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NASA NTRS2003

Toward a Dynamically Reconfigurable Computing and Communication System for Small Spacecraft

Future science missions will require the use of multiple spacecraft with multiple sensor nodes autonomously responding and adapting to a dynamically changing space environment. The acquisition of random scientific events will require rapidly changing network topologies, distributed processing power, and a dynamic resource management strategy. Optimum utilization and configuration of spacecraft communications and navigation resources will be critical in meeting the demand of these stringent mission requirements. There are two important trends to follow with respect to NASA's (National Aeronautics and Space Administration) future scientific missions: the use of multiple satellite systems and the development of an integrated space communications network. Reconfigurable computing and communication systems may enable versatile adaptation of a spacecraft system's resources by dynamic allocation of the processor hardware to perform new operations or to maintain functionality due to malfunctions or hardware faults. Advancements in FPGA (Field Programmable Gate Array) technology make it possible to incorporate major communication and network functionalities in FPGA chips and provide the basis for a dynamically reconfigurable communication system. Advantages of higher computation speeds and accuracy are envisioned with tremendous hardware flexibility to ensure maximum survivability of future science mission spacecraft. This paper discusses the requirements, enabling technologies, and challenges associated with dynamically reconfigurable space communications systems.

Kifle, Muli↗

NASA NTRSDate not supplied

NASA's Small Spacecraft and Distributed Systems: Development and Demonstration of Technologies Enabling Swarms and New Spacecraft Platforms with AI and Edge Computing

NASA’s Small Spacecraft & Distributed Systems (SSDS) within the Research and Technology Mission Directorate (RTMD) expands U.S. capability to execute unique missions through targeted investment, rapid development, and flight demonstration of small spacecraft technologies applicable to exploration, science and the commercial space sector. SSDS strategically invests in technology development and on-orbit demonstrations executed across NASA, other government agencies, industry, and academia. The program’s University SmallSat Technology Partnerships initiative awards academic researchers with the opportunity to collaborate with NASA to mature innovative technology. Capabilities aligned with RTMD’s technology shortfalls and interests - power, processing, propulsion, sensors, communications, autonomous navigation, architectures, and advanced applications like artificial intelligence (AI), machine learning, and edge computing - are prioritized in SSDS investments. These investments enable distributed, autonomous, and cooperative small spacecraft systems that support swarm missions extending beyond low Earth orbit into cislunar and deep space. This paper highlights representative SSDS flight demonstrations that mature these capabilities to enable a future operational infrastructure needed to support sustained exploration of the Moon and beyond. SSDS’s investment strategy emphasizes rapid development and on-orbit demonstration to validate spacecraft technologies required for swarms and distributed mission architectures. The Starling swarm technology demonstration mission exemplifies this approach by advancing distributed spacecraft autonomy, cooperative operations, and space situational awareness. Extended flight testing and ongoing studies of next generation swarm configurations and on-orbit space traffic monitoring and management continue to inform future swarm designs. DiskSat’s four-spacecraft demonstration mission represents SSDS’s strategic vision to expand the design space for future small spacecraft through its commitment to advance novel platform concepts that can impact how science is performed on orbit. Continuing to invest in future platforms, the notional PY12 concept is a 12-spacecraft swarm hosting neuromorphic processors and is envisioned as an on-orbit testbed for AI, edge computing, and positioning, navigation and timing technologies. SSDS also invests in single-spacecraft technology demonstrations that underpin the success of future swarm missions and accelerate the availability of validated technologies across the small spacecraft ecosystem. Examples of such demonstrations include Pathfinder Technology Demonstrator-3 (PTD-3), which performed high-rate optical communications; PTD-R, which demonstrated a camera capable of simultaneous ultraviolet and short-wave infrared optical sensing; and CAPSTONE, the Cislunar Autonomous Positioning System Technology and Operations Navigation Experiment, which validated autonomous navigation in cislunar space. Collectively, SSDS-funded demonstrations advance capabilities across swarms and illustrate a coordinated investment strategy to mature high-impact technologies required for autonomous, distributed, and cooperative small spacecraft systems for low Earth orbit, cislunar, and deep space applications. Technology demonstrations strengthen SSDS partnerships with industry, academia, and other government agencies, and promote small spacecraft community adoption of capabilities required to close technical gaps for swarm missions.

