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NASA NTRS1991

Description of the PMAD DC test bed architecture and integration sequence

NASA-LEWIS is responsible for the development, fabrication, and assembly of the electric power system (EPS) for the Space Station Freedom (SSF). The SSF power system is radically different from previous spacecraft power systems in both the size and complexity of the system. Unlike past spacecraft power systems, the SSF EPS will grow and be maintained on orbit and must be flexible to meet challenging user power needs. The SSF power system is also unique in comparison with terrestrial power systems because it is dominated by power electronic converters which regulate and control the power. A description is provided of the Power Management and Distribution DC Testbed which was assembled to support the design and early evaluation of the SSF EPS. A description of the integration process used in the assembly sequence is also given along with a description of the support facility.

Beach, R. F.↗

NASA NTRS2007

Spacecraft Power Systems Engineering: Solutions for NASA's Manned Space Program

An overview of spacecraft power systems is presented, with a focus on applications in the manned space program. The topics include: 1) History; 2) State-of-the-art; 3) Development directions; 4) Focus on applications in the manned space program led from JSC; 5) Power Systems Engineering Trade Space; 6) Power Generation and Energy Storage; 7) Power Distribution and Control; and 8) Actuation

Scott, John H.↗

NASA NTRS2023

Development of Two High-Energy Bus ‘Cores’ for Rapid Support of Low-TRL and Educational Payloads: A Software-Configured EPS Combined with Flexible C&DH

For several years the TechEdSat flight series (TES-n), developed by the Nano Orbital Workshop (NOW) group at NASA Ames, has relied upon an in-house developed unit to serve both EPS (Electrical Power System) and C&DH (Command and Data Handling) roles along with low data-rate telemetry functions, i.e., serving as the ‘core’ of the spacecraft bus. This ‘core’ has a considerable task given the rapid cadence of the TES program and the typically low-TRL of payloads; configurability and compatibility are key to prevent mission-specific hardware. However, at only 15 watts the current core has become insufficient to support the program’s growing missions and increasingly demanding payloads. To this end, the NOW program is developing new cores to support two TES mission classes: a single-PCB ‘MiniCore’ designed to support 80-watt missions 6U or smaller in LEO, and a three-PCB, radiation-tolerant ‘StackCore’ designed to support 6U and larger missions over 500 watts in LEO and beyond. The ‘MiniCore’ design consists of three main segments: a processor-agnostic C&DH, a software-configured EPS, and a backup low data-rate radio. The design philosophy was to enable rapid-manufacture in a turbulent supply chain, hence the design consists of COTS parts with a focus on those able to be drop-in replaced with radiation-tolerant versions when demanded by the mission. As a single PC-104 sized circuit board, power density and ease of integration also dominated design, demanding the use of modern features such as single-point USB-C for easy charging and monitoring of the spacecraft on the ground. The ‘MiniCore’ can support 80 watts of load, 140 watt-hours of storage, and over 20 watts of optimized solar generation with extensive power monitoring throughout. The ‘MiniCore’ supports one battery pack, six solar-panels, six loads, five actuators, Iridium SBD, and an internal 802.15.4 network. Additionally, the processor-agnostic design can accept any PJRC Teensy 3.x or Adafruit Feather microcontroller unit to enable processor scaling with mission requirements or environment. It is expected a development unit of this design will be completed before conference. The ‘StackCore’ design consists of three stacked PC-104 sized circuit boards: one dedicated to power generation and storage, one dedicated to power distribution, and one dedicated to C&DH tasks. This delineation is necessary to support the transition from highly integrated ICs to discrete analog circuitry, enabling a primarily analog control power system able to operate without software in a radiation environment with finer monitoring compared to the ‘MiniCore’ design. The planned base architecture supports over 500 watts of load, 250 watt-hours of storage, and over 80 watts of optimized solar generation. The power distribution board allows for the use of daughter cards hosting custom converters or interfaces for payloads, in addition to the software-configured supplies used on the ‘MiniCore’. This core stack will be managed by a Vorago ARM M4 microcontroller and support the same wireless communications as the ‘MiniCore’, with optional integration of a NOW S-band radio and attitude determination sensors for ‘black box’ functionality. It is expected the prototype will still be in development during conference.

