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NASA NTRS2006

BAE Systems Radiation Hardened SpaceWire ASIC and Roadmap

An Application Specific Integrated Circuit (ASIC) that implements the SpaceWire protocol has been developed in a radiation hardened 0.25 micron CMOS, technology. This effort began in March 2003 as a joint development between the NASA Goddard Space Flight Center (GSFC) and BAE Systems. The BAE Systems SpaceWire ASlC is comprised entirely of reusable core elements, many of which are already flight-proven. It incorporates a 4-port SpaceWire router with two local ports, dual PC1 bus interfaces, a microcontroller, 32KB of internal memory, -and a memory controller for additional external memory use. The SpaceWire ASlC is planned for use on both the Geostationary Operational Environmental Satellites (GOES)-R and the Lunar Reconnaissance Orbiter (LRO). Engineering parts have already been delivered to both programs. This paper discusses the SpaceWire protocol and those elements of it that have been built into the current SpaceWire reusable core. There are features within the core that go beyond the current standard that can be enabled or disabled by the user and these will be described. The adaptation of SpaceWire to BAE Systems' On Chip Bus (OCB) for compatibility with the other reusable cores will be discussed. Optional configurations within user systems will be shown. The physical imp!ementation of the design will be described and test results from the hardware will be discussed. Finally, the BAE Systems roadmap for SpaceWire developments will be discussed, including some products already in design as well as longer term plans.

Berger, Richard↗

NASA NTRS2007

A One Chip Hardened Solution for High Speed SpaceWire System Implementations. Session: Components

An Application Specific Integrated Circuit (ASIC) that implements the SpaceWire protocol has been developed in a radiation hardened 0.25 micron CMOS technology. This effort began in March 2003 as a joint development between the NASA Goddard Space Flight Center (GSFC) and BAE Systems. The BAE Systems SpaceWire ASIC is comprised entirely of reusable core elements, many of which are already flight-proven. It incorporates a router with 4 SpaceWire ports and two local ports, dual PC1 bus interfaces, a microcontroller, 32KB of internal memory, and a memory controller for additional external memory use. The SpaceWire cores are also reused in other ASICs under development. The SpaceWire ASIC is planned for use on the Geostationary Operational Environmental Satellites (GOES)-R, the Lunar Reconnaissance Orbiter (LRO) and other missions. Engineering and flight parts have been delivered to programs and users. This paper reviews the SpaceWire protocol and those elements of it that have been built into the current and next SpaceWire reusable cores and features within the core that go beyond the current standard and can be enabled or disabled by the user. The adaptation of SpaceWire to BAE Systems' On Chip Bus (OCB) for compatibility with the other reusable cores will be reviewed and highlighted. Optional configurations within user systems and test boards will be shown. The physical implementation of the design will be described and test results from the hardware will be discussed. Application of this ASIC and other ASICs containing the SpaceWire cores and embedded microcontroller to Plug and Play and reconfigurable implementations will be described. Finally, the BAE Systems roadmap for SpaceWire developments will be updated, including some products already in design as well as longer term plans.

Marshall, Joseph R.↗

NASA NTRS2013

Thermal Design and Analysis of the Supersonic Flight Dynamics Test Vehicle for the Low Density Supersonic Decelerator Project

The thermal design and analysis of the experimental Supersonic Flight Dynamics Test (SFDT) vehicle is presented. The SFDT vehicle is currently being designed as a platform to help demonstrate key technologies for NASA's Low Density Supersonic Decelerator (LDSD) project. The LDSD project is charged by NASA's Office of the Chief Technologist (OCT) with the task of advancing the state of the art in Mars Entry, Descent, and Landing (EDL) systems by developing and testing three new technologies required for landing heavier payloads on Mars. The enabling technologies under development consist of a large 33.5 meter diameter Supersonic Ringsail (SSRS) parachute and two different types of Supersonic Inflatable Aerodynamic Decelerator (SIAD) devices - a robotic class, SIAD-R, that inflates to a 6 meter diameter torus, and an exploration class, SIAD-E, that inflates to an 8 meter diameter isotensoid. As part of the technology development effort, the various elements of the new supersonic decelerator system must be tested in a Mars-like environment. This is currently planned to be accomplished by sending a series of SFDT vehicles into Earth's stratosphere. Each SFDT vehicle will be lifted to a stable float altitude by a large helium carrier balloon. Once at altitude, the SFDT vehicles will be released from their carrier balloon and spun up via spin motors to provide trajectory stability. An onboard third stage solid rocket motor will propel each test vehicle to supersonic flight in the upper atmosphere. After main engine burnout, each vehicle will be despun and testing of the deceleration system will begin: first an inflatable decelerator will be deployed around the aeroshell to increase the drag surface area, and then the large parachute will be deployed to continue the deceleration and return the vehicle back to the Earth's surface. The SFDT vehicle thermal system must passively protect the vehicle structure and its components from cold temperatures experienced during the ascent phase of the mission as well as from the extreme heat fluxes produced during the supersonic test phase by the main motor plume and aeroheating. The passive thermal design approach for the SFDT vehicle relies upon careful and complex bounding analysis of all three modes of heat transfer - conduction, convection, and radiation - coupled with a tightly managed transient power dissipation timeline for onboard electronics components throughout all mission phases.