Jan Stupl↗

NASA NTRSDate not supplied

NASA's Small Spacecraft and Distributed Systems: Development and Demonstration of Technologies Enabling Swarms and New Spacecraft Platforms with AI and Edge Computing

Jan Stupl↗

NASA NTRS1990

Flight elements subpanel

A summary of findings of the Flight Elements Subpanel is given in viewgraph form. Topics covered include advanced avionics architecture; processors; integrated Global Positioning System and guidance, navigation and control; displays and controls; telemetry; sensors and instrumentation; fault detection and management; electrical power distribution and control; spacecraft power supplies; and in-flight crew training.

Keckler, Claude R.↗

NASA NTRS2000

Power System Options Evaluated for the Radiation and Technology Demonstration Mission

The Radiation and Technology Demonstration (RTD) Mission is under joint study by three NASA Centers: the NASA Johnson Space Center, the NASA Goddard Space Flight Center, and the NASA Glenn Research Center at Lewis Field. This Earth-orbiting mission, which may launch on a space shuttle in the first half of the next decade, has the primary objective of demonstrating high-power electric thruster technologies. Secondary objectives include better characterization of Earth's Van Allen trapped-radiation belts, measurement of the effectiveness of the radiation shielding for human protection, measurement of radiation effects on advanced solar cells, and demonstration of radiation-tolerant microelectronics. During the mission, which may continue up to 1 year, the 2000-kg RTD spacecraft will first spiral outward from the shuttle-deployed, medium-inclination, low Earth orbit. By the phased operation of a 10-kW Hall thruster and a 10-kW Variable Specific Impulse Magneto-Plasma Rocket, the RTD spacecraft will reach a low-inclination Earth orbit with a radius greater than five Earth radii. This will be followed by an inward spiraling orbit phase when the spacecraft deploys 8 to 12 microsatellites to map the Van Allen belts. The mission will conclude in low Earth orbit with the possible retrieval of the spacecraft by the space shuttle. A conceptual RTD spacecraft design showing two photovoltaic (PV) array wings, the Hall thruster with propellant tanks, and stowed microsatellites is presented. Early power system studies assessed five different PV array design options coupled with a 120-Vdc power management and distribution system (PMAD) and secondary lithium battery energy storage. Array options include (1) state-of-the-art 10-percent efficient three-junction amorphous SiGe thin-film cells on thin polymer panels deployed with an inflatable (or articulated) truss, (2) SCARLET array panels, (3) commercial state-of-the-art, planar PV array rigid panels with 25-percent efficient, three-junction GaInP2/GaAs/Ge solar cells, (4) rigid panels with 25-percent efficient, three-junction GaInP2/GaAs/Ge solar cells, in a 2 -concentrator trough configuration, and (5) thin polymer panels with 25-percent efficient, three-junction GaInP2/GaAs/Ge solar cells deployed with an inflatable (or articulated) truss. To assess the relative merits of these PV array design options, the study group developed a dedicated Fortran code to predict power system performance and estimate system mass. This code also modeled Earth orbital environments important for accurately predicting PV array performance. The most important environmental effect, solar cell radiation degradation, was calculated from electron-proton fluence input from the industry standard AE8/AP8 trapped radiation models and the concept of damage equivalence. Power systems were sized to provide 10 kW of thruster power and approximately 1 kW of spacecraft power at end of life. Of the five PV array design options, the option 1 (thin-film cells) power system was the most massive 590 kg, whereas the option 4 (trough concentrator) power system was the lightest 260 kg. Arguably, the lowest cost would come from the option 3 (commercial array panels) power system with an acceptable, albeit greater, system mass of 320 kg. Predicted power system performance during the spiral-out mission phase is shown the preceding graph for the option 5 (flexible-panel) array. From the results, the radiation-induced power loss over time is evident as the spacecraft slowly spirals outward through the trapped proton belt. The importance of the spiral trip time is also evident in the two curves representing 74-day and 182-day spiral-out periods. The longer spiral time introduces a beginning-of-life power oversizing penalty greater than 1 kW. Future studies will analyze power system performance and mass with a 50-Vdc power management and distribution architecture favorable to the VASIMR thruster and longer missions.

