Spacecraft design with semi-rigid solar array
A spacecraft includes a semi-rigid solar array and a main body structure, the main body structure configured as a convex polyhedron and including an aft and a forward face disposed opposite to the aft face and at least four side faces disposed between and approximately orthogonal to the aft face and the forward face. The solar array includes a number of panels linked together with flexible couplings. In an undeployed configuration, panels of the solar array cover at least two adjacent side faces, the flexible couplings providing an articulable joint approximately aligned with a line along which the two adjacent side faces are joined and connecting a first panel of the solar array and a second panel of the solar array, the first panel being proximal to a first side face and the second panel being proximal to a second side face.
1. A spacecraft comprising:
a main body structure coupled with a structural interface adapter for mating to a launch vehicle, the main body structure configured as a convex polyhedron and including:
an aft face disposed proximate to the structural interface adapter;
a forward face disposed opposite to the aft face; and
a first side face disposed between and approximately orthogonal to the aft face and the forward face and a second side face disposed parallel and opposite to the first side face;
a third side face disposed between the first side face and second side face and approximately orthogonal to the aft face and the forward face and a fourth side face disposed parallel and opposite to the third side face; and
a first deployable solar array coupled with the first side face and comprising a first plurality of panels linked together with flexible couplings and a second deployable solar array coupled with the second side face and comprising a second plurality of panels linked together with flexible couplings, wherein:
the first plurality of panels comprises a first set of two or more panels disposed parallel to and outboard of the first side face, a second set of two or more panels disposed parallel to and outboard of the third side face, and a third set of two or more panels disposed parallel to and outboard of the fourth side face; and
the second plurality of panels comprises a fourth set of two or more panels disposed parallel to and outboard of the second side face, a fifth set of two or more panels disposed parallel to and outboard of the second set of panels, and a sixth set of two or more panels disposed parallel to and outboard of the fourth side face and inboard of the third set of panels.
2. The spacecraft of claim 1 , wherein the main body structure includes at least five side faces.
3. The spacecraft of claim 2 , wherein, in the undeployed configuration, panels of the solar array cover at least three of the side faces.
4. The spacecraft of claim 1 , wherein one or more of the at least two adjacent side faces are configured as thermal radiator panels.
5. The spacecraft of claim 4 , wherein the spacecraft is configured to undergo an orbit transfer operation with a substantial portion of the thermal radiator panels covered by panels of the solar array.
6. The spacecraft of claim 5 , wherein, during the orbit transfer operation, photovoltaic cells of at least two solar panels convert incident solar radiation to electricity.
7. The spacecraft of claim 1 , wherein each panel of the solar array has a thickness in the approximate range of 10-50 mils.
8. The spacecraft of claim 1 , wherein the flexible couplings are formed from an elastic material.
9. The spacecraft of claim 8 , wherein, in the undeployed configuration, the elastic material is in an unrelaxed condition that stores strain energy.
10. The spacecraft of claim 9 , wherein the solar array is configured to undergo a transition from the undeployed configuration to a deployed configuration.
11. The spacecraft of claim 10 , wherein the transition is a result, at least in part, of the strain energy causing the elastic material to change from the unrelaxed condition to a relaxed condition.
12. The spacecraft of claim 10 , wherein, in the deployed configuration, the panels are approximately co-planar.
13. The spacecraft of claim 10 , wherein the transition is a passive result only of the strain energy causing the elastic material to change from the unrelaxed condition to a relaxed condition.
14. A method of operating an orbiting spacecraft, the method comprising:
reconfiguring deployable solar arrays of the spacecraft from an undeployed configuration to a deployed configuration; wherein:
the spacecraft includes:
a main body structure coupled with a structural interface adapter for mating to a launch vehicle, the main body structure configured as a convex polyhedron and including:
an aft face disposed proximate to the structural interface adapter;
a forward face disposed opposite to the aft face; and
a first side face disposed between and approximately orthogonal to the aft face and the forward face and a second side face disposed parallel and opposite to the first side face;
a third side face disposed between the first side face and second side face and approximately orthogonal to the aft face and the forward face and a fourth side face disposed parallel and opposite to the third side face; and
a first deployable solar array coupled with the first side face and comprising a first plurality of panels linked together with flexible couplings and a second deployable solar array coupled with the second side face and comprising a second plurality of panels linked together with flexible couplings, the first plurality of panels comprising a first set of two or more panels disposed parallel to and outboard of the first side face, a second set of two or more panels disposed parallel to and outboard of the third side face, and a third set of two or more panels disposed parallel to and outboard of the fourth side face; and
the second plurality of panels comprising a fourth set of two or more panels disposed parallel to and outboard of the second side face, a fifth set of two or more panels disposed parallel to and outboard of the second set of panels, and a sixth set of two or more panels disposed parallel to and outboard of the fourth side face and inboard of the third set of panels.
15. The method of claim 14 , wherein one or more of the at least two adjacent side faces are configured as thermal radiator panels and further comprising executing an orbit transfer operation with a substantial portion of the thermal radiator panels covered by panels of the solar array.
16. The method of claim 15 , wherein the executing the orbit transfer operation occurs prior to a start of the reconfiguring a deployable solar array of the spacecraft from the undeployed configuration to the deployed configuration.
17. The method of claim 15 , wherein the executing the orbit transfer operation occurs subsequent to a start of the reconfiguring and prior to completion of the reconfiguring.
18. The method of claim 15 , wherein, during the orbit transfer operation, photovoltaic cells of at least two solar panels convert incident solar radiation to electricity.