thermal desktop↗

NASA NTRS2013

NASA Tech Briefs, Februrary 2013

Topics covered include: Measurements of Ultra-Stable Oscillator (USO) Allan Deviations in Space; Gaseous Nitrogen Orifice Mass Flow Calculator; Validation of Proposed Metrics for Two-Body Abrasion Scratch Test Analysis Standards; Rover Low Gain Antenna Qualification for Deep Space Thermal Environments; Automated, Ultra-Sterile Solid Sample Handling and Analysis on a Chip; Measuring and Estimating Normalized Contrast in Infrared Flash Thermography; Spectrally and Radiometrically Stable, Wideband, Onboard Calibration Source; High-Reliability Waveguide Vacuum/Pressure Window; Methods of Fabricating Scintillators With Radioisotopes for Beta Battery Applications; Magnetic Shield for Adiabatic Demagnetization Refrigerators (ADR); CMOS-Compatible SOI MESFETS for Radiation-Hardened DC-to-DC Converters; Silicon Heat Pipe Array; Adaptive Phase Delay Generator; High-Temperature, Lightweight, Self-Healing Ceramic Composites for Aircraft Engine Applications; Treatment to Control Adhesion of Silicone-Based Elastomers; High-Temperature Adhesives for Thermally Stable Aero-Assist Technologies; Rockballer Sample Acquisition Tool; Rock Gripper for Sampling, Mobility, Anchoring, and Manipulation; Advanced Magnetic Materials Methods and Numerical Models for Fluidization in Microgravity and Hypogravity; Data Transfer for Multiple Sensor Networks Over a Broad Temperature Range; Using Combustion Synthesis to Reinforce Berms and Other Regolith Structures; Visible-Infrared Hyperspectral Image Projector; Three-Axis Attitude Estimation With a High-Bandwidth Angular Rate Sensor Change_Detection.m; AGATE: Adversarial Game Analysis for Tactical Evaluation; Ionospheric Simulation System for Satellite Observations and Global Assimilative; Modeling Experiments (ISOGAME); An Extensible, User- Modifiable Framework for Planning Activities; Mission Operations Center (MOC) - Precipitation Processing System (PPS) Interface Software System (MPISS); Automated 3D Damaged Cavity Model Builder for Lower Surface Acreage Tile on Orbiter; Mixed Linear/Square-Root Encoded Single-Slope Ramp Provides Low-Noise ADC with High Linearity for Focal Plane Arrays; RUSHMAPS: Real-Time Uploadable Spherical Harmonic Moment Analysis for Particle Spectrometers; Powered Descent Guidance with General Thrust-Pointing Constraints; X-Ray Detection and Processing Models for Spacecraft Navigation and Timing; and Extreme Ionizing-Radiation-Resistant Bacterium

Source record↗

NASA NTRS2002

Ethernet for Space Flight Applications

NASA's Goddard Space Flight Center (GSFC) is adapting current data networking technologies to fly on future spaceflight missions. The benefits of using commercially based networking standards and protocols have been widely discussed and are expected to include reduction in overall mission cost, shortened integration and test (I&T) schedules, increased operations flexibility, and hardware and software upgradeability/scalability with developments ongoing in the commercial world. The networking effort is a comprehensive one encompassing missions ranging from small University Explorer (UNEX) class spacecraft to large observatories such as the Next Generation Space Telescope (NGST). Mission aspects such as flight hardware and software, ground station hardware and software, operations, RF communications, and security (physical and electronic) are all being addressed to ensure a complete end-to-end system solution. One of the current networking development efforts at GSFC is the SpaceLAN (Spacecraft Local Area Network) project, development of a space-qualifiable Ethernet network. To this end we have purchased an IEEE 802.3-compatible 10/100/1000 Media Access Control (MAC) layer Intellectual Property (IP) core and are designing a network node interface (NNI) and associated network components such as a switch. These systems will ultimately allow the replacement of the typical MIL-STD-1553/1773 and custom interfaces that inhabit most spacecraft. In this paper we will describe our current Ethernet NNI development along with a novel new space qualified physical layer that will be used in place of the standard interfaces. We will outline our plans for development of space qualified network components that will allow future spacecraft to operate in significant radiation environments while using a single onboard network for reliable commanding and data transfer. There will be a brief discussion of some issues surrounding system implications of a flight Ethernet. Finally, we will show an onboard network architecture for a proposed new mission using Ethernet for science data transport.

