A solar wing and a spacecraft comprising such solar wing

Abstract

A solar wing (1) for a spacecraft (100), the solar wing (1) being configured to be movable between a stowed position and a deployed position and comprising: − a solar array (10) comprising a plurality of solar panels (2) comprising solar cells for generating electrical current and configured to be movable between the stowed position and the deployed position, the solar panels being arranged adjacent to one another and along a longitudinal axis (X) in the deployed position, − a plurality of stiffening assemblies (3) extending in at least one line (L1 or L2), each stiffening assembly (3) comprising at least one stiffening panel (30), said stiffening panel (30) of each stiffening assembly (3) being mechanically and electrically linked to one of said solar panels (2), each stiffening assembly (3) being configured to be movable between the stowed position and the deployed position, each stiffening panel (30) forming a predetermined non-zero and non-planar angle (A) with said solar panel (2) in the deployed position, said stiffening assemblies (3) of said at least one line comprising a plurality of mechanical and electrical links (4) in the form of electrically conductive C-shaped tape springs (5), at least one pair of mechanical and electrical links (4) being fastened to adjacent stiffening panels (30) and extending therebetween, at least some of the mechanical and electrical links (4) extending parallel to said line in the deployed position, each mechanical and electrical link (4) being folded on itself in the stowed position, the mechanical and electrical links (4) being configured to transmit electrical current produced by the solar cells towards a spacecraft body (101).

Description

DescriptionTitle: A solar wing and a spacecraft comprising such solar wing Technical Field

[0001] This disclosure pertains to the field of spacecrafts such as satellites. More precisely, it concerns a solar wing for generating electricity to power the spacecraft and a spacecraft comprising such a solar wing.Background Art

[0002] In order to produce electricity for use in the spacecrafts such as satellites, it is known to provide solar array structures, that can be in a stowed condition in the launch vehicle, and that can be deployed for use in a deployed configuration.

[0003] US 4787580 provides a solar array structure for use with satellites requiring large areas of solar arrays. Each array is a single continuous structure having continuous longitudinal members which in a stowed condition enable the panel to be wrapped around the satellite within its launch vehicle. In a deployed configuration, the longitudinal members are reconfigured to provide substantially increased stiffness in the direction of winding.

[0004] EP 1 043228 discloses, in a solar panel, a first support member and a second support member overlying the first support member. The second support member is flexible whereas the first support member has a stiffness. There are also biasing means carried by the first support member for pushing opposite ends of the second support member toward one another to outwardly bow the second support member.

[0005] WO 2011 / 109436 discloses a deployable structure that may include a slit-tube longeron and a flat panel coupled with the slit-tube longeron. The slit-tube longeron may include a tubular member having a slit that runs along the longitudinal length of the slit-tube longeron. The deployable structure may be configured to couple with a satellite and to transform between a stowed state and a deployed state. The tubular member is substantially straight when the deployable structure is in the deployed state, and the tubular member is wrapped around the satellite when the deployable structure is in the stowed state.

[0006] US 2003 / 164186 relates to an apparatus and method for the design and manufacture of foldable integrated device array stiffeners.

[0007] EP 0884241 relates to a solar panel assembly.

[0008] US 2021 / 320619 relates to a stacked solar array.

[0009] US 2017 / 175806 relates to a hinge assembly for a space structure.

[0010] There is a need to benefit from a solar wing that is simpler than the prior art.Summary

[0011] This disclosure improves the situation.

[0012] It is proposed a solar wing for a spacecraft, the solar wing being configured to be movable between a stowed position and a deployed position and comprising a solar array comprising :- a plurality of solar panels comprising solar cells for generating electrical current and configured to be movable between the stowed position and the deployed position, the solar panels being arranged adjacent to one another and along a longitudinal axis in the deployed position,- a plurality of stiffening assemblies extending, in the deployed position, in at least one line, each stiffening assembly comprising at least one stiffening panel,characterized in that each stiffening panel is mechanically and electrically linked to one of said solar panels by at least two C-shaped tape springs placed between the solar panel and the stiffening panel for forming a hinge and protruding on said solar panel in the deployed position,stiffening panels in the same line being, successively, mechanically and electrically linked to one another by at least two C-shaped tape springs fastened to adjacent stiffening panels and extending therebetween parallel to said line in the deployed position and being folded on itself in the stowed position,each stiffening assembly being configured to be movable between the stowed position and the deployed position, each stiffening panel forming a predetermined non-zero and non-planar angle with said solar panel in the deployed position,C-shaped tape springs mechanically and electrically linked to stiffening panels being configured to transmit electrical current produced by the solar cells towards a spacecraft body.

