Deployable radiator for spacecraft

0001The technical field of the invention relates to such spacecraft satellites and probes with electronic or optical payloads. Such vehicles are typically equipped with means for generating, storing and distribution of electrical power to supply all embedded electronics. These electronic equipment in operation dissipate according to their own electrical efficiency a part of the electric power consumed into heat. This heat must be distributed and channeled conduction and must be evacuated from the spacecraft to the space vacuum radiation. Provided spacecraft comprising heat radiators to fulfill this function.

0002We know more fixed or deployable radiators configurations whose dimensions change with the size and the electric powers of satellites. A current trend télécommunicationsest have satellites sizes larger and larger, and the power generators more and more powerful, resulting in a growing need for radiating surfaces aboard satellites.

0003A radiator is more effective if it is not lit by the sun; so, for geostationary satellites, fixed radiators are generally preferably installed on the panels of the satellite structure exposed to North and South when the satellite is in its operational orbital position and that of the satellite body is oriented so as to align the antennas to the Earth's surface.

0004The deployable radiators are used when stationary surfaces inherent to the satellite structure are not sufficient to move the entire heat generated on board. The deployable radiators, like other deployable structures such as solar generators, antenna reflectors, et cetera, are folded against the satellite fixed structure in order to enroll in the volume available under the launcher fairing. After launch, the satellite separates from the launcher after ejection of the launcher fairing, and the various deployable structures are deployed as and when required to Mission by specifically tailored deployment mechanisms by each case.

0005When launching a geostationary satellite, the rocket launch typically causes the satellite to a so-called transfer orbit injection or "GTO", English acronym for "Geostationary Transfer Orbit."

0006Then board propulsion means on a satellite geostationary possible to transfer it from the injection orbit to its final orbital position on the geostationary arc says "GEO" Acronym English for "Geostationary Earth Orbit."

0007These propulsion means are, at present, most often consist of one or several chemical combustion engines integrated to the satellite.

0008In a future generation of satellites to increase capacity and the satellite's life, it is envisaged to expand the use of engines Power for the station acquisition and / or station keeping of satellites. Such motor is powered by renewable energy, namely electricity harvested by solar generators, which allows to limit the drawbacks of board chemical fuels such their mass and footprint storage.

0009So a new generation of electric powered satellites is track design and implementation. Regarding the sequencing of the last position, i.e. the transfer from its injection orbit to the position final orbital geostationary orbit, several options are possible.

0010It is necessary in the case of the use of electric propulsion for the transfer or a part thereof, harvesting of solar energy for conversion to electrical energy to power the electric motors. This can be achieved by partial or full deployment of the solar generators of the satellite. Or, in other scenarios, the satellite could perform its path between its orbit injection and the geostationary orbit or a part of the route with energy board at launch, either chemical or electrochemical (batteries) or nuclear.

0011In all cases, the energy available for the transfer is limited by the means harvesting or storage that need to be dimensioned with just.

0012Therefore in all cases, the available energy being dimensioned at the close, so do not waste as much as possible during the phase transfer.

0013It is known from document D1 = EP 0,780,304 A1 a spacecraft comprising radiators deployable, each having two substantially parallel surfaces, one of which first radiative face and a second insulating opposite face, so that that during the transfer phase GTO to GEO, the deployable radiators are in the stored position, pressed against the satellite structure, the faces radiant oriented satellite structure, and the insulation faces toward the vacuum space to reduce heating needs on board during this phase, and thus save energy on board. This system is illustrated in Figure 1 and be described in more detail below.

0014Other spacecraft with deployable radiator systems are known in the art earlier, for example the document WO 99/19212 = D2, the teaching will be described using FIG 2 of the present application. Here, it is deployable radiators have two radiating faces.

0015The deployable radiators of document D1 are very effective thermal insulation during the transfer phase into geostationary orbit from the injection orbit. However, their effectiveness as radiators heat when put in geostationary orbit position is compromised by that about half of the available area is insulating and not radiating.

0016The deployable radiators of document D2 are radiating on both faces completely, but provide no thermal insulation during phase transfer.

0017Spacecraft of the invention can overcome the drawbacks of the prior art. To this end, the invention provides a spacecraft comprising at least one deployable radiator having a stored pre-deployment position and a operational deployed position after deployment said radiator having two main faces, a first face which is stored towards the machine, and opposite second face which is oriented towards the space in the stored position; said first face being radiating at least for the major part of its surface; the least a portion of said second face being thermally insulating;<b>characterized in that</b>a substantial portion of said second face is thermally radiating, said portion being substantially hidden prior to deployment of said radiator by other elements of said spacecraft by screen during launch and at least a portion of the transfer path of said vehicle to its orbital position in geostationary orbit.

