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At least 271 records · Page 15
All sourcesNASA NTRSDOE OSTI
NASA NTRS1994

Common spaceborne multicomputer operating system and development environment

A preliminary technical specification for a multicomputer operating system is developed. The operating system is targeted for spaceborne flight missions and provides a broad range of real-time functionality, dynamic remote code-patching capability, and system fault tolerance and long-term survivability features. Dataflow concepts are used for representing application algorithms. Functional features are included to ensure real-time predictability for a class of algorithms which require data-driven execution on an iterative steady state basis. The development environment supports the development of algorithm code, design of control parameters, performance analysis, simulation of real-time dataflow applications, and compiling and downloading of the resulting application.

Craymer, L. G.↗
NASA NTRS2003

Physics of Regolith Impacts in Microgravity Experiment (PRIME)

Collisions between planetary ring particles and in some protoplanetary disk environments occur at low impact velocities (v less than 1 m/s) . In some regions of Saturn s rings, for example, the typical collision velocity inferred from observations by the Voyager spacecraft and dynamical modeling is a fraction of a centimeter per second. Although no direct observations of an individual ring particle exist, the abundance of dust in planetary rings and protoplanetary disks suggests that larger ring and disk particles are coated with a layer of smaller particles and dust - the "regolith". Because the ring particles and proto-planetesimals are small (cm to m-sized), the regolith is only weakly bound to the surface by gravity. Similarly, secondary impacts on asteroids by large blocks of ejecta from high velocity cratering events result in low velocity impacts into the asteroid regolith, which is also weakly bound by the asteroid s gravity. At the current epoch and throughout their history, low velocity collisions have played an important role in sculpting planetary systems. In a one-Earth-gravity environment, it is not possible to experimentally determine the behavior of impact eject from such low velocity collisions. Impacts typically occur at speeds exceeding the mutual escape velocity of the two bodies. Thus, impacts at speeds on the order of 10 m/sec or less involve objects that are tens of meters across, or smaller. This research program is an experimental study of such low velocity collisions in a microgravity environment. The experimental work builds on the Collisions Into Dust Experiment (COLLIDE), which has flown twice on the space shuttle. The PRIME experimental apparatus is a new apparatus designed specifically for the environment provided on the NASA KC- 135 reduced gravity aircraft.

Motil, Brian↗
NASA NTRS2019

A Comparison of ARTEMIS Data with the Lunar Plasma Design Environment for NASA Crewed Missions

NASA’s Gateway will provide the capability for sustaining a human presence in cis-lunar space. Operations of the Gateway will include spacecraft dockings, extra vehicular activities (EVA), and high-power solar arrays. NASA’s experience with the International Space Station highlighted the importance of evaluating spacecraft charging effects for such operations. For crewed spacecraft, which tend to employ the use of dielectric surfaces in this dynamic plasma environment, reliance on spacecraft charging simulation packages, such as the NASA/Air Force Spacecraft Charging Analyzer Program (Nascap-2k) [Mandell et al., 2006] and Spacecraft Plasma Interaction System (SPIS) [Roussel et al., 2008], is required to understand the risks to hardware and humans. The variability in the lunar plasma environment as the Moon revolves around the Earth, lunar wake effects, and a strong dependency on photoemission and secondary electron emission creates challenges for spacecraft charging analysis. The Design Specification for Natural Environments (DSNE) [NASA, MSFC] is the primary resource for space environments affecting NASA’s crewed missions, and the DSNE provides plasma environments in a standard form for input into simulation packages. NASA developed the existing lunar plasma environment using data from Geotail [Nishida, 1994] along with published lunar plasma wake models [Halekas et al., 2005] based on Lunar Prospector. Since 2011, NASA’s twin Acceleration Reconnection Turbulence & Electrodynamics of Moon’s Interaction with the Sun (ARTEMIS) satellites [Angelopoulos, 2010] have been collecting high resolution plasma and fields observations within the lunar plasma environment providing a much larger dataset of the plasma properties in cislunar space. This research compares the existing lunar plasma environment definition with ARTEMIS data and makes recommendations on the refinement of the environment definition for future lunar missions.