[0013] The solar wing provides a self-stiffening system providing rigidity in the deployed position and a small storage volume in the stowed position, in launch configuration. This makes it possible to have large surfaces of solar cells.

[0014] The following features can be optionally implemented, separately or in combination one with the others:

[0015] The predetermined non-zero angle may be comprised between 80° and 100°. The angle optimized value may be 90°. The stiffening panel is therefore opened by 90° from its stowed position. Nevertheless, this opening angle can be modified. The rear stiffeners opening angle could be between [-alpha, + alpha], alpha being determined in accordance with system consideration.

[0016] The pair of C-shaped tape springs of the mechanical and electrical link may be configured to have one end portion and one opposite end portion. Said one end portion may be configured to rotate relative to the opposite end portion between the stowed position and the deployed position with a rotating angle of about 180°.

[0017] The solar wing may comprise an electrical wiring configured to transmit the electrical current. The electrical wiring may comprise:- a plurality of C-shaped tape springs extending between the stiffening assemblies, parallel to said line in the deployed position,- a plurality of C-shaped tape springs extending between the solar panels and the stiffening assemblies and- a harness within the stiffening assemblies.

[0018] The harness may allow to electrically link the plurality of C-shaped tape springs extending between the stiffening assemblies and the plurality of C-shaped tape springs extending between the solar panels and the stiffening assemblies. The solar panels may comprise a wiring electrically connecting the solar cells to the plurality of C-shaped tape springs extending between the solar panels and the stiffening assemblies.

[0019] The solar wing may comprise a yoke and a pressure plate. The yoke may be configured to be fastened at one proximal end to the spacecraft body and fastened at one distal end to the pressure plate. The pressure plate may extend adjacent to one of the solar panels along the longitudinal axis in the deployed position.

[0020] The plurality of stiffening assemblies may comprise a stiffening assembly comprising a stiffening panel fastened to said pressure plate and a stiffening assembly comprising a stiffening panel fastened to said yoke. The stiffening assembly of the pressure plate and the stiffening assembly of the yoke may be configured to be movable between the stowed position and the deployed position. A plane including said stiffening panels may form the predetermined non-zero angle with the pressure plate and the yoke, respectively, in the deployed position.

[0021] In such case, the electrical wiring may also be provided between the stiffening assemblies so as to transfer electrical current to the spacecraft body.

[0022] The role of the electrical and mechanical links is not only a stiffening role but also a role of participating to the transfer of electrical current from the solar panels to the spacecraft body.

[0023] The solar panels may comprise a front face supporting the solar cells and a rear face that is opposite to the front face. The plurality of stiffening assemblies may be fastened to the rear face of the solar panels.

[0024] Each stiffening assembly may comprise a frame having elongated stiffeners. The stiffening panel may be in the form of said frame having elongated stiffeners. The stiffening frame is for example made up of an assembly of long beams.

[0025] The plurality of stiffening assemblies may extend in a single line, the line being parallel to the longitudinal axis.

[0026] The stiffening assemblies once deployed shall be aligned in order to provide the correct rigidity.

[0027] All the stiffening assemblies may be aligned in a single line. In such case the single line may be positioned at the center of the solar array.

[0028] The plurality of stiffening assemblies may extend in a plurality of lines. The plurality of lines may comprise two lines that are substantially parallel to one another, preferably extending at lateral ends of the solar panels. In another example, the plurality of lines comprises for example two lines diverging away from the spacecraft body in two diverging directions. In another example, the plurality of lines comprises for example two lines converging away from the spacecraft body in two converging directions.

[0029] Two lines of stiffening assemblies provide more rigidity to the solar array. They form stiffening lines. For larger solar arrays, three or more stiffening lines could be implemented if necessary. Lines of stiffening assemblies are for example parallel or not. For example, the lines of stiffening assemblies may diverge.

