Abstract:
A spacecraft architecture is defined that distinguishes components and sub-systems based on both functional and physical dependencies. On one side of the interface are kernel components that are both functionally and physically independent of the vehicle configuration and functionally and physically independent of the mission-specific system. On the other side of the interface are components that depend on either the spacecraft configuration or the mission-specific system. The kernel components can be included in a variety of spacecraft, independent of the spacecraft architecture and independent of the spacecraft mission. The kernel includes a communications system for communicating with an earth station, a command and data handling processor, and a power regulation and distribution system. The preferred kernel is extensible by allowing the selection of different capacity components within the kernel, each different capacity component utilizing the same standardized interface for communicating with the vehicle and mission-specific components. By providing a standardize interface and extensible kernel, design changes do not propagate beyond the standardized interface, thereby substantially damping the costly ripple effect typically associated with changes that are introduced late in the design cycle.

Description:
BACKGROUND OF THE INVENTION 
     1. Field of the Invention 
     This invention relates to the field of aerospace, and in particular to the field of spacecraft system architecture and design. 
     2. Description of Related Art 
     All spacecraft have substantially the same basic requirements: power, communications, guidance, navigation, control, and command and data handling. Conventionally, the design of a spacecraft, such as a satellite system, is effected by partitioning the spacecraft into two independent sub-systems: a payload system and a transport system. The payload system comprises the mission-specific equipment, such as a collection system that collects data in a research satellite, a relay system that retransmits signals in a communications satellite, and so on. The transport system, or “bus”, comprises the equipment required to effect the mission in space, including: the power generation and storage system, the attitude determination and control system, the command and data handling system, the communications system, and the infra-structure and super-structure to support each of the components of each system. 
     Although the functional partitioning of tasks between payload and transport systems provides the desired degree of functional independence for effective system design, the physical constraints inherent in spacecraft design often forces a structural dependence that minimizes the advantages that can be gained by this functional partitioning. For example, spacecraft missions often involve the collection of data. The arrangement of the solar panels that provide power to the spacecraft, the design of the attitude control system, and other spacecraft specific designs will be dependent upon the particulars of the collection equipment. If the mission is to visually collect data related to the earth&#39;s surface, for example, the solar panels must be arranged so as not to obscure the view of the earth, and the spacecraft must be controlled to orient the visual collection device toward the earth. Conversely, if the mission is to measure the effects of weightlessness on crystal growth, the solar panels can be placed anywhere on the exterior of the spacecraft, whereas the spacecraft propulsion and control system must be designed to minimize acceleration in any direction. 
     In like manner, the demands on spacecraft sub-systems, such as the communications system and the power generation systems, are substantially affected by mission-specific requirements. Typically, the payload and transport systems are designed using a specified allocation of power and bandwidth among the components. As the designs of the payload system and the transport system progress independently, issues arise when the actual requirements exceed the anticipated requirements. When such issues arise, a choice typically must be made between increasing the allocation of resources to the component requiring the additional resources, or decreasing the capabilities of the component to conform to the specified allocation. Increasing the allocation often requires a redesign of the transport system components that provide the resource, while decreasing the capabilities to conform to the specified allocation often requires a redesign of the payload system. Often, the determination of the actual requirements of each component or sub-system does not occur until a substantial portion of each system is designed. As is known in the art, the cost of design changes, in time, effort, and materials, typically increases exponentially with respect to the degree of completion of the design, and there is a very high cost associated with changes that occur late in the design cycle. 
     The overall structure of the transport system is also substantially dependent upon the payload requirements. The transport system typically provides the mechanical load-bearing structure to contain each of the components and sub-systems. As in the case of power and bandwidth allocation, space and weight are allocated among components. When an actual requirement exceeds the allocation, a redesign of the transport or payload system, or both, is typically required. 