Spacecraft↗

NASA NTRSDate not supplied

Development of Two High-Energy Bus ‘Cores’ for Rapid Support of Low-TRL and Educational Payloads: A Software-Configured EPS Combined with Flexible C&DH

Spacecraft↗

NASA NTRS2017

NASA's Space Launch System: Deep-Space Delivery for Smallsats

Designed for human exploration missions into deep space, NASA's Space Launch System (SLS) represents a new spaceflight infrastructure asset, enabling a wide variety of unique utilization opportunities. While primarily focused on launching the large systems needed for crewed spaceflight beyond Earth orbit, SLS also offers a game-changing capability for the deployment of small satellites to deep-space destinations, beginning with its first flight. Currently, SLS is making rapid progress toward readiness for its first launch in two years, using the initial configuration of the vehicle, which is capable of delivering 70 metric tons (t) to Low Earth Orbit (LEO). On its first flight test of the Orion spacecraft around the moon, accompanying Orion on SLS will be small-satellite secondary payloads, which will deploy in cislunar space. The deployment berths are sized for "6U" CubeSats, and on EM-1 the spacecraft will be deployed into cislunar space following Orion separate from the SLS Interim Cryogenic Propulsion Stage. Payloads in 6U class will be limited to 14 kg maximum mass. Secondary payloads on EM-1 will be launched in the Orion Stage Adapter (OSA). Payload dispensers will be mounted on specially designed brackets, each attached to the interior wall of the OSA. For the EM-1 mission, a total of fourteen brackets will be installed, allowing for thirteen payload locations. The final location will be used for mounting an avionics unit, which will include a battery and sequencer for executing the mission deployment sequence. Following the launch of EM-1, deployments of the secondary payloads will commence after sufficient separation of the Orion spacecraft to the upper stage vehicle to minimize any possible contact of the deployed CubeSats to Orion. Currently this is estimated to require approximately 4 hours. The allowed deployment window for the CubeSats will be from the time the upper stage disposal maneuvers are complete to up to 10 days after launch. The upper stage will fly past the moon at a perigee of approximately 100km, and this closest approach will occur about 5 days after launch. The limiting factor for the latest deployment time is the available power in the sequencer system. Several NASA Mission Directorates were involved in the development of programs for the competition, selection, and development of EM-1 payloads that support directorate priorities. CubeSat payloads on EM-1 will include both NASA research experiments and spacecraft developed by industry, international and potentially academia partners. The Human Exploration and Operations Mission Directorate (HEOMD) Advanced Exploration Systems (AES) Division was allocated five payload opportunities on the EM-1 mission. Near Earth Asteroid (NEA) Scout is designed to rendezvous with and characterize a candidate NEA. A solar sail, an innovation the spacecraft will demonstrated for the CubeSat class, will provide propulsion. Lunar Flashlight will use a green propellant system and will search for potential ice deposits in the moon's permanently shadowed craters. BioSentinel is a yeast radiation biosensor, planned to measure the effects of space radiation on deoxyribonucleic acid (DNA). Lunar Icecube, a collaboration with Morehead State University, will prospect for water in ice, liquid, and vapor forms as well as other lunar volatiles from a low-perigee, highly inclined lunar orbit using a compact