[0030] The solar panels may be arranged in a Z configuration in the stowed position, with the pressure plate surrounding the solar panels and being distant from the spacecraft body and with the solar panel that is at a free end of the solar wing being adjacent to the spacecraft body in said stowed position.

[0031] In such case, the solar wing may comprise spacers extending between the solar panels in the stowed position so that the solar panels are not directly in contact to one another in the stowed position.

[0032] In the stowed position, the solar panels may be in a preloaded state. All the C shape springs in folded configuration are for example preloaded, in order to provide the driving force to deploy the solar array. In deployed configuration, the C shape springs are no more preloaded. The energy released by the C shape springs from the folded to the deployed configuration represents the driving energy.

[0033] In another aspect, it is proposed a spacecraft, notably satellite, comprising a spacecraft body and a solar wing as recited above.Brief Description of Drawings

[0034] Other features, details and advantages will be shown in the following detailed description and on the figures, on which:

[0035] [Fig. 1 ] is a schematic view in perspective of an example of a solar wing in a deployed position, connected to a spacecraft body.

[0036] [Fig. 2] is a partial schematic view in perspective of an example of mechanical and electrical link, in the form of a double C shape spring, used in a solar wing, in a folded or stowed position.

[0037] [Fig. 3] is a partial schematic view in perspective of the example of mechanical and electrical link of figure 2, in an unfolded or deployed position.

[0038] [Fig. 4] is a schematic view in perspective of an example of solar wing in the deployed position.

[0039] [Fig. 5] is a schematic view in perspective of the solar wing of figure 4 in a stowed position.

[0040] [Fig. 6] show schematic views in perspective of the solar wing of figures 3 and 4 in various positions, from the stowed position shown in figure 4, through intermediate positions, to the deployed position shown in figure 3.

[0041] [Fig. 7] is a schematic view in perspective of another example of solar wing in the deployed position.

[0042] [Fig. 8] is a schematic cross-sectional view of an example of solar wing in the stowed position.

[0043] [Fig. 9] is a detail IX of figure 8.

[0044] [Fig. 10] is a detail X of figure 8.

[0045] [Fig. 11] is a partial schematic view in perspective of an example of solar wing in the stowed position.

[0046] [Fig. 12] is a partial schematic view in perspective of an example of solar wing in the stowed position.

[0047] [Fig. 13] is a partial schematic view in perspective of an example of solar wing in the deployed position.

[0048] [Fig. 14] show schematic views in perspective of an example of spacecraft comprising the solar wing of figure 7 in various positions, from the deployed position shown in figure 7, through intermediate positions, to the stowed position.Description of Embodiments

[0049] In the various figures, the same references designate identical or similar elements. For simplicity’s sake, only the elements that are useful for understanding the described example are shown in the figures and are described in detail below.

[0050] In the following description, when referring to terms qualifying absolute position, such as the terms “front”, “rear”, “top”, “bottom”, “left”, “right”, etc., or relative ones, such as the terms “above”, “below”, “upper”, “lower”, etc., or to qualifiers of orientation, such as “horizontal”, “vertical”, etc., reference is made, unless otherwise specified, to the orientation of the figures.

[0051] It is now referred to figure 1 showing an example of a spacecraft 100, which is a satellite, comprising a spacecraft body 101 and a solar wing 1 . The solar wing 1 is configured to be movable between a stowed position and a deployed position, this deployed position being shown in figure 1 . The solar wing 1 comprises a solar array 10 comprising a plurality of solar panels 2, also known as solar array panels, comprising solar cells for generating electrical current and configured to be movable between the stowed position and the deployed position, the solar panels being arranged adjacent to one another and along a longitudinal axis X in the deployed position, as shown. The solar panels 2 each comprise a front face 21 supporting the solar cells and a rear face 20 that is opposite to the front face 21 . The solar panels 2 may be made with a sandwich frame or photovoltaic assembly.

[0052] As shown in figure 1 , the solar wing 1 comprises a plurality of stiffening assemblies 3 extending in at least one line L. The plurality of stiffening assemblies 3 extends for example in aplurality of lines L1 , L2. The plurality of lines L1 , L2 comprises for example two lines L1 , L2. The lines L1 and L2 are for example parallel to one another and to the longitudinal axis X. Two lines L1 and L2 extend at lateral end portions 25, 26 of the solar panels 2.

