Abstract:
A spacecraft architecture and accompanying standard allows for the creation of a spacecraft using an assortment of modules that comply with the standard. The standard preferably includes both mechanical and electrical compatibility criteria. To assure physical/mechanical compatibility, the structure of each module is constrained to be compatible with any other compatible module. To minimize the interference among modules, the extent of each module in select dimensions is also constrained. To assure functional compatibility, a common communication format is used to interface with each module, and each public-function module is configured to respond to requests for function capabilities that it can provide to other functions. Each module is preferably designed to provide structural support to the assemblage of modules, and an anchor module is provided or defined for supporting the entire assemblage and coupling the assemblage to other structures, such as a launch vehicle.

Description:
[0001]     This application claims the benefit of U.S. Provisional Patent Application 60/579,231, filed 14 Jun. 2004. 
     
    
     BACKGROUND AND SUMMARY OF THE INVENTION  
       [0002]     This invention relates to the field of satellite and spacecraft design, and in particular to a design architecture that provides a cross-mission set of modules and design rules that minimize the delay time between the definition of requirements and the launch of a spacecraft that satisfies these requirements.  
         [0003]     There is an increasing need for rapid requirements-to-launch turn-around time for deploying spacecraft. In military applications, for example, dynamic changes in political or military situations often result in a need for surveillance and/or communications satellites in orbits having particular coverage areas, with different situations requiring different satellite capabilities. In research applications, universities or other agencies often require spacecraft platforms that are easy-to-configure to support particular research objectives. In each of these applications, and others, there is a need to launch a payload without incurring the substantial time and costs associated with the development of a spacecraft to support the payload.  
         [0004]     U.S. Pat. No. 6,283,416, “SPACECRAFT KERNEL”, issued 4 Sep. 2002 to Richard D. Fleeter and Scott A. McDermott, and incorporated by reference herein, teaches the advantages of designing and providing a spacecraft interface with kernel components on one side of the interface, and components that depend on either the spacecraft configuration or the mission-specific system on the other side of the interface. The kernel components are both functionally and physically independent of the vehicle configuration and function, and physically independent of the mission-specific system. The kernel components typically include communications equipment for communicating with an earth station, a power management system for receiving variable power input and providing regulated power output, and a processing system that receives commands from the earth station and provides corresponding commands to other subsystems, on both sides of the interface, in a standard format. By providing communications, power management, and command processing in a kernel that is independent of the spacecraft configuration and the spacecraft&#39;s mission, the same kernel can be used on multiple spacecraft, thereby reducing the time and cost associated with the design and testing of new spacecraft, as well as potentially reducing manufacturing cost.  
         [0005]     Although the spacecraft kernel design architecture provides a means to provide potentially re-usable designs, it does not necessarily provide for rapid development of a spacecraft. As noted above, the kernel components are independent of the space vehicle&#39;s configuration. However, the packaging of one set of kernel components for a given space vehicle&#39;s configuration may be unsuitable for use in another space vehicle. Correspondingly, a re-packaging of the kernel components for a given configuration typically requires substantial mechanical design and testing time, and may not be able to take advantage of prior tests or certifications conducted with the original configuration of the kernel components.  
         [0006]     In like manner, by limiting the re-usable kernel components to those items that are independent of the spacecraft configuration, the number of potentially re-usable designs is substantially reduced. Most spacecraft, for example, use solar panels to provide the energy required to operate the spacecraft and mission-specific components. However, because the spacecraft configuration typically dictates how and where such panels can be placed relative to the other components, most solar panel arrangements are custom-designed for each spacecraft.  
         [0007]     U.S. Pat. No. 6,260,804, “FUNCTIONALLY AND STRUCTURALY MODULAR PARALLELOGRAM-SHAPED SPACECRAFT”, issued 17 Jul. 2001 to Anderson et al., and incorporated by reference herein, discloses a modular spacecraft design wherein each functional module of the spacecraft can be independently manufactured. Each module has a similar cross-section, so that the modules can be assembled into a spacecraft by stacking them along the vertical axis. Flats are provided at the vertices of the parallelogram-shaped modules for connection to vertical channel members that serve to join the modules together. This referenced patent does not address the design of each module, and of particular note, does not address the functional interface among the modules, other than to note that the channel members facilitate the routing of electrical cables. That is, the modular mechanical design of this referenced patent provides a structure that facilitates the independent manufacture of functional components, and potentially the rapid assembly of these components, but does not address techniques for reducing the time to design a spacecraft based on a given set of functional requirements.  
         [0008]     It is an object of this invention to provide a spacecraft architecture that facilitates rapid requirements-to-launch turnaround time. It is a further object of this invention to provide a spacecraft architecture that facilitates the use of previously designed and tested modules in a variety of configurations. It is a further object of this invention to provide a standard for spacecraft design that optimizes the potential for the use of designed modules in other spacecraft.  
