System for deployment of a solar array of a spacecraft

Abstract

A solar panel array for a spacecraft including a plurality of solar panels, a plurality of intermediate hinges and a deployment system; the solar panels being arranged in an articulated structure in which adjacent solar panels are interconnected by a respective hinge from the plurality of intermediate hinges and are able to rotate relative to each other around a rotational axis of the respective hinge; the articulated structure in a first condition being folded with the solar panels stacked, the hinges alternately positioned at a first side and a opposite second side of the stack, and the articulated structure in a second condition being unfolded with the solar panels substantially linearly aligned.

Description

[0001] System for deployment of a solar array of a spacecraft

[0002] Field of the invention

[0003] The present invention relates to a solar panel array for a spacecraft comprising a deployment system. Also, the invention relates to a spacecraft comprising such a solar panel array.

[0004] Background

[0005] Solar arrays and other spacecraft appendages are generally deployed in space from a launch configuration into an on-orbit, or deployed configuration. A solar array usually consists of one or more solar panels connected with hinges in between the adjacent solar panels and in between the inner solar panel and the spacecraft body. Some further components may be present as well in the solar array structure. In a configuration, a deployable structure (yoke) can be present between the spacecraft body and the inner solar panel to extend the solar array at a certain distance from the spacecraft. The hinges allow a solar panel to rotate with respect to the adjacent panel, yoke, or spacecraft. Typically, the hinges contain a spring element that provides the energy to actuate the rotation and have an end stop to stop the rotation when the deployed position is reached and a locking element to prevent the structure from bouncing back and to maintain the structure in the deployed position.

[0006] A solar array is considered as an articulated mechanical multi-body structure consisting of multiple (i.e. , two or more) rigid solar panels which are able to rotate with respect to each other along the axis of a hinge which forms a rotational joint between each pair of solar panels. In the launch configuration, the solar panels are stacked on top of each other; the inner solar panel being the body located closest to the spacecraft and the outer solar panel being the body most distant from the spacecraft. The inner solar panel has a hinged connection to the spacecraft. In the deployed configuration, the hinges between adjacent solar panels have rotated typically about 180° thereby aligning all panels along one line perpendicular to the hinge axes. The hinge between the spacecraft body and the inner solar panel rotates 180° or less, depending on the configuration. The typical rotation angle of this hinge is 90°.

[0007] When such a multi-body structure is deployed from the spacecraft, it may be necessary to control the motion of the solar panels. In the launch configuration and in the deployed configuration, the positions and orientations of all solar panels are uniquely defined. However, the intermittent positions and orientations are not uniquely defined; the position and orientation of a solar panel depends on the rotation angles of the hinges between that solar panel and the spacecraft. If the motion of the solar panels is not controlled, the position and orientation of each solar panel is a function of dimensions, actuation torques and moments of inertia of all solar panels and further components, if any, with mutual dependencies. Small variations in one of these parameters, for example the actuation torque of an individual hinge, may result in a large variation in the deployment trajectory and of forces acting on individual solar panels in the array. Generally, the rotation angles of the hinges are synchronised in order to obtain a controlled and predictable envelope in which the structure deploys. It is an object of the present invention to overcome or mitigate one or more of the disadvantages from the prior art.

[0008] Summary of the invention

[0009] The object is achieved by a solar panel array in accordance with claim 1.

[0010] There is provided a solar panel array for a spacecraft; the solar panel array comprising a plurality of solar panels, a plurality of intermediate hinges and a deployment system; the solar panels being arranged in an articulated structure in which adjacent solar panels are interconnected by a respective hinge from the plurality of intermediate hinges and are able to rotate relative to each other around a rotational axis of the respective hinge, the articulated structure in a first condition being folded with the solar panels stacked, the hinges alternately positioned at a first side and a opposite second side of the stack, and the articulated structure in a second condition being unfolded with the solar panels substantially linearly aligned wherein the deployment system comprises one or more deployment mechanisms each arranged on an associated hinge of the plurality of intermediate hinges located at the first side of the stack, each deployment mechanism comprising a pin, a bracket element and a driving device; the pin mounted on or near the associated hinge and extending parallel thereto and arranged in an off- centred position relative to the rotational axis of the associated hinge such that the pin is rotatable along a trajectory around the rotational axis, the bracket element, to be mounted on a body of the spacecraft, configured for holding the pin while the adjacent solar panels are stacked; the driving device configured for causing the pin to rotate along the trajectory and to move from a first position restrained by the associated bracket element to a second position free from the associated bracket element.

