Abstract:
An antenna reflector for a spacecraft is elastically deformable so as to be capable of changing from a folded position, when inside the spacecraft, to a deployed position, outside the spacecraft, at least partly under the action of its own elasticity. The reflector is separated into at least two parts by first opposing slot edges forming a crossing slot, so that in the folded position, the first opposing edges move with respect to each other so as to allow a protruding peripheral part of the reflector to be arranged closer to a longitudinal end of the spacecraft with damaging any portion of the reflector. In one embodiment, the two parts may be further separated by a radial slot which is arranged substantially orthogonal to the crossing slot.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     The present invention relates to an elastically deformable antenna reflector for a spacecraft, such as an artificial satellite or space probe. 
     2. Description of the Related Art 
     It is known that the items of equipment, such as the antenna, solar panels, etc, associated with a spacecraft have to be able to be folded in order to be able to be accommodated in a launch vehicle (rocket, shuttle) and to be deployed after ejection out of said launch vehicle, so as to take up their operational configuration. 
     It is known, moreover, that such equipment has already been produced in such a way that it is elastically deformable, this equipment then being able to take up either a deployed state, or a folded state, elastically deformed. By way of example, mention may be made of: 
     the U.S. Pat. No. 3,521,290, which describes an antenna reflector in a single piece of an elastically deformable material provided with a rigid central base to which are linked a plurality of radial ribs integral with the convex face of said reflector and elastically articulated to said central base. Thus, said antenna reflector can take up a position folded into the shape of a tulip, which does not risk entailing a permanent deformation of said reflector, and the change from the folded position into the deployed position in the shape of a concave disk can be carried out under the action of the elastic energy stored during the folding of the antenna structure. Controllable retaining means, consisting of a belt with pyrotechnic bolts surrounding said folded reflector and arranged on the opposite side from said central base, are provided in order to hold said reflector and said radial ribs in the folded position under stress; 
     the U.S. Pat. No. 4,133,501, which describes a solar panel in a single piece for a spacecraft produced in an elastically deformable way in order to take up either a curved folded position under stress in which said solar panel matches the convex outer surface of said spacecraft, or a flat deployed position, clear of said outer surface, the change from the curved, folded position to the flat, deployed position being due to the elastic relaxation of said solar panel. In the curved folded position, the solar panel is held against the outer surface of said spacecraft by latches, carried by the spacecraft; and 
     the U.S. Pat. No. 4,926,181, which describes an antenna reflector in a single piece of an elastically deformable material, which can be rolled into a cylindrical shape and held in this shape by clamps. An underlying pliable structure can be deployed, in order to serve as a support on which said reflector can unroll and take up its deployed, operating shape, under the action of its elastic relaxation. 
     Moreover, as demonstrated, for example, by the U.S. Pat. No. 5,644,322, it is usual, for launching said spacecraft, to store it in an elongate casing, for example of cylindrical-conical shape, constituting, for example, the upper nose cone of the launch rocket, the reflector of the antenna or antennae of said spacecraft being arranged laterally with respect to the body of the latter in the peripheral space defined between said body and said casing. In this prior document, the antenna reflector consists of a central rigid base of large surface area surrounded by a peripheral frustoconical ring produced from an elastically deformable material. Hence, by virtue of such a structure, the size of the reflector, within said cylindrical-conical casing, can be slightly reduced by temporarily elastically deforming said peripheral ring, said reflector then taking up the shape, at least approximately, of a bowl laterally encasing said body. The reflector is kept in this bowl shape by a belt, loosening of which is controlled electrically and which surrounds said body and said reflector at the center of said base, this belt folding said elastically deformable ring down onto said body, bearing on two diametrally opposite points of said ring. After ejection in space, said reflector can resume its operating position, by removal of said belt and elastic return of said peripheral ring to its elastically relaxed, stable, deployed position. 