         [0009]     These objects, and others, are achieved by a spacecraft architecture and accompanying standard that allows for the creation of a spacecraft using an assortment of modules that comply with the standard. The standard preferably includes both mechanical and electrical compatibility criteria. To assure physical/mechanical compatibility, the structure of each module is constrained to be compatible with any other compatible module. To minimize the interference among modules, the extent of each module in select dimensions is also constrained. To assure functional compatibility, a common communication format is used to interface each module, and each public-function module is configured to respond to requests for function capabilities that it can provide to other modules. Each module is preferably designed to provide structural support to the assemblage of modules, and an anchor module is provided or defined for supporting the entire assemblage and coupling the assemblage to other structures, such as a launch vehicle. 
     
    
     BRIEF DESCRIPTION OF THE DRAWINGS  
       [0010]     The invention is explained in further detail, and by way of example, with reference to the accompanying drawings wherein:  
         [0011]      FIGS. 1A-1B  illustrate an example spacecraft comprising a stack of modules in accordance with this invention.  
         [0012]      FIGS. 2A-2D  illustrate details of example components of a spacecraft module in accordance with this invention.  
         [0013]      FIG. 3  illustrates an example block diagram schematic of the interconnection of components among the modules in accordance with this invention.  
         [0014]      FIGS. 4A-4C  illustrate example spacecrafts, with and without an exoskeleton in accordance with an aspect of this invention.  
         [0015]      FIGS. 5A-5B  and  6 A- 6 C illustrate example solar panel modules in accordance with an aspect of this invention.  
         [0016]      FIG. 7  illustrates an example block diagram of a location-determination system in accordance with this invention. 
     
    
       [0017]     Throughout the drawings, the same reference numerals indicate similar or corresponding features or functions. The drawings are included for illustrative purposes and are not intended to limit the scope of the invention.  
       DETAILED DESCRIPTION  
       [0018]     In the following description, for purposes of explanation rather than limitation, specific details are set forth such as the particular architecture, interfaces, techniques, etc., in order to provide a thorough understanding of the concepts of the invention. However, it will be apparent to those skilled in the art that the present invention may be practiced in other embodiments, which depart from these specific details. In like manner, the text of this description is directed to the example embodiments as illustrated in the Figures, and is not intended to limit the claimed invention beyond the limits expressly included in the claims. For purposes of simplicity and clarity, detailed descriptions of well-known devices, circuits, and methods are omitted so as not to obscure the description of the present invention with unnecessary detail.  
         [0019]      FIGS. 1A-1B  illustrate an example spacecraft comprising a stack of modules in accordance with this invention. The stack of modules includes an anchor module  110 , an upper module  120 , and a plurality of intermediate modules  200 ( a - d ) between the anchor and upper modules. Typically, the stack of modules will include a communications module, an attitude determination module, an attitude control module, a power supply module, and a solar panel module. Preferably, alternative configurations of each of these modules are available, and the spacecraft system designer selects from among these configurations to satisfy the given mission objective. For example, the attitude determination and control modules may be a momentum-biased earth-pointer, a three-axis stabilized star pointer, and so on, each having a particular accuracy and resolution, and each consuming different power. The solar panel and power supply modules may each have different capacity, power, voltage, and so on. The spacecraft system designer selects the appropriate combination of modules to satisfy the mission objectives, knowing that the modules are designed to be compatible with each other.  
         [0020]     Each of the intermediate modules  200  has a common cross-section profile, so that they can be interchangeably stacked. Each of the intermediate modules is also structured such that, when they are coupled together, and coupled to the anchor and upper modules, a structural integrity is formed that extends across the entirety of the modules. Preferably, the structure of each module does not limit its availability for placement anywhere within the stack of modules, so that the structural integrity of the stack of modules is not dependent upon the order of placement of the intermediate modules  200 .  
         [0021]     The anchor module  110  preferably includes a coupling device  101  that facilitates the coupling to a launch vehicle, or other structure that serves to facilitate the deployment of the spacecraft. A “lightband” coupling device, as described in U.S. Pat. No. 6,390,416, “REUSABLE, SEPARABLE, STRUCTURAL CONNECTOR ASSEMBLY”, issued 21 May 2002 to Walter Holemans, is particularly well suited to provide this coupling.  
         [0022]     If there are functions of the spacecraft that cannot be provided by the modules  200   a - d , a mission-specific payload  130  provides these functions. The upper module  130  is preferably configured to facilitate the coupling of such a payload to the spacecraft. Optionally, a payload module can be designed that conforms to the aforementioned structural constraints of the intermediate modules  200   a - d , and placed within the stack as another intermediate module, or as the anchor  110  or upper  120  module.  