Webb, Evan↗

NASA NTRS2009

Human Research Program Human Health Countermeasures Element Cardiovascular Risks Standing Review Panel (SRP)

The Cardiovascular Risk Standing Review Panel (SRP) evaluated several cardiovascular risks associated with space flight along with the ongoing and emerging plans to study these issues and potentially propose and/or develop countermeasures. The areas of focus included: 1) The risk of cardiac rhythm problems during prolonged space flight, and 2) Issues related to the risk of orthostatic intolerance during re-exposure to gravity. An emerging area of concern is radiation associated vascular injury. The risk of cardiac rhythm disturbances has emerged based on case reports only. No systematic study of this risk has been published. However, concerns about this risk are heightened by the age range of astronauts, the structural changes in the heart that occur during space flight, and the potential shifts in fluids and electrolytes. The current plan is to use prolonged Holter monitor EKG records made as part of the "Integrated Cardiovascular SMO" in space to determine more about the frequency and magnitude of this problem and to link this data to complementary data from the nutrition group on electrolytes. The SRP was supportive of this approach. The SRP also felt that any data related to cardiovascular risk in space should be better coordinated with the medical screening data that all astronauts undergo at regular intervals. Additionally, while there are potential privacy issues related to this suggestion, many of the current barriers to better coordination of experimental and clinical data appear to reflect longstanding cultural traditions at NASA that need rethinking. The risk of orthostatic intolerance during re-exposure to gravity was seen by the SRP as an area supported by a wealth of published physiological evidence. The SRP also felt that moving forward with the planned approach to countermeasures was reasonable and that extensive additional hypothesis testing on the physiology of orthostatic intolerance was not needed at this time. There was support for developing ground based models of limited (e.g. 1/6 th) G environments on Earth that generated a number of ideas for consideration by NASA investigators.

Joyner, Michael↗

NASA NTRS1986

Biological and Medical Experiments on the Space Shuttle, 1981 - 1985

This volume is the first in a planned series of reports intended to provide a comprehensive record of all the biological and medical experiments and samples flown on the Space Shuttle. Experiments described have been conducted over a five-year period, beginning with the first plant studies conducted on STS-2 in November 1981, and extending through STS 61-C, the last mission to fly before the tragic Challenger accident of January 1986. Experiments were sponsored within NASA not only by the Life Sciences Division of the Office of Space Science and Applications, but also by the Shuttle Student Involvement Program (SSIP) and the Get Away Special (GAS) Program. Independent medical studies were conducted as well on the Shuttle crew under the auspices of the Space Biomedical Research Institute at Johnson Space Center. In addition, cooperative agreements between NASA and foreign government agencies led to a number of independent experiments and also paved the way for the joint US/ESA Spacelab 1 mission and the German (DFVLR) Spacelab D-1. Experiments included: (1) medically oriented studies of the crew aimed at identifying, preventing, or treating health problems due to space travel; (2) projects to study morphological, physiological, or behavioral effects of microgravity on animals and plants; (3) studies of the effects of microgravity on cells and tissues; and (4) radiation experiments monitoring the spacecraft environment with chemical or biological dosimeters or testing radiation effects on simple organisms and seeds.