     The above noted interdependencies, and others, between the payload system and the transport system are often a major contributing factor to the high cost, in time, effort, and material, of conventional spacecraft development programs. Because of the interdependencies imposed between the payload and transport systems, costly redesigns are often required late in the development cycle, when actual requirements and dependencies become known. Because of the interdependencies imposed between the payload and transport systems, the re-use of systems or sub-systems among spacecrafts having different missions is a sought-after but often unachievable goal. 
     BRIEF SUMMARY OF THE INVENTION 
     It is an object of this invention to provide a spacecraft architecture that facilitates independent sub-system design and development. It is a further object of this invention to provide a method and apparatus that facilitates the reuse of spacecraft sub-system designs. It is a further object of this invention to provide a method and apparatus that facilitates the extension of a spacecraft sub-system design without introducing substantial system interdependencies. It is a further object of this invention to provide a mission-independent sub-system design that can be used on a variety of spacecraft. 
     These objects and others are achieved by providing a standard interface that is spacecraft and mission independent. This interface is structured to distinguish components and sub-systems based on both functional and physical dependencies. On one side of the interface are kernel components that are both functionally and physically independent of the vehicle configuration and functionally and physically independent of the mission-specific system. On the other side of the interface are components that depend on either the spacecraft configuration or the mission-specific system. In a preferred embodiment, the kernel components are organized and structured as a kernel sub-system that can be included in a variety of spacecraft, independent of the spacecraft architecture and independent of the spacecraft mission. In a preferred embodiment, the kernel includes a communications system for communicating with an earth station, a command and data handling processor, and a power regulation and distribution system. The preferred kernel is extensible to include, for example, low-level functions, such as clock signaling and data buffering, as well as high-level functions, such as a navigation and attitude information processing system, a propulsion control system, and other mission and spacecraft independent processors and control devices. The preferred kernel is also extensible by allowing the selection of different capacity components within the kernel, each different capacity component utilizing the same standardized interface for communicating with the vehicle and mission-specific components. By providing a standardize interface and extensible kernel, design changes do not propagate beyond the standardized interface, thereby substantially damping the costly ripple effect typically associated with changes that are introduced late in the design cycle. 
    
    
     BRIEF DESCRIPTION OF THE DRAWINGS 
     The invention is explained in further detail, and by way of example, with reference to the accompanying drawings wherein: 
     FIG. 1 illustrates an example block diagram of a spacecraft system in accordance with this invention. 
     FIG. 2 illustrates an example kernel and interface in accordance with this invention. 
    
    
     DETAILED DESCRIPTION OF THE INVENTION 
     FIG. 1 illustrates an example block diagram of a spacecraft system  100  in accordance with this invention. The spacecraft system  100  includes an interface  150  for the communication of information and other signals between a kernel system  200  and a vehicle/mission-specific system  300 . For ease of reference, the term spacecraft-specific is used herein to refer to either vehicle-specific or mission-specific items. The kernel system  200  includes components that are common among spacecraft, and are not dependent upon the physical configuration of the spacecraft. In a preferred embodiment of this invention, the kernel  200  includes a communications system  210 , a processor  220 , and a power management system  230 . As contrast to conventional spacecraft design, the logical partitioning of components in accordance with this invention includes physical configuration considerations. Note, for example, that the communications system  210  does not include the antenna system  310  that is conventionally associated with a communications system. In like manner, the power management system  230  does not include the power generation system  330 . By including physical considerations in the determination of logical partitioning, the kernel components can be readily used in a variety of spacecraft. 