         [0023]      FIG. 1B  illustrates a preferred cross-section profile of the modules  200 . As contrast to the four-sided profile of U.S. Pat. No. 6,260,804, referenced above, the preferred profile of the modules  200  is substantially hexagonal. Generally, spacecraft use solar panels that are deployed perpendicular to alternate surfaces or vertices of the spacecraft. Alternate surfaces or vertices are used so as to avoid a complete occlusion of “view” by the solar panels. However, limiting the view to two of four orientations often introduces design and operational constraints, particularly if the view from the un-occluded surface must include a particular target. By providing a polygon shape of six sides, or more, with solar panels deployed perpendicular to alternate sides, the feasibility of providing un-occluded views that encompass any target area is substantially increased. Consider, for example, an optical system that includes cameras at each un-occluded area. In order to include any target area from a four-sided spacecraft with alternate solar panels, each camera would require a full 180° field of view; otherwise, the spacecraft must be appropriately oriented to aim a camera with a smaller field of view toward the target. In a six sided system with alternate solar panels, each camera need only have a 120° field of view, thereby reducing the likelihood of requiring a reorientation of the spacecraft to capture a particular target. In like manner, an eight-sided system will allow four cameras with 90° fields of view to be used, and so on.  
         [0024]     Note that the sides of the spacecraft need not be flat, and the “shape” of the spacecraft as defined herein relates to the “functional” shape, rather than a strict geometric interpretation. For example, the shape of the profile in  FIG. 1B  is defined as being substantially hexagonal, even though the flats at each vertex of the hexagon shape result in the profile being twelve-sided, which does not satisfy the geometric definition of a hexagon. In like manner, if the walls  220  and supports  210  were curved, the strict geometric definition might be a circular shape, whereas for the purposes of this disclosure, the placement of the six supports  210  partition the perimeter into six segments, thereby providing a functional partitioning that is substantially hexagonal, as the term hexagonal is used herein. In general, the supporting structure (e.g. the arrangement of supports  220 ) of each component defines its functional shape, and in a preferred embodiment, the supporting structure of each component has a common form, varying only in height.  
         [0025]     A base plate  250  provides a mounting surface for the functional components  280  in each module  200 , and also provides the structural support to supports  210  arranged at the perimeter of the module  200 . The base plate  250  includes “V” shaped notches  252  at each vertex, to facilitate coupling to an exoskeleton, discussed further below with regard to  FIGS. 4B-4C .  
         [0026]     The supports  210  are provided at the vertices in each module  200 , and serve to provide a primary load path between the upper module  120  and the anchor module  110 .  FIG. 2A  illustrates an example support  210  positioned upon a base plate  250 . In a preferred embodiment, holes  215 ,  255  in the support  210  and base plate  250  allow for bolts and nuts to couple the modules  200  together. A recess  211  permits tool access to secure each nut and bolt. A similar coupling is used for coupling the supports to the anchor module and the upper module. Other coupling arrangements may be used; for example, threaded holes may be provided in the anchor and upper modules to receive bolts from the supports  210  in the lower and upper intermediate modules  200   a ,  200   d . In a preferred embodiment, the anchor module includes enhanced mechanical structure, to provide stiffness to the overall stack structure, thereby reducing the mechanical structural constraints placed on each intermediate module, and thereby potentially reducing the mass of each intermediate module.  
         [0027]     In a preferred embodiment of this invention, an electrical interface  230  is provided in each module  200 . The electrical interface is designed to electrically couple each module  200  to a set of common electrical busses, and to provide an electronic integrity that extends across the stack of modules  200 .  FIG. 2B  illustrates an example configuration of such an interface, and  FIG. 3  illustrates an example block-diagram schematic. In this example, a multi-pin plug  231  and socket  232  are pass-through connected, and selected signal lines  235  provide coupling from the interface  230  to the functional components within the module  200  (e.g.  280  in  FIG. 1B ). The plug  231  of a lower module  200  in the stack plugs into the socket  232  on an upper-adjacent module  200 , thereby providing a continuous electrical connection through the stack. To ease assembly, and to avoid “tolerance build-up”, the plug  231  and socket  232  are preferably mounted on the support post  210  with pliant coupling devices, wherein these pliant devices allow for movement of the plug and socket combination in multiple dimensions.  
         [0028]     Although a pass-through connection from the socket  232  to the plug  231  is illustrated, for convenience, one of ordinary skill in the art will recognize that functional elements may be coupled in series between the socket  232  and the plug  231 . For example, some or all of the modules  200  may provide a “repeater” or “signal conditioning” function, wherein the input from a set of pins on the plug or socket is processed to provide a reconditioned output to a corresponding set of pins on the opposite socket or plug. However, a pass-through connection is generally preferred, because it facilitates the isolation of a faulty component without affecting the operation of other components.  
