Thermal control arrangements for a geosynchronous spacecraft

• US 5,372,183 A
• Filed: 08/22/1991
• Issued: 12/13/1994
• Est. Priority Date: 08/22/1991
• Status: Expired due to Term

• Current Assignee: Lockheed Martin Corporation (Martin Marietta Corporation)
• Original Assignee: Martin Marietta Corporation
• Inventors: Strickberger, Harold P.
• Primary Examiner(s): Rivell, John
• Assistant Examiner(s): LEO, LEONARD R
• Application Number:
  US07/748,609
• Time in Patent Office: 1,209 Days
• Field of Search: 165/41, 165/904, 165/133, 244/158 R, 244/158 A, 244/163
• US Class Current: 165/41
• CPC Class Codes: B64G 1/503 Radiator panels
• CPC Class Codes: Y10S 165/904 Radiation

Abstract

A spacecraft adapted for operation in a low inclination angle earth orbit comprises north, south, east and west panels defining a spacecraft interior volume. The north and south panels are oppositely disposed with respect to each other and the east and west panels are oppositely disposed with respect to each other. The spacecraft interior volume generally and preferably lacks structural elements that substantially restrict thermal radiation among the panels. The north and south panels, to which spacecraft equipment is usually mounted, each include conductive heat pipes for reducing the temperature differences across each panel. The exterior surfaces of the north, south, east and west panels have a covering, preferably of optical solar reflectors (OSRs), for radiating thermal energy therefrom, wherein the OSRs have a solar absorptivity that is substantially less than their thermal emissivity. The interior surfaces of the north, south, east and west panels have a covering for effectively radiating thermal energy between and among the panels across the interior volume.

Claims

1. A thermal control arrangement for a spacecraft adapted for orbiting a body illuminated by a sun, said spacecraft comprising a plurality of external panels at least first and second of which are oppositely disposed on said spacecraft in locations not subject to substantial illumination by said sun, and at least third and fourth of said panels are oppositely disposed on said spacecraft and receive illumination from said sun in different portions of the orbiting of said spacecraft about said body, at least one of said first and second external panels including means for conducting heat therethrough in a manner for reducing temperature differences over substantially the entire area of said panel, said first and second panels each having external surfaces having a solar absorptivity substantially lower than its thermal emissivity, said third and fourth panels each having external surfaces having a solar absorptivity substantially lower than its thermal emissivity, and wherein said first, second, third and fourth panels have surfaces internal to said spacecraft adapted to radiate thermal energy therebetween.

2. The arrangement of claim 1 wherein said means for conducting heat includes a plurality of heat conducting devices within said one panel arranged for conducting heat over substantially the entire area of said one panel in said manner for reducing temperature differences over said one panel.

3. The arrangement of claim 2 wherein a first group of said heat conducting devices is arranged to conduct heat in a first direction and a second group of said heat conducting devices is arranged to conduct heat in a second direction transverse to said first direction.

4. The arrangement of claim 1 wherein the external surfaces of said first, second, third and fourth panels include optical solar reflectors.

5. A thermal control arrangement for a spacecraft adapted for orbiting a body illuminated by a sun, said spacecraft comprising a plurality of external panels at least first and second of which are oppositely disposed on said spacecraft in locations not subject to substantial illumination by said sun, and at least third and fourth of said panels are oppositely disposed on said spacecraft and receive illumination from said sun in different portions of the orbiting of said spacecraft about said body, said first and second panels each having external surfaces having a solar absorptivity substantially lower than its thermal emissivity, said third and fourth panels each having external surfaces having a solar absorptivity substantially lower than its thermal emissivity, and wherein said first, second, third and fourth panels have surfaces internal to said spacecraft adapted to radiate thermal energy therebetween without interference from substantial intervening structural elements of said spacecraft.

6. The arrangement of claim 5 wherein at least one of said first and second external panels includes means for conducting heat therethrough in a manner for reducing temperature differences over said panel.

7. The arrangement of claim 6 wherein said means for conducting heat includes a plurality of heat conducting devices within said one panel arranged for conducting heat over substantially the entire area of said one panel in said manner for reducing temperature differences over said one panel.

8. The arrangement of claim 7 wherein a first group of said heat conducting devices is arranged to conduct heat in a first direction and a second group of said heat conducting devices is arranged to conduct heat in a second direction transverse to said first direction.

9. The arrangement of claim 5 wherein the external surfaces of said first, second, third and fourth panels include optical solar reflectors.

10. A spacecraft adapted for operation in a low inclination angle earth orbit comprising:

north, south, east and west panels defining a spacecraft interior volume, said north and south panels being oppositely disposed with respect to each other and said east and west panels being oppositely disposed with respect to each other;

said north and south panels each including respective means for conducting heat thereacross for reducing the temperature differences on each of said panels;

a source of electrical energy;

utilization means mounted on interior surfaces of at least one of said north and south panels, wherein said utilization means dissipate at least a portion of said electrical energy;

said north, south, east and west panels having means covering surfaces thereof exterior to said spacecraft for radiating thermal energy therefrom, said means having a solar absorptivity that is substantially less than its thermal emissivity;

said north, south, east and west panels further having means covering surfaces thereof interior to said spacecraft for effectively radiating thermal energy between and among said north, south, east and west panels across said interior volume.

11. The spacecraft of claim 10 wherein said means for conducting heat includes a plurality of heat conducting devices within said one panel arranged for conducting heat over substantially the entire area of said one surface for reducing temperature differences over said panel.

12. The spacecraft of claim 11 wherein a first group of said heat conducting devices is arranged to conduct heat in a first direction and a second group of said heat conducting devices is arranged to conduct heat in a second direction transverse to said first direction.

13. The spacecraft of claim 10 wherein said north, south, east and west panels include optical solar reflectors.

14. The spacecraft of claim 10 wherein said orbit is equatorial.

15. The spacecraft of claim 14 wherein said orbit is equatorial at geosynchronous altitude.

16. The spacecraft of claim 10 wherein said spacecraft interior volume lacks structural elements that substantially restrict thermal radiation among said north, south, east and west panels.