     The power management system  230  receives unregulated power  331 ′ from the power generation system  330  via the standard interface  150 , and provides therefrom regulated power signals  231 . The regulated power signals  231  are communicated via the interface  150  as regulated power signals  231 ′ to each spacecraft component or sub-system that requires regulated power. The power management system  230  includes components for power balancing, voltage and current regulation, and overload protection, as well as a power storage sub-system, such as a battery sub-system, for auxiliary power when the power generation system  330  does not provide sufficient power  331 . The regulated power signals  231  include a variety of voltage and current outputs. In a minimal embodiment of the power management system  230 , the regulated power output  231  comprises power from the power generation system  330  that is regulated to be below a specified voltage or current level; in a preferred embodiment of the power management system  230 , the regulated power output  231  also includes regulated +3 and +5 volt outputs for powering commonly available electronic devices. Note that, in accordance with this invention, the components of the power management system  230  can be provided without the aforementioned considerations typically required to conform the power generation devices, such as solar panels, to the mission-specific physical constraints. As would be evident to one of ordinary skill in the art, the design of the power generation system  330  must be designed to conform to the mission-specific physical constraints, but, if designed in accordance with the principles of this invention, the power management system  230  is not affected by this design conformance. 
     The standard interface  150  in accordance with this invention includes a specification for the power generation system  330 . These specifications include, for example, a minimum power input as a function of the output  231  load, a maximum below-limit time as a function of the output  231  load and the capacity of the power storage sub-system, and other factors that affect the design of the power generation system  330 , common to one of ordinary skill in the art. In a preferred embodiment of this invention, alternative power management systems  230  are provided so as to minimize the cost and weight demands of the kernel  200  in dependence upon the specified output  231  load. In accordance with this invention, however, the particular choice of power management system  230  is transparent to the interface  150 , and transparent to the spacecraft-specific system  300 . Provided that the power generation system  330  conforms to the specifications corresponding to a required output  231  load, the power management system  230  supplies the required output  231  load independent of the particular configuration of solar cells, mission-specific equipment, and the like. 
     The communications system  210  includes a transmitter  216  and a receiver  218  for communicating with an earth station. Note, however, that the corresponding antenna system  310  is not included in the kernel  200 , because although the antenna system  310  is functionally independent of all components except the communications system  210 , the antenna system  310  is likely to be dependent upon the physical configuration of the spacecraft-specific components. In a preferred embodiment, the frequency of operation of the transmitter  216  and receiver  218  are presettable to specified frequencies in the X-band (7-12 GHz), and the standard interface  150  includes a specification for an antenna system compatible with this frequency band. Other frequency bands may also be used, with an accompanying change to the interface specification. The communications system  210  handles all command and control signaling for the spacecraft and mission-specific system  350 , as well as the data communications from, for example, a mission-specific collection device. In a preferred embodiment of this invention, alternative transmitters  216  and receivers  218  are available for use in the communications system  210 , based on bandwidth requirements for transmitting or receiving data. In accordance with this invention, however, the information  221  that is provided to the interface that corresponds to communications via the communication system  210  is standardized to be independent of the particular selected transmitter  216  or receiver  218 . 
     The processor  220  provides the resources for the conventional “housekeeping” tasks associated with a deployed spacecraft, such as controlling and monitoring each of the spacecraft sub-systems, activating routine operations in response to received stimuli, and so on. In accordance with this invention, the processor  220  also provides the network protocol and management functions to effect the standardized communication of data and commands via the interface  150 . One of the fundamental tasks of the processor  220  is the translation and routing of data and commands to the proper sub-systems, via the standard interface  150 . Generally, commands are received from an earth-station, and data is transmitted to the earth-station. The processor  220  processes the commands from the receiver  218  and produces the appropriate commands that are communicated to the intended sub-system via the standard interface  150 . The commands from the earth-station via the receiver  218  typically effect a modification to one or more parameters of the corresponding sub-system, to change, for example, the spacecraft&#39;s orientation, the criterion used in the mission-specific data collection devices, and so on. In like manner, the processor  220  receives communications from each of the spacecraft sub-systems via the standard interface  150  and produces the corresponding data that is transmitted to the earth station via the transmitter  216 . Note that in this manner, the spacecraft-specific components and sub-systems are isolated from the particular protocol or other parameters of the earth-to-spacecraft communications link, and changes to the details of the communications link will not affect the design of the spacecraft sub-system communications via the standard interface  150 . For example, the mission-specific system  350  may include a data collection system that communicates mission-data  221 ′ to the standard interface  150  in the standard format. A change to the particular means employed to communicate this data to the earth station may affect the communications  212  between the processor  220  and the transmitter  216 , but will not affect the components beyond the standard interface  150 . 