         [0029]     Of particular note, the electrical interface  230  is preferably configured such that the electrical integrity, i.e. the ability of the modules  200  to provide their electrically-dependent functions, is substantially independent of the order arrangements of the intermediate modules  200 . In a preferred embodiment, the “positional preference” of each module is taken into account when choosing a particular stacking order, but a lack of satisfaction of a module&#39;s preference does not preclude its ability to satisfy its functional requirements. For example, the MTBF (mean time between failure) of electronic components is generally known to be dependent upon the operating temperature of the components, and the operating temperature of modules  200  that are in proximity of solar panels is generally known to be higher than the operating temperature of modules  200  that are farther from the solar panels. Thus, modules  200  that have more components than other modules, and thus a potentially lower MTBF, are preferably placed farther from the solar panels. However, each module  200  in a preferred embodiment of this invention is designed to meet its operational MTBF criteria even if it is placed adjacent to the solar panels. In like manner, other module characteristics, such as preferences for unobstructed view, long or short moment arm relative to the vehicle center of mass, proximity to the payload for data exchange, and so on, can be used to provide a positional preference for determining relative or absolute positions in the stack, but each module  200  is preferably designed such that placement at any position within the stack will satisfy at least a minimal requirement for each of such preferences.  
         [0030]     In some instances, different “versions” of a functional module  200  may be provided, and these different versions may be formed by adding ‘sub-modules’ to a ‘basic’ module. The term “sub-module” is used herein to facilitate understanding. The sub-module will generally conform to all of the constraints of a module  200  as defined herein, such that a coupling of sub-module F to the basic module forms an integral module  200  for inclusion in the stack. With regard to thermal management, for example the sub-module may be an insulation module that provides additional insulation to the basic module when the module A happens to be placed at a ‘high-temperature location’, such as near solar panels. In like manner, a sub-module may be a module that contains “anchor” elements to allow the basic module to function as the anchor module, as discussed further below.  
         [0031]     In a preferred embodiment of this invention, the electrical interface  230  includes a plurality of lines/pins that provide power to the modules, one or more sets of lines/pins that provide communications among the modules, and one or more lines/pins that provide control and/or monitoring signals among the modules. Copending U.S. patent application ______, “SPACECRAFT NETWORK ARCHITECTURE”, filed concurrently for McDermott et al., Attorney Docket AA-040518, and incorporated by reference herein, discloses a spacecraft network architecture that is particularly well suited for use in a distributed-management and/or variable-resource spacecraft system. In this copending application, a component that requires a resource or service broadcasts a request for the resource or service; components that can provide the resource or service announce their availability, and the requesting component thereafter selects from among the available providers to receive the resource or service. The disclosed architecture also includes preferred power distribution and control schemes, as well as techniques for assembling functional modules based on functional requirements.  
         [0032]     The power pins of the interface  230  generally provide multiple power sources, with a regulated voltage for routine low-power elements, such as network-interface elements, and an unregulated current source, to allow each component to provide the appropriate voltages and degrees of regulation for its needs. In some embodiments, the ‘regulation’ of the voltage may merely include assuring that the voltage does not exceed the maximum voltage level specified for the interface  230 . As discussed further within, the anchor module  110  is particularly well suited as the module that provides the solar panels, and such an anchor module  110  may also include the voltage and power control components required to provide the power to the interface  230 . Each module  200  that is coupled to a power pin of the interface  230  preferably includes a current sensor that is configured to decouple the module from the electrical interface if current above a threshold value is detected, to prevent a short in one module from affecting the operation of the other modules. In a preferred embodiment, the power-supply component is configured to be able to terminate power to any component, or otherwise decouple the component from the interface  230 , to prevent anomalous behavior of the component from affecting other components.  
         [0033]     The communication pins of the interface  230  preferably provide at least a relatively low-speed communication channel, using, for example, the I 2 C standard and protocol. In a preferred embodiment, an inverse I 2 C channel is also provided, to provide common-mode noise rejection. Preferably, the I 2 C link layer is used with the inverse-channel principles of the CANbus physical layer to provide this complementary I 2 C channel. Additionally, an Ethernet, Firewire, or other high-speed communication channel will also be provided via the interface  230 .  
         [0034]     The control and monitoring pins of the interface  230  include, for example, a synchronization signal that facilitates synchronization among the components, a signal that indicates whether the spacecraft is in a ‘launch’ mode or a ‘deployed’ mode, and/or whether the spacecraft is operating under reduced power-available conditions.  
