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At least 343 records · Page 19
All sourcesNASA NTRSDOE OSTI
NASA NTRS1993

LDEF Materials Results for Spacecraft Applications

These proceedings describe the application of LDEF data to spacecraft and payload design, and emphasize where space environmental effects on materials research and development is needed as defined by LDEF data. The LDEF six years of exposure of materials has proven to be by far the most comprehensive source of information ever obtained on the long-term performance of materials in the space environment. The conference provided a forum for materials scientists and engineers to review and critically assess the LDEF results from the standpoint of their relevance, significance, and impact on spacecraft design practice. The impact of the LDEF findings on materials selection and qualification, and the needs and plans for further study, were addressed from several perspectives. Many timely and needed changes and modifications in external spacecraft materials selection have occurred as a result of LDEF investigations.

Whitaker, Ann F.↗
NASA NTRS2012

Space Environments and Spacecraft Effects Organization Concept

The National Aeronautics and Space Administration (NASA) is embarking on a course to expand human presence beyond Low Earth Orbit (LEO) while also expanding its mission to explore the solar system. Destinations such as Near Earth Asteroids (NEA), Mars and its moons, and the outer planets are but a few of the mission targets. Each new destination presents an opportunity to increase our knowledge of the solar system and the unique environments for each mission target. NASA has multiple technical and science discipline areas specializing in specific space environments disciplines that will help serve to enable these missions. To complement these existing discipline areas, a concept is presented focusing on the development of a space environments and spacecraft effects (SENSE) organization. This SENSE organization includes disciplines such as space climate, space weather, natural and induced space environments, effects on spacecraft materials and systems and the transition of research information into application. This space environment and spacecraft effects organization will be composed of Technical Working Groups (TWG). These technical working groups will survey customers and users, generate products, and provide knowledge supporting four functional areas: design environments, engineering effects, operational support, and programmatic support. The four functional areas align with phases in the program mission lifecycle and are briefly described below. Design environments are used primarily in the mission concept and design phases of a program. Engineering effects focuses on the material, component, sub-system and system-level selection and the testing to verify design and operational performance. Operational support provides products based on real time or near real time space weather to mission operators to aid in real time and near-term decision-making. The programmatic support function maintains an interface with the numerous programs within NASA, other federal government agencies, and the commercial sector to ensure that communications are well established and the needs of the programs are being met. The programmatic support function also includes working in coordination with the program in anomaly resolution and generation of lessons learned documentation. The goal of this space environment and spacecraft effects organization is to develop decision-making tools and engineering products to support all mission phases from mission concept through operations by focusing on transitioning research to application. Products generated by this space environments and effects application are suitable for use in anomaly investigations. This paper will describe the scope of the TWGs and their relationship to the functional areas, and discuss an organizational structure for this space environments and spacecraft effects organization.

Edwards, David L.↗
NASA NTRS2012

An Overview of the Space Environments and Spacecraft Effects Organization Concept

The National Aeronautics and Space Administration (NASA) is embarking on a course to expand human presence beyond Low Earth Orbit (LEO) while also expanding its mission to explore our Earth, and the solar system. Destinations such as Near Earth Asteroids (NEA), Mars and its moons, and the outer planets are but a few of the mission targets. Each new destination presents an opportunity to increase our knowledge on the solar system and the unique environments for each mission target. NASA has multiple technical and science discipline areas specializing in specific space environments fields that will serve to enable these missions. To complement these existing discipline areas, a concept is presented focusing on the development of a space environment and spacecraft effects (SESE) organization. This SESE organization includes disciplines such as space climate, space weather, natural and induced space environments, effects on spacecraft materials and systems, and the transition of research information into application. This space environment and spacecraft effects organization will be composed of Technical Working Groups (TWG). These technical working groups will survey customers and users, generate products, and provide knowledge supporting four functional areas: design environments, engineering effects, operational support, and programmatic support. The four functional areas align with phases in the program mission lifecycle and are briefly described below. Design environments are used primarily in the mission concept and design phases of a program. Environment effects focuses on the material, component, sub-system, and system-level response to the space environment and include the selection and testing to verify design and operational performance. Operational support provides products based on real time or near real time space weather to mission operators to aid in real time and near-term decision-making. The programmatic support function maintains an interface with the numerous programs within NASA, other federal government agencies, and the commercial sector to ensure that communications are well established and the needs of the programs are being met. The programmatic support function also includes working in coordination with the program in anomaly resolution and generation of lessons learned documentation. The goal of this space environment and spacecraft effects organization is to develop decision-making tools and engineering products to support all mission phases from mission concept through operations by focusing on transitioning research to application. Products generated by this space environments and effects application are suitable for use in anomaly investigations. This paper will describe the scope and purpose of the space environments and spacecraft effects organization and describe the TWG's and their relationship to the functional areas.