Kerslake, Thomas W.↗

NASA NTRS2019

Cold Survivable Distributed Motor Controller (CSDMC)

This paper presents the results of NASA’s COLDTECH development entitled “Cold Survivable Distributed Motor Controller (CSDMC)”. This work addresses the need to lower the mass, power and volume of the motor control electronics and its associated cabling. Landed payload mass of ocean world missions typically requires a spacecraft launch mass of 7-10x the landed mass due to the required propellant to get the payload to the surface. Reduction of landed mass leads to cheaper, more frequent missions and/or increased science return. This work addresses this need by developing a distributed electronics architecture, which places control and power electronics near or at actuators and instruments. The outcome of this effort will result in a 10X reduction in harness mass, enabling a significant increase in science payload which then enables more capable sample acquisition, delivery and analysis systems on these missions. Placing the control and power conversion electronics at or near the actuators or instruments is the cornerstone of our distributed architecture. To do this, we developed the technology necessary to distribute the electronics and place them on a shared interface and power bus. This enables a significant reduction in cable mass along with its associated complexity. This allows spacecraft designers to take advantage of volume at the extremities that would normally not be utilized. In this paper we discuss the technologies and system design to achieve these goals in support of ocean world missions. These technologies include the development of our motor control modules, a point of load regulator and isolated converter modules along with the packaging technology necessary to allow our electronics to survive the extreme temperatures.

Stell, Chris↗

NASA NTRS1982

High voltage distribution and grounding in high power spacecraft

Many space missions proposed for the time period from 1985 to 2000 will require large spacecraft to support the onboard loads. In some cases, large electrical power systems will be needed to supply the electrical/electronic equipment loads. These electronic systems will be used for communications, radar, and experimental equipment for aid to earth's overcrowded communication systems, exploration of new energy resources, space exploration, and eventually to supplement terrestrial electric power utilities. For the near term (1985-1990), some of these systems have power levels to 50 kW. The long-term programs, 1990 to post-2000, could possibly have demands in the order of multimegawatts. The problems which have to be solved to construct the required high-voltage power supply systems are considered. Data and conceptual designs generated are found to indicate that grounding and bonding for high power systems can be accomplished in spacecraft by using either manual or automatic joining of the structural members.

Dunbar, W. G.↗

NASA NTRSDate not supplied

Eleven Countries, an Integrated Spacecraft: the Story of International Collaboration that Built the Orion Spacecraft and Powered the Success of the Artemis I Mission

The quest to return humans to the Moon in the next step towards humanity's exploration of space is more alive than ever. After a great deal of achievements, failures, and lessons learned, the Artemis I mission set o to the Moon on November 16, 2022, with the goal of testing a new rocket, the Space Launch System, and a new spacecraft, Orion: designed, assembled, and tested across two continents, and 11 countries. Behind this mission, decades of experience with the International Space Station, Autonomous Transfer Vehicle operations, and many other program collaborations built the know-how on how to succeed together in the toughest environment | deep space. The Artemis I mission proved to be an incredible success, meeting 161 total mission objectives, including 21 developed during the flight based on outperforming spacecraft. It was also a case-study in international collaboration, given that ESA, NASA, and industry partners Airbus and Lockheed Martin for the first time had to design, build, test, and fly a fully integrated human-rated spacecraft, with most critical functions dependent and interconnected across U.S. and European systems. The U.S.-built Orion Crew Module and Crew Module Adapter and European-built European Service Module (ESM) shared critical interfaces and commodities, from propulsion, avionics, active/passive thermal, electrical power generation, storage and distribution to the software that managed it all. In this paper, we will describe relevant aspects of the integrated spacecraft design, providing context for the challenges that the team faced in all phases required to get Orion ready to fly, and provide a direct account of how the joint team formed, trained, and supported the operations of the successful Artemis I mission. We will also explore the evolution of the partnerships, given that these allow a multi-national e ort to sustain the program production, share costs, leverage a broader base of engineering expertise, and build more diverse capabilities over the long haul to support the Artemis goals and objectives. Lastly, we will cover critical lessons learned and how the Orion Program has implemented these in preparation of the next Artemis missions to repeat the success of Artemis I. The purpose of this paper is to document knowledge we gained and lessons we learned through the development of an integrated Orion spacecraft, since it is imperative we build on this now, at the dawn of the Artemis Program, an international endeavor to push human space exploration.