         [0035]     On most modules  200 , the relative location of the interface  230  is immaterial to the function of the module. However, in some modules, such as modules that include sensors on the perimeter of the module, the rotational-orientation of one module relative to another may affect the performance of the module. For example, a “viewing” sensor&#39;s field of view may be substantially limited if the sensor is mounted on the wall of the spacecraft that also includes a solar panel. For such modules, either redundant interfaces  230  are placed at multiple vertices of the module  200 , or the interface  230  on the module is configured to be mountable onto multiple alternative supports  210 . Preferably, the module  200  is configured to determine and report its rotational orientation, to facilitate verification tests before launch, and/or to determine the field of view of components within the module during operation. Example techniques for determining relative location and orientation are presented below, with reference to  FIG. 7 .  
         [0036]     Note that not all of the electrical connectivity of the modules  200  need be provided via the interface  230 . Some modules may have signal lines that have requirements that cannot be accommodated by the standard interface  230 . For example, the signal lines that couple an S-band radio to externally mounted antenna will typically require low-loss lines with RF-shielding, which may not be provided by the standard interface  230 . In a preferred embodiment of such modules, one or more sockets  282  are optionally mounted on the external surface  220 , and an external connection between the modules is provided after the stack is assembled. With specific regard to radio systems, the anchor module  110  is typically well suited for mounting antennas and other external components, on its lower surface area.  
         [0037]     The placement of functional components in the anchor  110  and upper  120  modules is illustrated by components  310  and  320  in  FIG. 3 . As also illustrated in  FIG. 3 , the interface  230 , or a subset thereof, is also provided via a connector  325  in the upper module to the payload ( 130  of  FIG. 1 ). Generally, because each spacecraft will have one lower-most, i.e. anchor, component, electrical or structural functions that require one instantiation of an item, such as line-termination elements, “pull-up” or “pull-down” elements, reference datum, launch vehicle interface, and so on, are preferably located at the anchor module  110 . Preferably, any component can be configured as the anchor component, either as a specifically designed ‘anchor version’ of the component, or via the addition of an anchoring sub-module that contains the aforementioned typical anchor-elements.  
         [0038]      FIG. 2C  illustrates an example perimeter plate  220 . The example perimeter plate  220  is configured to attach to the base plate ( 250  of  FIG. 2A ) using bolts that pass through holes  225  in the perimeter plate to threaded holes ( 226  in  FIG. 2A ) in the base plate  250 . Preferably, tabs  228  are provided for coupling the perimeter plate  220  to the support posts  210 , thereby providing additional lateral support to the posts  210 . Optionally, some or all of the perimeter plates  220  may include reinforcing ribs, to provide additional mechanical support. To facilitate alignment of the modules  200  in the stack, and to further enhance the structural integrity of the stack, each plate  220  has tabs  221  and corresponding recesses  222  that allow each plate  220  to couple to a base plate of an adjacent module  200 . With the example tabs  221  and recesses  222  illustrated, the upper module  120  is preferably configured to allow for coupling of the upper module  120  to the uppermost module  200  via the tabs  221 , and the anchor module  110  preferably includes tabs  221  for coupling to the base plate  250  of lowermost module  200 .  
         [0039]     Because thermal management is often one of the most challenging constraints in modular design, each module  200  in a preferred embodiment is designed to be thermally self-sufficient, such that the heat from a module does not affect adjacent modules. Thus, the coupling of the perimeter plates  220  to the adjacent base plates preferably includes insulated components to minimize thermal transfer. Preferably, each spacecraft module is designed to maintain its own thermal environment, as described further below, without a priori knowledge of the thermal characteristics of the other modules it may be placed with.  
         [0040]      FIG. 2D  illustrates an example configuration of a wall panel  220  that is configured to maintain thermal isolation between the wall panel  220  and a base plate  250 ′ of an adjacent module. Preferably an insulating gap is maintained between the wall panel  220  and the base plate  250 ′; a layer of insulation material may provide for this gap. The bolt  270  that couples the panel  220  to the adjacent base plate  250 ′, at tab  221 , is installed with insulation washers  272  that insulate the bolt head from the wall panel  220 , and insulate the panel  220  from the base plate  250 ′.  
         [0041]     Although the use of an insulation gap and an insulated coupling arrangement provides excellent thermal isolation, consideration must be made for electrical conduction along the walls and other chassis surfaces of the spacecraft, and particularly between the surfaces of adjacent modules. The vacuum of space causes potentially damaging discharges to occur if two neighboring metallic surfaces are allowed to go to different voltage potentials, and the radio performance of spacecraft is optimized if the spacecraft exterior surfaces appear as consistent equi-potential planes. Preferably, a thermally insulating but electrically conducting path is preferably provided between modules.  