[0053] Each stiffening assembly 3 comprises at least one stiffening panel 30. The stiffening panels 30 of each stiffening assembly 3 are mechanically and electrically linked to one of the solar panels 2.

[0054] Each stiffening assembly 3 is configured to be movable between the stowed position and the deployed position. Each stiffening panel 30 forms a predetermined non-zero and non-planar angle A with said solar panel 2 in the deployed position as shown in figure 1 . The predetermined non-zero and non-planar angle A is preferably comprised between 80° and 100°.This angle A has for example an optimized value of 90°, making the stiffening panels open by 90° from the stowed position to the deployed position. Angle A corresponds to the opening angle. Other values for this angle A are possible. The opening angle A could be between [-alpha, + alpha], alpha being determined in accordance with system consideration.

[0055] The stiffening assemblies 3 for example in the form of two lines L1 and L2, comprise a plurality of mechanical and electrical links 4 in the form of electrically conductive C-shaped tape springs 5. At least one mechanical and electrical link 4 is fastened to adjacent stiffening panels 30 and extend therebetween, in each line. The mechanical and electrical links 4 may be fastened to adjacent edges of adjacent stiffening panels.

[0056] Alternatively, a solar panel comprises for example a plurality of stiffening assemblies in the form of a single line. At least two mechanical and electrical links 4 are fastened to adjacent stiffening panels 30 and extend therebetween, in the line.

[0057] As shown in figure 4, at least some of the mechanical and electrical links 4 extend parallel to said line L, i.e. to line L1 or line L2 in this example, in the deployed position.

[0058] Each mechanical and electrical link 4 is folded on itself in the stowed position. As shown in figures 2 and 3, the mechanical and electrical link 4 comprises a pair of C-shaped tape springs 5 which is configured to have one end portion 7 and one opposite end portion 8, said one end portion 7 being configured to rotate relative to the opposite end portion 8 between the stowed position and the deployed position with a rotating angle. The rotary angle is for example of 180 degrees for longitudinal C-shaped spring or 90 degrees for transverse C-shaped spring. The pair of C-shaped tape springs 5 tends to move into the deployed position, i.e. from the folded position shown in figure 2 to the folded position shown in figure 3.

[0059] The mechanical and electrical links 4 are also configured to transmit electrical current produced by the solar cells towards the spacecraft body 101 .

[0060] As shown in Figure 1 , the plurality of stiffening assemblies 3 is for example linked to the rear face 20 of the solar panels 2. The stiffening assemblies 3 extend from the rear face 20. Each stiffening panel 30 is for example linked to one solar panel 2 by one or several C-shaped tape spring(s). Each C-shaped tape spring is either in the form of said mechanical and electrical link or inthe form of a mechanical link. In the first case, mechanical connection is ensured, and also electrical current is transmitted. In the second case, only the mechanical connection is ensured.

[0061] As shown in figure 13, the C-shaped tape springs protrude for example on each solar panel 2 in the deployed position and constitute a hinge 6.

[0062] As shown in the figure 13, the stiffening panels 30 are for example connected to the solar panels 2 by hinges 6 and are connected to each other by mechanical and electrical links 4.

[0063] As shown in figure 1 , the solar wing 1 comprises, for example, a yoke 15 and a pressure plate 16. The yoke 15 is for example in the form of a U or a V or a Y. The yoke is for example hinged at one point with the body of the spacecraft. The yoke is also hinged to the solar wing or to the pressure plate placed between the solar wing and the yoke. The yoke is for example deployable. The yoke is for example orientable.

[0064] The pressure plate 16 extends adjacent to one of the solar panels 2 along the longitudinal axis X in the deployed position. The plurality of stiffening assemblies 3 comprises a stiffening assembly comprising a stiffening panel 31 fastened to said pressure plate 16 and a stiffening assembly comprising a stiffening panel 32 fastened to said yoke 15. The stiffening assembly 3 of the pressure plate 16 and the stiffening assembly 3 of the yoke 15 are configured to be movable between the stowed position and the deployed position, said deployed position being shown in figure 1. A plane including these stiffening panels 31 , 32 forms the predetermined non-zero and non-planar angle A with the plane of the pressure plate 16 and the yoke 15, respectively, in the deployed position. Thus, as shown in figure 1 , in the deployed position, the stiffening panels 30, 31 , 32 are aligned substantially in a same plane forming the angle A with the plane including the yoke 15, the pressure plate 16, and the solar panels 2.

