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Next generation space-based robotics systems will be constructed using distributed architectures where electronics capable of working in the extreme environments of the planets of the solar system are integrated with the sensors and actuators in plug-and-play modules and are connected through common multiple redundant data and power buses.
As power levels of advanced spacecraft climb above 25 kW, higher solar array operating voltages become attractive. Even in today s satellites, operating spacecraft buses at 100 V and above has led to arcing in GEO communications satellites, so the issue of spacecraft charging and solar array arcing remains a design problem. In addition, micrometeoroid impacts on all of these arrays can also lead to arcing if the spacecraft is at an elevated potential. For example, tests on space station hardware disclosed arcing at 75V on anodized A1 structures that were struck with hypervelocity particles in Low Earth Orbit (LEO) plasmas. Thus an understanding of these effects is necessary to design reliable high voltage solar arrays of the future, especially in light of the Vision for Space Exploration of NASA. In the future, large GEO communication satellites, lunar bases, solar electric propulsion missions, high power communication systems around Mars can lead to power levels well above 100 kW. As noted above, it will be essential to increase operating voltages of the solar arrays well above 80 V to keep the mass of cabling needed to carry the high currents to an acceptable level. Thus, the purpose of this paper is to discuss various solar array approaches, to discuss the results of testing them at high voltages, in the presence of simulated space plasma and under hypervelocity impact. Three different types of arrays will be considered. One will be a planar array using thin film cells, the second will use planar single or multijunction cells and the last will use the Stretched Lens Array (SLA - 8-fold concentration). Each of these has different approaches for protection from the space environment. The thin film cell based arrays have minimal covering due to their inherent radiation tolerance, conventional GaAs and multijunction cells have the traditional cerium-doped microsheet glasses (of appropriate thickness) that are usually attached with Dow Corning DC 93-500 silicone adhesive. In practice, these cover glasses and adhesive do not cover the cell edges. Finally, in the SLA, the entire cell and cell edges are fully encapsulated by a cover glass that overhangs the cell perimeter and the silicone adhesive covers the cell edges providing a sealed environment. These three types of blanket technology have been tested at GRC and Auburn. The results of these tests will be described. For example, 15 modules composed of four state-of-the-art 2x4 cm GaAs solar cells with 150 pm cover glasses connected in two-cell series strings were tested at high voltage, in plasma under hypervelocity impact. A picture of one of the modules is shown in figure 1. These were prepared by standard industry practice from a major supplier and had efficiencies above 18%. The test results and other fabrication factors that influenced the tests will be presented. In addition, results for SLA segments tested under the same conditions will be presented. Testing of thin film blankets at GRC will also be presented. Figure 1 : Typical GaAs Solar Cell Module These results will show significant differences in resistance to arcing that are directly related to array design and manufacturing procedures. Finally, the approaches for mitigating the problems uncovered by these tests will be described. These will lay the foundation for future higher voltage array operation, even including voltages above 300-600 V for direct drive SEP applications.
As humans venture farther from Earth for longer durations, it will become essential for those on the journey to have significant control over the scheduling of their own activities as well as the activities of their companion systems and robots. However, the crew will not do all the scheduling; timelines will be the result of collaboration with ground personnel. Emerging technologies such as in-space message buses, delay-tolerant networks, and in-space internet will be the carriers on which the collaboration rides. Advances in scheduling technology, in the areas of task modeling, scheduling engines, and user interfaces will allow the crew to become virtual scheduling experts. New concepts of operations for producing the timeline will allow the crew and the ground support to collaborate while providing safeguards to ensure that the mission will be effectively accomplished without endangering the systems or personnel.
As humans venture farther from Earth for longer durations, it will become essential for those on the journey to have significant control over the scheduling of their own activities as well as the activities of their companion systems and robots. However, the crew will not do all the scheduling; timelines will be the result of collaboration with ground personnel. Emerging technologies such as in-space message buses, delay-tolerant networks, and in-space internet will be the carriers on which the collaboration rides. Advances in scheduling technology, in the areas of task modeling, scheduling engines, and user interfaces will allow the crew to become virtual scheduling experts. New concepts of operations for producing the timeline will allow the crew and the ground support to collaborate while providing safeguards to ensure that the mission will be effectively accomplished without endangering the systems or personnel.