     It can easily be understood that, in such a device, the saving in size of said reflector in the folded position, by comparison with the deployed position, is limited. This is because, on the one hand, because of the large diameter of said rigid central base, the lateral compression of the reflector can be applied only to the peripheral ring, such that the saving in the lateral size is relatively small. Moreover, this lateral compression not only exerts no reducing action on the longitudinal dimension of said reflector, but also increases said dimension due to the fact that it entails the straightening of the upper part of said peripheral ring outwards. The longitudinal size of the reflector, in the folded position, is thus greater than that of its deployed position. Because of its dimensions, said reflector generally overshoots the upper longitudinal end of the body of said vehicle housed in the cylindrical part of the casing and has to be extended into the conical part thereof. This conical shape thus imposes a limitation on the diameter of the reflector. However, for obvious reasons of performance, it would be advantageous for the reflector to be able to have as large a diameter as necessary, and to match the convergent shape of the conical part of the casing. 
     Moreover, it will be noted that, in the folded position, the reflector of the U.S. Pat. No. 5,644,322 is not held firmly, such that it is subject to the vibrations induced during the launch. This can result in difficulties of dynamic balancing and of damping of the vibration of said reflector, and even damage to the reflector or to the surrounding objects. 
     SUMMARY OF THE INVENTION 
     The object of the present invention is to remedy these drawbacks, while making it possible to increase the dimensions of said antenna reflector. 
     To this end, according to the invention, the antenna reflector for a spacecraft having to be stored in a casing of cylindrical-conical shape that is elongate along an axis, in such a way that said reflector is arranged laterally with respect to the body of said spacecraft, in the peripheral space defined between said body and the cylindrical part of said casing, and that said reflector exhibits a peripheral part projecting longitudinally into the conical part of said casing from the longitudinal end of said body, said reflector being at least partly elastically deformable in such a way that: 
     outside said casing, said reflector can take up a stable, deployed state without elastic stress, corresponding to its functional shape; 
     within said casing, said reflector, by elastic folding around said axis of the casing, can take up a folded state allowing it laterally to envelop said body, said reflector being held in this folded state by virtue of controllable retaining means; and 
     the change by said reflector from its folded state to its deployed state being due at least in part to the release of the energy stored in said reflector when it was elastically folded in order to make it change from its deployed state to its folded state, 
     is noteworthy: 
     in that said reflector is separated into two parts by opposing slot edges; and 
     in that, in the folded position of said reflector, said opposing edges of said reflector parts move with respect to one another so as to bring said protruding peripheral part of the reflector closer to said longitudinal end of said body. 
     Moreover, in accordance with another important feature of the present invention, provision is made for said controllable retaining means, in the folded position of the reflector, to secure said reflector parts to the body of said spacecraft. 
     It is thus easily seen that, according to the present invention, the abovementioned problems of longitudinal size and of vibration are solved. 
     The arrangement of said the slot edges with respect to the reflector can be varied, provided that it allows the relative movement of said edges entailing the bringing of the protruding peripheral part of the reflector closer to the longitudinal end of the body of the spacecraft. 
     However, two particular arrangements, capable, moreover, of being taken in combination, are especially advantageous. 
     According to the first arrangement, said opposing slot edges delimit a slot crossing said reflector from one edge to the other, the general direction of said slot being at least approximately orthogonal to said axis of the casing. Said reflector parts may be linked to one another near the middle of said slot and said opposing edges of said reflector parts may overlap so as to bring said protruding peripheral part of the reflector closer to said longitudinal end of said body. Said slot may be diametral, such that said reflector parts each consist of a half of said reflector, said reflector halves being integral with one another near the center of said reflector. In a variant, said slot may be off-center with respect to said reflector, such that said reflector parts are unequal, said reflector parts being linked to one another by means of an articulation. According to yet another variant, said reflector parts are linked to one another near the periphery of said reflector, for example by articulations (ball joints, pin joints, elastic parts, etc) arranged at the ends of said slot, and said opposing edges of said reflector parts move apart from one another in order to bring said protruding peripheral part of the reflector closer to said longitudinal end of said body. Here again, the reflector parts bounded by said slot may be equal or unequal. 
     According to the second arrangement (which may supplement the first or be supplemented by it), said opposing slot edges delimit a radial slot, the general direction of said radial slot being at least approximately parallel to said axis of the casing. 