[0011] Advantageously, the deployment system provides that during unfolding from the stack, an unfolding solar panel rotates around the associated hinge, while the hinge remains at its position in the stacked position until the pin has rotated and moved into a position where the pin is released by the bracket element. In this manner the deployment mechanism provides that the solar panel is controllably released by keeping the solar panels in the stack remaining their stacked position while the unfolding solar panel rotates to its unfolded position.

[0012] By arranging between a predetermined number of solar panels in the stack a hinge with the deployment mechanism according to the invention, a controllable release sequence for the full solar array can be provided.

[0013] According to an aspect, the invention provides a deployment mechanism as described above, wherein the deployment mechanism is arranged on at least one hinge in the stack.

[0014] According to an aspect, the invention provides a solar panel array as described wherein above the driving device is configured for causing the associated hinge to unfold around the rotational axis thereof.

[0015] According to an aspect, the invention provides a solar panel array as described above, wherein the driving device is configured to exert an outward directed torque on an outer one of the associated solar panels connected to the hinge to rotate the outer of the associated solar panels relative to an inner one of the associated solar panels in the stack around the rotational axis of the hinge.

[0016] According to an aspect, the invention provides a solar panel array as described above, wherein the driving device comprises a spring element mounted on the hinge, the spring element in the first condition being loaded for exerting torque on the hinge.

[0017] According to an aspect, the invention provides a solar panel array as described above, wherein the bracket element comprises an upright to be attached to the body of the spacecraft, and extending perpendicularly therefrom; the upright provided with a holding beam attached to the upright, wherein the holding beam has a holding surface parallel to the body at a position along the upright, spaced apart from the body, corresponding with a position of the pin of the respective hinge in the folded stack at the first side for obstructing motion of the pin.

[0018] According to an aspect, the invention provides a solar panel array as described above, wherein the holding surface of the holding beam is either flat or curved.

[0019] According to an aspect, the invention provides a solar panel array as described above, wherein the holding surface of the bracket is directed parallel to the body or is inclined relative to the body.

[0020] According to an aspect, the invention provides a solar panel array as described above, wherein during the first condition of the solar panel array, the pin is configured to be in between the holding beam and the body. According to an aspect, the invention provides a solar panel array as described above, wherein the pin is mounted off-centred on a disk rotatable around the rotational axis of the associated hinge.

[0021] According to an aspect, the invention provides a solar panel array as described above, wherein the hinge comprises a first leg connecting to a first of a pair solar panels adjacent in the stack, a second leg connecting to a second of the pair adjacent solar panels, the first leg and second leg rotatably connected by the rotational axis; the first and second legs each at one end connected to the rotational axis and at the other end thereof connected to the respective one of the pair adjacent solar panels connected by the respective hinge; the first leg further comprising a secondary leg with a free end thereof extending over the rotational axis, and the pin being mounted on the free end of the secondary leg in the off-centred position.

[0022] According to an aspect, the invention provides a solar panel array as described above, wherein the pin is spaced apart from the associated hinge along a direction of the rotational axis, and the bracket element is located at a position corresponding with a position of the pin relative to the associated hinge along the direction of the rotational axis .

[0023] According to an aspect, the invention provides a solar panel array as described above, wherein the pin is spaced apart from the rotational axis of the hinge at a relatively larger distance from the holding surface of the holding beam than a distance between the rotational axis of the hinge and the holding surface.

[0024] According to an aspect, the invention provides a solar panel array as described above, further comprising a hold-and-release device wherein the hold-and-release device is arranged between the body and the stack of solar panels for controllably releasing an outer one of the stacked solar panels from the stowed arrangement in a direction away from the body.

[0025] According to an aspect, the invention provides a spacecraft consisting of a body and at least one solar panel array as described above attached thereto.

[0026] Advantageous embodiments are further defined by the dependent claims.

[0027] Brief description of drawings

[0028] Embodiments of the present invention will be described hereinafter, by way of example only, with reference to the accompanying drawings which are schematic in nature and therefore not necessarily drawn to scale.

[0029] In the drawings, identical or similar elements are indicated by the same reference sign. Figures 1a, 1b schematically show a spacecraft provided with a solar panel array in stowed arrangement and deployed arrangement, respectively;

[0030] Figure 2 schematically shows a perspective detailed view of the solar panel array in stowed arrangement, provided with a deployment mechanism according to an embodiment;

[0031] Figure 3 schematically shows a perspective detailed view of the solar panel array while deploying, provided with a deployment mechanism according to an embodiment; Figure 4 shows a schematic view of a hinge with a deployment mechanism in accordance with an embodiment, and

[0032] Figure 5 shows a schematic view of a hinge with a deployment mechanism in accordance with an embodiment.