ARTEMIS mission↗
NASA NTRS2019

Enabling Assurance in the MBSE Environment

A number of specific benefits that fit within the hallmarks of effective development are realized with implementation of model-based approaches to systems and assurance. Model Based Systems Engineering (MBSE) enabled by standardized modeling languages (e.g., SysML®) is at the core. These benefits in the context of spaceflight system challenges can include: Improved management of complex development, Reduced risk in the development process, Improved cost management, Improved design decisions. With appropriate modeling techniques the assurance community can improve early oversight and insight into project development. NASA has shown the basic constructs of SysML in an MBSE environment offer several key advantages, within a Model Based Mission Assurance (MBMA) initiative.

Evans, John W.↗
NASA NTRS2020

Enabling Assurance in the MBSE Environment

A number of specific benefits that fit within the hallmarks of effective development are realized with implementation of model-based approaches to systems and assurance. Model Based Systems Engineering (MBSE) enabled by standardized modeling languages (e.g., SysML®) is at the core. These benefits in the context of spaceflight system challenges can include [1]: • Improved management of complex development • Reduced risk in the development process • Improved cost management • Improved design decisions With appropriate modeling techniques the assurance community also can improve early oversight and insight into project development. NASA has shown the basic constructs of SysML in an MBSE environment offer several key advantages, within a Model Based Mission Assurance (MBMA) initiative [2, 3]. These include the following: • Model viewpoints that promote rapid and systematic assessment of requirements coverage, hazard tagging and risk management • Embedded safety assessments for launch vehicles • Deployment of model assisted development of reliability products - Failure Modes and Effects Analyses (FMEAs) and Fault Trees • Test Planning • Validation and Verification of complex functions • Support of Assurance Case development for complex systems In addition, while there are benefits to be harvested, there is a realization that these do not come without effort and cost. Enabling model-based approaches requires structure, not only in an organizational context, but in a modeling context as well. There can be a steep learning curve and costs associated to train skilled modelers. But, on the other hand, not all of the assurance community need to be modelers. Models themselves must conform to ontologies that enable assurance. This places constraints upon the models and modelers. Optimums have yet to be developed where resources and constraints on modeling must be traded off in the organization and modeling efforts for projects. A number of barriers need to be overcome, as well, which pose challenges to the developers of the software that supports MBSE/MBMA. Information and data must be made to flow seamlessly through the life cycle. Because there is a wide variety of tools used in the community, to avoid the problems of the past of silos, delays, and diverging interests, information should flow among these tools to support the “single source of truth” paradigm of MBSE. This will greatly facilitate MBMA and advancement of assurance functions.

Evans, John W.↗
NASA NTRS2011

Testing and Failure Mechanisms of Ice Phase Change Material Heat Exchangers

Phase change materials (PCM) may be useful for thermal control systems that involve cyclical heat loads or cyclical thermal environments such as specific spacecraft orientations in Low Earth Orbit (LEO) and low beta angle Low Lunar Orbit (LLO). Thermal energy can be stored in the PCM during peak heat loads or in adverse thermal environments. The stored thermal energy can then be released later during minimum heat loads or in more favorable thermal environments. One advantage that PCM s have over evaporators in this scenario is that they do not use a consumable. The use of water as a PCM rather than the more traditional paraffin wax has the potential for significant mass reduction since the latent heat of formation of water is approximately 70% greater than that of wax. One of the potential drawbacks of using ice as a PCM is its potential to rupture its container as water expands upon freezing. In order to develop a space qualified ice PCM heat exchanger, failure mechanisms must first be understood. Therefore, a methodical experimental investigation has been undertaken to demonstrate and document specific failure mechanisms due to ice expansion in the PCM. A number of ice PCM heat exchangers were fabricated and tested. Additionally, methods for controlling void location in order to reduce the risk of damage due to ice expansion were investigated. This paper presents the results of testing that occurred from March through September of 2010 and builds on testing that occurred during the previous year.

