Thermal control arrangements for a geosynchronous spacecraft

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The present invention relates to thermal control arrangements for spacecraft and, in particular, for spacecraft intended for certain generally equatorial orbits.

A spacecraft in space would be subject to a wide range of temperature conditions if it were not for various arrangements of thermally controlling devices operating cooperatively to control the flow of heat from the sources thereof within the spacecraft, and the various heat gains and heat losses. Heat is lost from a spacecraft by its being radiated into space and heat is gained by a spacecraft from input from the sun. Solar energy produces temperature rise in two ways. Principally, solar energy impinging upon the spacecraft is absorbed to some degree and produces direct heating. Secondarily, the electrical energy produced by the solar cell arrays as a result of solar energy impinging thereon is utilized by the various equipment on a spacecraft and is thereby generally converted to heat energy. It is noted that the spacecraft equipment would similarly produce heat energy to be dissipated if the electrical energy were produced by an alternative source, such as a battery, a fuel cell or a nuclear-powered generator.

Various arrangements have been developed in the prior art to address the foregoing need for a thermal control and management system for spacecraft. One arrangement employs all sides of the spacecraft as a thermal radiator and includes an arrangement of highly thermally conductive paths for transferring heat from the location where it is generated to the cooler side of the spacecraft, as well as from the hotter portions of the body to the cooler portions thereof. Once such arrangement is disclosed in U.S. Pat. No. 4,880,050 (Nakamura et al.) which describes a thermal management system wherein a plurality of T-shaped pallets each have a radiator panel and a mounting panel thermally coupled by L-shaped external heat pipes. Each mounting panel is coupled via L-shaped internal heat pipes to a closed-loop heat pipe by which they are thermally coupled to each other. The heat then is transferred to the radiator panels facing away from the sun via the closed-loop heat pipe, the internal heat pipes, the mounting panels, and the external heat pipes. It is submitted that the Nakamura et al. arrangement is extremely complex and expensive, and requires an elaborate network of heat pipes which have a significant mass thereby reducing the mass available in the spacecraft to be devoted to a useful payload. It is further submitted that this arrangement may be more suitable to a spacecraft that is spinning at a sufficiently high rate that the effect of thermal energy input from the sun tends to be averaged over the surface of the spacecraft due to such rotation.

A different arrangement is disclosed in the prior art in relation to spacecrafts that are not spinning. Consider, for example, an earth-orbiting spacecraft generally in the shape of a rectangular solid having six surfaces usually referred to as panels. Generally, the spacecraft is oriented with respect to the earth so that one surface faces the earth and is generally referred to as the earth-facing panel or nadir panel. The surface opposite the earth-facing panel is generally referred to as an anti-earth or anti-nadir panel because it always faces away from the earth. In a generally equatorial orbit, i.e. that of relatively low inclination with respect to the equatorial plane, the earth-facing and anti-earth panels are perpendicular to the orbital plane and to a radius of the orbit. Two other surfaces of the spacecraft are oriented perpendicular to the orbital plane and opposite each other on the spacecraft; one faces in the direction of spacecraft travel and the other faces away from that direction, and are generally referred to as the east panel and the west panel, respectively. These panels are generally oriented perpendicular to the orbital plane as well as to the velocity vector of the spacecraft, which is generally tangential to the orbit. The remaining two surfaces of the spacecraft are oriented parallel to the orbital plane and are generally referred to as the north panel and south panel, in correspondence to the poles of the body about which the spacecraft is orbiting, here the earth.

Each time the spacecraft makes an orbit about the earth the sun illuminates the east panel, then the anti-earth facing panel, then the west panel, and then the earth facing panel. Each panel is so illuminated for approximately one-third of the orbit, there being two panels illuminated at any given time. Thus, each of these panels is a source of solar heating of the spacecraft. It should be pointed out that if the spacecraft is in a relatively low orbit, say between 100 and 1,000 miles above the earth's surface, there is a substantial eclipse period in each orbit during which there is no solar illumination of the spacecraft. As a result, the earth-facing panel receives much less solar energy input than do the other panels. On the other hand, the earth-facing panel is able to radiate a relatively lesser amount of thermal energy away from the spacecraft because the thermal temperature of the earth is relatively high as compared to that of space. The east, west, and anti-earth facing panels can radiate substantial amounts of thermal energy when facing away from the sun into the thermally cold abyss of space. The north and south panels, however, face space during the entire orbit and generally receive relatively little solar energy input and so are advantageously used as the principal thermal radiating surfaces of the spacecraft.

