Aeronautics and astronautics – Spacecraft – Attitude control
Reexamination Certificate
1998-10-07
2001-03-06
Eldred, J. Woodrow (Department: 3641)
Aeronautics and astronautics
Spacecraft
Attitude control
C244S158700
Reexamination Certificate
active
06196501
ABSTRACT:
BACKGROUND OF THE INVENTION
1. Field of the Invention
The present invention relates generally to a system for dissipating waste heat from heat generating equipment such as electronic components borne by satellites when in orbit. While the invention will be described in the context of a geosynchronous satellite, it will be understood that the teachings of the invention are applicable to any kind of a satellite whether in earth orbit or in some other orbit.
2. Description of the Prior Art
Communication satellite payload power requirements continue to rise. The increase in satellite payload in turn requires that the satellite allow for (a) increased equipment mounting area, and (b) increased waste heat thermal dissipation. In addition it is highly desirable that this increased capability be accommodated with the minimum increase in the satellite mass and size. The former requirement is obvious as launch cost tends to be proportional to mass. The latter requirement is based on the same issue in that larger satellite require in turn larger fairings which are heavier, have greater aerodynamic drag (reduced lift capability), and are more costly.
The two conventional approaches for addressing the above issues are
1. increase of the satellite body size thereby increasing its equipment mounting area and thermal radiating area which in turn increases the dissipation capability; the problem with this approach is that the larger satellite requires a larger fairing hence reducing launch mass per the above discussion; and
2. use of deployable thermal radiators to increase the total satellite radiating area; this approach does not address the mounting area requirements and adds mass for the deployable radiator.
The ideal solution would be one in which equipment mounting and thermal dissipation areas are increased in a satellite of small body size. The problem is identification of available surface area. In present satellites about 80% of the North-South panel area is already occupied with payload equipment. The East-West faces are typically divided between the battery radiator and the output multiplexer (OMUX) for a communication satellite, a non-electronic component capable of operating at high temperatures and capable of usefully dissipating heat even when its mounting surface is subject to direct solar illumination. The earth deck is increasingly filled with communication receiving equipment. Only the anti-earth deck remains largely unoccupied, however, as this surface is subject to 12 hours of direct solar illumination every day and therefore is not generally useful by itself as a thermal radiator.
It was with knowledge of the foregoing state of the technology that the present invention has been conceived and is now reduced to practice.
SUMMARY OF THE INVENTION
The present invention relates to a satellite comprising first and second modules. The first module includes a first assembly of components and the second module includes a second assembly of components and a thermal radiator. The second module is movable between a stowed position proximate the first module whereat the thermal radiator is only partially operative and a deployed position distant from the first module whereat the thermal radiator means is fully operative for dissipating waste heat to deep space for at least one of the first and second assemblies of components. In one embodiment, the first module includes a housing defining an interior cavity, the second module including a casing so shaped and dimensioned as to be telescopically received within the internal cavity of the first module and the second module being translationally movable along a deploying axis between the stowed position and the deployed position but incapable of mutual rotation about the deploying axis. In another embodiment, a satellite includes a plurality of discrete faces and comprises an equipment panel including an assembly of heat generating components which require cooling, a thermal radiator, and a connecting hinge mounting the equipment panel for pivotal movement between a stowed position proximate one of said faces whereat the thermal radiator is only partially operative and a deployed position distant from the face whereat said thermal radiator means is fully operative for dissipating waste heat from the assembly to deep space.
The basic concept of the present invention, then, is to combine electrical equipment into modules which are stowed within the satellite during launch and deployed from the satellite after launch. This concept is most applicable to the satellite bus equipment as the connections between this equipment and the satellite are non-rigid, generally cables or the like. A similar concept is also disclosed for the EPC (electric power converter) to TWTA (travelling wave tube amplifier) connection which is, once again, non-rigid.
Body stabilized satellites are typically right rectangular prisms with equipment mounted on the inside face surfaces. The interior volume of the satellite is principally occupied with fuel tanks, however, such utilization is typically not complete. According to the present invention, the deployed surface area of the satellite is increased by nesting one or more internal modules within the main satellite body. These internal modules are stowed during launch but deployed on orbit. The concept is enabled by (a) the low profile of equipment mounted to satellite surfaces, (b) the partial volumetric utilization of the present satellite, and (c) the fact that the connection between the bus equipment mounted on the deployable module and the satellite is adequately flexible. Furthermore, as the bus components tend to be individually severable, the concept is tailorable to the overall satellite, that is, if insufficient internal volume is available for stowage of the entire bus, then a partial implementation is possible where only selected elements of the bus (e.g., the battery) are configured on the deployable module.
An example of the concept is illustrated in
FIGS. 1 and 2
.
FIGS. 1 and 2
illustrate a satellite after deployment of a bus module which includes a battery and an equipment deck accommodating most of the electronic bus equipment. In this embodiment, the prismatic bus module is attached via deployment rails to the central cylinder of the satellite. The center of the module is a hollow cylinder which allows it to be stowed in a nested configuration with the central cylinder. Once the satellite has reached its operating orbit the module is lowered via deployment rails past the anti-earth face of the main satellite body thereby increasing the total surface area and hence heat dissipation of the satellite. The bus module required surface area may be more or less freely adjusted by changing the height of the module and or its radiator extension panels. The length and width of the module are relatively more constrained by the need to need to allow for mounting of equipment on the inside surfaces of the main satellite body. The result is an increase in the heat rejection capability of the satellite and hence payload in direct proportion to the achieved increase in radiator area as a result of deployment of the internally mounted bus module.
An alternative example of a deployable bus module is provided in FIG.
3
. In this instance, the module is intended for a satellite architecture m which fuel is stored in an array of tanks located near the aft end of the satellite, thereby lowering satellite CG (center of gravity) as opposed tanks stacked inside the central cylinder. The deployable module containing the very heavy satellite batteries and bus equipment is stowed immediately below the fuel tanks thereby further enhancing the lowering of the satellite CG.
As illustrated in
FIG. 4
, the deployment approach for the non central cylinder module is to hinge the module from the lower edges of the north and south satellite faces. In this instance, the module radiator area appears to be fixed by the length and thickness of the satellite main body. In fact, this is not strictly the case as the hinged modules may
Gelon Walter S.
Hall John C.
Helmer Robert E.
Eldred J. Woodrow
Perman & Green LLP
Space Systems Loral, Inc.
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