0018According to an advantageous characteristic, said deployable radiator is connected to the craft structure by a holding mechanism and deployment having a hinge and a hinge pin; <b>characterized in that</b>said axis is positioned with an optionally non-zero angle relative to main axes of the structure of said vehicle. According to a particular embodiment, said hinge is fixed on a portion of said spacecraft called "module Communication "module dedicated primarily to on-board equipment payload. According to an alternative embodiment, said hinge is attached to said part of said spacecraft "service module" dedicated module mainly embedded devices of the platform are in Support the good functioning of the equipment payload.

0019In a preferred embodiment, said radiator has, at position deployed determined at least one degree of freedom about at least one axis. according to another embodiment, said radiator has, established deployed position, the least two degrees of freedom about at least two axes. According to one embodiment particular, the main axis of said radiator in the deployed position is not parallel to a main axis of said gear. In another particular embodiment, the plane said main radiator in the deployed position is not parallel to a plane Master said vehicle.

0020Other features and advantages of the invention and its principal variants appear on reading the detailed description that follows, with its accompanying figures, of which<ul><li>Figure 1, already mentioned, shows deployable radiators known from the prior art, having a thermally insulating face;</li><li>Figure 2, already mentioned, shows deployable radiators known from the prior art, having both fully radiating faces;</li><li>Figures 3a and 3b, which show schematically and in plan first embodiment of the invention deployable radiators in stored position against the structure of the spacecraft in Figure 3a, and 3b deployed;</li><li>4a and 4b, which show two perspective views of a spacecraft equipped with deployable radiators according to Figure 3, but in deployed position;</li><li>Figure 5, which shows another embodiment of a deployable radiator according to the invention in stored position;</li><li>Figure 6, showing the deployable radiator of Figure 5 deployed position and its relation to a field instrument onboard;</li><li>Figure 7, showing the deployable radiator of Figure 5 deployed position and its relationship with the jet nozzles chemical propulsion of the vehicle;</li><li>Figures 8a and 8b, which show two deployable radiators Figure 5 in the extended position, in perspective in two directions of different view, and their relationship with mission antennas;</li><li>Figure 9 show the evolution of a machine such as that of Figures 8a and 8b in its orbit around the earth;</li><li>10 shows a mechanism for rotation about two axes.</li></ul>

0021In all the figures, the same references refer to the same elements. The scale is not always respected for reasons of clarity and readability.

0022Figure 1 shows a known satellite 10 of the prior art by the document D1. The satellite 10 is composed mainly of a service module 14 and a communication module 16. Deployable radiators 28, 30, 32, 34 are shown in the deployed position.

0023The heat to be removed, from the fixed equipment within the structure 12 of the satellite 10, is conducted to the radiators via heat pipes (not shown) having a flexible portion at the at least one hinge (s) of attachment deployable radiators on the structure 12 of the satellite.

0024Different propulsion equipment is intended to operate the satellite position, as well during the satellite transfer phases a orbit to another, during the spacecraft's mission to ensure station keeping. Figure 1 shows two pairs of two motors 24 to chemical propulsion on the ends of the panel "anti-land" of the module service 14, and a pair of electrical ion thrusters 26. Multiple engines of both types are provided, but are not all visible in this figure because of the perspective. It can also have a large central engine (Not shown) for combustion of liquid or powder propellants for the transfer of orbit injection into geostationary orbit.

0025In this figure 1 shows other equipment deployed such as solar panels 18 and the large antenna reflectors 20. In the example of Figure 1, the large reflectors 20 are facing the primary reflectors 22, themselves facing primary antenna sources.

0026To retain this figure 1, for a better appreciation of the invention, are the different natures of the two faces of each deployable radiator. In Indeed, each deployable radiator 28, 30, 32, 34 of this example of the prior art has a first radiant face shown by reference numeral "R" and a second face opposite insulating indicated by reference numeral "I" following the reference numeral (28, 30, 32, 34) of each radiator, for example 28R, 30R, 321, 341. The properties radiating and insulating are obtained in conventional manner using coatings well known to those skilled in the art, for example multiple layers Insulation (MLI, English acronym for "multi-layer insulation") on the face insulation; and radiant and reflective coatings on the radiating faces (OSR, English acronym for "optical solar reflector").