Halstead, Thora W.↗

NASA NTRS1999

Development Unit Configuration and Current Status of the MIP/MAAC Experiment

The Mars In-Situ Propellant Production (ISPP) Precursor (MIP) experiment package is planned for inclusion on the Mars 2001 Lander. This experiment package consists of five experiments whose purpose is to demonstrate the performance of various ISPP processes in-situ on Mars. The demonstrated ability to produce propellant for Mars Return Vehicles (MRV) is considered to be a necessary precursor to any future manned mission to Mars. The Mars Atmosphere Acquisition and Compression (MAAC) experiment is part of the MIP package and is intended to demonstrate that, by using a sorption compressor, CO2 can be preferentially adsorbed at about 6 torr from the Mars atmosphere during the night when the bed is cold then subsequently compressed to about 800 torr by heating the bed and desorbing CO2 during the day. The compressed CO2 produced by MAAC is to be fed to the Oxygen Generator Subsystem (OGS) where pure oxygen is to be produced. Pure oxygen is considered to be one of the primary constituents of a future manned MRV propellant system. A MAAC Development Unit (DU) has been fabricated and tested at JPL. The MAAC DU consists of 1) a sorption bed filled with a CO2 selective sorbent material, 2) a purge system to be used to periodically backflush non-CO2 gases from the sorbent bed during adsorption, 3) a JPL-developed gas-gap heat switch that allows heat transfer to a radiator for heat removal from the bed during the night time adsorption period and that impedes heat transfer during the day time desorption period, 4) a radiator to radiate heat to the night sky during the adsorption period, 5) a set of three isolation valves and connecting tubing. 6) two pressure transducers and several thermocouples for monitoring the MAAC operating conditions, and command and data handling electronics. This paper will describe the operational theory and the configuration of the MAAC DU and will discuss the current status of the MAAC experiment development including some selected results of performance testing that has been completed prior to the ISRU III meeting.

Karlmann, P. B.↗

NASA NTRS1999

Development Unit Configuration Status of the MIP/MAAC Experiment

The Mars In-Situ Propellant Production (ISPP) Precursor (MIP) experiment package is planned for inclusion on the Mars 2001 Lander. This experiment package consists of five experiments whose purpose is to demonstrate the performance of various ISPP processes in-situ on Mars. The demonstrated ability to produce propellant for Mars Return Vehicles (MRV) is considered to be a necessary precursor to any future manned mission to Mars. The Mars Atmosphere Acquisition and Compression (MAAC) experiment is part of the MIP package and is intended to demonstrate that, by using a sorption compressor, CO2 can be preferentially adsorbed at about 6 torr from the Mars atmosphere during the night when the bed is cold then subsequently compressed to about 800 torr by heating the bed and desorbing C02 during the day. The compressed CO2 produced by MAAC is to be fed to the Oxygen Generator Subsystem (OGS) where pure oxygen is to be produced. Pure oxygen is considered to be one of the primary constituents of a future manned MRV propellant system. A MAAC Development Unit (DU) has been fabricated and tested at JPL. The MAAC DU consists of: (1) a sorption bed filled with a CO2 selective sorbent material; (2) a purge system to be used to periodically backflush non-CO2 gases from the sorbent bed during adsorption; (3) a JPL-developed gas-gap heat switch that allows heat transfer to a radiator for heat removal from the bed during the night time adsorption period and that impedes heat transfer during the day time desorption period; (4) a radiator to radiate heat to the night sky during the adsorption period; (5) a set of three isolation valves and connecting tubing; (6) two pressure transducers and several thermocouples for monitoring the MAAC operating conditions, and command and data handling electronics. This paper will describe the operational theory and the configuration of the MAAC DU and will discuss the current status of the MAAC experiment development including some selected results of performance testing that has been completed prior to the ISRU III meeting.