[0033] Detailed description of embodiments

[0034] Figures 1a, 1b schematically show a spacecraft provided with a solar array in a stowed arrangement and deployed arrangement, respectively

[0035] A spacecraft 100 is shown which comprises a body 10 and a solar panel array 20. The body typically comprises equipment configured for use during space flight.

[0036] The solar panel array 20 is arranged for providing electrical power to the equipment in the body 10 and is configured as an articulated structure that comprises a plurality of solar panels 25 that are linked to each other, by hinges 30, 301 and by electrical wiring (not shown). The solar panel array 20 is hingedly connected to the body 10 by a hinged joint or yoke 40. Each solar panel comprises a plurality of solar cells arranged adjacent to each other on a surface of the solar panel. Depending on the configuration the solar panel may be configured to have a carrier panel with solar cell devices on one surface of the panel or on two opposite surfaces for receiving solar energy, during operation.

[0037] In Figure 1a, the spacecraft 100 is shown in a stowed arrangement which is applied when the spacecraft is stowed as payload in a rocket during launch. In this arrangement the articulated structure of the solar panel array 20 is in a folded position where the solar panels 25 are piled onto each other in which surfaces of two adjacent solar panels are facing each other.

[0038] Between each pair of adjacent solar panels 25 a hinge 30 is arranged at a connected side 44 of the stack located at the hinged joint 40 to the body, and at a free side 46 of the stack opposite to the connected side 44. The free side 46 of the stack is unattached to the body 10 of the spacecraft. In Figure 1b, the spacecraft 100 is shown in the deployed configuration in which the solar panel array 20 is in a deployed and unfolded position in which solar panels are positioned with their light receiving surfaces adjacent to each other oriented in a same direction, with the adjacent solar panels connected to each other by one of the hinges 30. The solar panel array is connected to the body 10 of the spacecraft by the hinged joint 40

[0039] Figure 2 schematically shows a perspective detailed view of the solar panel array 20 in stowed arrangement, provided with a deployment mechanism according to an embodiment.

[0040] A stack of solar panels 25 is in stowed arrangement on a surface 12 of the body 10 of the spacecraft 100.

[0041] The hinge comprises a pair of legs L1 , L2 which are configured to rotate relatively to each other around a rotational axis 31 of the hinge 30. One leg L1 of the hinge is connected with one of the pair solar panels. The other leg L2 of the hinge is connected with the other solar panel of the pair. Additionally, the hinge comprises a rotational spring 34 that is configured to be in loaded condition while the pair of solar panels is in stacked position and to exert torque on the hinge to rotate one of the solar panels relative to the other one of the pair around the rotational axis of the hinge and bring the pair of solar panels into a deployed or unfolded position.

[0042] A deployment mechanism is provided on one or more hinges on the free side 46 of the stack to assist the deployment of the solar panels in a controlled manner by guiding the solar panels while unfolding. The deployment mechanism comprises a pin 36 and a bracket element 37.

[0043] The pin 36 can be mounted on or at the hinge 30 in a manner which will be described in more detail below with reference to figure 4. The pin is positioned radially off- centred relative to the rotational axis 31 of the hinge 30 and lengthwise extending substantially parallel to the rotational axis of the hinge.

[0044] The pin is positioned at a same or larger distance from the edge of the stack than a distance of the rotational axis 31 of the hinge from the edge 27.

[0045] The bracket element 37 is arranged along the edge direction of the solar panels at a position corresponding with the position of the pin 36. The bracket element 37 comprises an upright 38 attached to the body 10 of the spacecraft, extending perpendicularly therefrom. Further, the bracket element 37 comprises a holding beam 39 substantially traverse to the upright 38 and directed towards the edge 27 of the stacked solar panels. The holding beam 39 is positioned at larger height from the surface 12 of the body 10 than the height of the pin 36 from the surface 12 of the body.

[0046] In the stowed position, the pin 36 is located under the holding beam 39, when viewed in a projection perpendicular to the body 10. This arrangement causes that the hinge 30 of the pair of solar panels on which the pin is mounted, is in fixed position.

[0047] In the exemplary embodiment of Figure 2, a stack of six solar panels is shown with three hinges 30 that are each provided with a deployment mechanism 35 comprising a respective pin on the free side 46 of the stack. On the connected side 44 of the stack the solar panels are in similar manner coupled to each other by spring loaded hinges 301 which may be of the same type as the hinges on the free side 46.