[0065] The yoke 15 is for example fastened to the pressure plate 16 only in an indirect manner, firstly through hinges which connect the yoke 15 to the corresponding stiffening panel 32 and the pressure plate 16 to the corresponding stiffening panel 31 , and secondly through mechanical and electrical links fastened to the stiffening panels 31 and 32. Hinges and electrical and mechanical links between panels are for example realized by C-shaped tape springs.

[0066] Preferably adjacent solar panels are for example electrically and mechanically linked to each other only in an indirect manner, through hinges 6 that connect the solar panels 2 to corresponding stiffening panels 3, and through mechanical and electrical links 4 between adjacent stiffening panels.

[0067] As shown in figure 1 , one stiffening panel 31 linked to the yoke and one stiffening panel 32 linked to the pressure plate, are for example aligned with stiffening panels linked to solar panels, such that they deploy together.

[0068] The solar wing 1 is a self-stiffening structure providing a solar array rigidity in the deployed position using soft links formed by C-shaped tape springs. No synchronization system is needed in this solar wing 1 . Indeed, the C-shaped tape springs can be used as a function of the deployment torque that is needed.

[0069] The C-shaped tape springs 5 or the C-shaped tape springs forming the hinge 6 may comprise steel or carbon fiber skins.

[0070] As shown in figure 4, the solar wing 1 in the deployed position comprises, for example, . the stiffening panels 30 linked to lateral ends 27 and 28 of the solar panels 2. The solar wing 1 comprises for example four solar panels 2.

[0071] Figure 5 shows the solar wing 1 of figure 4 in the stowed position. The solar panels 2 are for example arranged in a Z configuration in the stowed position.

[0072] Figure 6 shows an example of transitional configurations of the solar wing 1 of figures 4 and 5 from the stowed position to the deployed position via intermediate positions, three of which are shown here.

[0073] In the stowed position, the solar panels 2 are for example in a preloaded state. The stiffening panels 30 are folded between the solar panels 2. The mechanical and electrical links 4 are for example protruding laterally, in the folded position.

[0074] During the deployment phase, after a release the mechanical and electrical links 4 deploy. As a consequence, the stiffening panels together with the corresponding solar panel deploy, while the stiffening panels are still folded against their corresponding solar panels 2.

[0075] Then, at the end of the deployment, the stiffening panels 30 are deployed, relative to the corresponding solar panel, thanks to the hinges 6 which open automatically, forming the opening A angle. As a consequence, the stiffening assemblies 3 provide stiffness, electrical link and automatic positioning of the solar array.

[0076] As shown in Figure 7 the solar wing 1 comprises for example, in the deployed position, ten solar panels 2 and one pressure plate 16. The stiffening assemblies comprise for example stiffening panels forming the orientation yoke 15. The stiffening assemblies comprise stiffening panels linked to solar panels. The stiffening assemblies comprise for example stiffening panels linked to the pressure plate.

[0077] As shown in Figure 7, the plurality of stiffening assemblies 3 is arranged for example in two lines L1 and L2, that are linked at the base of the yoke and that diverge away from each other at the end of the solar wing. The stiffening panels 30, 31 are fastened through hinges to the solar panels 2 or pressure plate 16 at different positions of the rear face 20 of the solar panels 2 or pressure plate 16, near the center thereof for the pressure plate 16 and near the lateral ends for the more distant solar panel, reference 2x, at the free end 29. The plurality of stiffening assemblies 30 extends for example in a plurality of lines L1 , L2, the plurality of lines L1 , L2 comprising two lines diverging away from the spacecraft body 101 in two diverging directions D1 , D2.

[0078] Alternatively, when using a yoke, the lines of stiffening panels are not parallel and converge relative to one another away from the spacecraft body.

[0079] Alternatively, the plurality of stiffening assemblies 30 extends in a single line. This line is for example parallel to the longitudinal axis X of the solar wing.

[0080] In a same line of stiffening assemblies, the C-shaped tape springs of the hinge are for example aligned in order to provide rigidity. In a same line of stiffening assemblies, the C-shaped tape springs between stiffening panels are for example aligned in order to provide rigidity. In a same line of stiffening assemblies, the stiffening panels are for example aligned in order to provide rigidity.

