connected to the first and second radiators, for example, via heat pipes also not shown. The first radiator 22, of parallelepipedal general shape, has four lateral faces 26, an inner main face 28 fixed on the face -Y 16, an outer main face 30 opposite to the inner main face and located on the outer space side. the spacecraft. The inner main faces 28 and outer 30 extend in the plane (X, Y). The second radiator 24 is identical to the first radiator 22. It will not be described in detail. It is fixed on the face + Y 18.
    [0011] The spacecraft 2 comprises two support arms 32 extending substantially in the z direction, two solar panels 20 each fixed to a distal end of a support arm 32, and one or two unrepresented geared motors capable of rotating in rotation. each support arm 32 at a speed substantially equal to the speed of rotation of the sun around the spacecraft. Thus, the solar panels 20 are directed towards the sun all day long. The spacecraft 2 further comprises two assemblies 34 of lamellae 36, each attached to a support arm 32 between the body 4 and a solar panel 20.
    [0012] Each assembly 34 of slats is stationary relative to the support arm 32 which carries it. Each assembly 34 of slats is rotated by the support arm 32 which carries it. The lamellae 36 of the lamellae assemblies are calorifiers. They are adapted to absorb the solar radiation at the equinox and to transfer, by radiation, the heat generated by this solar radiation to the first radiator 22 and the second radiator 24, as explained below. The heating slats 36 are plates of very small thickness of the order of millimeters or less. They are, for example, made from a sheet and stiffening rods fixed on one side of the sheet. The sheet is made, for example, by a polyimide film designated by the trademark "Kapton" or by graphite layers. The rods are, for example, made of carbon. As shown in Figure 3, the heating slats 36 are fixed one above the other. They are superimposed along the z direction. A free space e is arranged between two heating slats 36 superimposed one on the other. The heating slats 36 of the assembly are advantageously fixed in a frame 38. The heating slats 36 have a flat main face 40, called the absorption face 40 and disposed opposite the outer face 30 of the first radiator or the second radiator. radiator 24, an outer face 42 opposite to the absorption face 40. The absorption face 40 of the assembly 34 of lamellae mounted on the face -Y 16 is inclined relative to the outer main face 30 of the first radiator d a dihedral angle of about 23.5 °. In the same way, the absorption face 40 of the assembly of lamellae mounted on the + Y face 18 is inclined relative to the outer main face 30 of the second radiator by a dihedral angle of approximately equal to 23.5. °. The opening of said dihedral angle extends towards the direction of incident solar radiation Is.
    [0013] As can be seen in FIG. 4, the value of this dihedral angle corresponds to the maximum angle formed between the outer main face 30 of the first radiator 22 and the direction of the incident solar radiation Is during the winter solstice. In the same way, this value corresponds to the maximum angle formed between the outer main face 30 of the second radiator 24 and the direction of incident solar radiation Is during the summer solstice. In operation, during the summer solstice, the direction of solar radiation is parallel to the heating slats 36. As the heating slats have a low thickness, their presence does not lead to loss of thermal rejection capacity for the first radiator 22 Almost all of the incident solar radiation reaches the first radiator 22. The support arm 32 pivots about its longitudinal axis (around the z direction) at a speed equal to the speed of rotation of the sun around the spacecraft. . Thus, the direction of solar radiation Is is parallel to the heating slats throughout the day.
    [0014] During the winter solstice, the direction of the solar radiation is parallel to the heating slats of the assembly 34 of slats attached to the support arm located on the side of the face + Y 18. The presence of the heating slats does not lead to heat loss for the second radiator 24. This situation, similar to the situation of the winter solstice, was not represented.
    [0015] With reference to FIG. 5, during the equinox, the direction of the incident solar radiation extends in the plane (X, Y). This direction forms it at an angle of about 23.5 ° with the median plane of the heating slats 36. Consequently, the absorption faces 40 of the heating slats which are facing the outer face 30 of the first radiator and the faces of absorption 40 of the heating slats which are facing the external face 30 of the second radiator absorb the UV portion of the incident solar radiation 1 and heat up. The heating slats 36 can thus reach a temperature of about 90 ° C. When the heating slats 36 are hot, they radiate this heat to the radiator located on the side of the support arm carrying the heating assembly. Thus, during the equinox, the first radiator 22 and the second radiator 24 heat up and warm the + Y faces 18 and -Y 16 of the body. As for the summer solstice, the support arm 32 pivots about its longitudinal axis at a speed equal to the speed of rotation of the sun around the spacecraft so that the incident solar radiation warms the absorption faces. 40 heat slats throughout the day. A temperature rise of about 15 ° C. may, for example, be achieved by means of an assembly of twenty-five heat slats of a size of 2300 by 50 millimeters, two adjacent lamellae being spaced apart from each other. a gap of 19.9 millimeters. Thus, the temperature variations between the equinoxes and the solstices, are diminished. Preferably, the absorption faces (40) of the heating slats 36 have a coating with high solar absorptivity. to absorb a maximum of solar radiation. They are, for example, black. The skilled person of the spatial thermic craft chooses a material adapted according to multiple criteria apart from the absorptivity. In general, the materials and coating chosen will have a solar absorptivity greater than 0.6. Preferably, the first radiator 22 and the second radiator 24 are coated with a nonspecular or non-specular material so that the radiators do not heat the heat sipes 36 during the solstices by reflection of the solar flux. Preferably, the heating slats 36 are slabs less than one millimeter thick. Note that unlike the louvers (English "louvers"), the heating slats 36 are not mobile about an axis contained in their median plane or parallel to their median plane. They are fixed relative to the frame and the support arm 32 which carries them. As a variant, the heating slats 36 have a median plane inclined with respect to the outer main face 30 of the first and second radiators with a dihedral angle of between 20 ° and 26 °.