Infrared spectrometer. Skyfire, a partnership with Lockheed Martin, is a technology demonstration mission that will perform a lunar flyby, collecting spectroscopy, and thermography data to address questions related to surface characterization, remote sensing, and site selection. NASA's Space Technology Mission Directorate (STMD) was allocated three payload opportunities on the EM-1 mission. These slots will be filled via the Centennial Challenges Program, NASA's flagship program for technology prize competitions, which directly engages the public, academia, and industry in open prize competitions to stimulate innovation. The NASA Science Mission Directorate (SMD) was allocated two payload opportunities on the EM-1 mission. The CubeSat Mission to Study Solar Particles (CuSP) payload will study the sources and acceleration mechanisms of solar and interplanetary particles in near-Earth orbit, support space weather research by determining proton radiation levels during Solar Energetic Particle (SEP) events and identifying suprathermal properties that could help predict geomagnetic storms. The LunaH-Map payload will help scientists understand the quantity of H-bearing materials in lunar cold traps (~10 km), determine the concentration of H-bearing materials with 1m depth, and constrain the vertical distribution of H-bearing materials. The final three payload opportunities for the EM-1 mission were allocated for NASA's international space agency counterparts. The flight opportunities are intended to benefit the international space agency and NASA as well as further the collective space exploration goals. ArgoMoon is sponsored by ESA/ASI and will fly along with the ICPS on its disposal trajectory to perform proximity operations with the ICPS post-disposal, take external imagery of engineering and historical significance, and perform an optical communications demonstration. EQUULEUS, sponsored by JAXA, will fly to a libration orbit around the Earth-Moon L2 point and demonstrate trajectory control techniques within the Sun-Earth- Moon region for the first time by a nano spacecraft. The mission will also contribute to the future human exploration scenario by understanding the radiation environment in geospace and deep space, characterizing the flux of impacting meteors on the far side of the moon, and demonstrating the future deep space exploration scenario using the "deep space port" at Lagrange points. OMOTENASHI, also sponsored by JAXA, will land the smallest lunar lander to date on the lunar surface to demonstrate the feasibility of the hardware for distributed cooperative exploration system. Small landers will enable multi-point exploration, which is complimentary with large-scale human exploration. Once on the lunar surface, the OMOTENASHI spacecraft will observe the radiation and soil environments of the lunar surface by active radiation measurements and soil shear measurements. Following EM-1, Space Launch System will evolve to the more-powerful Block 1B configuration, which uses a new Exploration Upper Stage to increase the vehicle's LEO payload capability from 70 t to 105 t. With that transition, the Orion Stage Adapter, which will carry the secondary payloads on EM-1, will be phased out, and a new Universal Stage Adapter will be introduced, creating opportunities for flying larger secondary payloads. This paper will provide a brief status of SLS progress toward first launch; an overview of smallsat accommodations, integration, and operations on EM-1; information about the specific payloads flying on that launch; and a discussion of future accommodations and opportunities for secondary payloads on SLS for Exploration Mission-2 and beyond.