[0081] The number of lines is for example increased for setting the rigidity, according to the needs. Two lines of stiffeners provide for example more rigidity to the solar array than a single line. A single line is for example positioned at the center of the solar array. For larger solar arrays, three or more stiffening lines are for example positioned on the solar wing.

[0082] The solar wing power is for example be 3,5 kW. The performance of each solar panel 2 may be 110 W / m2, requiring a total surface of about 32 m2. The mass of the solar wing 1 may be under 50 kg. The mass of a solar panel 2 may be about 700 g / m2. The panel surface may be about 3 m2for each panel.

[0083] As shown in Figure 8, in the stowed position, the pressure plate 16 surrounds for example the solar panels 2 away from the spacecraft body 101 and the solar panel 2x that is at the free end 29 of the solar wing 1 is adjacent to the spacecraft body 101 .

[0084] As shown in Figure 8, the solar wing 1 comprises for example spacers 38 extending between the solar panels 2 in the stowed position so that the solar panels 2 are not directly in contact to one another in the stowed position. The spacers 38 may be every 200 mm, for example.

[0085] In the stowed position, the stiffening panels 30 are for example stowed within the volume of the solar panels 2. The solar panels 2 may be slightly in tension.

[0086] As shown in figures 9 and 10, in the stowed position, the solar panels 2 are for example superposed and the stiffening panels 30, 31 and 32 are interposed between the solar panels 2. The C-shaped tape springs forming the mechanical and electrical links 4 are for example protruding in the stowed position.

[0087] In the stowed position, the thickness of the yoke 15 may be equal to 20 mm, the thickness of the pressure plate 16 may be equal to 20 mm, the thickness of each solar panel 2 may be equal to 8 mm, the thickness of the stiffening panels 30 may be equal to 6 mm and the thickness of each C-shaped tape spring may be equal to 0,6 mm, so that the total thickness may be equal to 130 mm, for a configuration corresponding to figure 7.

[0088] As shown on figure 11 , the pressure plate 16 may form a grid with hollow parts 40. The hollow parts are for example square. The pressure plate is for example monolithic.

[0089] As shown in figure 13, each stiffening panel comprises for example a hollow frame 35 forming the stiffening panel 31 and comprising tubes or rods forming elongated stiffeners 36. The frame 35 comprises for example five elongated stiffeners 36, four being arranged as the edges of a rectangle and one forming a diagonal of the rectangle linking two opposite rectangle corners. A frame forming for example the stiffening panel at the end of the line is in the form of a triangle or a trapeze. Thelarge edge is for example facing the adjacent stiffening panel. The small edge of the trapeze is for example linked to the body of the spacecraft.

[0090] As shown in figure 13, the solar wing 1 comprises for example an electrical wiring 9 configured to transmit the electrical current. The electrical wiring 9 comprises for example a plurality of C-shaped tape springs 11 extending between the stiffening assemblies 3 parallel to the line or lines L1 , L2, for example, in the deployed position. The C-shaped tape springs 11 extend for example between the elongated stiffeners 36 of adjacent frames 35. The electrical wiring 9 also comprises for example a plurality of C-shaped tape springs 12 extending between the solar panels 2 and the stiffening assemblies 3 and a harness 13 within the stiffening assemblies 3.

[0091] As shown in figure 13, the link between a stiffening panel and the pressure plate is for example similar to the link between a stiffening panel and a solar panel, even if the pressure plate does not produce electricity, the electrically conductive function being not exploited with the pressure plate.

[0092] The C-shaped tape springs 11 , 12 provide for example mechanical fastening and rigidity, and also electrical link.

[0093] Each stiffening panel comprises for example a harness 13 for connecting some of the C-shaped tape springs together, in order to transfer the electrical current from the solar panels to the body of the spacecraft. The elongated stiffeners 36 are for example tubes comprising the harness.

[0094] All the links between the stiffening panels and the solar panels or between two adjacent stiffening panels comprise for example C-shaped tape springs. Some of the C-shaped tape springs, for example the mechanical and electrical links 4 and hinges 6, may be strong so as to ensure a mechanical stiffening and connection. Some of the C-shaped tape springs, for example the C-shaped tape springs 11 and the C-shaped tape springs 12, may be soft only to transfer electrical current.