Deep Space Exploration↗

NASA NTRS2006

Modular, Reconfigurable, High-Energy Technology Development

The Modular, Reconfigurable High-Energy (MRHE) Technology Demonstrator project was to have been a series of ground-based demonstrations to mature critical technologies needed for in-space assembly of a highpower high-voltage modular spacecraft in low Earth orbit, enabling the development of future modular solar-powered exploration cargo-transport vehicles and infrastructure. MRHE was a project in the High Energy Space Systems (HESS) Program, within NASA's Exploration Systems Research and Technology (ESR&T) Program. NASA participants included Marshall Space Flight Center (MSFC), the Jet Propulsion Laboratory (JPL), and Glenn Research Center (GRC). Contractor participants were the Boeing Phantom Works in Huntsville, AL, Lockheed Martin Advanced Technology Center in Palo Alto, CA, ENTECH, Inc. in Keller, TX, and the University of AL Huntsville (UAH). MRHE's technical objectives were to mature: (a) lightweight, efficient, high-voltage, radiation-resistant solar power generation (SPG) technologies; (b) innovative, lightweight, efficient thermal management systems; (c) efficient, 100kW-class, high-voltage power delivery systems from an SPG to an electric thruster system; (d) autonomous rendezvous and docking technology for in-space assembly of modular, reconfigurable spacecraft; (e) robotic assembly of modular space systems; and (f) modular, reconfigurable distributed avionics technologies. Maturation of these technologies was to be implemented through a series of increasingly-inclusive laboratory demonstrations that would have integrated and demonstrated two systems-of-systems: (a) the autonomous rendezvous and docking of modular spacecraft with deployable structures, robotic assembly, reconfiguration both during assembly and (b) the development and integration of an advanced thermal heat pipe and a high-voltage power delivery system with a representative lightweight high-voltage SPG array. In addition, an integrated simulation testbed would have been developed containing software models representing the technologies being matured in the laboratory demos. The testbed would have also included models for non-MRHE developed subsystems such as electric propulsion, so that end-to-end performance could have been assessed. This paper presents an overview of the MRHE Phase I activities at MSFC and its contractor partners. One of the major Phase I accomplishments is the assembly demonstration in the Lockheed Martin Advanced Technology Center (LMATC) Robot-Satellite facility, in which three robot-satellites successfully demonstrated rendezvous & docking, self-assembly, reconfiguration, adaptable GN&C, deployment, and interfaces between modules. Phase I technology maturation results from ENTECH include material recommendations for radiation hardened Stretched Lens Array (SLA) concentrator lenses, and a design concept and test results for a hi-voltage PV receiver. UAH's accomplishments include Supertube heatpipe test results, which support estimates of thermal conductivities at 30,000 times that of an equivalent silver rod. MSFC performed systems trades and developed a preliminary concept design for a 100kW-class modular reconfigurable solar electric propulsion transport vehicle, and Boeing Phantom Works in Huntsville performed assembly and rendezvous and docking trades. A concept animation video was produced by SAIC, wllich showed rendezvous and docking and SLA-square-rigger deployment in LEO.