0027The spacecraft 10 of the prior art taught by document D2 and shown in Figure 2 comprises a structure 12 having solar panels 18 and deployable radiators 48 which are connected to the structure 12 of the instrument 10 by a deployment mechanism having a hinge 29 and an axis of hinge, said axis being positioned with a non-zero angle relative to the axes main structure 12 of device 10. As shown in FIG 2, both sides of the deployable radiators are radiating and are reported such as by the "R" mark on both sides, following the mark 48 for radiators. Once deployed, the main plane of the radiators is in the plane of faces North or South facing fixed satellite, the main axis of the radiator having a non-zero angle to the axis satellite-earth.

00283a shows schematically and in plan a first embodiment according to the invention of a deployable radiator 5 to spacecraft 10 in position stored. The antenna reflectors 2 are folded against the body of the satellite, on signs east and west. The antenna sources 3 are visible at the top of the figure, which is the land side by convention. The deployable radiators 5 are folded, in this example, against the north and south faces of the satellite 10. In the embodiment of these Figures 3a and 3b, these deployable radiators are partially hidden by solar panels 18, themselves folded against the satellite in the storage position. Some of these deployable radiators 5, which is not hidden by another deployable equipment is insulating and this is indicated by numeral "I" following the mark "5" for "insulation". The part hidden, in turn, is radiant, but not visible in this drawing position storage.

0029Figure 3b shows the same device as Figure 3a, but in position deployed. We see the solar panels 18 according to their edges in this view. When normal is almost perpendicular to the panels to land management - satellite, radiators 5 can be deployed. It is seen that these radiators have a fully radiant face, bearing the reference "R" following the landmark "5".

0030We can easily understand the main advantage of the invention over a Prior art Figure 1, because according to the teaching of D1, for radiator surface deployable given, half of the surface of said radiator is insulating, so only half is radiant; while a radiator deployable according to the invention can have a radiating surface much more greater than the half. The radiator of the invention is therefore more efficient, has equal surface to remove heat the radiator D1.

0031We can easily understand the main advantage of the invention over a Prior art Figure 2 because the deployable radiators of document D2 are fully radiating on both sides. This means that during the phase launch and transfer, the machine is exposed to the cold of space without insulation. Part of the energy of edge will therefore be devoted to heating the machine during its transfer path, the expense of the energy necessary for the implementation of some among other propulsion means needed to transfer.

0032The invention therefore optimizes energy use simultaneously board during the transfer phase, and the radiating capacity radiators made once the machine is in position, and for the duration of his mission.

0033Figures 4a and 4b diagrammatically show the same satellite that Figures 3a and 3b, but in perspective. In Fig 4a, we look at the north face satellite, and in Figure 4b, seeing the face is. We see the same elements in the two figures: the antenna sources 3, reflector 2 antennas, fixed radiators 4R, and deployable radiators 5 with a first fully radiative face 5R, which will be folded back against the fixed radiator 4R in the stored position, and a second face of which a part 51 is insulating, and A second important part 5R is radiant heat. When phase transfer orbit, the second radiant 5R part of the second face said radiator is hidden by another deployable equipment, e.g. by solar panels (not shown), or by reflectors deployable antennas, or any other equipment that would not need to be deployed during the transfer phase.

00345 shows another embodiment of a radiator deployable according to the invention, in the stored position. Here a radiator more greater than in the previous figures but having the same characteristics A first radiative face, not visible in this drawing, is pressed against the face Northern communication module 16. The second face of the radiator comprises a first thermally insulating part 51 and a second portion thermally radiating 5R, this second radiating portion being substantially hidden by other deployable elements (not shown) shielding during the craft transfer path. The radiator is connected to the communication module 16 by a hinge 6 with a hinge axis. As shown in the drawing, the hinge axis is positioned with a non-zero angle relative to the main axes of the machine. The non-zero angle allows positioning the radiator, in deployed position, outside the field of view of sensors, outside the paths of the chemical or ion jet propellants, and optionally, outside the radiation field of antennas, their Reflectors and their sources. This arrangement is shown more explicitly in the following figures.