Robinson, Kimberly F.↗

NASA NTRS2017

NASA's Space Launch System: Deep-Space Opportunities for SmallSats

Designed for human exploration missions into deep space, NASA's Space Launch System (SLS) represents a new spaceflight infrastructure asset, enabling a wide variety of unique utilization opportunities. While primarily focused on launching the large systems needed for crewed spaceflight beyond Earth orbit, SLS also offers a game-changing capability for the deployment of small satellites to deep-space destinations, beginning with its first flight. Currently, SLS is making rapid progress toward readiness for its first launch in two years, using the initial configuration of the vehicle, which is capable of delivering 70 metric tons (t) to Low Earth Orbit (LEO). On its first flight test of the Orion spacecraft around the moon, accompanying Orion on SLS will be small-satellite secondary payloads, which will deploy in cislunar space. The deployment berths are sized for "6U" CubeSats, and on EM-1 the spacecraft will be deployed into cislunar space following Orion separate from the SLS Interim Cryogenic Propulsion Stage. Payloads in 6U class will be limited to 14 kg maximum mass. Secondary payloads on EM-1 will be launched in the Orion Stage Adapter (OSA). Payload dispensers will be mounted on specially designed brackets, each attached to the interior wall of the OSA. For the EM-1 mission, a total of fourteen brackets will be installed, allowing for thirteen payload locations. The final location will be used for mounting an avionics unit, which will include a battery and sequencer for executing the mission deployment sequence. Following the launch of EM-1, deployments of the secondary payloads will commence after sufficient separation of the Orion spacecraft to the upper stage vehicle to minimize any possible contact of the deployed cubesats to Orion. Currently this is estimated to require approximately 4 hours. The allowed deployment window for the cubesats will be from the time the upper stage disposal maneuvers are complete to up to 10 days after launch. The upper stage will fly past the moon at a perigee of approximately 100km, and this closest approach will occur about 5 days after launch. The limiting factor for the latest deployment time is the available power in the sequencer system. Several NASA Mission Directorates were involved in the development of programs for the competition, selection, and development of EM-1 payloads that support directorate priorities. CubeSat payloads on EM-1 will include both NASA research experiments and spacecraft developed by industry, international and potentially academia partners. The Human Exploration and Operations Mission Directorate (HEOMD) Advanced Exploration Systems (AES) Division was allocated five payload opportunities on the EM-1 mission. Near Earth Asteroid (NEA) Scout is designed to rendezvous with and characterize a candidate NEA. A solar sail, an innovation the spacecraft will demonstrated for the CubeSat class, will provide propulsion. Lunar Flashlight will use a green propellant system and will search for potential ice deposits in the moon's permanently shadowed craters. BioSentinel is a yeast radiation biosensor, planned to measure the effects of space radiation on deoxyribonucleic acid (DNA). Lunar Icecube, a collaboration with Morehead State University, will prospect for water in ice, liquid, and vapor forms as well as other lunar volatiles from a low-perigee, highly inclined lunar orbit using a compact Infrared spectrometer. Skyfire, a partnership with Lockheed Martin, is a technology demonstration mission that will perform a lunar flyby, collecting spectroscopy, and thermography data to address questions related to surface characterization, remote sensing, and site selection. NASA's Space Technology Mission Directorate (STMD) was allocated three payload opportunities on the EM-1 mission. These slots will be filled via the 2 Centennial Challenges Program, NASA's flagship program for technology prize competitions, which directly engages the public, academia, and industry in open prize competitions to stimulate innovation. The NASA Science Mission Directorate (SMD) was allocated two payload opportunities on the EM-1 mission. The CubeSat Mission to Study Solar Particles (CuSP) payload will study the sources and acceleration mechanisms of solar and interplanetary particles in near-Earth orbit, support space weather research by determining proton radiation levels during Solar Energetic Particle (SEP) events and identifying suprathermal properties that could help predict geomagnetic storms. The LunaH-Map payload will help scientists understand the quantity of H-bearing materials in lunar cold traps (~10 km), determine the concentration of H-bearing materials with 1m depth, and constrain the vertical distribution of H-bearing materials. The final three payload opportunities for the EM-1 mission were allocated for NASA's international space agency counterparts. The flight opportunities are intended to benefit the international space agency and NASA as well as further the collective space exploration goals. ArgoMoon is sponsored by ESA/ASI and will fly along with the ICPS on its disposal trajectory to perform proximity operations with the ICPS post-disposal, take external imagery of engineering and historical significance, and perform an optical communications demonstration. EQUULEUS, sponsored by JAXA, will fly to a libration orbit around the Earth-Moon L2 point and demonstrate trajectory control techniques within the Sun-Earth-Moon region for the first time by a nano spacecraft. The mission will also contribute to the future human exploration scenario by understanding the radiation environment in geospace and deep space, characterizing the flux of impacting meteors on the far side of the moon, and demonstrating the future deep space exploration scenario using the "deep space port" at Lagrange points. OMOTENASHI, also sponsored by JAXA, will land the smallest lunar lander to date on the lunar surface to demonstrate the feasibility of the hardware for distributed cooperative exploration system. Small landers will enable multi-point exploration, which is complimentary with large-scale human exploration. Once on the lunar surface, the OMOTENASHI spacecraft will observe the radiation and soil environments of the lunar surface by active radiation measurements and soil shear measurements. Following EM-1, Space Launch System will evolve to the more-powerful Block 1B configuration, which uses a new Exploration Upper Stage to increase the vehicle's LEO payload capability from 70 t to 105 t. With that transition, the Orion Stage Adapter, which will carry the secondary payloads on EM-1, will be phased out, and a new Universal Stage Adapter will be introduced, creating opportunities for flying larger secondary payloads. This paper will provide a brief status of SLS progress toward first launch; an overview of smallsat accommodations, integration, and operations on EM-1; information about the specific payloads flying on that launch; and a discussion of future accommodations and opportunities for secondary payloads on SLS for Exploration Mission-2 and beyond.