     It will be noted that: 
     the U.S. Pat. No. 3,176,303 already describes an antenna reflector comprising a rigid central base, from which a plurality of petals, in the form of sectors, extend radially, constituting the reflector of the antenna, each of said petals being produced from an elastically deformable rigid material, in order to be able to take up either a bent position under stress, which corresponds to the folded storage position of said reflector, or a position spontaneously deployed in the form of a concave disk, which corresponds to the deployed operating position of said reflector, the change from the folded position to the deployed position being due solely to the elastic relaxation of said petals, previously elastically bent. In the folded position, the opposing edges of adjacent petals overlap in pairs, and controllable retaining means, of the peripheral belt type, are provided to keep said petals together in the folded position of the reflector. 
     In such a known antenna reflector, the number of petals is very high, such that, in the folded position, said reflector exhibits the shape of a tulip, which prevents it being arranged laterally with respect to the body of the spacecraft, in the peripheral space defined between said body and said casing and of laterally enveloping said body. 
     Hence the U.S. Pat. No. 3,176,303 is completely foreign to the technical problems of longitudinal size and of vibration, which are solved by the present invention; 
     the U.S. Pat. No. 5,574,472 and the patent EP-A-0 534 110 describe an antenna reflector in a single piece of an elastically deformable material, which can take up a folded position in the form of a bowl by virtue of a controllable-breaking tension link arranged between two diametrally opposed points on the periphery on said reflector. It will be noted that, in this bowl-shaped folded position, the upper peripheral edge of the reflector, protruding with respect to the body of the spacecraft, is straightened outwards and cannot therefore be accommodated in the conical part of the casing. 
     Hence it will be noted that these latter two patents cannot suggest the present invention. Moreover, it will be noted that said tension link constitutes an obstacle, or at the very least an impediment, for arranging the body of the spacecraft in the concavity of the reflector in the folded position, and that the production of said reflector in a single piece allows neither precise control of the shape of the reflector in the folded position, nor optimal enveloping of the body of the spacecraft. 
     By comparison with the prior art reviewed above, it is thus seen that: 
     the first special feature of the present invention consisting in producing said elastically deformable reflector in a limited number of parts which can be partially superposed along their opposing edges makes it possible: 
     to increase the dimensions of said reflector; 
     to increase the accommodation capacity of said casing; 
     better to control the shape of the reflector in the folded position, by concentrating the major part of the elastic deformation in limited areas; 
     the second special feature of the present invention consisting in fixing said reflector in the folded position on the body of said spacecraft makes it possible: 
     better to control the vibration of said reflector in the folded position; 
     better to control the shape of said reflector in the folded position; 
     to use retaining mechanisms known for other uses. 
     In the reflector in accordance with the present invention, said controllable retaining means may secure said opposing edges together and fix them onto the body of said spacecraft. They may, in a variant, secure the reflector parts independently of one another onto the body of said spacecraft. 
     It will be noted that, in the case in which the two arrangements of slot edges mentioned above are combined, the possibility of tightening said reflector about the body of said spacecraft is enhanced, with the spacecraft being better enveloped. It is then advantageous to provide floating controllable retaining means in order to keep the edges of said radial slot together, in an overlapping position. 
     Moreover, in order to maintain the relative positions of said reflector parts in the deployed position of the reflector, linking means are provided allowing relative movement of said edges in the folded position of said reflector, but linking said parts together when said reflector is in the deployed position. In the case of edges which are capable of overlapping, such linking means may comprise self-rigid tapes, which however are flexible in compression, arranged transversely to said slots and anchored to said reflector, on either side of the slots. In the case of edges which are capable of being spaced apart, such linking means may include at least one tension spring. 
    
    
     The Figures of the attached drawing will give a good understanding of how the invention can be produced. In these Figures, identical references designate similar elements. 
     BRIEF DESCRIPTION OF THE DRAWINGS 
     FIG. 1 is a view in diagrammatic perspective, from the rear, of an embodiment of the antenna reflector in accordance with the present invention, in the deployed position. 
     FIG. 2 is a view in diagrammatic perspective, from the rear, of the embodiment of the antenna reflector of FIG. 1, in the folded position. 
     FIGS. 3A and 3B, in views corresponding respectively to FIGS. 1 and 2, illustrate a flexible linking device preventing relative movement of the reflector halves in the deployed position. 
     FIG. 4 diagrammatically shows the reflector in accordance with FIGS. 1 and 2 arranged around a satellite, under the nose cone of a launcher. 
     FIGS. 5A and 5B illustrate, in its locked and unlocked positions respectively, a device for retaining said reflector of FIG. 4 on the body of said satellite, along the line V—V of this latter Figure. 