0035Figure 6 shows schematically and in plan the deployable radiator 5R of FIG 5, in the deployed position. The hinge 6 is mounted on the module communication 16, and the axis of the hinge 6 is positioned at an angle non-zero relative to the main axes of the machine. The field of view of a sensor is shown in the drawing with the marker 11. This is such a position sensor, such as a star sensor. It is seen that the positioning the hinge axis with a non-zero angle relative to the principal axes of the apparatus allows the deployment of the radiator outside the field of view 11 of position sensor. Deployment bringing the main plane of a radiator the placing position in a non-collinear with still shots of faces North South of the satellite also contributes to this spot.

0036Furthermore, to facilitate industrialization in many gear having deployable radiators according to the invention, it is desirable to designing the vehicle with the modularity of these components, for example a module service 14, and a communication module 16. In the case the most often met, the heat sources on board will be concentrated in the module Communication 16; so it makes sense to plan to fix the hinge 6 of this communication module 16, and to provide for the conduction of heat from the communication module to the deployable radiator via one or more Flexible heat pipes or fluid loops (not shown) at the hinge 6.

0037In another embodiment, not shown, may be preferred to fix the hinge 6 of the service module 14 which may be of the same design for a variety of different communication modules in a range satellites.

00387 schematically shows a detail of deployable radiators, in the deployed position, the anti-earth side of the machine. Here we see, in the foreground, 5R radiant face of a first deployable radiator around the hinges 6 of the opposite east or west of the vehicle, as in the example of Figure 6, with an axis hinges positioned with a non-zero angle relative to the main axes of the machine, and a part of a second deployable radiator fixed on the face opposite of the craft (north or south respectively) the second radiator having an insulating portion and a radiating portion 5l 5R, as seen in this Figure 7. This figure also shows the envelopes paths 23 and 25 effluent chemical jets of propellant 24. It is noted that the form 5R radiator, as well as the positioning of the latter in deployed position relative to the machine, avoids the reception on the radiator product combustion.

0039In this figure 7, the jet 23 is represented by volume, and one sees his geometrical relationship with the radiator. This jet can be completely avoided by radiator so that a main plane of said radiator in the deployed position, which is not parallel with a main plane of said vehicle, as shown in FIG 7.

00408a and 8b show schematically and in perspective, respectively for north face and is facing a satellite with radiators deployable as shown in Figures 5, 6, and 7, with radiators deployed position. All the pins are identical with those of Figures above, the detailed description will not be repeated here. We notice especially in Figure 8b the main planes of the radiator in position deployed are parallel to the main planes of the machine, for the reasons already mentioned above.

0041Figure 9 show the evolution of an instrument 10 such as that of Figures 8a and 8b in its orbit around the earth O E. We see that the sun S is not in the geostationary orbital plane O of the instrument 10 around the earth E. This will result that the angle of the sunlight falling on the satellite 10 therefore these 5R radiators varies depending on the position of the machine 10 of orbit O and season.

0042The effectiveness of a radiator is much greater than its sunshine, and thus the absorption of solar energy is low. Absorption of solar energy varies depending on the angle of incident solar rays with the radiating surface 4R, 5R. This variation may be expressed as a function ns scalar product of a normal vector n to the surface radiant and 5R vector s parallel to sunlight. With an engine of a radiator deployable around at least one axis of rotation, one can minimize this product scalar ns and minimize the absorption of solar energy by radiator therefore optimize heat rejection efficiency. With an engine of a deployable radiator around at least two axes, one can ensure that the scalar product ns is kept zero, for maximum efficiency of the radiator.

004310 shows a mechanism allowing rotation around two axes. This kind of mechanism is well known to the skilled artisan. He understands a first fixing lug bottom 41 containing an annular motor (not shown) for obtaining a rotation about a first axis 42 by a second intermediate leg 43 containing a second annular motor (not shown) for obtaining a second rotation about a second axis 44 a third tab interface with equipment to turn. The wiring 46 can convey telemetry signals of the position of the two motors, and the necessary food for their rotation. Such mechanism two axes arranged between the satellite structure 10 and the radiators dépoyables or in the deployable radiator (deployment arm interface - panels radiative) provides a scalar product ns zero over the entire orbit O.

0044The invention described with the aid of these several embodiments and their explanatory figures is not limited to these specific examples. By example, the shape of the radiator is shown in the shape of an "L", but could easily imagine other shapes for optimal arrangement on the machine, for example a triangular or other shape. The arrangement of radiators was illustrated on the banks of the satellite structure, or on the face anti-land in the deployed position, but one can easily imagine the position elsewhere on the craft structure. The examples given relate mainly satellites, but may also concern the probes interplanetary manned flights, or any other spacecraft requiring additional capacity of thermal discharge.