Leimkuehler, Thomas O.↗
NASA NTRS1999

Proton Effects and Test Issues for Satellite Designers: Displacement Effects - Section 4

Microelectronic and photonic systems in the natural space environment are bombarded by a variety of charged particles including electrons, trapped protons, cosmic rays, and solar particles (protons and other heavy ions). These incident particles cause both ionizing and non-ionizing effects when traversing a device, and the effects can be either transient or permanent. The vast majority of the kinetic energy of an incident proton is lost to ionization, creating the single event effects (SEES) and total ionizing dose (TID) effects described in section IVA. However, the small portion of energy lost in non-ionizing processes causes atoms to be removed from their lattice sites and form permanent electrically active defects in semiconductor materials. These defects, i.e., "displacement damage," can significantly degrade device performance. In general, most of the displacement damage effects in the natural space environment can be attributed to protons since they are plentiful and extremely energetic (and therefore not readily shielded against). For this reason, we consider only proton induced displacement damage in this course. (Nevertheless, we identify solar cells as an important example of a case where both electron and proton damage can be important since only very light shielding is feasible.) The interested reader is encouraged to explore the three previous NSREC and RADECS short courses [Srou88a, Summ92, Hopk97] which also treat displacement damage issues for satellite applications. Part A of this segment of the short course introduces the space environment, proton shielding issues, and requirements specifications for proton-rich environments. In order to exercise the displacement damage analysis tools for on-orbit performance predictions, the requirements document must provide the relevant proton spectra in addition to the usual total ionizing dose-depth curves.

Marshall, Cheryl J.↗
NASA NTRS1999

Proton Effects and Test Issues for Satellite Designers

Microelectronic and photonic systems in the natural space environment are bombarded by a variety of charged particles including electrons, trapped protons, cosmic rays, and solar particles (protons and other heavy ions). These incident particles cause both ionizing and non-ionizing effects when traversing a device, and the effects can be either transient or permanent. The vast majority of the kinetic energy of an incident proton is lost to ionization, creating the single event effects (SEES) and total ionizing dose (TID) effects. However, the small portion of energy lost in non-ionizing processes causes atoms to be removed from their lattice sites and form permanent electrically active defects in semiconductor materials. These defects, i.e., "displacement damage," can significantly degrade device performance. In general, most of the displacement damage effects in the natural space environment can be attributed to protons since they are plentiful and extremely energetic (and therefore not readily shielded against). For this reason, we consider only proton induced displacement damage in this course. (Nevertheless, we identify solar cells as an important example of a case where both electron and proton damage can be important since only very light shielding is feasible.) The interested reader is encouraged to explore the three previous NSREC and RADECS short courses which also treat displacement damage issues for satellite applications. Part A of this segment of the short course introduces the space environment, proton shielding issues, and requirements specifications for proton-rich environments. In order to exercise the displacement damage analysis tools for on-orbit performance predictions, the requirements document must provide the relevant proton spectra in addition to the usual total ionizing dose-depth curves. Ion-solid interactions and the nature of the displacement damage they generate have been studied extensively for over half a century, yet they still remain a subject of investigation. In this section, a description of the mechanisms by which displacement damage is produced will be followed by a summary of the major consequences for device performance in a space environment. Often the degradation of a device parameter can be characterized by a damage factor (measured in a laboratory using monoenergetic protons) that is simply the change in a particular electrical or optical parameter per unit proton fluence. In addition, we will describe the concept of a non-ionizing energy loss rate (NIEL) which quantifies that portion of the energy lost by an incident ion that goes into displacements. It has been calculated as a function of proton energy, and is analogous to (and has the same units as) the linear energy transfer (LET) for ionizing energy. We will discover that, to first order, the calculated NIEL describes the energy dependence of the measured device damage factors. This observation provides the basis for predicting proton induced device degradation in a space environment based on both the calculated NIEL and relatively few laboratory test measurements. The methodology of such on-orbit device performance predictions will be described, as well as the limitations. Several classes of devices for which displacement damage is a significant (if not the dominant) mode of radiation induced degradation will be presented.

