A spacecraft battery system

Abstract: A spacecraft battery system arranges battery cells (34) in two planes with a common mounting plate (54) between them allowing for a more compact arrangement so the battery assembly can be removed from a limited mounting area resource of the spacecraft such as the north/south faces and placed remotely. The system includes a plurality of heat pipes (56) embedded in the planar mounting panel (54) and thermally connected with each of the battery cells (34) for distributing heat uniformly across the planar mounting panel (54). The system also desirably includes a thermal radiator (58) for radiating heat to deep space.

Description: The present invention relates generally to spacecraft battery systems and, more particularly, to a technique resulting in a more compact arrangement of battery cells in a spacecraft. Conventional two dimensional battery packaging methods need more spacecraft resources (mounting space, mass and heat pipes) to handle the increased number of battery cells required for higher power. Double stacking cells (arranging them three dimensionally as recommended in this disclosure) greatly reduces these resource demands. Present spacecraft place all battery cells in one plane (a two dimensional arrangement) on a panel which is directly mounted to a thermal radiator with embedded heat-pipes. Demands for greater eclipse power increase the number of cells so the two dimensional area devoted to batteries must also increase. Present spacecraft already use up all the north/south panel mounting area reserved for payload but conventional two dimensional batteries will require some of this area as payload power increases. The mounting area saved by three dimensionally arranged battery cells can now be devoted to adding payload. Modem geosynchronous satellite batteries comprise a series connection of cells (most commonly nickel hydrogen cells at the present time) and include a suitable construction for distributing heat within the battery and removing excess heat. The series connections and thermal subsystems are typically carried out by distributing the individual cells across a two dimensional planar surface. Fig. 1 illustrates a typical location of a battery assembly 20 adjoining a bus equipment assembly 22 for a spacecraft 24 (such as a geosynchronous spacecraft) where a pair of opposed faces are generally kept away from direct solar energy impingement. In a geosynchronous communications satellite, these faces are called the north and south faces 26, 28, respectively, and are the prime panel mounting areas and thermal radiators for the bus equipment, batteries and communications payload equipment. There is limited area for mounting all equipment directly onto these thermal radiator panels so if the batteries require more mounting area, then the communications payload equipment mounting area and or bus equipment mounting area must be reduced to compensate. Fig. 2 illustrates a conventional battery technique, a two dimensional arrangement in which a plurality of battery cells 34 for a battery 36 mounted on a panel 38 are spread in a single plane over the prime battery panel mounting area. There may be embedded heat pipes (not shown) in this battery radiator panel which are used to isothermalize all the battery cells to produce a more efficient battery. In this instance, the battery assembly shares real estate with the bus equipment panel. However, higher power spacecraft of the future will require larger batteries which will result in more prime panel mounting area devoted to batteries and less to the equipment the customer really desires, such as the bus equipment. Fig. 3 illustrates a representative application of the single plane (two dimensional) technology to an advanced spacecraft with higher power requirements for eclipse, depicting a modified enlarged battery assembly 20A and a smaller bus equipment assembly 22A, for example. Fig. 4 illustrates a perspective view of a conventional battery mounting arrangement. Figs. 3 and 4 illustrate that as the number of battery cells 34 increases, the prime panel mounting area devoted to batteries increases proportionately so that the battery assembly 20 impinges into areas previously reserved for the bus equipment assembly 22 which then must be transferred to the communications payload equipment panel mounting area 40 as seen in Fig. 5. The problem is that higher power is being sought so as to increase the payload capability of the spacecraft 41 but the negative result is a decrease in the prime area available for payload mounting below what is provided today. One way to mitigate the loss of panel mounting area for higher power spacecraft is to develop battery cells with higher capacity but the same mounting footprint per cell. These larger cells can then yield the required higher total watt-hour capability while maintaining the same battery panel mounting area as a lower power spacecraft. However, larger battery cell developments are costly and time consuming and entail program schedule risks. The simplest implementation of higher power is to increase the number of battery cells. Another way to maintain payload equipment mounting area is to move the single plane (two dimensional) of battery cells 34 internal to the spacecraft 42 as illustrated in Fig. 6. The problem with this implementation is that it occupies a substantial volume. It is very difficult to isothermalize (i.e., maintain isothermal conditions within) a battery spread over this large area and there are many possible leakage paths for heat to degrade the low temperature environment favored by the batteries. Also, since the batteries must have their own separate and unique temperature environment they must be heavily blanketed and blocked from viewing the very hot communications payload equipment. But effectively accomplishing this thermal isolation also blocks crossbody thermal radiation paths that are needed for cooling the propulsion fuel as well as all of the electronics for the spacecraft. US 5,634,612 shows a spacecraft having battery cells mounted on a mounting plane with the long axes of the cells parallel to the plane. According to the present invention, a spacecraft battery system comprises: a planar mounting panel (54) having first and second opposed mounting surfaces; a plurality of battery cells (34) in electrical continuity comprising a first battery assembly (50) mounted on and spread in a single plane over said first mounting surfaces of said planar mounting panel; and, a plurality of battery cells (340 in electrical continuity comprising a second battery assembly (52) mounted on and spread in a single plane over said second mounting surface of said planar mounting panel, the plane of said second battery assembly being spaced from the plane of said first battery assembly; characterised in that each of the plurality of battery cells (34) is mounted on and projects away from the respective mounting surface of said planar mounting panel. The present invention relates to a spacecraft battery system according to which battery cells are arranged in two planes with a common mounting plate between them allowing for a more compact arrangement. By so doing, the battery assembly can be removed from a limited mounting area resource of the spacecraft such as the north/south faces and placed remotely. Examples of the present invention will now be described with reference to the accompanying drawings in which: Fig. I is a diagrammatic perspective view of a satellite of known construction incorporating a battery assembly and a bus equipment assembly on a north or south face; Fig. 2 is a plan view of the battery assembly and bus equipment assembly on the north or south face of the satellite illustrated in Fig. 1; Fig. 3 is a plan view of the battery assembly and bus equipment assembly on the north or south face of a known advanced satellite with higher power requirements for eclipse; Fig. 4 is a perspective view of the advanced satellite of Fig. 3; Fig. 5 is a plan view of a mounting panel of a known satellite illustrating the encroachment of the battery assembly and bus equipment assembly into payload mounting areas in order to support higher power requirements; Fig. 6 is a plan view of a mounting panel of a known satellite illustrating the mounting area required for a large battery if the cells are spread into a single plane configuration; Fig. 7 is a side elevation view illustrating a battery mounting configuration according to the present invention; Fig. 8 is an end elevation view of the battery mounting configuration illustrated in Fig. 7; Fig. 9 is plan view of a three-dimensional two-plane battery mounting configuration having two separate packs constructed as four modules; Fig. 10 is a plan view of an alternate three-dimensional configuration constructed as four packs and eight modules; Fig. 11 is a perspective view of a three-dimensional two-plane battery constructed in four packs, with heat pipes connected to the north and south battery thermal radiators; and, Fig. 12 is a perspective view of a three-dimensional two-plane battery designed for implementation as east and west batteries. Referring to the drawings and, initially, to Fig. 7 which generally illustrates the arrangement of two cell planes, specifically, first and second battery assemblies 50, 52 with a common battery mounting panel 54 between them. This mounting panel 54 for both planes of cells 34 is the same and has embedded heat pipes 56 which simultaneously isothermalize both planes and are used to transport heat generated in the battery to a separate remote thermal radiator 58 as shown in a side view in Fig. 8. The resulting battery structure is not only more compact but it can be located away from its thermal radiator 58 to allow more choices for effective integration and more room for communications payload mounting. In addition, the battery assemblies 50, 52 require fewer expensive heat pipes 56 to attain the desired isothermal condition. Fig. 6, discussed above, illustrates the magnitude of the mounting area required for a pair of large battery assemblies 44, 46 each comprised of 72 cells if they were spread in a single two dimensional plane. By way of contrast, Fig. 9 illustrates a pair 59, 60 of the same size battery assemblies and cells 34 with a three dimensional or two plane battery mounting implementation having two separate packs and built as four modules extending into the plane of the paper. This resulting three dimensional two plane battery pack in each instance has the obvious advantage of being more compact so it will be easier to handle and integrate. Fig. 10 illustrates an alternate three dimensional implementation built as four packs or assemblies 62, 64, 66, 68 and eight modules (four being unseen and into the plane of the paper) which will also yield flexibility in integration and location of a large battery on the spacecraft 70. Fig. 11 illustrates a spacecraft 72 with a three dimensional two plane battery assembly 74, built in 4 packs, with heat pipes 56 (see Figs. 7 and 8) in mounting panel 54 connected to north and south battery thermal radiators 58. This implementation allows the bus electronics equipment 80 to have the same mounting area available as in the lower power applications. There may be some loss in battery capacity in this example because the battery thermal radiator area can be made smaller than the equivalent prime panel mounting area but this loss can be made up with the addition of a few more battery cells. Fig. 12 illustrates a three dimensional two plane battery assembly 82 for a spacecraft 84 designed for east and west battery implementation of the type disclosed in commonly assigned patent application Serial No. 07/985,452 filed by applicants herein on December 5, 1997 and entitled "Satellite With East and West Battery Radiators". This figure illustrates the compact nature of a large battery when packaged in three dimensions and requiring no resources from the prime north and south mounting and thermal radiator panels.