The Deep Space Avionics (DSA) Project is developing a new generation of power system building blocks. Using application specific integrated circuits (ASICs) and power switching modules a scalable power system can be constructed for use on multiple deep space missions including future missions to Mars, comets, Jupiter and its moons. The key developments of the DSA power system effort are five power ASICs and a mod ule for power switching. These components enable a modular and scalab le design approach, which can result in a wide variety of power syste m architectures to meet diverse mission requirements and environments . Each component is radiation hardened to one megarad) total dose. The power switching module can be used for power distribution to regular spacecraft loads, to propulsion valves and actuation of pyrotechnic devices. The number of switching elements per load, pyrotechnic firin gs and valve drivers can be scaled depending on mission needs. Teleme try data is available from the switch module via an I2C data bus. The DSA power system components enable power management and distribution for a variety of power buses and power system architectures employing different types of energy storage and power sources. This paper will describe each power ASIC#s key performance characteristics as well a s recent prototype test results. The power switching module test results will be discussed and will demonstrate its versatility as a multip urpose switch. Finally, the combination of these components will illu strate some of the possible power system architectures achievable fro m small single string systems to large fully redundant systems.
Power-processing unit uses AC buses (30, 32) to supply all current dependent needs such as connections (54, 56) to an ion thruster through an inductor (88) and the primary of a transformer (90), to assure limited currents to such loads. Where temperature control is also required, such as to the main discharge vaporizer heater connection (36, 38), switches (100, 102) are serially connected with inductor (96) and the primary of transformer (98). Temperature sensor (104) controls the switches (100, 102) for temperature regulation.
A document discusses the replication of the functionality of the onboard space-shuttle general-purpose computers (GPCs) in field-programmable gate arrays (FPGAs) and application-specific integrated circuits (ASICs). The purpose of the replication effort is to enable utilization of proven space-shuttle flight software and software-development facilities to the extent possible during development of software for flight computers for a new generation of launch vehicles derived from the space shuttles. The replication involves specifying the instruction set of the central processing unit and the input/output processor (IOP) of the space-shuttle GPC in a hardware description language (HDL). The HDL is synthesized to form a "core" processor in an FPGA or, less preferably, in an ASIC. The core processor can be used to create a flight-control card to be inserted into a new avionics computer. The IOP of the GPC as implemented in the core processor could be designed to support data-bus protocols other than that of a multiplexer interface adapter (MIA) used in the space shuttle. Hence, a computer containing the core processor could be tailored to communicate via the space-shuttle GPC bus and/or one or more other buses.
A paper reports the results of a Hybrid Diesel Vehicle Project focused on a parallel hybrid configuration suitable for diesel-powered, medium-sized, commercial vehicles commonly used for parcel delivery and shuttle buses, as the missions of these types of vehicles require frequent stops. During these stops, electric hybridization can effectively recover the vehicle's kinetic energy during the deceleration, store it onboard, and then use that energy to assist in the subsequent acceleration.
A high power density, modular and state-of-the-art Power System Electronics (PSE) has been developed for the Lunar Reconnaissance Orbiter (LRO) mission. This paper addresses the hardware architecture and performance, the power handling capabilities, and the fabrication technology. The PSE was developed by NASA s Goddard Space Flight Center (GSFC) and is the central location for power handling and distribution of the LRO spacecraft. The PSE packaging design manages and distributes 2200W of solar array input power in a volume less than a cubic foot. The PSE architecture incorporates reliable standard internal and external communication buses, solid state circuit breakers and LiIon battery charge management. Although a single string design, the PSE achieves high reliability by elegantly implementing functional redundancy and internal fault detection and correction. The PSE has been environmentally tested and delivered to the LRO spacecraft for the flight Integration and Test. This modular design is scheduled to flight in early 2009 on board the LRO and Lunar Crater Observation and Sensing Satellite (LCROSS) spacecrafts and is the baseline architecture for future NASA missions such as Global Precipitation Measurement (GPM) and Magnetospheric MultiScale (MMS).