Robinson, Kimberly F.↗

NASA NTRS2015

Standardized Modular Power Interfaces for Future Space Explorations Missions

Earlier studies show that future human explorations missions are composed of multi-vehicle assemblies with interconnected electric power systems. Some vehicles are often intended to serve as flexible multi-purpose or multi-mission platforms. This drives the need for power architectures that can be reconfigured to support this level of flexibility. Power system developmental costs can be reduced, program wide, by utilizing a common set of modular building blocks. Further, there are mission operational and logistics cost benefits of using a common set of modular spares. These benefits are the goals of the Advanced Exploration Systems (AES) Modular Power System (AMPS) project. A common set of modular blocks requires a substantial level of standardization in terms of the Electrical, Data System, and Mechanical interfaces. The AMPS project is developing a set of proposed interface standards that will provide useful guidance for modular hardware developers but not needlessly constrain technology options, or limit future growth in capability. In 2015 the AMPS project focused on standardizing the interfaces between the elements of spacecraft power distribution and energy storage. The development of the modular power standard starts with establishing mission assumptions and ground rules to define design application space. The standards are defined in terms of AMPS objectives including Commonality, Reliability-Availability, Flexibility-Configurability and Supportability-Reusability. The proposed standards are aimed at assembly and sub-assembly level building blocks. AMPS plans to adopt existing standards for spacecraft command and data, software, network interfaces, and electrical power interfaces where applicable. Other standards including structural encapsulation, heat transfer, and fluid transfer, are governed by launch and spacecraft environments and bound by practical limitations of weight and volume. Developing these mechanical interface standards is more difficult but an essential part of defining physical building blocks of modular power. This presentation describes the AMPS projects progress towards standardized modular power interfaces.

Spacecraft Power↗

NASA NTRS1978

Power management and control for space systems

Power management and control technology for the large, high-power spacecraft of the 1980's is discussed. Systems weight optimization that indicate a need for higher bus voltages are shown. Environmental interactions that are practical limits for the maximum potential on exposed surfaces are shown. A dual-voltage system is proposed that would provide the weight savings of a high-voltage distribution system and take into account the potential environmental interactions. The technology development of new components and circuits is also discussed.

Finke, R. C.↗

NASA NTRS2009

A High Power Density Power System Electronics for NASA's Lunar Reconnaissance Orbiter

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).

Hernandez-Pellerano, A.↗

NASA NTRS1992

The changing nature of spacecraft operations: From the Vikings of the 1970's to the great observatories of the 1990's and beyond

Four trends in spacecraft flight operations are discussed which will reduce overall program costs. These trends are the use of high-speed, highly reliable data communications systems for distributing operations functions to more convenient and cost-effective sites; the improved capability for remote operation of sensors; a continued rapid increase in memory and processing speed of flight qualified computer chips; and increasingly capable ground-based hardware and software systems, notably those augmented by artificial intelligence functions. Changes reflected by these trends are reviewed starting from the NASA Viking missions of the early 70s, when mission control was conducted at one location using expensive and cumbersome mainframe computers and communications equipment. In the 1980s, powerful desktop computers and modems enabled the Magellan project team to operate the spacecraft remotely. In the 1990s, the Hubble Space Telescope project uses multiple color screens and automated sequencing software on small computers. Given a projection of current capabilities, future control centers will be even more cost-effective.

Ledbetter, Kenneth W.↗

NASA NTRS2002

Design and architecture of the Mars relay network planning and analysis framework

In this paper we describe the design and architecture of the Mars Network planning and analysis framework that supports generation and validation of efficient planning and scheduling strategy. The goals are to minimize the transmitting time, minimize the delaying time, and/or maximize the network throughputs. The proposed framework would require (1) a client-server architecture to support interactive, batch, WEB, and distributed analysis and planning applications for the relay network analysis scheme, (2) a high-fidelity modeling and simulation environment that expresses link capabilities between spacecraft to spacecraft and spacecraft to Earth stations as time-varying resources, and spacecraft activities, link priority, Solar System dynamic events, the laws of orbital mechanics, and other limiting factors as spacecraft power and thermal constraints, (3) an optimization methodology that casts the resource and constraint models into a standard linear and nonlinear constrained optimization problem that lends itself to commercial off-the-shelf (COTS)planning and scheduling algorithms.