Karlmann, P. B.↗

NASA NTRS2013

Thermal Performance of ATLAS Laser Thermal Control System Demonstration Unit

The second Ice, Cloud, and Land Elevation Satellite mission currently planned by National Aeronautics and Space Administration will measure global ice topography and canopy height using the Advanced Topographic Laser Altimeter System {ATLAS). The ATLAS comprises two lasers; but only one will be used at a time. Each laser will generate between 125 watts and 250 watts of heat, and each laser has its own optimal operating temperature that must be maintained within plus or minus 1 degree Centigrade accuracy by the Laser Thermal Control System (LTCS) consisting of a constant conductance heat pipe (CCHP), a loop heat pipe (LHP) and a radiator. The heat generated by the laser is acquired by the CCHP and transferred to the LHP, which delivers the heat to the radiator for ultimate rejection. The radiator can be exposed to temperatures between minus 71 degrees Centigrade and minus 93 degrees Centigrade. The two lasers can have different operating temperatures varying between plus 15 degrees Centigrade and plus 30 degrees Centigrade, and their operating temperatures are not known while the LTCS is being designed and built. Major challenges of the LTCS include: 1) A single thermal control system must maintain the ATLAS at 15 degrees Centigrade with 250 watts heat load and minus 71 degrees Centigrade radiator sink temperature, and maintain the ATLAS at plus 30 degrees Centigrade with 125 watts heat load and minus 93 degrees Centigrade radiator sink temperature. Furthermore, the LTCS must be qualification tested to maintain the ATLAS between plus 10 degrees Centigrade and plus 35 degrees Centigrade. 2) The LTCS must be shut down to ensure that the ATLAS can be maintained above its lowest desirable temperature of minus 2 degrees Centigrade during the survival mode. No software control algorithm for LTCS can be activated during survival and only thermostats can be used. 3) The radiator must be kept above minus 65 degrees Centigrade to prevent ammonia from freezing using no more than 135 watts of heater power. 4) The LHP reservoir control heater power is limited to 15 watts with a 70 percent duty cycle. 5) The voltage of the power supply can vary between 26 volts direct current and 34 volts direct current during the spacecraft lifetime. A design analysis shows that a single LTCS can satisfy these requirements. However, shutdown of· the LHP is particularly challenging and the shutdown heater must be wired in series with two reservoir thermostats and two CCHP thermostats at different set points. An LTCS demonstration unit has been tested to verify these performance characteristics experimentally prior to proceeding to the final LTCS design and fabrication. Test results showed that the LHP shutdown scheme would be able to shut down the LHP as designed and the reservoir control heater can maintain the ATLAS mass simulator within the plus or minus 1 degrees Centigrade accuracy under various combinations of the heat load, sink temperature, and power supply voltage.

Ku, Jentung↗

NASA NTRSDate not supplied

Investigating Molecular Responses to Space Radiation for Biological Missions Beyond Low Earth Orbit

As we plan crewed missions to the Moon, Mars, and beyond, it is essential to understand how persistent exposure to deep space radiation affects biology. Unlike on the International Space Station (ISS), where crew support and sample return are possible, experiments for long-duration missions require autonomous systems with no sample return. Human cells would be ideal biosensors, but limitations in culture methods, extended prelaunch storage, and long flight durations make it difficult to keep human cells alive. Unlike other model systems, yeast can survive the constraints of long-duration spaceflight. Despite a billion years of evolution separating yeast from humans, we share homology in hundreds of genes important for basic cell function, including responses to DNA damage. Thus, yeast are excellent biosensors for detecting types/extent of damage induced by space radiation. BioSentinel is NASA’s latest biological CubeSat, and first interplanetary space bioscience mission. BioSentinel is launching on Artemis 1, the first flight of NASA’s Space Launch System, in 2022. The BioSensor payload within BioSentinel contains two yeast strains. The wild type serves as a control for health and normal DNA damage repair (DDR). The rad51 deletion mutant is defective for DDR and will undergo alterations to growth and metabolism as it accumulates radiation damage. Changes in growth and metabolic activity will be measured using a 3-color LED detection system and the metabolic redox dye alamarBlue®. Preliminary tests indicate a significant change in alamarBlue responses to space-like, low-dose ionizing radiation. We will discuss these findings in five parts – Introduction to NASA’s biological CubeSats and BioSentinel (presented by Sergio Santa Maria), analysis of flight data from the ISS mission (presented by Kylie Akiyama), preliminary molecular responses to space radiation (presented here), a deeper dive into those pathways (presented by Kyra Keenan), and characterizing stress response through redox potential data (presented by Diana Gentry).

Lauren Courtney Liddell↗

NASA NTRS1993

LDEF archival system plan

The Long Duration Exposure Facility (LDEF) has provided the first significant opportunity to extensively study the space environment and its effects upon spacecraft systems and materials. The long-term value of the data obtained from LDEF, which is applicable to a wide range of areas including space environment definition, space environmental effects, and spacecraft design, will depend upon the system developed to archive and retrieve the data. Therefore, in addition to the large effort undertaken to analyze LDEF data, a substantial effort is also necessary in order to retain and disseminate LDEF resources for future research and design. W. J. Schafer Associates, Inc., has a task subcontract to develop the LDEF archival system. The LDEF resources include data, hardware, photographic records, and publications which cover the 20-year history of LDEF from concept design through data analysis. Chronologically, prelaunch resources include documentation of facility and experiment development, testing integration, and operation. Post-retrieval resources are the observations, testing, analysis, and publications since the January 1990 retrieval of LDEF. A third set of resources is the experiment and facility hardware and specimens, including more than 10,000 test specimens flown on LDEF and subsequently divided and distributed among investigators at numerous laboratories. Many valuable science and technology investigations were undertaken with LDEF experiments and hardware, and many more investigations are being identified in areas not yet explored. LDEF data applications encompass primarily low-Earth orbit spacecraft and structures. The nearly six-year space exposure of LDEF has provided data to evaluate materials, systems, and living specimens exposed to radiation, meteoroids, debris, and other constituents of the low-Earth environment. Structural, mechanical, electrical, optical, and thermal systems were studied, and materials with applications in all aspects of space systems were exposed to the space environment. The objectives of the LDEF archival system are to maintain the existing LDEF hardware, data, analysis, publications, and photographs as a long term resource, and to provide a quick and simple mechanism by which LDEF resources can be identified, located, and applied.