A phased-array system comprised of custom-made and commercially available microelectromechanical system (MEMS) silicon microphones and custom ancillary hardware has been developed for use in aeroacoustic testing in hard-walled and acoustically treated wind tunnels. Recent advances in the areas of multi-channel signal processing and beam forming have driven the construction of phased arrays containing ever-greater numbers of microphones. Traditional obstacles to this trend have been posed by (1) the high costs of conventional condenser microphones, associated cabling, and support electronics and (2) the difficulty of mounting conventional microphones in the precise locations required for high-density arrays. The present development overcomes these obstacles. One of the hallmarks of the new system is a series of fabricated platforms on which multiple microphones can be mounted. These mounting platforms, consisting of flexible polyimide circuit-board material (see left side of figure), include all the necessary microphone power and signal interconnects. A single bus line connects all microphones to a common power supply, while the signal lines terminate in one or more data buses on the sides of the circuit board. To minimize cross talk between array channels, ground lines are interposed as shields between all the data bus signal lines. The MEMS microphones are electrically connected to the boards via solder pads that are built into the printed wiring. These flexible circuit boards share many characteristics with their traditional rigid counterparts, but can be manufactured much thinner, as small as 0.1 millimeter, and much lighter with boards weighing as much as 75 percent less than traditional rigid ones. For a typical hard-walled wind-tunnel installation, the flexible printed-circuit board is bonded to the tunnel wall and covered with a face sheet that contains precise cutouts for the microphones. Once the face sheet is mounted, a smooth surface is established over the entire array due to the flush mounting of all microphones (see right side of figure). The face sheet is made from a continuous glass-woven-fabric base impregnated with an epoxy resin binder. This material offers a combination of high mechanical strength and low dielectric loss, making it suitable for withstanding the harsh test section environment present in many wind tunnels, while at the same time protecting the underlying polyimide board. Customized signal-conditioning hardware consisting of line drivers and antialiasing filters are coupled with the array. The line drivers are constructed using low-supply-current, high-gain-bandwidth operational amplifiers designed to transmit the microphone signals several dozen feet from the array to external acquisition hardware. The anti-alias filters consist of individual Chebyshev low-pass filters (one for each microphone channel) housed on small printed-circuit boards mounted on one or more motherboards. The mother/daughter board design results in a modular system, which is easy to debug and service and which enables the filter characteristics to be changed by swapping daughter boards with ones containing different filter parameters. The filter outputs are passed to commercially- available acquisition hardware to digitize and store the conditioned microphone signals. Wind-tunnel testing of the new MEMS microphone polyimide mounting system shows that the array performance is comparable to that of traditional arrays, but with significantly less cost of construction.
A method is provided for a radiationhardened (rad-hard) solid-state drive for space mission memory applications by combining rad-hard and commercial off-the-shelf (COTS) non-volatile memories (NVMs) into a hybrid architecture. The architecture is controlled by a rad-hard ASIC (application specific integrated circuit) or a FPGA (field programmable gate array). Specific error handling and data management protocols are developed for use in a rad-hard environment. The rad-hard memories are smaller in overall memory density, but are used to control and manage radiation-induced errors in the main, and much larger density, non-rad-hard COTS memory devices. Small amounts of rad-hard memory are used as error buffers and temporary caches for radiation-induced errors in the large COTS memories. The rad-hard ASIC/FPGA implements a variety of error-handling protocols to manage these radiation-induced errors. The large COTS memory is triplicated for protection, and CRC-based counters are calculated for sub-areas in each COTS NVM array. These counters are stored in the rad-hard non-volatile memory. Through monitoring, rewriting, regeneration, triplication, and long-term storage, radiation-induced errors in the large NV memory are managed. The rad-hard ASIC/FPGA also interfaces with the external computer buses.
Software featuring a multilevel architecture is used to control the hardware on the K9 Rover, which is a mobile robot used in research on robots for scientific exploration and autonomous operation in general. The software consists of five types of modules: Device Drivers - These modules, at the lowest level of the architecture, directly control motors, cameras, data buses, and other hardware devices. Resource Managers - Each of these modules controls several device drivers. Resource managers can be commanded by either a remote operator or the pilot or conditional-executive modules described below. Behaviors and Data Processors - These modules perform computations for such functions as planning paths, avoiding obstacles, visual tracking, and stereoscopy. These modules can be commanded only by the pilot. Pilot - The pilot receives a possibly complex command from the remote operator or the conditional executive, then decomposes the command into (1) more-specific commands to the resource managers and (2) requests for information from the behaviors and data processors. Conditional Executive - This highest-level module interprets a command plan sent by the remote operator, determines whether resources required for execution of the plan are available, monitors execution, and, if necessary, selects an alternate branch of the plan.