    [0016] FIG. 6 represents a spacecraft 50 according to a second embodiment of the invention. This spacecraft 50 is identical or similar to the spacecraft 2 according to the first embodiment of the invention with the exception that the assemblies 34 of heat sinkers carried by the support arms have been replaced by assemblies of 52 cones attached to the first 22 and second 24 radiators. The technical elements of the spacecraft 2 according to the first embodiment identical or similar to the technical elements of the spacecraft 50 according to the second embodiment will not be described a second time.
    [0017] The cone assemblies 52 are fixed on the portion of the first 22 and the second 24 radiators located above the solar panels 20, when they are folded for launching. They have a small diameter D so that they cover a small part of the radiators. For example, cones 54 having circular edges 56 with a diameter of 150 mm can be used.
    [0018] The assembly of cones 52 has, moreover, a low height H so as to be housed in the space under the cap. The cone assemblies 52 have the function of absorbing the maximum amount of solar radiation at the time of the equinox and of transferring this heat to the faces Y 16 and Y 18 and absorbing the minimum of solar radiation during the solstices. The cone assemblies 52 fixed on the first radiator 22 are identical and fixed in the same way as the cone assemblies 52 fixed on the second radiator 24. To simplify the description, only the cone assemblies 52 fixed on the first radiator are described. below.
    [0019] The exemplary embodiment of a cone assembly 52, illustrated in FIG. 6, comprises six cones 54 each provided with an orifice 56 at its top 58 and a central heat pipe 60 housed in the orifice 56 of each cone 54. The central heat pipe 60 is substantially perpendicular to the outer main face 30 of the first radiator. It is attached to each cone 54 as well as to the first radiator 22.
    [0020] The cones 54 of the cone assembly are arranged one above the other with their apex 58 located in line with the outer face 30 of the first radiator. The circular edges 56 of two adjacent cones 54 are separated by a free space e. The cones 54 have a cone angle 13 substantially equal to 23, 50. The cones 54 are made of a light and heat conducting material. A material such as graphoil can for example be used.
    [0021] The outer face 62 of the cones, that is to say the face facing the outer face 30 of the first radiator is covered with a coating 64 which absorbs the UV portion of the incident solar radiation. This coating 64 has, in addition, a low emissivity. For example, a black polished nickel coating 64 may be used. This coating 64 is attached to the central heat pipe 60.
    [0022] The inner face 66 of the cones is made of a conductive material. For example, graphoil can be used. The inner face 66 is not thermally connected to the central heat pipe 60. An insulating mat is arranged between the inner face 66 and the outer face 62. Advantageously, the inner face 66 is painted in black color. Solar fluxes are absorbed and re-emitted in the IR.
    [0023] As the outer face of the cone above has low emissivity, this radiation is radiated to space. In operation, during the equinox, the outer faces 62 of each cone absorb solar radiation and transmit the heat captured, by thermal conduction, the central heat pipe 60. The central heat pipe 60 transmits this heat to the first radiator 22. During from the solstice, the inner face 66 of the cones absorbs the incident solar radiation and re-emits it by infra-red radiation to the space. Advantageously, the cone assemblies 52 can be easily mounted and disassembled on new satellites or already existing satellites. Advantageously, cone assemblies 52 are passive systems that operate without electrical power. Advantageously, the cone assemblies 52 have a low mass and a low cost.
    权利要求:
    Claims (10)
    [0001]
    CLAIMS.-Spatial machine (2, 50) comprising: - a box (4) having a + Y face (18) and a -Y face (16) opposite to the + Y face (18); a radiator (22, 24) carried by at least one of the + Y face (18) and the Y face (16); said radiator (22, 24) having an outer face (30); characterized in that it comprises an assembly (34, 52) having at least one absorption face (40, 62) adapted to absorb a portion of the incident solar radiation (1c) during the spring and autumn equinoxes, said assembly (34, 52) being adapted to transmit said absorbed heat to the radiator (22, 24); said absorption face (40, 62) being inclined with respect to the outer face (30) of the radiator (22, 24) by an angle (a, p) substantially between 20 ° and 26 °.
    [0002]
    2. A spacecraft (2) according to claim 1, wherein said assembly (34) comprises at least one lamella (36) having a flat main face (40) disposed in line with the radiator outer face (30), this planar main face (40) being the absorption face (40), the absorbed heat being transmitted to the radiator (22, 24) by radiation.