[0048] In correspondence with the number of hinges equipped with the deployment mechanism, the bracket element therefore is configured on the upright 38 with multiple holding beams. For each hinge 30, an individual holding beam is positioned on the upright 38 at a corresponding height relative to the position of the respective pin.

[0049] Figure 3 schematically shows a perspective detailed view of the solar array of Figure 2 while deploying, provided with deployment mechanisms according to an embodiment. During deployment, an outermost one of pair of adjacent solar panels 251 , 252 will be released by a hold and release mechanism (not shown). By the torque exerted by the rotational spring 34 attached to the hinge 30, said outermost solar panel 251 will rotate accordingly.

[0050] As a result, since the pin 36 of the deployment mechanism 35 is coupled to the hinge 30 but radially off-centre over a distance d1 from the rotational axis 31 , the pin will follow a circular path R under the holding beam when the hinge is rotated.

[0051] At the same time, the spring-loaded outermost hinge 301 at the connected side 44 of the stack will exert torque on the other of the pair of solar panels and push the pin 36 upward to the holding beam 39. When the pin during its rotation reaches an open end O of the holding beam 39, the pin 36 is no longer blocked by the holding beam and the hinge 30 will be released to rotate around the spring-loaded outermost hinge 301. The pair of adjacent solar panels 251 , 252 connected by the hinge 30 in between can now unfold to their deployed arrangement.

[0052] The open end O of the holding beam may be provided with a rounded edge to guide the pin during release.

[0053] It will be appreciated that in the solar array the number of hinges 30 may be different from the example embodiment shown in Figures 2 and 3. Further it is noted that not all hinges may necessarily be equipped with the deployment mechanism 37: the number of hinges 30 equipped with the deployment mechanism may be less than the overall number of hinges in the solar array at the free side 46 of the stack.

[0054] Figure 4 shows a schematic view of a hinge 30 with a deployment mechanism in accordance with an embodiment.

[0055] In this exemplary embodiment the hinge 30 is provided with a disk 28 on which the pin 36 is positioned radially off-centred with respect to the rotational axis 31, at a radial distance d1 from the axis.

[0056] The hinge has two legs L1 , L2 that are rotatable relative to each other. On one leg L1 the hinge 30 is connected to an outer solar panel 251 and to an inner solar panel 252 on the other leg L2. At the side of the hinge 30 facing away from the solar panels 251 , 252, the bracket element 37 is arranged with an upright 38 attached on the body 10 of the spacecraft. On the upright 38 a holding beam 39 is mounted that is directed towards the solar panels 251 , 252. The holding beam 39 may extend over or below the rotational axis 31 towards an open end O that is at a distance d2 less than the radial distance d1 from the rotational axis 31.

[0057] In the stowed arrangement the pin 36 is positioned under the holding beam 39 at the side of the upright 38. Caused by torque exerted by a second spring loaded hinge 301 attached to the inner solar panel 252 on the connected side 44 of the stacked solar panels, the pin 36 is forced upwards (arrow Fup) towards the holding beam 39.

[0058] When the outer solar panel 251 starts deploying by rotating , the pin 36 attached to the hinge 30 will follow a portion of a circular trajectory (R) due to the rotation of the hinge 30.

[0059] At the same time the force exerted by the second spring loaded hinge 301 will push the pin 36 to the holding beam 39, a motion of the pin 36 along the holding beam being the resultant of the rotation of the outer solar panel 251.

[0060] When the pin 36 reaches the open end O, the pin 36 under upward force, is no longer confined by the holding beam 39 and released. As a result, the inner solar panel 252 can now rotate relative to the second spring loaded hinge 301 and deploy together with the outer solar panel 251.

[0061] As will be appreciated by the skilled in the art, the rotating angle of the outer solar panel 251 before the inner solar panel 252 is released, depends on the initial position of the pin when the outer and inner solar panels are in stowed arrangement, the position of the open end 34 of the holding beam and also on the inclination of the surface of the holding beam along which the pin will move.

[0062] The holding beam 39 can be embodied with a flat surface or a curved surface for obstructing the pin when pushed upwards and away from the body. Both the shape of the surface of the holding beam and the inclination will affect the motion of the outer and inner solar panel during the initial phase of deployment before the pin is released. Figure 5 shows a schematic view of a hinge 30 with a deployment mechanism 35 in accordance with an embodiment. In this embodiment the pin 36 of the deployment mechanism is arranged on a leg L3 extending substantially radially from the outer solar panel 251 across the rotational axis 31 of the hinge 30, such that the pin is positioned between the rotational axis 31 and the upright 38 of the bracket element 37. The deployment mechanism functions in a similar manner as has been described with reference to figure 4. In this arrangement the deployment mechanism 35 may be located at some distance from the hinge mechanism (along the direction of the rotational axis). The invention has been described with reference to some preferred embodiments. Obvious modifications and alterations will occur to the skilled in the art upon reading and understanding the preceding detailed description. It is intended that the invention be construed as including all such modifications and alterations insofar as they come within the scope of the appended claims, only limited by the definitions presented therein.