Mars relay network planning analysis↗

NASA NTRSDate not supplied

IceCube: Demonstration of an 883 GHz Radiometer for Ice Cloud Remote Sensing

IceCube was a technology demonstration of an 883 GHz heterodyne radiometer on a 3U CubeSat for ice cloud characterization. The project was a collaboration between Goddard Space Flight Center, Virginia Diodes Inc., and Wallops Flight Facility. IceCube was launched to the International Space Station (ISS) in April 2017, and was deployed to the orbit in May 2017. The radiometer measured ice cloud emissions from an ISS orbit for over 15 months. IceCube generated the first 883 GHz cloud map over a large operation temperature range of (5 ºC—37 ºC). Cloud ice plays a major role in the cloud precipitation process and Earth’s energy budget. Ice clouds are used in global circulation models as tuning parameters to achieve model agreement with observation at the top of the atmosphere in the radiation budget and at the bottom for precipitation, however, due to a lack of accurate ice cloud measurements large uncertainties exist in these models. Submillimeter wave remote sensing is capable of addressing this issue by measuring cloud ice mass and microphysical properties in the middle-to-upper troposphere. This fills the sensitivity gap not covered by the visible/infrared and microwave sensors [1]. The goal of IceCube was to increase the TRL of a heterodyne 883 GHz radiometer (using commercial parts) from 5 to 7 by validating the performance in a relevant spaceflight environment. The design of the radiometer was driven by frequency of operation, bandwidth, calibration, available power, and thermal environment requirements. The design included a 15 mm aperture off-axis parabolic reflector with a Potter feed horn, an 883 GHz 2nd-harmonic mixer that is fed by a local oscillator chain with a 24.3 GHz dielectric resonator (MLA), followed by a 6 GHz bandwidth centered at 9 GHz intermediate frequency assembly (IFA), receiver interface card, and power distribution unit. The IFA included an internal noise diode calibration to separate the MLA performance from the rest of the system. For this the radiometer had four operational states: antenna, antenna + noise, reference, and reference + noise; each state’s duration was 10 ms. The total power dissipation of the instrument was 5.6 W. The spacecraft had spinning capabilities to provide a cold sky view for calibration. We present the instrument design, ground test results, challenges, and highlight some of the flight measurements.

N Ehsan↗

NASA NTRS2000

Vehicle Dynamics due to Magnetic Launch Propulsion

The field of Magnetic Levitation Lind Propulsion (MagLev) has been around for over 30 years, primarily in high-speed rail service. In recent years, however, NASA has been looking closely at MagLev as a possible first stage propulsion system for spacecraft. This approach creates a variety of new problems that don't currently exist with the present MagLev trains around the world. NASA requires that a spacecraft of approximately 120,000 lbs be accelerated at two times the acceleration of gravity (2g's). This produces a greater demand on power over the normal MagLev trains that accelerate at around 0.1g. To be able to store and distribute up to 3,000 Mega Joules of energy in less than 10 seconds is a technical challenge. Another problem never addressed by the train industry and, peculiar only to NASA, is the control of a lifting body through the acceleration of and separation from the MagLev track. Very little is understood about how a lifting body will react with external forces, Such as wind gusts and ground effects, while being propelled along on soft springs such as magnetic levitators. Much study needs to be done to determine spacecraft control requirements as well as what control mechanisms and aero-surfaces should be placed on the carrier. Once the spacecraft has been propelled down the track another significant event takes place, the separation of the spacecraft from the carrier. The dynamics involved for both the carrier and the spacecraft are complex and coupled. Analysis of the reaction of the carrier after losing, a majority of its mass must be performed to insure control of the carrier is maintained and a safe separation of the spacecraft is achieved. The spacecraft angle of attack required for lift and how it will affect the carriage just prior to separation, along with the impacts of around effect and aerodynamic forces at ground level must be modeled and analyzed to define requirements on the launch vehicle design. Mechanisms, which can withstand the expected forces and allow the spacecraft to separate at the desired time must be developed. Modeling and analysis of the full-scale system must be developed to derive the requirements for design of the separation mechanisms. A description of several MagLev demonstrators will be presented and the testing and analysis to model the demonstrators will be included.

Galaboff, Zachary J.↗

NASA NTRS1980

Large space systems technology electronics: Data and power distribution

The development of hardware technology and manufacturing techniques required to meet space platform and antenna system needs in the 1980s is discussed. Preliminary designs for manned and automatically assembled space power system cables, connectors, and grounding and bonding materials and techniques are reviewed. Connector concepts, grounding design requirements, and bonding requirements are discussed. The problem of particulate debris contamination for large structure spacecraft is addressed.