Carrington, Connie↗

NASA NTRS1990

Development of an automated electrical power subsystem testbed for large spacecraft

The NASA Marshall Space Flight Center (MSFC) has developed two autonomous electrical power system breadboards. The first breadboard, the autonomously managed power system (AMPS), is a two power channel system featuring energy generation and storage and 24-kW of switchable loads, all under computer control. The second breadboard, the space station module/power management and distribution (SSM/PMAD) testbed, is a two-bus 120-Vdc model of the Space Station power subsystem featuring smart switchgear and multiple knowledge-based control systems. NASA/MSFC is combining these two breadboards to form a complete autonomous source-to-load power system called the large autonomous spacecraft electrical power system (LASEPS). LASEPS is a high-power, intelligent, physical electrical power system testbed which can be used to derive and test new power system control techniques, new power switching components, and new energy storage elements in a more accurate and realistic fashion. LASEPS has the potential to be interfaced with other spacecraft subsystem breadboards in order to simulate an entire space vehicle. The two individual systems, the combined systems (hardware and software), and the current and future uses of LASEPS are described.

Hall, David K.↗

NASA NTRS2012

A Spacecraft Housekeeping System-on-Chip in a Radiation Hardened Structured ASIC

Housekeeping systems are essential to health monitoring of spacecraft and instruments. Typically, sensors are distributed across various sub-systems and data is collected using components such as analog-to-digital converters, analog multiplexers and amplifiers. In most cases programmable devices are used to implement the data acquisition control and storage, and the interface to higher level systems. Such discrete implementations require additional size, weight, power and interconnect complexity versus an integrated circuit solution, as well as the qualification of multiple parts. Although commercial devices are readily available, they are not suitable for space applications due the radiation tolerance and qualification requirements. The Housekeeping System-o n-A-Chip (HKSOC) is a low power, radiation hardened integrated solution suitable for spacecraft and instrument control and data collection. A prototype has been designed and includes a wide variety of functions including a 16-channel analog front-end for driving and reading sensors, analog-to-digital and digital-to-analog converters, on-chip temperature sensor, power supply current sense circuits, general purpose comparators and amplifiers, a 32-bit processor, digital I/O, pulse-width modulation (PWM) generators, timers and I2C master and slave serial interfaces. In addition, the device can operate in a bypass mode where the processor is disabled and external logic is used to control the analog and mixed signal functions. The device is suitable for stand-alone or distributed systems where multiple chips can be deployed across different sub-systems as intelligent nodes with computing and processing capabilities.

Suarez, George↗

NASA NTRS2001

Flight Computer Design for the Space Technology 5 (ST-5) Mission

As part of NASA's New Millennium Program, the Space Technology 5 mission will validate a variety of technologies for nano-satellite and constellation mission applications. Included are: a miniaturized and low power X-band transponder, a constellation communication and navigation transceiver, a cold gas micro-thruster, two different variable emittance (thermal) controllers, flex cables for solar array power collection, autonomous groundbased constellation management tools, and a new CMOS ultra low-power, radiation-tolerant, +0.5 volt logic technology. The ST-5 focus is on small and low-power. A single-processor, multi-function flight computer will implement direct digital and analog interfaces to all of the other spacecraft subsystems and components. There will not be a distributed data system that uses a standardized serial bus such as MIL-STD-1553 or MIL-STD-1773. The flight software running on the single processor will be responsible for all real-time processing associated with: guidance, navigation and control, command and data handling (C&DH) including uplink/downlink, power switching and battery charge management, science data analysis and storage, intra-constellation communications, and housekeeping data collection and logging. As a nanosatellite trail-blazer for future constellations of up to 100 separate space vehicles, ST-5 will demonstrate a compact (single board), low power (5.5 watts) solution to the data acquisition, control, communications, processing and storage requirements that have traditionally required an entire network of separate circuit boards and/or avionics boxes. In addition to the New Millennium technologies, other major spacecraft subsystems include the power system electronics, a lithium-ion battery, triple-junction solar cell arrays, a science-grade magnetometer, a miniature spinning sun sensor, and a propulsion system.