Conventionally, the north and south panels are designed to be effective radiators to space, whereas the other panels are insulated so that they will not absorb thermal energy from the sun; therefore, they cannot radiate energy at other times. Such an arrangement is shown in U.S. Pat. No. 3,749,156 (Fletcher et al.) which discloses a thermal control system for a spacecraft modular housing wherein north and south oriented walls are designed to facilitate controlled transfer of heat from the interior of the module to space. The three sides that are subject to direct exposure to the sun's rays during orbital flight are constructed in a manner containing superinsulating material to prevent the transfer of heat therethrough. All significant heat transfer is through the north-south walls. The preferred coating for north and south walls is an optical solar reflector; multilayer, coated thin films (Mylar.RTM. or Kapton.RTM. films) are preferred superinsulating materials for the other walls. A heat pipe system within the module provides a rapid and efficient means of transferring heat to the north-south walls as well as uniformly across those walls.

An arrangement of like design is French Pat. No. 2 463 058 (Dornier System) which describes an installation for heat removal and temperature stabilization in which heat PV.sub.1 (where PV.sub.n apparently is used as a designation for heat) from a component on the earth panel is carried by one or several heat pipes to the north radiator or to the south radiator or to both. The north and south radiators include several heat pipes and are "subject to the variable action of the sun." "If the surfaces of the North and South radiators of a spacecraft do not suffice for the removal of PV.sub.1 . . . PV.sub.3, one must use for this removal one or several additional radiators. This result can be arrived at on the earth panel itself or by radiators on the West or East side." (Page 4, last paragraph.)

Both of these conventional arrangements have the same disadvantages as the aforementioned Nakamura et al. system in that a complex, expensive, and heavy network of internal thermally conductive heat pipes is required to transfer heat from one panel to another panel.

SUMMARY OF THE INVENTION

In a spacecraft adapted for orbiting a body illuminated by a sun, the spacecraft comprises 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. The thermal control arrangement therefor comprises the first and second panels each having external surfaces having a solar absorptivity substantially lower than its thermal emissivity, and the third and fourth panels also each having external surfaces having a solar absorptivity substantially lower than its thermal emissivity. The first, second, third and fourth panels each have surfaces internal to said spacecraft that are adapted to radiate thermal energy between and among the panels.

IN THE DRAWING

FIG. 1 is a diagram of a prior art spacecraft;

FIGS. 2 and 4 are diagrams of a spacecraft according to the present invention;

FIG. 3 is a graph showing relative performance of two types of thermal systems; and

FIGS. 5 and 6 are diagrams representing portions of the spacecraft of FIG. 4.

The prior art spacecraft 10 of FIG. 1 is viewed from the anti-earth side with the anti-earth panel removed. This spacecraft avoids the problems of the other prior art spacecraft described above and provides satisfactory spacecraft thermal control and management for low and medium power communications spacecrafts operating in geosynchronous earth orbit. Spacecraft 10 includes north panel 12, south panel 14, east panel 16, and west panel 18. Interior to the spacecraft is a cylindrical central support column 20 and radially mounted support panels 22, 24, 26, 28, 30, and 32 on which equipment may be mounted. Because north and south panels 12 and 14 face into space and can therefore be effectively utilized as heat radiating surfaces, it is advantageous to mount equipment that dissipates substantial energy (thus producing substantial heat) directly thereon where it can be conveniently and effectively thermally radiated into space. North and south panels 12 and 14 include heat pipes 34 and 36, and 38 and 40, respectively, for spreading the heat generated by the items of equipment mounted on those panels over respective substantial portions of the areas thereof. It is noted that although heat pipes 34, 36, 38, and 40 are diagrammatically shown spaced apart from the panels 12 and 14 for illustrative purposes, they are, in fact, embedded within the panels and in intimate thermal contact with the equipment and the radiators thereon. North and south panels 12 and 14 are covered on their outer surfaces with optical solar reflectors (OSRs), also known as optical surface radiators or second surface mirrors, which are thermal control coatings (described in more detail below) which serve as efficient thermal radiators. East and west panels 16 and 18 are covered with multilayer insulating (MLI) blankets 42 and 44, respectively, which reduce the thermal energy absorbed by such panels from the sun to an insignificant amount and which likewise reduce the amount of thermal energy radiated from the east and west panels.

Because the arrangement of FIG. 1 only utilizes a portion of the available surfaces of only the north and south panels to radiate excess heat to space, and because the internal structural elements comprising the central support cylinder and radial equipment panels are generally inefficient to transfer heat among the panels 12, 14, 16, and 18 by conduction, and would be inefficient to transfer heat among such panels by radiative coupling even if other radiating surfaces were available, this arrangement is less well suited to spacecraft employing higher power dissipating equipment.

Accordingly, there is a need for a thermal control arrangement for a spacecraft which can provide the greater thermal radiating surface area required by higher power payloads while avoiding the undesirable complexity, excess weight, and high cost of the arrangements disclosed in the prior art patent documents.