     To ease the interconnection task, a preferred embodiment of this invention comprises a serial interface, such as RS-485 (also termed multi-drop RS-422) for communicating data, using embedded addresses within each data packet to route each packet. Common networking protocols, such as CANbus, IEEE-1394, I 2 C, Mil-Std 1553/1773, and the like, can also be used. Alternative embodiments of the kernel  200  use different protocols, to facilitate different communications schemes. An embodiment of the kernel  200  that includes the CANbus, for example, provides an easy to use interface for rapid broadcast communications among elements, but is not well suited for large quantities of data communications. An embodiment of the kernel  200  that includes IEEE-1394 is more difficult to interface with, but provides for very high speed data communications. The processor  220  in a preferred embodiment provides the network and protocol functions required to support the network operation via the standard interface  150 . The processor  220  in a preferred embodiment also provides ancillary signaling, such as a standard common clock signal to facilitate a synchronization among spacecraft sub-systems, as required, and discrete digital input/output ports for interface signaling, via the standard interface  150 . As required, alternative processors  220  may be provided, depending upon the processing speed and bandwidth requirements of the overall spacecraft system  100 . In accordance with this invention, however, given the selected protocol, the choice of an alternative processor  220  will be transparent to the interface  150 , and transparent to the vehicle and mission-specific sub-system  300 . 
     By providing the above standard interface  150  and kernel  200 , the mission-specific system  350  can be designed and verified based on a loosely specified allocation of resources. The particular communications equipment  210 , processor  220 , and power management  230  components are selected after the design of the mission-specific system  350  is refined to the extent required to determine actual resource requirements. Thereafter, the remaining sub-systems can be designed or selected, based on fairly well defined sub-system requirements. Note that by providing a well defined interface  150 , the design of the mission-specific sub-system  350  can progress without interference from the dependencies, for example, of the particular choice of communications equipment  210 . If an alternative receiver  218  becomes available that is less costly, for example, it can be utilized without impacting the mission-specific system, because, in accordance with this invention, the communications  221  via the standard interface  150  remain the same. In like manner, an alternative power generation system can be developed without impacting any other component, provided that it provides power in conformance with the aforementioned specifications associated with the interface  150 . That is, by providing a standard interface  150  in accordance with this invention, the effect of a design change in a particular component or sub-system is substantially dampened by the standard interface  150 . Note also that by providing an interface  150  that partitions components and sub-systems based on physical as well as functional dependencies, the interface  150  and kernel sub-system  200  can be embodied in a variety of spacecraft, independent of the spacecraft&#39;s structure or particular mission. 
     Illustrated in FIG. 1 are a variety of components and sub-systems that form the vehicle and mission-specific sub-system  300 . As mentioned above, the antenna system  310  provides the communications  211 ′ to and from an earth station (not shown) from and to the communications system  210  via the standard interface  150 . As in the case of the power generation system  330 , the requirements for the antenna system  310  are contained in a specification that is associated with the standard interface  150 . The specification includes, for example, the maximum allowable routing distance from the interface  150  at a given frequency band, the required signal to noise ratio at the receiver  218 , a minimum and maximum power output from the transmitter  216 , the input or output impedances, and so on. 
