Spacecraft radiator system using crossing heat pipes

Details Spacecraft radiator system using crossing heat pipes Spacecraft radiator system using crossing heat pipes

1999-08-19

2004-08-17

Atkinson, Christopher (Department: 3753)

Heat exchange

With vehicle feature

C165S104330, C165S168000

Reexamination Certificate

active

06776220

ABSTRACT:

BACKGROUND
The present invention relates generally to satellites or spacecraft, and more specifically, to a spacecraft or satellite radiator system comprising a crossing heat pipe system that utilizes heat pipes to thermally couple north and south radiator panels on a satellite or spacecraft.
The assignee of the present invention manufactures and deploys spacecraft or satellites into geosynchronous and low earth orbits. Such spacecraft use heat pipes that are used to dissipate heat. The heat pipes transfer thermal energy to spacecraft radiator panels where it is radiated into space.
Conventional communication satellite radiator panels must be sized individually to reject both the payload and solar dissipation during the solstice seasons. Conventional solutions to this problem are disclosed in U.S. Pat. No. 3,749,156 issued to Fletcher, U.S. Pat. No. 5,351,746 issued to Mackey, and U.S. Pat. No. 5,806,803 issued to Watts.
U.S. Pat. No. 3,749,156 discloses the use of two spacecraft panels with heat pipes in one direction (lateral). The lateral heat pipes are coupled together via a third heat pipe on the interior of the spacecraft.
U.S. Pat. No. 5,351,746 discloses three spacecraft panels, each panel constructed with J-shaped heat pipes coupled with straight heat pipes at the short part of the J. The panels are coupled together at intersecting edges with straight heat pipes.
U.S. Pat. No. 5,806,803 discloses a heat pipe network that is constructed from formed and straight heat pipes connected together with a highly conductive material, such as Grafoil, that is used as an interface gasket. The heat pipe network is interconnected to subnadir and auxiliary panels, and thermally couples to North and South radiator panels. The electronic components are mounted to heat pipes on panels attached to the main spacecraft radiator panels. The heat pipes transfer the thermal energy to the radiator panels where it is radiated into space.
The design disclosed in U.S. Pat. No. 5,806,803 does not use header heat pipes in their spacecraft panel. Instead the design of U.S. Pat. No. 5,806,803 couples the lateral heat pipes on the two spacecraft radiator panels to an intermediate set of transverse heat pipes which in turn are thermally coupled by a set of curved heat pipes on the transverse panel.
Accordingly, it is an objective of the present invention to provide for heat dissipating apparatus comprising a spacecraft or satellite radiator system comprising a crossing heat pipe system that utilizes heat pipes to thermally couple north and south radiator panels on a satellite or spacecraft.
SUMMARY OF THE INVENTION
To accomplish the above and other objectives, the present invention provides for a spacecraft or satellite radiator system comprising a crossing heat pipe system that utilizes heat pipes to thermally couple north and south radiator panels embedded in radiator panels disposed on the spacecraft. The spacecraft radiator system comprises first and second radiator panels each comprising an outer panel faceskin, an inner panel faceskin, and a heat pipe matrix disposed between the outer and inner panel faceskins. A honeycomb core may be disposed between the inner and outer panel faceskins that surrounds the heat pipe matrix. A transverse panel interconnects the first and second radiator panels. A plurality of crossing heat pipes extend between and thermally coupled to the heat pipe matrices of the first and second radiator panels. The plurality of crossing heat pipes extend outside the plane of the transverse panel.
By thermally coupling the first and second radiator panels together, the crossing heat pipe system allows the panels to share the solar load during solstice seasons. This increases the thermal dissipation capability of the spacecraft by approximately 11 percent.


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patent: 5806803 (1998-09-01), Watts
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patent: 0083586 (1988-04-01), None

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