Wilson, Brenda K.↗

NASA NTRS2014

SLS-Derived Lab: Precursor to Deep Space Human Exploration

Plans to send humans to Mars are in work and the launch system is being built. Are we ready? Robotic missions have successfully demonstrated transportation, entry, landing and surface operations but for human missions there are significant, potentially show-stopping issues. These issues, called Strategic Knowledge Gaps (SKGs) are the unanswered questions concerning long-duration exploration beyond low-earth-orbit. The gaps represent a risk of loss of life or mission and because they require extended exposure to the weightless environment outside earth's protective geo-magnetic field they cannot be resolved on the earth or on the International Space Station (ISS). Placing a laboratory at the relatively close and stable lunar Distant Retrograde Orbit (DRO) provides an accessible location with the requisite environmental conditions for conducting SKG research and testing mitigation solutions. Configurations comprised of multiple 3 meter and 4.3 meter diameter modules have been studied but the most attractive solution uses elements of the human Mars launch vehicle or Space Launch System (SLS) for a Mars proving ground laboratory. A shortened version of an SLS hydrogen propellant tank creates a Skylab-like pressure vessel that flies fully outfitted on a single launch. This not only offers significant savings by incorporating SLS pressure vessel development costs but avoids the expensive ISS approach using many launches with substantial on-orbit assembly before becoming operational. One of the most challenging SKGs is crew radiation protection; this is why SKG laboratory research is combined with Mars transit Habitat systems development. Fundamentally, the two cannot be divorced because using the habitat systems for protection requires actual hardware geometry and material properties intended to contribute to shielding effectiveness. The SKGs are difficult problems, solutions are not obvious, and require integrated, iterative, and multi-disciplinary development. A lunar DRO lab built from the launch system elements enables an early and representative transit habitat test bed necessary for closing gaps before sending humans on a 1000 day Mars mission.

Griffin, Brand↗

NASA NTRS2015

SLS-Derived Lab- Precursor to Deep Space Human Exploration

Plans to send humans to Mars are in the works and the launch system is being built. Are we ready? Transportation, entry, landing, and surface operations have been successfully demonstrated for robotic missions. However, for human missions, there are significant, potentially show-stopping issues. These issues, called Strategic Knowledge Gaps (SKGs), are the unanswered questions concerning long duration exploration Beyond low Earth Orbit (BEO). The gaps represent a risk of loss of life or mission and because they require extended exposure to the weightless environment outside of earth's protective geo-magnetic field, they cannot be resolved on Earth or on the International Space Station (ISS). Placing a laboratory at a relatively close and stable lunar Distant Retrograde Orbit (DRO) provides an accessible location with the requisite environmental conditions for conducting SKG research and testing mitigation solutions. Configurations comprised of multiple 3 m and 4.3 m diameter modules have been studied but the most attractive solution uses elements of the human Mars launch vehicle or Space Launch System (SLS) for a Mars proving ground laboratory. A shortened version of an SLS hydrogen propellant tank creates a Skylab-like pressure vessel that flies fully outfitted on a single launch. This not only offers significant savings by incorporating SLS pressure vessel development costs but avoids the expensive ISS approach using many launches with substantial on-orbit assembly before becoming operational. One of the most challenging SKGs is crew radiation protection; this is why SKG laboratory research is combined with Mars transit habitat systems development. Fundamentally, the two cannot be divorced because using the habitat systems for protection requires actual hardware geometry and material properties intended to contribute to shielding effectiveness. The SKGs are difficult problems. The solutions to these problems are not obvious; they require integrated, iterative, and multi-disciplinary development. A lunar DRO lab built from SLS elements enables an early and representative transit habitat test bed necessary for closing gaps before sending humans on a 1,000-day Mars mission.

