Articulation techniques for a spacecraft solar array
Techniques for two-axis articulation of a deployed spacecraft solar array are disclosed. In one aspect, an arrangement mechanically coupling a solar array with a sidewall of a body of a spacecraft includes a proximal appendage, a distal appendage rotatably coupled with the proximal appendage by way of a hinge, and a closed cable loop (CCL) system coupled with the proximal appendage and the distal appendage. In an on-orbit configuration, a long axis of the proximal appendage defines an α axis that is substantially orthogonal to the sidewall. The hinge includes CCL disengagement mechanism configured to de-couple the CCL system from the proximal appendage and the distal appendage and a rotation driving mechanism configured to cause the distal appendage to rotate about a β axis when the proximal appendage is in the on-orbit configuration, the β axis being not aligned with the α axis.
1. A spacecraft comprising:
a body;
a solar array; and
an arrangement mechanically coupling the solar array with a sidewall of the body, the arrangement including:
a proximal appendage having a proximal end connected with the body, the proximal appendage configured to be reconfigurable from a launch configuration to an on-orbit configuration, wherein, in the on-orbit configuration, a long axis of the proximal appendage defines an α axis that is substantially orthogonal to the sidewall;
a distal appendage rotatably coupled with the proximal appendage by way of a hinge connecting a distal end of the proximal appendage with a proximal end of the distal appendage, the hinge being configured such that the distal appendage is rotatable with respect to the proximal appendage about a β axis that is not aligned with the α axis; and
a closed cable loop (CCL) system coupled with the proximal appendage and the distal appendage and configured to synchronize deployment of at least a portion of the solar array wherein:
the hinge includes a rotation driving mechanism configured to cause the distal appendage to rotate about the β axis when the proximal appendage is in the on-orbit configuration and a CCL disengagement mechanism configured to de-couple the CCL system from the proximal appendage and the distal appendage.
2. The spacecraft of claim 1 , the arrangement further including a restraint mechanism; wherein:
the rotation driving mechanism is configured to cause the distal appendage to rotate about the β axis in a first direction to a first orientation and the restraint mechanism is configured to cause the distal appendage to be fixed in the first orientation, and
the rotation driving mechanism is configured to cause the distal appendage to rotate about the β axis in a second direction from the first orientation to a second orientation and the restraint mechanism is configured to cause the distal appendage to be fixed in the second orientation.
3. The spacecraft of claim 2 , wherein:
the rotation driving mechanism includes a drive assembly coupled with the proximal appendage, the drive assembly including a motor having a central shaft connected with the proximal end of the distal appendage by way of a plurality of rods positioned radially outwardly from the central shaft; and
the restraint mechanism includes interlocking interfaces connected with each of the rods, each of the interlocking interfaces configured to interlock with a corresponding interlocking surface of the drive assembly.
4. The spacecraft of claim 3 , wherein the interlocking interfaces include: one or more of a friction pad interface, a pin type interlock interface, and a mechanical interlock interface.
5. The spacecraft of claim 2 , the spacecraft further comprising a cam coupled with the distal appendage, the cam having a plurality of slots, each slot corresponding to one of a plurality of positions; wherein:
the restraint mechanism includes a latch roller connected with the proximal appendage, the latch roller configured to rest in the slots, and
the rotation driving mechanism includes a motor connected with a fork configured to drive the latch roller from a first one of the slots to a second one of the slots.
6. The spacecraft of claim 5 , wherein the slots have a semi-obround cross-section and the latch roller has a semi-circular cross-section.
7. The spacecraft of claim 1 , wherein the CCL disengagement mechanism includes:
a pulley connected with the distal appendage and the proximal appendage and a restraining member coupled with the distal end of the proximal appendage, wherein:
the pulley is configured to transmit torque from the CCL system to the proximal appendage and the distal appendage when the pulley is in a compressed position;
the pulley is configured to not transmit torque from the CCL system to the proximal appendage and the distal appendage when the pulley is in an expanded position; and
the restraining member is configured to press a tab coupled with the pulley such that the pulley is in the compressed position when the restraining member presses the tab and the pulley is in the expanded position when the restraining member does not press the tab, the tab being configured such that the restraining member presses the tab only during deployment of the distal appendage.
8. A system comprising:
an arrangement mechanically coupling a solar array with a sidewall of a body of a spacecraft, the arrangement including:
a proximal appendage having a proximal end connected with the body, the proximal appendage configured to be reconfigurable from a launch configuration to an on-orbit configuration, wherein, in the on-orbit configuration, a long axis of the proximal appendage defines an α axis that is substantially orthogonal to the sidewall;
a distal appendage rotatably coupled with the proximal appendage by way of a hinge connecting a distal end of the proximal appendage with a proximal end of the distal appendage, the hinge being configured such that the distal appendage is rotatable with respect to the proximal appendage about a β axis that is not aligned with the α axis; and
a closed cable loop (CCL) system coupled with the proximal appendage and the distal appendage and configured to synchronize deployment of at least a portion of the solar array wherein:
the hinge includes a rotation driving mechanism configured to cause the distal appendage to rotate about the β axis when the proximal appendage is in the on-orbit configuration and a CCL disengagement mechanism configured to de-couple the CCL system from the proximal appendage and the distal appendage, and
the hinge includes a restraint mechanism; and
a controller configured to provide control signals to cause:
rotating, using the rotation driving mechanism, the distal appendage about the β axis.