An extender system for MIL-STD-1553 buses transparently couples bus components at multiple developer sites. The bus network extender is a relatively inexpensive system that minimizes the time and cost of integration of avionic systems by providing a convenient mechanism for early testing without the need to transport the usual test equipment and personnel to an integration facility. This bus network extender can thus alleviate overloading of the test facility while enabling the detection of interface problems that can occur during the integration of avionic systems. With this bus extender in place, developers can correct and adjust their own hardware and software before products leave a development site. Currently resident at Johnson Space Center, the bus network extender is used to test the functionality of equipment that, although remotely located, is connected through a MILSTD- 1553 bus. Inasmuch as the standard bus protocol for avionic equipment is that of MIL-STD-1553, companies that supply MIL-STD-1553-compliant equipment to government or industry and that need long-distance communication support might benefit from this network bus extender
A high-performance design, and fabrication and growth processes to implement the design, have been devised for encapsulating a hydrated salt in a container that both protects the salt and provides thermal conductance between the salt and the environment surrounding the container. The unitary salt/container structure is known in the art as a salt pill. In the original application of the present design and processes, the salt is, more specifically, a hydrated paramagnetic salt, for use as a refrigerant in a very-low-temperature adiabatic demagnetization refrigerator (ADR). The design and process can also be applied, with modifications, to other hydrated salts. Hydrated paramagnetic salts have long been used in ADRs because they have the desired magnetic properties at low temperatures. They also have some properties, disadvantageous for ADRs, that dictate the kind of enclosures in which they must be housed: Being hydrated, they lose water if exposed to less than 100-percent relative humidity. Because any dehydration compromises their magnetic properties, salts used in ADRs must be sealed in hermetic containers. Because they have relatively poor thermal conductivities in the temperature range of interest (<0.1 K), integral thermal buses are needed as means of efficiently transferring heat to and from the salts during refrigeration cycles. A thermal bus is typically made from a high-thermal-conductivity met al (such as copper or gold), and the salt is configured to make intimate thermal contact with the metal. Commonly in current practice (and in the present design), the thermal bus includes a matrix of wires or rods, and the salt is grown onto this matrix. The density and spacing of the conductors depend on the heat fluxes that must be accommodated during operation.
A document describes a high-performance, modular, and state-of-the-art Command and Data Handling (C&DH) system developed for use on the Lunar Reconnaissance Orbiter (LRO) mission. This system implements a complete hardware C&DH subsystem in a single chassis enclosure that includes a processor card, 48 Gbytes of solid-state recorder memory, data buses including MIL-STD-1553B, custom RS-422, SpaceWire, analog collection, switched power services, and interfaces to the Ka-Band and S-Band RF communications systems. The C&DH team capitalized on extensive experience with hardware and software with PCI bus design, SpaceWire networking, Actel FPGA design, digital flight design techniques, and the use of VxWorks for the real-time operating system. The resulting hardware architecture was implemented to meet the LRO mission requirements. The C&DH comprises an enclosure, a backplane, a low-voltage power converter, a single-board computer, a communications interface board, four data storage boards, a housekeeping and digital input/output board, and an analog data acquisition board. The interfaces between the C&DH and the instruments and avionics are connected through a SpaceWire network, a MIL-STD-1553 bus, and a combination of synchronous and asynchronous serial data transfers over RS-422 and LVDS (low-voltage differential-signaling) electrical interfaces. The C&DH acts as the spacecraft data system with an instrument data manager providing all software and internal bus scheduling, ingestion of science data, distribution of commands, and performing science operations in real time.
Over the past few years interest in the development and use of small satellites has rapidly gained momentum with universities, commercial, and government organizations. In a few years we may see networked clusters of dozens or even hundreds of small, cheap, easily replaceable satellites working together in place of the large, expensive and difficult-to-replace satellites now in orbit. Standards based satellite buses and deployment mechanisms, such as the CubeSat and Poly Pico-satellite Orbital Deployer (P-POD), have stimulated growth in this area. The use of small satellites is also proving to be a cost effective capability in many areas traditionally dominated by large satellites, though many challenges remain. Currently many of these small satellites undergo very little testing prior to flight. As these small satellites move from technology demonstration and student projects toward more complex operational assets, it is expected that the standards for verification and validation will increase.
Over the past few years interest in the development and use of small satellites has rapidly gained momentum with universities, commercial, and government organizations. In a few years we may see networked clusters of dozens or even hundreds of small, cheap, easily replaceable satellites working together in place of the large, expensive and difficult-to-replace satellites now in orbit. Standards based satellite buses and deployment mechanisms, such as the CubeSat and Poly Pico-satellite Orbital Deployer (P-POD), have stimulated growth in this area. The use of small satellites is also proving to be a cost effective capability in many areas traditionally dominated by large satellites, though many challenges remain. Currently many of these small satellites undergo very little testing prior to flight. As these small satellites move from technology demonstration and student projects toward more complex operational assets, it is expected that the standards for verification and validation will increase.