Speer, David↗

NASA NTRS2001

ISS And Space Environment Interactions Without Operating Plasma Contactor

The International Space Station (ISS) will be the largest, highest power spacecraft placed in orbit. Because of this the design of the electrical power system diverged markedly from previous systems. The solar arrays will operate at 160 V and the power distribution voltage will be 120 V. The structure is grounded to the negative side of the solar arrays so under the right circumstances it is possible to drive the ISS potential very negative. A plasma contactor has been added to the ISS to provide control of the ISS structure potential relative to the ambient plasma. The ISS requirement is that the ISS structure not be greater than 40 V positive or negative of local plasma. What are the ramifications of operating large structures with such high voltage power systems? The application of a plasma contactor on ISS controls the potential between the structure and the local plasma, preventing degrading effects. It is conceivable that there can be situations where the plasma contactor might be non-functional. This might be due to lack of power, the need to turn it off during some of the build-up sequences, the loss of functionality for both plasma contactors before a replacement can be installed, similar circumstances. A study was undertaken to understand how important it is to have the contactor functioning and how long it might be off before unacceptable degradation to ISS could occur. The details of interaction effects on spacecraft have not been addressed until driven by design. This was true for ISS. If the structure is allowed to float highly negative impinging ions can sputter exposed conductors which can degrade the primary surface and also generate contamination due to the sputtered material. Arcing has been known to occur on solar arrays that float negative of the ambient plasma. This can also generate electromagnetic interference and voltage transients. Much of the ISS structure and pressure module surfaces exposed to space is anodized aluminum. The anodization thickness is very thin to provide the required solar absorptance and emittance. For conditions where ISS structure can charge negative a large percentage of the array voltage, the dielectric strength of this layer is low, and dielectric breakdown (arcing) can occur. The energy stored capacitively in the structure can be delivered to the arc. The mechanisms by which this energy is delivered and how much of the energy is available hasn't been fully quantified. Questions have been raised regarding the possibility of whether a sustained arc might result due to current collected by the solar arrays from local plasma. It was postulated that even if dielectric breakdown didn't occur, impacts due to micrometeoroids and space debris could penetrate thin layers of dielectric on ISS and initiate an arc due to the coupling provided by the dense local plasma produced by the impact. This was proven in experiments conducted jointly by MSFC and Auburn University. A target chamber with a simulated ionospheric plasma and a biased, anodized aluminum plate and a 1-microfarad capacitor was used. The plate was then impacted by 75-micron particles accelerated to orbital velocity. Arc discharges were sustained for higher voltages but a threshold appears below which no discharge was initiated. Most items without an exposed power system will float electrically near the local plasma potential. This is true of the Space Shuttle, an Astronaut on EVA, and similar items. The structure of ISS might be at a large negative voltage. Therefore, capacitively stored energy can be transferred during docking, installing external boxes and equipment and Astronaut contact with ISS structure. The circumstances of when this can happen and the resulting effects are evaluated in this study. Also, a crewmember on EVA might be in the vicinity of an arc. All safety aspects of such an encounter including charging, molten particles from the arc site and EMI have been evaluated. This paper will report on the total results of this study focussed on the 4A configuration, scheduled to be complete in November, 2000. Interactions such as arcing, debris induced arcs, sustained arcs, sputtering, contamination from sputtering and arcing, docking interactions and Astronaut safety issues will all be addressed.

Carruth, M. R., Jr.↗

NASA NTRS2017

Distributed Timing and Localization (DiGiTaL)

The Distributed Timing and Localization (DiGiTaL) system provides nano satellite formations with unprecedented,centimeter-level navigation accuracy in real time and nanosecond-level time synchronization. This is achieved through the integration of a multi-constellation Global Navigation Satellite System (GNSS) receiver, a Chip-Scale Atomic Clock (CSAC), and a dedicated Inter-Satellite Link (ISL). In comparison, traditional single spacecraft GNSS navigation solutions are accurate only to the meter-level due to the sole usage of coarse pseudo-range measurements. To meet the strict requirements of future miniaturized distributed space systems, DiGiTaL uses powerful error-cancelling combinations of raw carrier-phase measurements which are exchanged between the swarming nano satellites through a decentralized network. A reduced-dynamics estimation architecture on board each individual nano satellite processes the resulting millimeter-level noise measurements to reconstruct the fullformation state with high accuracy.