     A processor  320  facilitates communications  221 ′ between the mission-specific system  350  and the kernel  200 , via the standard interface  150 . As noted above, by providing the standard interface  150 , the communications system  210  in the kernel  200  can be designed independent of the communications  351  and protocol of the mission-specific system  350 . As also noted above, alternative versions of the protocol used by the standard interface  150  may be provided, each requiring a different level of complexity for translating to and from the standard interface  150 . Thus, the complexity of the processor  320  can range from a mere serial data interface device to a high speed IEEE-1394 (“Firewire”) interface device. Note, however, that the choice of the protocol of the standard interface  150  is loosely based on the quantity of data that is expected to be transmitted, and is preferably made at the commencement of the mission-specific design program. 
     An attitude determination and control system  340  controls the orientation and trajectory of the spacecraft. In a preferred embodiment of this invention the attitude determination and control system  340  is designed to communicate directly  221 ″ with the kernel system  200  via the standard interface  150 , so as to obviate the need for a processor similar to the processor  320  to transform communications to and from different protocols. The selected communications protocol used for data transfer via the standard interface  150  allows for a multiplexing and routing of communications from the earth station to each spacecraft sub-system, such as the attitude determination and control system  340 , and the mission-specific system  350 , using, for example, a packet protocol with destination addressing. 
     Other spacecraft or mission-specific sub-systems  360  are similarly configured to communicate with the kernel system  200 , using the aforementioned standard protocol of the interface  150 , or using a separate processor, or the processor  320 , to effect a communications protocol translation to the standard protocol of the standard interface  150 . These other sub-systems  360  receive their power  231 ′ from the power management system  230 , via the standard interface  150 . 
     Illustrated in FIG. 1 are other kernel task processors  240 . Such processors  240  may include, for example, position and attitude determination aids, processors for determining propulsion parameters, and so on. Other processors  240  provide general purpose services to the spacecraft, and may include, for example, a memory management processor with associated memory, providing each component of the spacecraft  100  a means for storing and retrieving data as required. Such processors  240  in a preferred embodiment are configured to communicate via the protocols established for the standard interface  150 , and will receive power  231 ′ from the standard interface  150 . 
     Because the interface  150  and kernel  200  are well defined, and independent of spacecraft and mission-specific components at both a physical and functional level, a kernel module can be provided that includes one or more of the above defined kernel components. FIG. 2 illustrates an example kernel module  200 P with an integrated standard interface  150 P. In FIG. 2, the “P” suffix on the reference numerals indicate a physical embodiment of the corresponding reference items in FIG.  1 . As discussed above, because the kernel  200  and interface  150  are spacecraft and mission independent, the physical embodiment  200 P,  150 P illustrated in FIG. 2 can be used in a variety of spacecraft, independent of the spacecraft&#39;s configuration and mission. In a preferred embodiment, the embodiment  200 P,  150 P can be made available at the commencement of a spacecraft development process, thereby facilitating the development of spacecraft and mission-specific components and systems with minimal interdependencies and minimal ripple-effect delays and costs. 
     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, as experience is gained in the use and benefits provided by the use of kernel components that are both physically and functionally independent of the spacecraft and mission, other tasks will be partitioned so as conform to this paradigm, and additional kernel processes will be identified and appropriate standards established for accessing these processes via the standard interface  150 . 
     The particular functional partitionings in the figures are presented for illustrative purposes, and alternative partitionings will be evident to one of ordinary skill in the art. For example, the processor  220  may include the functional control components of the power management system  230 . In like manner, if excess capacity is available on the processor  220 , tasks from other spacecraft sub-systems, on either side of the interface  150 , can be effected within the processor  220 . In a preferred embodiment, each of these additional tasks are structured and maintained as independent tasks from the kernel tasks discussed above. Similarly, the various components and sub-systems may be embodied in hardware, software, or a combination of both. For example, the network management task of the processor  220  may be effected via a gate-array device that is preprogrammed to effect the network management, while the interface from the processor  220  to the transmitter  216  may be via a program that is run on a general purpose computing device, and so on. Such modifications and extensions to the concepts presented herein will be evident to one of ordinary skill in the art in light of this disclosure, and within the scope of the following claims.