9. The system of claim 8 , the controller further configured to provide control signals to cause:
rotating, using the rotation driving mechanism, the distal appendage about the β axis in a first direction to a first orientation;
fixing, using the restraint mechanism, the distal appendage in the first orientation;
rotating, using the rotation driving mechanism, the distal appendage about the β axis in a second direction from the first orientation to a second orientation; and
fixing, using the restraint mechanism, the distal appendage in the second orientation.
10. The system of claim 8 , wherein:
the rotation driving mechanism includes a drive assembly coupled with the proximal appendage, the drive assembly including a motor having a central shaft connected with the proximal end of the distal appendage by way of a plurality of rods positioned radially outwardly from the central shaft; and
the restraint mechanism includes interlocking interfaces connected with each of the rods, each of the interlocking interfaces configured to interlock with a corresponding interlocking surface of the drive assembly.
11. The system of claim 10 , wherein the interlocking interfaces include: one or more of a friction pad interface, a pin type interlock interface, and a mechanical interlock interface.
12. The system of claim 8 , the spacecraft further comprising a cam coupled with the distal appendage, the cam having a plurality of slots, each slot corresponding to one of a plurality of positions, wherein:
the restraint mechanism includes a latch roller connected with the proximal appendage, the latch roller configured to rest in the slots, and
the rotation driving mechanism includes a motor connected with a fork configured to drive the latch roller from a first one of the slots to a second one of the slots.
13. The system of claim 12 , wherein the slots have a semi-obround cross-section and the latch roller has a semi-circular cross-section.
14. The system of claim 8 , wherein the CCL disengagement mechanism includes:
a pulley connected with the distal appendage and the proximal appendage and a restraining member coupled with the distal end of the proximal appendage, wherein:
the pulley is configured to transmit torque from the CCL system to the proximal appendage and the distal appendage when the pulley is in a compressed position;
the pulley is configured to not transmit torque from the CCL system to the proximal appendage and the distal appendage when the pulley is in an expanded position; and
the restraining member is configured to press a tab coupled with the pulley such that the pulley is in the compressed position when the restraining member presses the tab and the pulley is in the expanded position when the restraining member does not press the tab, the tab being configured such that the restraining member presses the tab only during deployment of the distal appendage.
15. A method comprising:
causing an arrangement mechanically coupling a solar array with a sidewall of a body of a spacecraft to transition from a launch configuration to an on-orbit configuration wherein, the arrangement includes:
a proximal appendage having a proximal end connected with the body, the proximal appendage configured to be reconfigurable from a launch configuration to an on-orbit configuration, wherein, in the on-orbit configuration, a long axis of the proximal appendage defines an α axis that is substantially orthogonal to the sidewall;
a distal appendage rotatably coupled with the proximal appendage by way of a hinge connecting a distal end of the proximal appendage with a proximal end of the distal appendage, the hinge being configured such that the distal appendage is rotatable with respect to the proximal appendage about a β axis that is not aligned with the α axis; and
a closed cable loop (CCL) system coupled with the proximal appendage and the distal appendage and configured to synchronize deployment of at least a portion of the solar array wherein:
the hinge includes a rotation driving mechanism configured to cause the distal appendage to rotate about the β axis when the proximal appendage is in the on-orbit configuration and a CCL disengagement mechanism configured to de-couple the CCL system from the proximal appendage and the distal appendage;
decoupling, using the CCL disengagement mechanism, the CCL system from the CCL system from the proximal appendage and the distal appendage; and
rotating, using the rotation driving mechanism, the distal appendage about the β axis.
16. The method of claim 15 , further comprising:
rotating, using the rotation driving mechanism, the distal appendage about the β axis in a first direction to a first orientation;
fixing the distal appendage in the first orientation;
rotating, using the rotation driving mechanism, the distal appendage about the β axis in a second direction from the first orientation to a second orientation; and
fixing the distal appendage in the second orientation.
17. The method of claim 16 , wherein:
rotating, the distal appendage to about the β axis includes activating a drive assembly coupled with the proximal appendage, the drive assembly including a motor having a central shaft connected with the proximal end of the distal appendage by way of a plurality of rods positioned radially outwardly from the central shaft; and
fixing the distal appendage includes causing interlocking interfaces connected with each of the rods to interlock with a corresponding interlocking surface of the drive assembly.
18. The method of claim 17 , wherein the interlocking interfaces include: one or more of a friction pad interface, a pin type interlock interface, and a mechanical interlock interface.
19. The method of claim 15 , wherein:
rotating the distal appendage to about the β axis includes activating a motor connected with a fork to drive a latch roller from a first one of a plurality of slots on a surface of a cam coupled with the distal appendage to second one of the slots; and
fixing the distal appendage includes causing the latch roller to rest in the slots.
20. The method of claim 15 , wherein the CCL disengagement mechanism includes:
a pulley connected with the distal appendage and the proximal appendage and a restraining member coupled with the distal end of the proximal appendage, wherein:
the pulley is configured to transmit torque from the CCL system to the proximal appendage and the distal appendage when the pulley is in a compressed position;
the pulley is configured to not transmit torque from the CCL system to the proximal appendage and the distal appendage when the pulley is in an expanded position; and
the restraining member is configured to press a tab coupled with the pulley such that the pulley is in the compressed position when the restraining member presses the tab and the pulley is in the expanded position when the restraining member does not press the tab, the tab being configured such that the restraining member presses the tab only during deployment of the distal appendage.