[0095] A “soft” C shaped tape spring is for example made up of only one C shape spring. This “soft” C shape spring provides a small driving torque from folded to deployed configuration. The deployed final position is considered stable as long as the C shaped tape spring does not buckle due to local bending moment in the tape.

[0096] A “strong” C shaped tape is for example made up of an assembly of several soft C shaped tape springs facing each other. This “strong” C shaped tape provides a driving torque from the folded to the deployed configuration. The deployed final position is much more stable than with the soft spring and is considered stable as long as the C shaped tape does not buckle due to local bending moment in the tape, but this buckling moment threshold is much higher than in the soft spring configuration, thanks to the several C shaped tape springs and to the distance between them.

[0097] Soft or strong C shaped tape springs are chosen as a function of stability needs in the different parts of the Solar array, depending of loads induced during orbit maneuvers. A soft shaped tape spring represents for example one electrical line and a “strong” C shaped tape represents for example two electrical lines.

[0098] The C-shaped tape springs may be single, double, triple or more. A double C-shaped tape spring means that the link comprises a pair of C-shaped tape springs that are parallel to each other or face each other, and interact with each other within the link, as shown in figures 2 and 3.

[0099] A hinge 6 comprises for example a strong single C-shape tape spring for deployment and wiring.

[0100] A single C-shape tape spring is for example linked between two adjacent stiffening panels and is configured to rotate from a stowed position at 180 degrees during the deployment.

[0101] Alternatively the mechanical and electrical link 4 between adjacent stiffening panels, comprises at least one pair of strong C-shaped tape springs 5.

[0102] As shown in Figure 13, the hinge 6 and the mechanical and electrical link 4 between adjacent stiffening panels are for example positioned at the nearest of the corners of the frames 35. The C-shaped tape springs are for example organized in two groups placed according to the two corners respectively.

[0103] The C-shaped tape springs 11 between stiffening panels are for example double soft C-shaped tape springs for electrical wiring.

[0104] The C-shaped tape springs 12 linked to a solar panel are for example single soft C-shaped tape springs for electrical wiring.

[0105] The number of C-shaped tape springs 11 between stiffening panels and the number of C-shaped tape springs 12 linked to a solar panel depends on the position of the solar panel 2 within the solar array 10 as they are sometimes dedicated to transfer electricity from one single solar panel 2 and sometimes from several solar panels in series. For distal solar panels 2, there are few C-shaped tape springs 11 conducting electricity between adjacent stiffening panels and for proximal solar panels 2, there are many C-shaped tape springs 11 , conducting electricity between adjacent stiffening panels, depending on the number of solar panels 2.

[0106] As shown in Figure 14 the solar wing may also be moved from the deployed position to the stowed position. First, the stiffening panels 30, 31 and 32 are folded against the solar panels and eventually against the pressure panel or the yoke panel. This means that the hinges 6 are folded from around 90 degrees to zero degrees. Then, the solar wing 1 may be folded in a Z-configuration for example using cables. This means that the mechanical and electrical links 4 are folded. When all the solar panels 2 are stacked, for example against the pressure panel, the group of solar panels is rotated against the yoke 15. The stack of solar panels together with the yoke are then rotated against the body of the spacecraft, in the stowed position. The stack is for example pressured between the pressure plate and the body of the spacecraft, the pressure plate being on the top of the stack.

Claims

Claims

1. A solar wing (1) for a spacecraft (100), the solar wing (1) being configured to be movable between a stowed position and a deployed position and comprising a solar array (10) comprising :- a plurality of solar panels (2) comprising solar cells for generating electrical current and configured to be movable between the stowed position and the deployed position, the solar panels being arranged adjacent to one another and along a longitudinal axis (X) in the deployed position, - a plurality of stiffening assemblies (3) extending, in the deployed position, in at least one line (L1 or L2), each stiffening assembly (3) comprising at least one stiffening panel (30),characterized in that each stiffening panel (30) is mechanically and electrically linked to one of said solar panels (2) by at least two C-shaped tape springs placed between the solar panel and the stiffening panel for forming a hinge (6) and protruding on said solar panel (2) in the deployed position, stiffening panels in the same line being, successively, mechanically and electrically linked to one another by at least two C-shaped tape springs fastened to adjacent stiffening panels (30) and extending therebetween parallel to said line in the deployed position and being folded on itself in the stowed position,each stiffening assembly (3) being configured to be movable between the stowed position and the deployed position, each stiffening panel (30) forming a predetermined non-zero and non-planar angle (A) with said solar panel (2) in the deployed position,C-shaped tape springs (5) mechanically and electrically linked to stiffening panels being configured to transmit electrical current produced by the solar cells towards a spacecraft body (101 ).