Griffin, Brand M.↗

NASA NTRS2001

Satellite Remote Sensing of Tropical Precipitation and Ice Clouds for GCM Verification

This project, supported by the NASA New Investigator Program, has primarily been funding a graduate student, Darren McKague. Since August 1999 Darren has been working part time at Raytheon, while continuing his PhD research. Darren is planning to finish his thesis work in May 2001, thus some of the work described here is ongoing. The proposed research was to use GOES visible and infrared imager data and SSM/I microwave data to obtain joint distributions of cirrus cloud ice mass and precipitation for a study region in the Eastern Tropical Pacific. These joint distributions of cirrus cloud and rainfall were to be compared to those from the CSU general circulation model to evaluate the cloud microphysical amd cumulus parameterizations in the GCM. Existing algorithms were to be used for the retrieval of cloud ice water path from GOES (Minnis) and rainfall from SSM/I (Wilheit). A theoretical study using radiative transfer models and realistic variations in cloud and precipitation profiles was to be used to estimate the retrieval errors. Due to the unavailability of the GOES satellite cloud retrieval algorithm from Dr. Minnis (a co-PI), there was a change in the approach and emphasis of the project. The new approach was to develop a completely new type of remote sensing algorithm - one to directly retrieve joint probability density functions (pdf's) of cloud properties from multi-dimensional histograms of satellite radiances. The usual approach is to retrieve individual pixels of variables (i.e. cloud optical depth), and then aggregate the information. Only statistical information is actually needed, however, and so a more direct method is desirable. We developed forward radiative transfer models for the SSM/I and GOES channels, originally for testing the retrieval algorithms. The visible and near infrared ice scattering information is obtained from geometric ray tracing of fractal ice crystals (Andreas Macke), while the mid-infrared and microwave scattering is computed with Mie scattering. The radiative transfer is performed with the Spherical Harmonic Discrete Ordinate Method (developed by the PI), and infrared molecular absorption is included with the correlated k-distribution method. The SHDOM radiances have been validated by comparison to version 2 of DISORT (the community "standard" discrete-ordinates radiative transfer model), however we use SHDOM since it is computationally more efficient.

Evans, K. Franklin↗

NASA NTRS2013

Orion GNC Mitigation Efforts for Van Allen Radiation

The Orion Crew Module (CM) is NASA's next generation manned space vehicle, scheduled to return humans to lunar orbit in the coming decade. The Orion avionics and GN&C architectures have progressed through a number of project phases and are nearing completion of a major milestone. The first unmanned test mission, dubbed "Exploration Flight Test One" (EFT-1) is scheduled to launch from NASA Kennedy Space Center late next year and provides the first integrated test of all the vehicle systems, avionics and software. The EFT-1 mission will be an unmanned test flight that includes a high speed re-entry from an elliptical orbit, which will be launched on an expendable launch vehicle (ELV). The ELV will place CM and the ELV upper stage into a low Earth orbit (LEO) for one revolution. After the first LEO, the ELV upper stage will re-ignite and place the combined upper stage/CM into an elliptical orbit whose perigee results in a high energy entry to test CM response in a relatively high velocity, high heating environment. While not producing entry velocities as high as those experienced in returning from a lunar orbit, the trajectory was chosen to provide higher stresses on the thermal protection and guided entry systems, as compared against a lower energy LEO entry. However the required entry geometry with constraints on inclination and landing site result in a trajectory that lingers for many hours in the Van Allen radiation belts. This exposes the vehicle and avionics to much higher levels of high energy proton radiation than a typical LEO or lunar trajectory would encounter. As a result, Van Allen radiation poses a significant risk to the Orion avionics system, and particularly the Flight Control Module (FCM) computers that house the GN&C flight software. The measures taken by the Orion GN&C, Flight Software and Avionics teams to mitigate the risks associated with the Van Allen radiation on EFT-1 are covered in the paper. Background on the Orion avionics subsystem is provided, as well as an overview of the GN&C software architecture. The measures taken to handle radiation induced failure of the one or both of the FCM's are presented, and finally simulation and actual hardware-in-the-loop (HWIL) results are shown confirming the validity of the implementation. The paper presents an overview of the Orion avionics architecture describing the GNC sensors, onboard data network as well as the flight control computers and their planned restart capabilities. GN&C sensors include two Orion Inertial Measurement Units (OIMU's), a Vision Processing Unit (VPU) to process camera images, three barometric altimeters and a single GPS receiver. All of the sensors communicate to one of two Power and Data Units (PDU's). The PDU's multiplex analog and serial data from the sensors and write the data to the Orion Data Network (ODN). The OIMU s write measurement messages directly as onto the ODN, but they are routed through PDU network switches.