Claims

Claims1 . A solar panel array for a spacecraft comprising: a plurality of solar panels, a plurality of intermediate hinges and a deployment system; the solar panels being arranged in an articulated structure in which adjacent solar panels are interconnected by a respective hinge from the plurality of intermediate hinges and are able to rotate relative to each other around a rotational axis of the respective hinge; the articulated structure in a first condition being folded with the solar panels stacked, the hinges alternately positioned at a first side and a opposite second side of the stack, and the articulated structure in a second condition being unfolded with the solar panels substantially linearly aligned; wherein the deployment system comprises one or more deployment mechanisms each arranged on an associated hinge of the plurality of intermediate hinges located at the first side of the stack, each deployment mechanism comprising a pin, a bracket element and a driving device; the pin mounted on or near the associated hinge and extending parallel thereto and arranged in an off-centred position relative to the rotational axis of the associated hinge such that the pin is rotatable along a trajectory around the rotational axis, the bracket element, to be mounted on a body of the spacecraft, configured for holding the pin while the adjacent solar panels are stacked; the driving device configured for causing the pin to rotate along the trajectory and to move from a first position restrained by the associated bracket element to a second position free from the associated bracket element.

2. The solar panel array according to claim 1 , wherein the deployment mechanism is arranged on at least one hinge in the stack.

3. The solar panel array according to claim 1 or 2, wherein the driving device is configured for causing the associated hinge to unfold around the rotational axis thereof.

4. The solar panel array according to claim 3, wherein the driving device is configured to exert an outward directed torque on an outer one of the associated solar panels connected to the hinge to rotate the outer of the associated solar panels relative to an inner one of the associated solar panels in the stack around therotational axis of the hinge.

5. The solar panel array according to any one of claims 1 — 4, wherein the driving device comprises a spring element mounted on the hinge, the spring element in the first condition being loaded for exerting torque on the hinge.

6. The solar panel array according to any one of claims 1 - 5, wherein the bracket element comprises an upright an upright to be attached to the body of the spacecraft, and extending perpendicularly therefrom; the upright provided with a holding beam attached to the upright, wherein the holding beam has a holding surface parallel to the body at a position along the upright, spaced apart from the body, corresponding with a position of the pin of the respective hinge in the folded stack at the first side for obstructing motion of the pin.

7. The solar panel array according to claim 6, wherein the holding surface of the holding beam is either flat or curved.

8. The solar panel array according to claim 6, wherein the holding surface of the bracket is directed parallel to the body or is inclined relative to the body.

9. The solar panel array according to claim 6, wherein during the first condition of the solar panel array, the pin is configured to be in between the holding beam and the body.

10. The solar panel array according to any one of claims 1 - 9, wherein the pin is mounted off-centred on a disk rotatable around the rotational axis of the associated hinge.

11. The solar panel array according to any one of claims 1 - 9, wherein the hinge comprises a first leg connecting to a first of a pair solar panels adjacent in the stack, a second leg connecting to a second of the pair adjacent solar panels, the first leg and second leg rotatably connected by the rotational axis; the first and second legs each at one end connected to the rotational axis and at the other end thereof connected to the respective one of the pair adjacent solar panels connected by the respective hinge; the first leg further comprising a secondary leg with a free end thereof extending overthe rotational axis, and the pin being mounted on the free end of the secondary leg in the off-centred position.

12. The solar panel array according to any one of the preceding claims 1 - 11, wherein the pin is spaced apart from the associated hinge along a direction of the rotational axis, and the bracket element is located at a position corresponding with a position of the pin relative to the associated hinge along the direction of the rotational axis .

13. The solar panel array according to claim 7, wherein the pin is spaced apart from the rotational axis of the hinge at a relatively larger distance from the holding surface of the holding beam than a distance between the rotational axis of the hinge and the holding surface.

14. The solar panel array according to any one of the preceding claims, further comprising a hold-and-release device wherein the hold-and-release device is to be arranged between the body and the stack of solar panels for controllably releasing an outer one of the stacked solar panels from the stowed arrangement in a direction away from the body.

15. A spacecraft consisting of a body and at least one solar panel array according to any one of the preceding claims 1 - 14 attached thereto.

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