2. The solar wing (1 ) according to claim 1 , wherein the predetermined non-zero angle (A) is comprised between 80° and 100°.

3. The solar wing (1 ) according to any preceding claim, wherein the pair of said C-shaped tape springs (5) of the mechanical and electrical link (4) is configured to have one end portion (7) and one opposite end portion (8), said one end portion (7) being configured to rotate relative to the opposite end portion (8) between the stowed position and the deployed position with a rotating angle of about 180°.

4. The solar wing (1) according to any preceding claim, comprising an electrical wiring (9) configured to transmit the electrical current, wherein the electrical wiring (9) comprises:- a plurality of C-shaped tape springs (11 ) extending between the stiffening assemblies (3) parallel to said line, in the deployed position,- a plurality of C-shaped tape springs (12) extending between the solar panels (2) and the stiffening assemblies (3) and- a harness (13) within the stiffening assemblies (3).

5. The solar wing (1) according to any preceding claim, comprising a yoke (15) and a pressure plate (16), the yoke (15) being configured to be fastened at one proximal end (17) to the spacecraft body (101 ) and fastened at one distal end (18) to the pressure plate (16), and the pressure plate (16) extending adjacent to one of the solar panels (2) along the longitudinal axis (X) in the deployed position, wherein the plurality ofstiffening assemblies (3) comprises a stiffening assembly comprising a stiffening panel (31) fastened to said pressure plate (16) and a stiffening assembly comprising a stiffening panel (32) fastened to said yoke (15), the stiffening assembly of the pressure plate (16) and the stiffening assembly of the yoke (15) being configured to be movable between the stowed position and the deployed position in which a plane including said stiffening panels (31 , 32) forms the predetermined non-zero angle (A) with the pressure plate (16) and the yoke (15), respectively, in the deployed position.

6. The solar wing (1) according to any preceding claim, wherein the solar panels (2) comprise a front face (21) supporting the solar cells and a rear face (20) that is opposite to the front face, wherein the plurality of stiffening assemblies (30) is fastened to the rear face (20) of the solar panels (2).

7. The solar wing (1 ) according to any preceding claim, wherein each stiffening assembly (3) comprises a frame (35) having elongated stiffeners (36).

8. The solar wing (1) according to any preceding claim, wherein the plurality of stiffening assemblies (3) extends in a single line, the line being parallel to the longitudinal axis (X).

9. The solar wing (1 ) according to the any of claims 1 to 8, wherein the plurality of stiffening assemblies (3) extends in a plurality of lines (L1 , L2), the plurality of lines (L1 , L2) comprising two lines (L1 , L2) that are substantially parallel to one another, preferably extending at lateral ends (27, 28) of the solar panels (2).

10. The solar wing (1) according to any of claims 1 to 8, wherein the plurality of stiffening assemblies (3) extends in a plurality of lines (L1 , L2), the plurality of lines (L1 , L2) comprising two lines diverging away from the spacecraft body (101 ) in two diverging directions (D1 , D2).

11. The solar wing (1) according to claim 6, wherein the solar panels (2) are arranged in a Z configuration in the stowed position, with the pressure plate (16) surrounding the solar panels (2) and being distant from the spacecraft body (101 ) and with the solar panel (2x) that is at a free end (29) of the solar wing (1) being adjacent to the spacecraft body (101) in said stowed position.

12. The solar wing (1) according to the preceding claim, comprising spacers (38) extending between the solar panels (2) in the stowed position so that the solar panels (2) are not directly in contact to one another in the stowed position.

13. The solar wing (1 ) according to the preceding claim, wherein the solar panels (2) are in a preloaded state.

14. Spacecraft (100), notably satellite, comprising a spacecraft body (101) and a solar wing (1) according to any preceding claim.

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