    [0003]
    3. Spatial vehicle (2) according to any one of claims 1 and 2, which comprises: - a support arm (32) carrying said assembly (34), said support arm (32) extending substantially perpendicular to the outer face (30) of the radiator, and - at least one own drive motor for driving the support arm (32) in rotation along a longitudinal axis of the support arm, the speed of rotation of the support arm (32) being substantially equal to the speed of rotation of the sun around the spacecraft (2).
    [0004]
    4. Spacecraft (2) according to any one of claims 1 to 3, wherein said assembly (34) comprises at least two lamellae (36) fixed one above the other, and separated the one from the other by a free space (e).
    [0005]
    5. Spacecraft (2) according to any one of claims 3 and 4, which comprises at least one solar panel (20), and wherein said support arm (32) supports said at least one solar panel (20) .
    [0006]
    6. A spacecraft (2) according to any one of claims 2 to 5, wherein said strips (36) comprise a polyimide film and rods adapted to stiffen said polyimide film.
    [0007]
    7. A spacecraft (2) according to any one of claims 1 to 6, wherein said at least one absorption face (40) has a coating having a high solar absorptivity, for example a black color coating.
    [0008]
    8. A spacecraft (2) according to any one of claims 1 to 7, wherein the outer face (30) of said radiator has a non-specular coating.
    [0009]
    9. A spacecraft (2) according to any one of claims 2 to 8, wherein each slat (36) has a thickness less than one millimeter.
    [0010]
    The spacecraft (50) according to claim 1, wherein the assembly (52) comprises at least two cones (54) each having an absorption face (62) and at least one thermally coupled heat pipe (60). absorption faces (62) and to the radiator (30), the heat absorbed by the absorption face (62) being transmitted by said heat pipe (60) to said radiator (30).
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    同族专利:
    公开号 | 公开日
    US20170341781A1|2017-11-30|
    FR3030457B1|2018-06-01|
    WO2016097575A1|2016-06-23|
    EP3218266A1|2017-09-20|
    EP3218266B1|2017-12-06|
    US9828116B1|2017-11-28|
    引用文献:
    公开号 | 申请日 | 公开日 | 申请人 | 专利标题
    JPH03109999U|1990-02-26|1991-11-12|
    US20080257525A1|2007-04-20|2008-10-23|Japan Aerospace Exploration Agency|Thermal control apparatus|
    KR101240434B1|2011-12-30|2013-03-11|한국항공우주연구원|Thermal control method for electrical heater of geostationart satellites by sun sensor|
    US20140299714A1|2013-04-09|2014-10-09|Lockheed Martin Corporation|Heat generating transfer orbit shield|
    US5954298A|1997-04-28|1999-09-21|Motorola, Inc.|Active spacecraft thermal control system and method|
    US6073888A|1998-12-02|2000-06-13|Loral Space & Communications, Ltd.|Sequenced heat rejection for body stabilized geosynchronous satellites|
    US7270302B1|2003-04-22|2007-09-18|Lockheed Martin Corporation|Scalable thermal control system for spacecraft mounted instrumentation|
    US9352856B1|2013-12-04|2016-05-31|Space Systems/Loral, Llc|Axially grooved crossing heat pipes|FR3030458B1|2014-12-18|2017-01-27|Airbus Defence & Space Sas|SPACE ENGINE|
    US10392135B2|2015-03-30|2019-08-27|Worldvu Satellites Limited|Satellite radiator panels with combined stiffener/heat pipe|
    FR3040045B1|2015-08-10|2017-09-08|Airbus Defence & Space Sas|ARTIFICIAL SATELLITE|
    US10780998B1|2017-03-22|2020-09-22|Space Systems/Loral, Llc|Spacecraft design with multiple thermal zones|
    法律状态:
    2015-12-30| PLFP| Fee payment|Year of fee payment: 2 |
    2016-06-24| PLSC| Publication of the preliminary search report|Effective date: 20160624 |
    2016-12-20| PLFP| Fee payment|Year of fee payment: 3 |
    2017-12-20| PLFP| Fee payment|Year of fee payment: 4 |
    2018-12-18| PLFP| Fee payment|Year of fee payment: 5 |
    2019-12-13| PLFP| Fee payment|Year of fee payment: 6 |
    2020-12-21| PLFP| Fee payment|Year of fee payment: 7 |
    2021-11-30| PLFP| Fee payment|Year of fee payment: 8 |
    优先权:
    申请号 | 申请日 | 专利标题
    FR1462565A|FR3030457B1|2014-12-17|2014-12-17|SPACE ENGINE|
    FR1462565|2014-12-17|FR1462565A| FR3030457B1|2014-12-17|2014-12-17|SPACE ENGINE|
    PCT/FR2015/053502| WO2016097575A1|2014-12-17|2015-12-15|Spacecraft|
    US15/535,900| US9828116B1|2014-12-17|2015-12-15|Spacecraft|
    EP15821119.3A| EP3218266B1|2014-12-17|2015-12-15|Spacecraft|
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