     FIG. 6, in a view similar to that of FIG. 4, shows a variant arrangement of the reflector under the nose cone of the launcher. 
     FIG. 7 shows, in the locked position, a device for retaining the reflector of FIG. 6 on the body of said satellite, along the line VII—VII of this latter Figure. 
     FIG. 8, in a view similar to FIG. 1, shows a variant embodiment of the antenna reflector in accordance with the present invention, in the deployed position. 
     FIG. 9 shows a flying retaining device for the antenna reflector of FIG.  8 . 
     FIGS. 10 and 11 are diagrammatic views in perspective, from the rear, of embodiment variants of the antenna reflector in accordance with the present invention, in the deployed position. 
     FIG. 12 is a diagrammatic view in perspective, from the front, of the variant embodiment of FIG. 11, in position enveloping a satellite. 
     FIG. 13 illustrates an articulation of the peripheral edges of the two parts of the reflector of FIGS.  11  and  12 . 
    
    
     DESCRIPTION OF THE PREFERRED EMBODIMENTS 
     The antenna reflector  1 , in accordance with the present invention and illustrated diagrammatically in FIGS. 1 and 2, exhibits the shape, at least approximately, of a concave disk and is formed of two halves  1 A and  1 B, separated from one another by a diametral slot  2  crossing said disk from one edge to the other. At the center of the reflector  1  a rigid base  3  is provided, linked on the rear side—that is to say on the convex side of said reflector—to a linking arm  4 , the end of which remote from said base  3  is intended to be articulated, in a known way which is not represented, to the body of a spacecraft. In the example represented in FIG. 1, the linking arm  4  is orthogonal to said diametral slot  2 , when the reflector  1  is deployed. 
     Each half  1 A and  1 B of the reflector  1  is produced from an elastically deformable material, for example from a carbon fiber fabric. If appropriate, stiffening rings (not represented) are arranged on the convex rear face of said reflector  1 . 
     In the deployed position of the reflector (FIG.  1 ), the diametral slot  2  is delimited by edges  5 A,  5 B respectively, and the relative positions of said edges  5 A and  5 B are maintained by flexible ties  6  transverse to said slot. 
     As FIG. 3A shows in more detail, each tie  6 , which consists for example of a flexible elastic strip of curved cross-section in the manner of the tape of a measuring tape, is anchored at its ends, respectively to the half of the reflector  1 A by a stud  7  and to the half of the reflector  1 B by a stud  8 . 
     Moreover, on either side of the slot  2 , eyelets  9  and  10  are provided in the halves of the reflector  1 A and  1 B, respectively. 
     Hence, as FIGS. 2 and 3B show, the reflector  1  can take up a position folded about the base  3  with discontinuity of curvature, for which the edges  5 A and  5 B overlap in an overlapping region  11 , which widens out from the center towards the periphery of said reflector. In this folded position, the ties  6  are relaxed and the eyelets  9  and  10  are superposed. 
     As FIG. 4 diagrammatically illustrates, the reflector  1  can be stored in an elongate casing  12  with longitudinal axis X—X, for example the nose cone of a space launcher including a cylindrical part  12 A surmounted by a conical part  12 B, the reflector  1  being arranged in the peripheral lateral space  13  delimited between the body  14  of a satellite and the cylindrical part  12 A of said casing  12 . As is usual (not visible in FIG.  4 ), the reflector  1  is linked to the satellite body  14  by the arm  4 , which is articulated to the lower part of said body. 
     In this storage position, the edges  5 A and  5 B of said reflector halves  1 A and  1 B overlap (region  11 ) and said reflector halves are moreover elastically folded about the axis X—X orthogonal to the slot  2 , so as to envelop the body  14 . 
     By virtue of the breaking of the reflector about the slot  2 , it is seen that the upper part  1 S of said reflector  1 , protruding by comparison with the satellite body  14 , is brought towards the center of the casing  12  and the upper face  14 S of said body, and can thus be housed in the conical part  12 B of the casing. 
     In the storage position of FIG. 4, the reflector  1  is, moreover, held by pyrotechnic studs  15  passing through the opposing eyelets  9  and  10  of the reflector halves  1 A and  1 B (see FIG. 5A) and integral with the body  14  of the satellite. 