In FIG. 2, spacecraft 50, viewed in like manner to that of FIG. 1, has north and south payload panels 52 and 54, each of which includes a network of embedded heat pipes 62 and 64, respectively, which spread the heat load generated by equipment (not shown) mounted on such panel substantially over the entire panel area. In a communications satellite of the sort suitable for operation in geosynchronous earth orbit, the highest power dissipating equipment is the communications payloads, particularly the transmitting equipment. This equipment is preferably mounted on the north and south panels. North and south panels 52 and 54 include an external thermal control surfaces comprising optical solar reflectors. East and west panels 56 and 58 also have external thermal control surfaces comprising optical solar reflectors. These panels enclose an internal volume 60 of spacecraft 50 which is substantially free of structural or other elements impeding the radiation of thermal energy between and among the north, south, east and west panels 52, 54, 56, and 58. This inventive arrangement takes advantage of the fact that over a complete orbit, more thermal energy is lost via radiation from each of the east and west panels with optical solar reflectors than is gained by absorption of energy from the sun by such panels. Accordingly, the average temperature of the spacecraft will be lower with this inventive arrangement than with the prior art arrangement of FIG. 1, for spacecraft of equivalent physical size and internal power dissipation (heat generation).

FIG. 3 is a graph comparing the thermal performance of the prior art spacecraft of FIG. 1 to that of the spacecraft of FIG. 2 embodying the present invention. This overall increase in heat rejection results because one of the east and west panels 56 and 58 is always able to radiate heat energy to the cold of space for substantially more than one-half of each orbit, while the solar energy absorbed on the other of these panels is minimized via the beneficial characteristics of the optical solar reflector (which is described below) as well as the fact that solar energy impinges thereon for substantially less than one-half of each orbit. The heat pipes 62 and 64 contribute to this increase by reducing the east-west thermal gradients on the north and south panels 52 and 54. The analysis used to produce these comparative parametric curves assumed two spacecraft according to the FIGS. 1 and 2, respectively, having identical electrical power dissipation on each of the panels thereof and equivalent characteristics for the optical solar reflectors thereon. The graph of FIG. 3 shows the dramatic advantage of the present invention. This advantage may be realized in a spacecraft as either a reduction in panel temperature for a predetermined area of the north and the south panels, or as reduced height (and therefore area) of the required payload radiator panels, as well as the potential for a somewhat lesser reduction of both the area and temperature of each, when employing the thermal control arrangement of the present invention. The slightly higher temperature on the south panel as compared to that on the north panel in each configuration reflects the seasonal variation of solar intensity. The analysis is for the winter season when the earth is closest to the sun thereby to produce the hottest spacecraft condition. Due to inclination of the equatorial orbital plane of the geosynchronous orbit with respect to the solar plane, the sun illuminates the south panel at a high angle of incidence which causes the payload equipment mounted on the south panel to operate at a slightly higher temperature as compared to that on the north, notwithstanding equal payload equipment power dissipation on each of those two panels.

Although it is preferred that the interior of spacecraft 50 be free of structure and other elements that impede thermal radiation, the present invention is suitable where such elements are employed even though some internal radiative advantage is reduced.

Thus, it is seen that the present invention provides more desirable thermal control characteristics for spacecraft having high power payloads than does that of the prior art arrangements. A spacecraft embodying the present invention further provides a dramatic reduction in spacecraft weight owing to the avoidance of the complex and heavy internal heat pipe transfer networks required by the prior art patents, the elimination of the internal structural elements of the prior-art, lower-power spacecraft, as well as that obtainable due to the reduced height (and therefore area) of the payload panels, as illustrated by FIG. 3.