Dunbar, W. G.↗

NASA NTRS1982

Application of multi-function display and control technology

The NASA orbiter spacecraft incorporates a complex array of systems, displays, and controls. The incorporation of discrete dedicated controls into a multifunction display and control system (MFDCS) offers the potential for savings in weight, power, panel space, and crew training time. Technology identified as applicable to a MFDCS is applied to the orbiter orbital maneuvering system (OMS) and the electrical power distribution and control system (EPDCS) to derive concepts for a MFDCS design. Several concepts of varying degrees of performance and complexity are discussed and a suggested concept for further development is presented in greater detail. Both the hardware and software aspects and the human factors considerations of the designs are included.

Spiger, R. J.↗

NASA NTRS1999

An Overview of Future NASA Missions, Concepts, and Technologies Related to Imaging of the World's Land Areas

In the near term NASA is entering into the peak activity period of the Earth Observing System (EOS). The EOS AM-1 /"Terra" spacecraft is nearing launch and operation to be followed soon by the New Millennium Program (NMP) Earth Observing (EO-1) mission. Other missions related to land imaging and studies include EOS PM-1 mission, the Earth System Sciences Program (ESSP) Vegetation Canopy Lidar (VCL) mission, the EOS/IceSat mission. These missions involve clear advances in technologies and observational capability including improvements in multispectral imaging and other observing strategies, for example, "formation flying". Plans are underway to define the next era of EOS missions, commonly called "EOS Follow-on" or EOS II. The programmatic planning includes concepts that represent advances over the present Landsat-7 mission that concomitantly recognize the advances being made in land imaging within the private sector. The National Polar Orbiting Environmental Satellite Series (NPOESS) Preparatory Project (NPP) is an effort that will help to transition EOS medium resolution (herein meaning spatial resolutions near 500 meters), multispectral measurement capabilities such as represented by the EOS Moderate Resolution Imaging Spectroradiometer (MODIS) into the NPOESS operational series of satellites. Developments in Synthetic Aperture Radar (SAR) and passive microwave land observing capabilities are also proceeding. Beyond these efforts the Earth Science Enterprise Technology Strategy is embarking efforts to advance technologies in several basic areas: instruments, flight systems and operational capability, and information systems. In the case of instruments architectures will be examined that offer significant reductions in mass, volume, power and observational flexibility. For flight systems and operational capability, formation flying including calibration and data fusion, systems operation autonomy, and mechanical and electronic innovations that can reduce spacecraft and subsystem resource requirements. The efforts in information systems will include better approaches for linking multiple data sets, extracting and visualizing information, and improvements in collecting, compressing, transmitting, processing, distributing and archiving data from multiple platforms. Overall concepts such as sensor webs, constellations of observing systems, and rapid and tailored data availability and delivery to multiple users comprise and notions Earth Science Vision for the future.

Salomonson, Vincent V.↗

NASA NTRS1984

Saturn's rings - Voyager 1 radio occultation experiment results

The Voyager 1 encounter with Saturn in 1980 included the first radio occultation measurements of the ring system. Opacity of and scattering by Saturn's rings were measured at 3.6- and 13-cm wavelengths using transmissions from the spacecraft to the earth. Detailed models of ring structure have been derived from these data at radial resolutions as small as 1 km, depending on the radio opacity of the region probed. Inverse power-law-type distributions of ring particle sizes with indices of about 3 and upper size cutoffs near 5 m are in good agreement with the data, but parameters vary throughout the rings. Occultation measurements have also provided information on ring dynamics and have led to a more precise definition of Saturn's pole vector.

Simpson, R. A.↗

NASA NTRS1987

Photovoltaic power modules for NASA's manned space station

The capability and the safety of manned spacecraft are largely dependent upon reliable electric power systems. Two similar space power systems able to survive the low Earth orbit environment, are being considered for NASA's Manned Space Station (SS), scheduled to begin operation in the mid 1990's. The Space Station Electric Power System (EPS) is composed of Photovoltaic (PV) Power Modules, Solar Dynamic (SD) Power Modules, and the Power Management and Distribution (PMAD) System. One EPS configuration will deliver 37.5 kW of PV based, utility grade, ac power to SS users. A second 75 kWe PV based EPS option is also being considered for SS deployment. The two EPS options utilize common modules and differ only in the total number of PV Power Modules used. Each PV Power Module supplies 18.75 kWe of ac power and incorporates its own energy storage and thermal control. The general requirements and the current preliminary design configuration of the Space Station PV Power Modules are examined.

Tatro, Charles A.↗