Marshall, Cheryl J.↗
NASA NTRS1992

Ionizing radiation exposure of LDEF (pre-recovery estimates)

The long duration exposure facility (LDEF), launched into a 258 nautical mile orbit with an inclination of 28.5 degrees, remained in space for nearly 6 yr. The 21,500 lb NASA satellite was one of the largest payloads ever deployed by the Space Shuttle. LDEF completed 32,422 orbits and carried 57 major experiments representing more than 200 investigators from 33 private companies, 21 universities and nine countries. The experiments covered a wide range of disciplines including basic science, electronics, optics, materials, structures and power and propulsion. A number of the experiments were specifically designed to measure the radiation environment. These experiments are of specific interest, since the LDEF orbit is essentially the same as that of the Space Station Freedom. Consequently, the radiation measurements on LDEF will play a significant role in the design of radiation shielding of the space station. The contributions of the various authors presented here attempt to predict the major aspects of the radiation exposure received by the various LDEF experiments and therefore should be helpful to investigators who are in the process of analyzing experiments which may have been affected by exposure to ionizing radiation. The paper discusses the various types and sources of ionizing radiation including cosmic rays, trapped particles (both protons and electrons) and secondary particles (including neutrons, spallation products and high-LET recoils), as well as doses and LET spectra as a function of shielding. Projections of the induced radioactivity of LDEF are also discussed.

Flight Experiment↗
NASA NTRS1993

Marshall Space Flight Center's Virtual Reality Applications Program 1993

A Virtual Reality (VR) applications program has been under development at the Marshall Space Flight Center (MSFC) since 1989. Other NASA Centers, most notably Ames Research Center (ARC), have contributed to the development of the VR enabling technologies and VR systems. This VR technology development has now reached a level of maturity where specific applications of VR as a tool can be considered. The objectives of the MSFC VR Applications Program are to develop, validate, and utilize VR as a Human Factors design and operations analysis tool and to assess and evaluate VR as a tool in other applications (e.g., training, operations development, mission support, teleoperations planning, etc.). The long-term goals of this technology program is to enable specialized Human Factors analyses earlier in the hardware and operations development process and develop more effective training and mission support systems. The capability to perform specialized Human Factors analyses earlier in the hardware and operations development process is required to better refine and validate requirements during the requirements definition phase. This leads to a more efficient design process where perturbations caused by late-occurring requirements changes are minimized. A validated set of VR analytical tools must be developed to enable a more efficient process for the design and development of space systems and operations. Similarly, training and mission support systems must exploit state-of-the-art computer-based technologies to maximize training effectiveness and enhance mission support. The approach of the VR Applications Program is to develop and validate appropriate virtual environments and associated object kinematic and behavior attributes for specific classes of applications. These application-specific environments and associated simulations will be validated, where possible, through empirical comparisons with existing, accepted tools and methodologies. These validated VR analytical tools will then be available for use in the design and development of space systems and operations and in training and mission support systems.

Hale, Joseph P., II↗
NASA NTRS2008

Defining the Natural Atmospheric Environment Requirements for the NASA Constellation Program

The National Aeronautics and Space Administration began developing a new vehicle under the Constellation Program to replace the Space Shuttle. The Ares-1 launch vehicle and the Orion capsule will be used to ferry crew and some payloads to the International Space Station and will also be used for new missions to the moon, As development of this new vehicle begins, the Natural Environments Branch at Marshall Space Flight Center has been tasked with defining the natural environments the vehicle will encounter and working with the program to develop natural environmental requirements for the vehicles' elements. An overview of the structure of the program is given, along with a description of the Constellation Design Specification for Natural Environments and the Constellation Natural Environments Definition for Design documents and how they apply to the Ares-I and Orion vehicles.