Claims: A spacecraft battery system comprising: a planar mounting panel (54) having first and second opposed mounting surfaces; a plurality of battery cells (34) in electrical continuity comprising a first battery assembly (50) mounted on and spread in a single plane over said first mounting surface of said planar mounting panel; and, a plurality of battery cells (34) in electrical continuity comprising a second battery assembly (52) mounted on and spread in a single plane over said second mounting surface of said planar mounting panel, the plane of said second battery assembly being spaced from the plane of said first battery assembly; characterised in that each of the plurality of battery cells (34) is mounted on and projects away from the respective mounting surface of said planar mounting panel. A spacecraft battery system according to claim 1, comprising isothermalizing means (56) for distributing heat from said battery cells in each of said first and second battery assemblies uniformly across said planar mounting panel (54). A spacecraft battery system according to claim 2, in which said isothermalizing means includes a plurality of heat pipes (56) embedded in said planar mounting panel (54) and thermally connected with said battery cells. A spacecraft battery system according to claim 1, comprising: thermal radiator means (58) for radiating heat to deep space; and, heat conducting means (56) thermally connecting said planar mounting panel (54) to said thermal radiator means (58) for drawing heat away from said planar mounting panel and to said thermal radiator means. A spacecraft battery system comprising: a first battery assembly having battery cells (34) arranged in two planes with a common mounting plate (54) between them; and, a second battery assembly having battery cells (34) arranged in two planes with a common mounting plate (54) between them; wherein the mounting plates (54) of the first and second battery assemblies are spaced apart in parallel planes; characterised in that each of the plurality of battery cells (34) is mounted on and projects away from the respective mounting surface of said planar mounting panel. A spacecraft including a spacecraft battery system according to any preceding claim.

Inventors: GELON WALTER S, HALL JOHN C, GELON, WALTER S., HALL, JOHN C.
Applicants: LORAL SPACE SYSTEMS INC, SPACE SYSTEMS / LORAL INC.
Application year: 1999
Main classification: H01M2
IPC codes: B64G1, H01M2, H01L31, H01M10