         [0042]      FIG. 2D  illustrates an example embodiment that provides electrical conductivity, with substantially less thermal conductivity. A conductive shelf  260  is provided along the interior of the wall panel  220 . This shelf  260  is configured to contact a flexible thin-wire interface  257 ′ on the baseplate  250 ′ of the module above. Using flexible thin-wire, such as a spiral shield, creates a good electrical contact through many thin metal conductors spaced relatively widely apart, but creates a poor thermal contact because the total amount of material connecting the two modules is minimal. The baseplate  250  of each module is configured to contain the thin-wire interface  257 , and the shelf  260  is positioned such that the interface  257 ′ of an adjacent baseplate  250 ′ is under compression when the wall panel  220  is coupled to the adjacent baseplate  250 ′. Because the wall plates  220  and the baseplate  250  are metallic, and not insulated from each other when assembled, the coupling of each wall panel  220  to each adjacent baseplate  250 ′ provides for a substantially uniform potential on each surface of each module.  
         [0043]     Conventionally, iriditing is used to make metallic (especially Aluminum) surfaces conductive for space use. Although iridited coatings are easily scratched and corroded, iriditing is an efficient process that is well suited for conventional spacecraft. Conventional spacecraft are typically assembled once for launch, and the coupled iridited surfaces are not subject to abrasion or corrosion. However, because repeated compressive contact with a thin-wire interface  257  will be abrasive to the shelf  260 , and because modules may be assembled and disassembled multiple times for testing, payload integration, and pre-launch module replacement (to install a fresh battery pack immediately prior to launch, for example); and further because spacecraft modules of this type may be stored in inventory for extended periods awaiting use, iriditing is not suitable. In a preferred embodiment, the shelf  260  is coated with a harder surface coating, such as nickel plating.  
         [0044]     Preferably, to maintain a desired thermal environment in a module  200 , the base plate  250  and perimeter plates  220  conform to the principles and techniques disclosed in copending U.S. patent application ______, “SPACECRAFT MODULE WITH ENHANCED THERMAL TRANSFER CAPABILITY”, filed concurrently for Barton et al., Attorney Docket AA040519, and incorporated by reference herein. In accordance with this copending application, the base plate includes uniformly placed thermal channels that couple the heat generated in the center area of the base plate to thermal-conductive perimeter plates that are uniformly distributed about the perimeter of the spacecraft, thereby allowing for efficient thermal transfer to the external environment regardless of the spacecraft&#39;s orientation. Preferably, the thermal-conductive perimeter plates are placed at locations that are not within the reflective field of the solar panels. Because the length of the perimeter plates is constrained by the distance between vertices of the module  200 , a sufficient thermal-transfer area is achieved by adjusting the height of the module  200 .  
         [0045]     In a preferred embodiment of this invention, each module  200  is constrained so as not to extend beyond the vertical extent (height) of the module  200 . That is, any external components that are attached to a module  200  may not extend into the vertical-space of any other module  200 . This constraint further facilitates the placement of each module  200  at any location in the stack, knowing that it will not interfere with, or be interfered by, some other module  200  in the stack. Consistent with this constraint, the upper module  120  is preferably constrained to not extend downward into the vertical-space of the stack of modules  200 . As noted above, the height of a module  200  is generally determined based on the thermal-transfer requirements for the module, and this constraint generally affects the placement of the external component. However, if a module  200  includes an external component that is larger than the height of the module, or otherwise infringes on the vertical space of another module, the height of the module should be adjusted accordingly.  
         [0046]     The anchor module  110  is also preferably constrained to not extend upward into the vertical-space of the stack of modules  200 . However, this constraint may be relaxed to accommodate the arrangement of solar panels in a launch configuration, as illustrated in  FIG. 4A .  
         [0047]      FIG. 4A  illustrates an arrangement of a spacecraft  100  using conventional fold-away solar panels  410 . In a conventional design, the design of each module must be coordinated, so as to avoid interference with the solar panels, or other external components. To facilitate modular design, the constraints placed on modules  200  preferably include a distance dl beyond which external elements of the module  200  are not permitted to extend. If all of the modules  200  conform to this constraint, an anchor module with fold-away solar panels  410  can be designed independently of the particular modules  200  that form the stack.  
         [0048]     As is well known in the art, however, the use of fold-away solar panels requires a substantial amount of design effort, because these external panels  410  introduce a substantial moment-arm factor that must be taken into account in the mechanical design of the entire spacecraft. In a modular design, each module  200  must be capable of supporting the remainder of the stack of modules in the presence of the forces that the panels  410  introduce during launch, and during deployment.  
         [0049]     In a preferred embodiment, the solar panels are stored within the module stack during launch, thereby avoiding the mechanical difficulties introduced by external panels. Techniques for providing such internally-stored panels are presented below with regard to  FIGS. 5A-5B  and  6 A- 6 C.  