Small Spacecraft↗

NASA NTRS1990

Power system state estimation for a spacecraft power system

An application of the maximum likelihood state estimator to a space-based power system is presented. The state estimator uses current and voltage measurements to generate estimates of node voltages for an electrical power distribution system for the Space Shuttle. Preliminary results on the effect of noisy measurements on estimated parameters are reported. The software used in generating these results is part of an overall package being developed at Louisiana Tech University. Intended applications of this package include the analysis of power systems and real-time parallel processing on the Space Shuttle.

Berry, F. C.↗

NASA NTRS2000

ISS and Space Environment Interactions in Event of Plasma Contactor Failure

The International Space Station (ISS), illustrated in Figure 1, will be the largest, highest power spacecraft placed in orbit. Because of this the design of the electrical power system diverged markedly from previous systems. The solar arrays will operate at 160 V and the power distribution voltage will be 120 V. The structure is grounded to the negative side of the solar arrays so under the right circumstances it is possible to drive the ISS potential very negative. A plasma contactor has been added to the ISS to provide control of the ISS structure potential relative to the ambient plasma. The ISS requirement is that the ISS structure not be greater than 40 V positive or negative of local plasma. What are the ramifications of operating large structures with such high voltage power systems? The application of a plasma contactor on ISS controls the potential between the structure and the local plasma, preventing degrading effects. It is conceivable that there can be situations where the plasma contactor might be non-functional. This might be due to lack of power, the need to turn it off during some of the build-up sequences, the loss of functionality for both plasma contactors before a replacement can be installed, and similar circumstances. A study was undertaken to understand how important it is to have the contactor functioning and how long it might be off before unacceptable degradation to ISS could occur.

Carruth, M. R., Jr.↗

NASA NTRS1992

Project APEX: Advanced Phobos Exploration. Manned mission to the Martian moon Phobos

The manned exploration of Mars is a massive undertaking which requires careful consideration. A mission to the moon of Mars called Phobos as a prelude to manned landings on the Martian surface offers some advantages. One is that the energy requirements, in terms of delta 5, is only slightly higher than going to the Moon's surface. Another is that Phobos is a potential source of water and carbon which could be extracted and processed for life support and cryogenic propellants for use in future missions; thus, Phobos might serve as a base for extended Mars exploration or for exploration of the outer planets. The design of a vehicle for such a mission is the subject of our Aerospace System Design course this year. The materials and equipment needed for the processing plant would be delivered to Phobos in a prior unmanned mission. This study focuses on what it would take to send a crew to Phobos, set up the processing plant for extraction and storage of water and hydrocarbons, conduct scientific experiments, and return safely to Earth. The size, configuration, and subsystems of the vehicle are described in some detail. The spacecraft carries a crew of five and is launched from low Earth orbit in the year 2010. The outbound trajectory to Mars uses a gravitational assisted swing by of Venus and takes eight months to complete. The stay at Phobos is 60 days at which time the crew will be engaged in setting up the processing facility. The crew will then return to Earth orbit after a total mission duration of 656 days. Both stellar and solar observations will be conducted on both legs of the mission. The design of the spacecraft addresses human factors and life science; mission analysis and control; propulsion; power generation and distribution; thermal control; structural analysis; and planetary, solar, and stellar science. A 0.5 g artificial gravity is generated during transit by spinning about the lateral body axis. Nuclear thermal rockets using hydrogen as fuel are selected to reduce total launch mass and to shorten the duration of the mission. The nuclear systems also provide the primary electrical power via dual mode operation. The overall spacecraft length is 110 meters and the total mass departing from low Earth orbit is 900 metric tons.

Source record↗

NASA NTRS1990

Implementation strategies for load center automation on the space station module/power management and distribution testbed

The Space Station Module/Power Management and Distribution (SSM/PMAD) testbed was developed to study the tertiary power management on modules in large spacecraft. The main goal was to study automation techniques, not necessarily develop flight ready systems. Because of the confidence gained in many of automation strategies investigated, it is appropriate to study, in more detail, implementation strategies in order to find better trade-offs for nearer to flight ready systems. These trade-offs particularly concern the weight, volume, power consumption, and performance of the automation system. These systems, in their present implementation are described.

Watson, Karen↗