     Hence, during the launch of the satellite  14 , the reflector is in the nose cone  12 , as represented in FIG. 4, held rigidly in its folded shape. After said nose cone  12  is jettisoned, and the satellite  14  is ejected, the pyrotechnic studs  15  are activated and they release the reflector halves  1 A and  1 B from the body of the satellite  14  (see FIG.  5 B). Next, under the action of its own elasticity, the reflector  1  relaxes so as to spontaneously to take up its deployed state of FIG. 1, the arm  4  tilting (in a known way which is not represented) so as to free said reflector from the body of the satellite  14 . 
     In FIG. 6, a variant for holding the reflector  1  in the folded position has been represented. In this case (see also FIG.  7 ), the reflector halves  1 A and  1 B are held integral with the body of the satellite  14  by individual pyrotechnic studs  16  which are integral with the body and pass through eyelets  17 . The individual pyrotechnic studs  16  are then outside the overlap region  11 . 
     FIG. 8 shows a variant embodiment 1.1 of the reflector  1  in accordance with the present invention, which is capable of fitting even better into the conical part  12 B of the nose cone  12 . In this variant, the reflector half  1 A is itself separated into two separate parts  1 A 1  and  1 A 2  by a radial slot  18 , along opposing edges  19 A 1  and  19 A 2  capable of overlapping in the manner of the edges  5 A and  5 B. In the folded position, this embodiment variant can thus even further envelop the body of the satellite  14 . 
     Linking means  6 ,  7 ,  8  are also provided for holding the edges  19 B 1  and  19 B 2  in position. Moreover, said parts  1 A 1  and  1 A 2  include respective eyelets  20  and  21  coming opposite one another when said edges  19 A 1  and  19 A 2  overlap (see FIG. 9) so as to be able to be traversed by a pyrotechnic stud  22 . The pyrotechnic stud  22  may be floating, that is to say not linked to the body of the satellite  14 . 
     Clearly, the pyrotechnic stud  22  is also activated when the reflector  1  of FIG. 8 is deployed. 
     In the variant embodiment 1.2 shown diagrammatically by FIG. 10, the concave disk of the reflector is formed by two unequal parts  23 A and  23 B, separated from one another by a non-diametral slot  24  crossing said disk from one edge to the other, off center with respect to said disk and arranged orthogonally to the axis X—X. The large part  23 B is linked to the rigid base  3 , which is itself carried by the arm  4 . The small part  23 A is provided with a rigid median base  25 , linked to the rigid base  3  by an articulation  26 . It can easily be envisaged that, by virtue of the slot  24  and of the articulation  26 , the embodiment 1.2 can take up a folded position similar to that shown by FIGS. 2,  4  and  6  in the case of embodiment 1, the only difference here being that the slot  24  is not diametral and that the parts  23 A and  23 B are unequal. 
     In the variant embodiment 1.3 shown diagrammatically by FIGS. 11 and 12, just as in variant 1.2 of FIG. 10, a non-diametral slot  27  crossing the concave disk of the reflector from one edge to the other divides it into two unequal parts  28 A and  28 B, the larger of which is linked to the rigid base  3  and to the articulated arm  4 . Said reflector parts  28 A and  28 B are linked to one another close to the periphery of said reflector by articulations  29 ,  30 , so that the opposing edges of the slot  27  can move apart from one another (see reference  31  in FIG. 12) so as to bring the protruding peripheral part  1 S of the reflector  1 . 3  back in the direction of the longitudinal end  14 S of the body  14 . A tension-spring device  32 , the action of which opposes the opening of the slot  27 , makes it possible to fix the parts  28 A and  28 B rigidly together in the deployed position. 
     As FIG. 13 shows, each of the peripheral articulations  29 ,  30  can be formed by a clevis provided with a rotation pin  33 . 
     Although only one single reflector  1 ,  1 . 1 ,  1 . 2  or  1 . 3  associated with the body of the satellite  14  has been mentioned above, it goes without saying that the present invention is not limited to such an embodiment. In fact, in the usual way, two such opposed reflectors can be combined jointly on said satellite body  14 , as is illustrated, for example, in the U.S. Pat. No. 5,644,322. Corresponding pyrotechnic studs  15 ,  16 ,  22  are then provided.