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Modular thermal management system for spacecraft

Modular thermal management system for spacecraft description/claims

This application: (1) is a continuation-in-part of co-pending U.S. patent application Ser. No. 11/743,555, filed on May 2, 2007; (2) claims the benefit of co-pending U.S. Provisional Patent Application Ser. No. 60/954,550, filed on Aug. 7, 2007; and (3) claims the benefit of co-pending U.S. Provisional Patent Application Ser. No. 61/042,205, filed on Apr. 3, 2008.

1. The Field of the Invention

This invention relates to heat transfer and, more particularly, to novel systems and methods for combined structural and heat transfer panels for spacecraft.

2. The Background Art

Heat transfer is fundamental to many processes. Engines releasing energy from fuel must routinely reject heat. Meanwhile, electrical and electronic devices that consume electricity likewise require cooling. One particularly demanding environment occurs with spacecraft. Spacecraft contain many electronic instruments. Weight is at a premium, as is space. Meanwhile, the loads required of any structure during launch may be substantial. Thus, the combination of light weight and high strength is difficult to achieve.

In heat exchange, metals have comparatively high thermal conductivities. By contrast, composite materials such as bonded composites, fiber-reinforced layers, and the like typically have comparatively poor thermal conductivity. Design and optimization of size, weight, strength, stiffness, location, packaging, and heat transfer systems is typically unique to every space craft. Thus lead times are long, and reuse or salvage value of any design is minimal. A modular system is needed in which standard components may be assembled to provide a stock spacecraft suitable for hosting equipment for a particular mission. Reusable components would also be an advance in the art.

What is needed is an apparatus and method for modular, lightweight, structurally strong and stiff, dual purpose panels providing both structure and heat exchange for a spacecraft without additional structures for that purpose. It would be an advance in the art to provide heat exchangers that also provide structural functionality to the apparatus. Moreover, it would be an advance in the art to provide a mechanism for maintaining as nearly as possible an isothermal belt about the circumference or other perimeter of a satellite in order to provide temperature stability and better dwell times or “time on target” for an observation platform to maintain its orientation without having to change in order to avoid solar radiation loads.

In view of the foregoing, in accordance with the invention as embodied and broadly described herein, a method and apparatus are disclosed in one embodiment of the present invention as including a modular, dual-function, structural-thermal-integrated spacecraft system.

An integrated structural and heat transfer panel relies on the two-phase thermal panels, and variations thereof, the basic operational and configuration details thereof having already been disclosed in U.S. patent application Ser. No. 11/743,555 filed May 2, 2007, U.S. Provisional Patent Application Ser. No. 60/954,550, filed on Aug. 7, 2007; and U.S. Provisional Patent Application Ser. No. 61/042,205, filed on Apr. 3, 2008, all of which are incorporated herein by reference. Panels in accordance with the invention integrate thermal management as well as structural requirements, particularly for satellite and other spacecraft systems.

In one embodiment, the panels are manufactured in a modular configuration, rather than in specialized, unique configurations. Thus, several panels may be used to create an assembly that approximates a cube or other rectangular structure. Each panel of the system may function independently from other panels, but may yet be capable of maintaining thermal contact with other panels at a sufficiently robust rate to substantially maintain a large portion, and even the entire perimeter of the satellite at substantially a single temperature.

For example, temperature differentials ranging from tens to hundreds of degrees exist to drive heat in various heat transfer mechanisms and devices. By contrast, panels in accordance with the invention may maintain temperature differentials of as little as two degrees Fahrenheit or less between a heat sink location receiving heat from a device mounted thereto and a location of heat rejection by the panel to the environment. Thus, individual panels may be capable of maintaining a temperature throughout within such a narrow limit. Meanwhile, whether or not a particular panel is viewing the sun or black space, interconnected panels may stabilize and reduce temperature differentials while efficiently moving heat about the connected system of panels.

Moreover, individual panels may be provided with doors having a thermally reflective side and a thermally emissive side. Thus, when a door is closed over a thermal panel, the exposed area reflects thermal radiation received from the sun or other objects. Meanwhile, when the door opens, the emissive side of the door operates like one of the fixed panels, while the now exposed, fixed panel also operates as an emissive panel.

Flexible thermal links, pivoting thermal links, or the like, thermally connecting doors to fixed panels forming the periphery of the satellite or other spacecraft may provide a comparatively high rate of heat transfer therebetween, thus providing significant radiation augmentation by the door panels.

In one embodiment, each panel has at least one face or surface that has a high emissivity finish so as to function effectively as a space radiator. Panels may be connected at their edges, or may themselves be formed to turn a corner and provide an additional segment, something like a foot extending from the main leg of the principal portion of the panel. Thus, the foot of a panel may provide, and be designed to provide, a sufficient area to provide substantial heat transfer at minimal temperature differential between the working fluid within one panel and the working fluid of another adjacent panel attached thereto.

In certain embodiments, panels need not be thermally connected to one another. Thus, each individual panel may operate independently. In yet other embodiments, the doors of a panel may provide a reflective surface, a solar electric device, or both when exposed to a radiation load, such as the sun. The doors may be opened to provide an emissive surface when exposed to the cold blackness space (e.g., when shaded). Thus, thermal loads being absorbed by panels in accordance with the invention may be transferred to another place on the same panel or to another connected panel to be rejected according to the orientation and exposure of the satellite or other spacecraft.

Doors may be configured in various sizes. For example, a pair of doors may have a central opening about half way across a surface of a fixed panel being covered. Thus, the two doors together may effectively cover a fixed panel of a satellite. When the two doors are opened fully to be substantially flat and coplaner with respect to the previously covered fixed panel, the effective radiation space or surface area may effectively be double that of the fixed panel alone. By contrast, when the doors are closed against the fixed panel, the missive surfaces of the door panels are basically re-reflecting back into the fixed panel, while the outer surfaces of the doors are reflecting solar radiation, or collecting radiation if a solar electric device is present. Thus, the effective radiating ability may be controlled, and may be multiplied or diminished, even to approach negligible values. When exposed to solar radiation, a fixed panel may be covered by the door panels presenting a reflective surface, solar electric cell array, or both. When shaded or facing black space, a fixed panel may be exposed by an open door for radiating heat away from the satellite, and may be augmented by additional heat transfer radiating from the emissive surfaces of any door.

In another embodiment, multiple doors may cover a fixed panel, each covering it entirely. For example, from one side of the panel, a door may swing toward a fixed panel to cover it. Meanwhile, from the opposite edge of the fixed panel, another door may close across substantially the entire radiative area of the same fixed panel. Thus, an outer door covers an inner door, which in turn directly covers the fixed panel. Thus, the exposed, radiating, surface area exposed to a cold or shaded space environment might be about triple the area of a single fixed panel alone. Four such doors may conceivably be folded over a rectangular, fixed panel to provide about five times the radiating area of the panel alone.

Nevertheless, in certain embodiments, a satellite or other spacecraft may simply contain, for example, six sides of a cube, none shielded from incoming radiation and all having the same emissive properties and substantially the same areas. In such an embodiment, heat loads from the sun may be driven by very low temperature differentials from the irradiated surfaces of the spacecraft to the shaded sides thereof to be re-radiated to the coldness of space.

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