         [0050]     To further reduce the mechanical/structural requirements of each module, the modules are preferably configured to facilitate the use of an exoskeleton  400 , as illustrated in  FIGS. 4B and 4C . The exoskeleton is configured to provide for the allowable extent d 1  of external components of the modules, except at the vertices, where the exoskeleton couples to the spacecraft  100 . As noted above with regard to  FIG. 2A , the modules  200  are constrained to allow access to the elements that couple the modules at each vertex, and this same clearance area is used as the area for coupling the exoskeleton.  
         [0051]     As illustrated in  FIG. 1B , the baseplate  250  of each module includes V-shaped notches  252  at each vertex, which are also indicated in  FIGS. 4B-4C . These V-shape notches may also or alternatively be provided via a cover-piece that covers the opening  211  in the support post  210 . The exoskeleton  400  includes columns  450  that include extendable elements  452  that are configured to fit into the V-shaped notch  252  of each baseplate. As the elements  452  are extended radially inward, they serve to align the modules within the exoskeleton, and with continued extension, clamp the spacecraft in place, so that the structural support required during launch will be provided by the exoskeleton.  
         [0052]     In a preferred embodiment, the column  450  and elements  452  provide the required radially inward clamping force by using a technique similar to that used by conventional wedge clamps, such as defined in U.S. Pat. No. 4,354,770, “WEDGE ASSEMBLY”, issued 19 Oct. 1982 to Sheldon A. Block, and incorporated by reference herein. In an example embodiment, the column  450  includes a channel within which opposite-facing trapezoidal elements (wedge clamps) are placed, some of these elements forming the elements  452  that extend beyond the column  450 . A screw assembly provides a compressive force to the stack of trapezoidal elements, causing them to slide away from each other, thereby causing the elements  452  to extend further beyond the column  450 . With continued compressive force on the stack, the elements  452  engage the V-shaped notch  252  in each of the baseplates and exert the desired radially-inward force onto these baseplates. Other techniques for providing a radially-inward force from a column  450  to the spacecraft  100  may be used, but the use of wedge clamps has been shown to be particularly effective and efficient for this task, because it requires only the turning of a screw at each vertex to provide the coupling, and allows each module to move freely in two directions prior to clamping, thereby compensating for misalignments and assuring proper contact during clamping.  
         [0053]     By providing the structural support via the exoskeleton, each module  200  need not be designed to provide the support, and therefore can be designed using less massive materials. When the spacecraft  100  is deployed, the exoskeleton  400  is shed, thereby substantially reducing the mass of the deployed spacecraft and allowing for easier attitude control and less energy-consuming maneuvers in orbit. Additionally, because the exoskeleton  400  is shed, it need not be designed to last for the entire mission-life of the spacecraft  100 , and can also be configured to use less costly and less massive materials, particularly the materials used for the outer skin, if any.  
         [0054]      FIGS. 5A-5B  and  6 A- 6 C illustrate two example alternatives for providing solar panels that do not exceed the diameter of the stack during launch.  
         [0055]     In  FIG. 5A , the solar panels  520  comprise a flexible material that is rolled or folded into a compact form within a module  510 , typically the anchor module of a stack. The flexible material may include a spring structure that is held under tension when stored for launch, and released to deploy the panel  520 , or it may include an inflatable member which, when inflated, extends the panel  520  to its deployed position.  FIG. 5B  illustrates the panels  520  in a deployed state.  
         [0056]      FIGS. 6A-6C  illustrate an alternative embodiment of a solar panel module  610 .  FIGS. 6A and 6B  illustrate top and side views of the module  610  when the solar panels  620  are stored, and  FIG. 6B  illustrates the module  610  when the solar panels  620  are deployed. Rigid solar panels  620  are arranged on tracks  621  (not illustrated in  FIGS. 6A-6B , for clarity) within the module  610 , and are configured to be extended and unfolded when deployed. The use of rigid panels allows for the use of conventional solar panel material.  
         [0057]     The panels  620  are illustrated as being deployed using a linear screw mechanism  626 , illustrated in  FIG. 6C , although other conventional techniques for lateral deployment of a device can be used, such a scissor-jack, a compressed spring, and so on. In a preferred embodiment, the deployment mechanism is also configured to provide control of the orientation of the solar panels after deployment, thereby avoiding the need to provide a separate orientation-control mechanism. In the example of  FIG. 6C , the screw mechanism  626  is configured to provide a latching action, such that, after deployment, the orientation of the solar panels  620  can be adjusted by controlling/turning the screw.  
         [0058]     Note that  FIGS. 5A-5B  and  6 A- 6 C are provided for illustrative purposes, and one of ordinary skill in the art will recognize other alternatives for containing solar panels within a given diametric constraint in view of this disclosure. One of ordinary skill in the art will also recognize that the concepts of  FIGS. 5A-5B  and  6 A- 6 C are not limited to solar panels, and may be used for other components, such as antennas. In like manner, the use of folded wings  625  on each panel  620  to increase the surface area is optional, and is not limited to the example structure of  FIGS. 6A-6C . Wings and other shapes that increase surface area may be used, for example, in the example flexible structure of  FIGS. 5A-5B .  