Roberts, Barry C.↗
NASA NTRS2018

Flight Plasma Diagnostics for High-Power, Solar-Electric Deep-Space Spacecraft

NASA’s Asteroid Redirect Robotic Mission (ARRM) mission concept plans included a set of plasma and space environment instruments, the Plasma Diagnostic Package (PDP), to fulfill ARRM requirements for technology extensibility to future missions. The PDP objectives were divided into the classes of 1) Plasma thruster dynamics, 2) Solar array-specific environmental effects, 3) Plasma environmental spacecraft effects, and 4) Energetic particle spacecraft environment. A reference design approach and interface requirements for ARRM’s PDP was generated by the PDP team at JPL and GRC. The reference design consisted of redundant single-string avionics located on the ARRM spacecraft bus as well as solar array, driving and processing signals from multiple copies of several types of plasma, effects, and environments sensors distributed over the spacecraft and array. The reference design sensor types were derived in part from sensors previously developed for USAF Research Laboratory (AFRL) plasma effects campaigns such as those aboard TacSat-2 in 2007 and AEHF-2 in 2012. ARRM project leadership also encouraged the PDP team to convene a team of topical subject matter experts from across the country to review and confirm the reference design and to consider effective alternatives and/or enhancements to the reference design. This activity was proposed and accepted as an interactive, informal JPL “A-team” study and a cadre of 25 participants gathered in early 2017 for discussions. The outcome of the two-day A-team study was that the PDP reference design would allow the most important induced-environment unknowns to be measured in the appropriate space environment. Another outcome addressed technology developments of new or improved space plasma environmental sensors. The A-team study concluded that selected developments would lead to improved measurements that could efficiently provide important and otherwise unavailable information about plasma thruster operation in the space environment as well as the plasma induced spacecraft environment. Specifically, the A-team group recommended greater sensor diversity, by inclusion of deployed sensor capabilities in the thruster plume, or by occasional gimbaling of the thruster(s) toward the sensor arrays. The A-team also recommended developing high-speed probes, optical plasma probes, energy selective probes, and direct erosion/deposition sensors, among others; and recommended the inclusion of cameras as well as, since ARRM was to be recovered in cis-lunar orbit by a crewed mission, astronaut assessments of thruster induced environments and collection of sample coupons. Overall, the PDP A-team study provided a clear, consensus supported validation of the reference design PDP approach and pointed out important directions for future flight plasma sensor development.

Choi, Maria↗
NASA NTRS1989

Requirements and designs for Mars Rover RTGs

A typical Rover mission, its requirements, the environment it imposes on the radionuclide thermoelectric generator (RTG), and a design approach for making the RTG operable in such an environment are described. Specific RTG designs for various thermoelectric element alternatives are presented. The results of studies show that current multifoil-insulated GPHS (general-purpose heat source) RTG and Mod-RTG designs can be modified to operate in an environment with an external atmosphere (e.g., Mars). This can be done while the helium generated by the fuel's alpha decay is vented to the external atmosphere. The use of novel selective vents and high-capacity getters is not required. The Rover RTGs can be built from standard, proven GPHS modules and standard SiGe unicouples or SiGe/GaP multicouples, using demonstrated thermoelectric material performance parameters. Rover's 500-W power requirement can be satisfied with two 250-W or four 125-W RTGs, whose sizes are compatible with currently envisaged Rover designs. An auxiliary cooling loop (e.g., water and antifreeze) will be required to cool the RTG while it is within the Rover's aeroshell during launch and transit to Mars.

Schock, A.↗
NASA NTRS1989

IDEF3 and IDEF4 automation system requirements document and system environment models

The requirements specification is provided for the IDEF3 and IDEF4 tools that provide automated support for IDEF3 and IDEF4 modeling. The IDEF3 method is a scenario driven process flow description capture method intended to be used by domain experts to represent the knowledge about how a particular system or process works. The IDEF3 method provides modes to represent both (1) Process Flow Description to capture the relationships between actions within the context of a specific scenario, and (2) Object State Transition to capture the allowable transitions of an object in the domain. The IDEF4 method provides a method for capturing the (1) Class Submodel or object hierarchy, (2) Method Submodel or the procedures associated with each classes of objects, and (3) the Dispath Matching or the relationships between the objects and methods in the object oriented design. The requirements specified describe the capabilities that a fully functional IDEF3 or IDEF4 automated tool should support.