FIG. 4 is a diagram of a complete spacecraft 400 intended for communications spacecraft service in C-band and Ku-band from geosynchronous earth orbit. Spacecraft 400 comprises body 410 which is oriented in space so that panel 420 faces the earth throughout every orbit. Body 410 contains the required payload and housekeeping equipment. Extended outwardly from the north panel 422 and south panel 424 (not visible) of body 410 are solar cell arrays 402 and 404, each of which comprise four rectangular panels having solar cells mounted on one side thereof. These arrays 402 and 404 rotate about the axis defined by boom 406 with respect to spacecraft body 410 at a rate of one revolution per orbit so that the orientation of the solar cell sides of arrays 402 and 404 are facing the sun as the orientation of the earth-facing panel 420 of body 410 remains toward the earth. Housekeeping equipment, such as that for telemetry tracking and control systems, power systems, attitude control systems, and the like, are contained in the generally rectangular housekeeping module 412, the interior of which includes a central cylinder and six radial structural panels in arrangement similar to that shown for the prior art spacecraft of FIG. 1. U.S. Pat. Nos. 4,009,851, entitled "Spacecraft Structure" and 4,682,744, entitled "Spacecraft Structure" describing structures employing central support cylinders and radial panels are incorporated herein by reference. The payload equipment, such as communications receivers, communications transmitters, and communications antennas and the like, are arranged within a generally rectangular payload module 414. North panel 422 thereof is covered with an optical solar reflector as are substantial portions 440, 442, and 444 of the west panel 428 thereof and corresponding portions of the east panel 426 (not visible) thereof. Both the north and south payload panels 422 and 424 contain an arrangement of heat pipes therein, described below, for spreading the heat generated by equipment mounted thereon over a substantial portion thereof, for radiation into space. It is noted that while the west payload panel 428, as well as the east payload panel 426, could be generally planar, it may be convenient to provide a sloped section so that the Ku-band beam forming networks 462 may have a clear, unobstructed view of the Ku-band antenna reflector 460 for receiving and transmitting signals. In the sloped section, center subpanel 444 is covered with OSR as are the left and right subpanels 440 and 442, and center subpanel 446 and triangular panel 448 are covered with MLI. Likewise, a similar arrangement may be used on the east panel so that the C-band feed horn assemblies 472 may have a clear, unobstructed view of the C-band antenna reflector 470. Alternatively, the entire east and west payload panels 426 and 428 could be sloped as are the central portions 444 and 446 thereof in FIG. 4. Deployable omnidirectional antenna 408 provides reception and transmission of signals for command, ranging, tracking and telemetry functions.

The foregoing arrangement cooperates with other thermal control features of spacecraft 400 pertaining to thermal control. Earth-facing panel 420 and anti-earth panel 430 are both covered with multilayer insulation as are substantial portions of the surfaces of housekeeping module 412. Both Ku-band antenna reflector 460 and C-band antenna reflector 470 are covered on their front faces by reflective membranes and on their rear surfaces by multilayer insulation. Ku-band beam forming network 462 has optical solar reflectors on its north and south surfaces and is otherwise covered with multilayer insulation. C-band feed horn assembly 472 is covered with a tent made of multilayer insulation material.

FIG. 5 is a diagram of spacecraft body 410 having each surface thereon divided into subsections for showing further detail of the thermal control surfaces thereon. Earth-facing panel 420 and anti-earth panel 430 (not visible) are covered with multilayer insulating material as are the east panel 436 and west panel 438 (not visible) of the housekeeping module 412. North payload panel 422 and south payload panel 424 (not visible) of payload module 414 are panels containing embedded heat pipes (described below) and are covered on their external surface with optical solar reflectors.

An optical solar reflector is a thin transparent cover which is coated by a highly reflective material. Fused silica OSRs available from Optical Coating Laboratory, Inc. (OCLI) of Santa Rosa, Calif. are about 6 mils thick and are covered with a deposited silver coating on the side that is bonded to the spacecraft and with an indium tin oxide conductive coating on the exposed surface. CMX glass OSRs available from Pilkington Space Technology of Bodelwyddan, Rhyl, Clwyd, United Kingdom, are about 3 mils thick and have a deposited silver coating on the side that is bonded to the spacecraft and an indium tin oxide conductive coating on the surface exposed to space. Both are available in 1.65 inch by 1.65 inch squares and have solar absorptivity .alpha. of about 0.08 to 0.10 and thermal emissivity .epsilon. of about 0.80 to 0.81. The low ratio of absorptivity to emissivity .alpha./.epsilon. is indicative of the fact that the OSR is a very efficient radiator of thermal energy in the infrared band in which thermal energy is principally radiated and has a low thermal absorptivity in the band in which energy contained in solar illumination is found. Protection against electrostatic discharge of plasma-induced electrostatic charge on the spacecraft is provided by the front to back surface resistance of about 200K .OMEGA. of the OSR in combination with the conductive indium oxide or indium tin oxide coatings on one surface thereof, the silver coating on the other surface thereof and their being bonded and bounded with an adhesive that is not an electrical insulator, such as Solithane.RTM. 113 adhesive (with carbon black filler) available from Morton-Thiokol, Inc., Morton Chemical Division of Chicago, Ill., or RTV 566 adhesive available from General Electric Company of Waterford, N.Y., or a mixture of RTV 566 and RTV 567 adhesives (with graphite fiber filler).

Although it is desirable that the external coating material have an .alpha. as close to zero as possible and an .epsilon. as close to one as possible, the invention may be advantageously employed with substantially different values, for example, .alpha..apprxeq.0.4 and .epsilon..apprxeq.0.6, as might be obtained from a thermal control paint that has experienced degradation by exposure to ultraviolet (UV) and ionizing radiation.