King, Ellis T.↗

NASA NTRS1999

Overview of the Landsat-7 Mission

Landsat-7 is scheduled for launch on April 15 from the Western Test Range at Vandenberg Air Force Base, Calif., on a Delta-H expendable launch vehicle. The Landsat 7 satellite consists of a spacecraft bus being provided by Lockheed Martin Missiles and Space (Valley Forge, Pa.) and the Enhanced Thematic Mapper Plus instrument built by Raytheon (formerly Hughes) Santa Barbara Remote Sensing (Santa Barbara, Calif.). The instrument on board Landsat 7 is the Enhanced Thematic Mapper Plus (ETM+). ETM+ improves upon the previous Thematic Mapper (TM) instruments on Landsat's 4 and 5 (Fig. la and lb). It includes the previous 7 spectral bands measuring reflected solar radiation and emitted thermal emissions but, in addition, includes a new 15 in panchromatic (visible-near infrared) band. The spatial resolution of the thermal infrared band has also been improved to 60 m. Both the radiometric precision and accuracy of the sensor are also improved from the previous TM sensors. After being launched into a sun-synchronous polar orbit, the satellite will use on-board propulsion to adjust its orbit to a circular altitude of 438 miles (705 kilometers) crossing the equator at approximately 10 a.m. on its southward track. This orbit will place Landsat 7 along the same ground track as previous Landsat satellites. The orbit will be maintained with periodic adjustments for the life of the mission. A three-axis attitude control subsystem will stabilize the satellite and keep the instrument pointed toward the Earth to within 0.05 degrees. Later this year, plans call for the NASA Earth Observation System (EOS) Terra (AM-1) observatory and the experimental EO-1 mission to closely follow Landsat-7's orbit to support synergistic research and applications from this new suite of terrestrial sensor systems. Landsat is the United States' oldest land-surface observation satellite system, with satellites continuously operating since 1972. Although the program has scored numerous successes in scientific and resource-management applications, Landsat has had a tumultuous history of management and funding changes over its nearly 27-year history. Landsat-7 marks a new direction in the program to reduce the cost of data and increase systematic global coverage for use in global change research as well as commercial and regional applications. With the passage of the Land Remote Sensing Policy Act in 1992, oversight of the Landsat program began to shift from the commercial sector to the federal government. NASA integrated Landsat-7 into its EOS science program in 1994. Landsat-7 is managed and operated jointly by NASA and U.S. Geological Survey (USGS). As a result, the costs of acquiring observations from

Williams, Darrel↗

NASA NTRS2008

Design and Development of an In-Space Deployable Sun Shield for the Atlas Centaur

The Centaur, by virtue of its use of high specific-impulse (Isp) LO2/LH2 propellants, has initial mass-to-orbit launch requirements less than half of those upper stages using storable propellants. That is, for Earth escape or GSO missions the Centaur is half the launch weight of a storable propellant upper stage. A drawback to the use of Liquid oxygen and liquid hydrogen, at 90 K and 20 K respectively, over storable propellants is the necessity of efficient cryogen storage techniques that minimize boil-off from thermal radiation in space. Thermal blankets have been used successfully to shield both the Atlas Centaur and Titan Centaur. These blankets are protected from atmospheric air loads during launch by virtue of the fact that the Centaur is enclosed within the payload fairing. The smaller Atlas V vehicle, the Atlas 400, has the Centaur exposed to the atmosphere during launch, and therefore, to date has not flown with thermal blankets shielding the Centaur. A design and development effort is underway to fly a thermal shield on the Atlas V 400 vehicle that is not put in place until after the payload fairing jettisons. This can be accomplished by the use of an inflatable and deployable thermal blanket referred to as the Centaur Sun Shield (CSS). The CSS design is also scalable for use on a Delta upper stage, and the technology potentially could be used for telescope shades, re-entry shields, solar sails and propellant depots. A Phase I effort took place during 2007 in a partnership between ULA and ILC Dover which resulted in a deployable proof-of-concept Sun Shield being demonstrated at a test facility in Denver. A Phase H effort is underway during 2008 with a partnership between ULA, ILC, NASA Glenn Research Center (GRC) and NASA Kennedy Space Center (KSC) to define requirements, determine materials and fabrication techniques, and to test components in a vacuum chamber at cold temperatures. This paper describes the Sun Shield development work to date, and the future plans leading up to a flight test in the 2011 time frame.

Dew, Michael↗