The main purpose of the Small Space-Based Geosynchronous Earth orbiting (GEO) satellite is to provide a space link to the user mission spacecraft for relaying data through ground networks to user Mission Control Centers. The Small Space Based Satellite (SSBS) will provide services comparable to those of a NASA Tracking Data Relay Satellite (TDRS) for the same type of links. The SSBS services will keep the user burden the same or lower than for TDRS and will support the same or higher data rates than those currently supported by TDRS. At present, TDRSS provides links and coverage below GEO; however, SSBS links and coverage capability to above GEO missions are being considered for the future, especially for Human Space Flight Missions (HSF). There is also a rising need for the capability to support high data rate links (exceeding 1 Gbps) for imaging applications. The communication payload on the SSBS will provide S/Ka-band single access links to the mission and a Ku-band link to the ground, with an optical communication payload as an option. To design the communication payload, various link budgets were analyzed and many possible operational scenarios examined. To reduce user burden, using a larger-sized antenna than is currently in use by TDRS was considered. Because of the SSBS design size, it was found that a SpaceX Falcon 9 rocket could deliver three SSBSs to GEO. This will greatly reduce the launch costs per satellite. Using electric propulsion was also evaluated versus using chemical propulsion; the power system size and time to orbit for various power systems were also considered. This paper will describe how the SSBS will meet future service requirements, concept of operations, and the design to meet NASA users' needs for below and above GEO missions. These users' needs not only address the observational mission requirements but also possible HSF missions to the year 2030. We will provide the trade-off analysis of the communication payload design in terms of the number of links looking above and below GEO; the detailed design of a GEO SSBS spacecraft bus and its accommodation of the communication payload, and a summary of the trade study that resulted in the selection of the Falcon 9 launch vehicle to deploy the SSBS and its impact on cost reductions per satellite. ======================================================================== Several initiatives have taken place within NASA1 and international space agencies2 to create a human exploration strategy for expanding human presence into the solar system; these initiatives have been driven by multiple factors to benefit Earth. Of the many elements in the strategy one stands out: to send robotic and human missions to destinations beyond Low Earth Orbit (LEO), including cis-lunar space, Near-Earth Asteroids (NEAs), the Moon, and Mars and its moons.3, 4 The time frame for human exploration to various destinations, based on the public information available,1,4 is shown in Figure 1. Advance planning is needed to define how future space communications services will be provided in the new budget environment to meet future space communications needs. The spacecraft for these missions can be dispersed anywhere from below LEO to beyond GEO, and to various destinations within the solar system. NASA's Space Communications and Navigation (SCaN) program office provides communication and tracking services to space missions during launch, in-orbit testing, and operation phases. Currently, SCaN's space networking relay satellites mainly provide services to users below GEO, at Near Earth Orbit (NEO), below LEO, and in deep space. The potential exists for using a space-based relay satellite, located in the vicinity of various solar system destinations, to provide communication space links to missions both below and above its orbit. Such relays can meet the needs of human exploration missions for maximum connectivity to Earth locations and for reduced latency. In the past, several studies assessed the ability of satellite-based relays working above GEO in conjunction with Earth ground stations. Many of these focused on the trade between space relay and direct-to-Earth station links5,6,7. Several others focused on top-level architecture based on relays at various destinations8,9,10,11,12. Much has changed in terms of microwave and optical technology since the publication of the referenced papers; Ka-band communication systems are being deployed, optical communication is being demonstrated, and spacecraft buses are becoming increasingly more functional and operational. A design concept study was undertaken to access the potential for deploying a Small Space-Based Satellite (SSBS) relay capable of serving missions between LEO and NEO. The needs of future human exploration missions were analyzed, and a notional relay-based architecture concept was generated as shown in Fig. 1. Relay satellites in Earth through cis-Lunar orbits are normally located in stable orbits requiring low fuel consumption. Relay satellites for Mars orbit are normally selected based on the mission requirement and projected fuel consumption. Relay satellites have extreme commonalities of functions between them, differing only in the redundancy and frequencies used; therefore, the relay satellite in GEO was selected for further analysis since it will be the first step in achieving a relay-based architecture for human exploration missions (see Fig.Figure 2). The mission design methodology developed by the Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) team13 was used to produce the satellite relay design and to perform various design trades. At the start of the activity, the team was provided with the detailed concept of the notional architecture and the system and communication payload drivers.