         [0059]     As noted above, the module that contains the solar panels  520  may also be configured to provide power to the common electrical interface  230 .  FIGS. 5A-5B  illustrate a component  580  that receives power from the solar panels  520  and provides power to the interface  230 ; component  580  may also include batteries that are configured to provide power in the absence of power from the solar panels and to store power received from the solar panels. In a preferred embodiment, the component  580  provides multiple power sources, optionally with differing degrees of regulation. The module  610  of  FIGS. 6A-6C  may similarly be configured with components to provide power to the common electrical interface  230 , but such components are omitted from  FIGS. 6A-6C  for clarity.  
         [0060]      FIG. 7  illustrates an example block diagram of a location-determining system for determining a relative location of each module in the stack. In the example of  FIG. 7 , the network interconnection includes a line in which each module  200   a - 200   n  inserts a series resistor Ra-Rn. Modules  110 ,  120  provide a constant current for this series-resistance path, so that a measure of the voltage Va-Vn at each module  200   a - 200   n  provides an indication of the module&#39;s position in the stack. If all of the resistors Ra-Rn are equal, these voltages will be indicative of the order in which the module is located on the stack. For example, if a 1 ma current is provided, and each resistor is 1000 ohms, there will be a one volt increment at each level of the stack.  
         [0061]     In a preferred embodiment, each module resistor Ra-Rn is sized to be proportional to the height of the module  200   a - 200   n , so that a measure of the voltage provides a measure of the actual vertical position of each module. For example, each hundred ohms could represent one centimeter, so that with a 1 ma current source, if module  200   a  is 8 cm tall, Ra will be 800 ohms, and voltage Vb will be 0.8 volts, indicating that the base of module B is 8 cm above the lower module  110 . Optionally, the lower module  110  may also include a resistor (not shown) that provides a measure of the height of the first module  100   a  relative to the vertical origin of the spacecraft. Other location-determining techniques will be evident to one of ordinary skill in the art in view of this disclosure.  
         [0062]     Similarly, modules that allow for placement of the connector  230  ( FIG. 1B ) on alternative vertices, to allow the sides of the module  200  to be re-oriented relative to the location of the backbone connection (and hence relative to other modules on the stack), preferably contain a means of determining which vertex of the module corresponds to the backbone of the stack. Each vertex that is permitted to be aligned with the backbone, for example, can be configured with a means for sensing the presence of the connector  230 , such as a microswitch, a conductive ‘finger’ that contacts a part of the connector  230 , and so on. Alternatively, each vertex may be configured to provide a unique signal on one of the wires  235  that couple the connector  230  to the components  280 . Other orientation-determining techniques will be evident to one of ordinary skill in the art in view of this disclosure.  
         [0063]     The foregoing merely illustrates the principles of the invention. It will thus be appreciated that those skilled in the art will be able to devise various arrangements which, although not explicitly described or shown herein, embody the principles of the invention and are thus within its spirit and scope. For example, although the example embodiments are presented in the context of a spacecraft bus that hosts a payload, one of ordinary skill in the art will recognize that the principles of this invention can be used to provide payload systems or other systems that may comprise a number of selectable modules. For example, a variety of payload systems may include the need for a detection function (such as a camera, IR sensor, and so on), a memory function (to store the detected information), and a steering function (to track the source of the detected information). Different missions may require different detection functions, but the memory and steering functions can be provided in pre-designed and tested modules in accordance with this invention. These and other system configuration and optimization features will be evident to one of ordinary skill in the art in view of this disclosure, and are included within the scope of the following claims.  
         [0064]     In interpreting these claims, it should be understood that: 
        a) the word “comprising” does not exclude the presence of other elements or acts than those listed in a given claim;     b) the word “a” or “an” preceding an element does not exclude the presence of a plurality of such elements;     c) any reference signs in the claims do not limit their scope;     d) several “means” may be represented by the same item or hardware or software implemented structure or function;     e) each of the disclosed elements may be comprised of hardware portions (e.g., including discrete and integrated electronic circuitry), software portions (e.g., computer programming), and any combination thereof;     f) hardware portions may be comprised of one or both of analog and digital portions;     g) any of the disclosed devices or portions thereof may be combined together or separated into further portions unless specifically stated otherwise;     h) no specific sequence of acts is intended to be required unless specifically indicated; and     i) the term “plurality of” an element includes two or more of the claimed element, and does not imply any particular range of number of elements; that is, a plurality of elements can be as few as two elements.