Blinn, Thomas M.↗
NASA NTRS2011

Modeling Mission-Specific Worst-Case Solar Energetic Particle Environments

To plan and design safe and reliable space missions, it is necessary to take into account the effects of the space radiation environment. The environment during large solar energetic particle events poses the greatest challenge to missions. As a starting point for planning and design, a reference environment must be specified representing the most challenging environment to be encountered during the mission at some confidence level. The engineering challenge is then to find plans and mission design solutions that insure safe and reliable operations in this reference environment. This paper describes progress toward developing a model that provides such reference space radiation environments at user-specified confidence levels.

Adam, James H., Jr.↗
NASA NTRS1990

Life sciences research using a lunar laboratory

The necessity for life sciences research on the lunar surface in order to determine the consequences of returning from extended missions in various low gravity environments and of transiting through high multiple gravity forces during decelerations is discussed. The functions of a lunar gravitational biology laboratory are outlined. Lunar science objectives include investigations in developmental biology including the evaluation of the capacity of diverse organisms to undergo normal development and the evaluation of the use of the lunar environment to study specific developmental phenomena in ways that cannot be accomplished by earth-based research. The need for musculoskeletal studies to examine the dynamics of osteoclast and osteoblast formation and breakdown and to address bone and demineralization problems is discussed. Biological adaptation to hypogravic environments and the effects of radiation and electromagnetic environmental factors are also considered.

Cipriano, Leonard F.↗
NASA NTRS2007

Progress in Spacecraft Environment Interactions: International Space Station (ISS) Development and Operations

The set of spacecraft interactions with the space flight environment that have produced the largest impacts on the design, verification, and operation of the International Space Station (ISS) Program during the May 2000 to May 2007 time frame are the focus of this paper. In-flight data, flight crew observations, and the results of ground-based test and analysis directly supporting programmatic and operational decision-making are reported as are the analysis and simulation efforts that have led to new knowledge and capabilities supporting current and future space explorations programs. The specific spacecraft-environment interactions that have had the greatest impact on ISS Program activities during the first several years of flight are: 1) spacecraft charging, 2) micrometeoroids and orbital debris effects, 3) ionizing radiation (both total dose to materials and single event effects [SEE] on avionics), 4) hypergolic rocket engine plume impingement effects, 5) venting/dumping of liquids, 6) spacecraft contamination effects, 7) neutral atmosphere and atomic oxygen effects, 8) satellite drag effects, and 9) solar ultraviolet effects. Orbital inclination (51.6deg) and altitude (nominally between 350 km and 460 km) determine the set of natural environment factors affecting the performance and reliability of materials and systems on ISS. ISS operates in the F2 region of Earth s ionosphere in well-defined fluxes of atomic oxygen, other ionospheric plasma species, solar UV, VUV, and x-ray radiation as well as galactic cosmic rays, trapped radiation, and solar cosmic rays. The micrometeoroid and orbital debris environment is an important determinant of spacecraft design and operations in any orbital inclination. The induced environment results from ISS interactions with the natural environment as well as environmental factors produced by ISS itself and visiting vehicles. Examples include ram-wake effects, hypergolic thruster plume impingement, materials out-gassing, venting and dumping of fluids, and specific photovoltaic (PV) power system interactions with the ionospheric plasma. Vehicle size (L) and velocity (v), combined with the magnitude and direction of the geomagnetic field (B) produce operationally significant magnetic induction voltages (VxB.L) in ISS conducting structure during high latitude flight (>+/- 45deg) during each orbit. In addition, ISS is a large vehicle and produces a deep wake structure from which both ionospheric plasma and neutrals species are largely excluded. ISS must fly in a very limited number of approved flight attitudes, so that exposure of a particular material or system to environmental factors depends upon: 1) location on ISS, 2) ISS flight configuration, 3) ISS flight attitude, and 4) variation of solar exposure (Beta angle), and hence thermal environment, with time. Finally, an induced ionizing radiation environment is produced by trapped radiation and solar/cosmic ray interactions with the relatively massive ISS structural shielding.

Koontz, Steve↗