Edwards, David L.↗
NASA NTRS1972

Ultraviolet and charged particle irradiation of proposed solar cell coverslide materials and conductive coatings for the Helios spacecraft

Coverslide materials consisting of Corning 7940 fused silica, multilayers of titanium and manganese oxides (blue reflector), and indium oxide (conductive-coating) were exposed to 16 UVSC up to 800 EUVSH in vacuum. Slight changes in optical transmittance and optical absorptance were found in the (200-360) millimicron regions of the fused silica and conductive coating respectively. Exposure to 4 KeV protons and 4.5 KeV electrons in vacuum, produced decreases of several percent in transmittance, (200-360) millimicron region in the fused silicas after total fluxes less or = 10 to the 14th power particles/sq cm. Sheet resistance of the conductive coating increased above 1.0 kilo-ohm/square after a total flux less or = 10 to the 14th power particles/sq cm. Solar cells with coverglasses utilizing the indium oxide conductive coating were exposed to 1 Mev electrons and 1 Mev protons in air and in vacuum. Total fluxes ranged from 10 to the 11th power particles/sq cm to 10 to the 15th power particle/sq cm. There was no appreciable degradation in the resistance of the conductive coating during or after these tests.

Fry, J.↗
NASA NTRS1991

First LDEF Post-Retrieval Symposium abstracts

The LDE facility was designed to better understand the environments of space and the effects of prolonged exposure in these environments on future spacecraft. The symposium abstracts presented here are organized according to the symposium agenda into five sessions. The first session provides an overview of the LDEF, the experiments, the mission, and the natural and induced environments the spacecraft and experiments encountered during the mission. The second session presents results to date from studies to better define the environments of near-Earth space. The third session addresses studies of the effects of the space environments on spacecraft materials. The fourth session addresses studies of the effects of the space environments on spacecraft systems. And the fifth session addresses other subjects such as results of the LDEF life science and crystal growth experiments.

Levine, Arlene S.↗
NASA NTRS1991

The interactions of atmospheric cosmogenic radionuclides with spacecraft surfaces

The discovery of the cosmogenic radionuclide Be-7 on the front surface of the Long Duration Exposure Facility (LDEF) has opened new opportunities to study several unexplored regions of space science. The experiments have shown that the Be-7 found was concentrated in a thin surface layer of spacecraft material. The only reasonable source of the isotope is the atmosphere through which the spacecraft passed. It is expected that the uptake of Be in such circumstances will depend on the chemical form of the Be and the chemical nature of the substrate. It was found that the observed concentration of Be-7 does differ between metal surfaces and organic surfaces such as PTFE (Teflon). It is noted however, that (1) organic surfaces are etched by the atomic oxygen found under these orbital conditions, and (2) the relative velocity of the species is 8 km/s relative to the surface and the interaction chemistry and physics may differ from the norm. Be-7 is formed by disintegration of O and N nuclei under cosmic ray proton bombardment. Many other isotopes are produced by cosmic ray reactions, and some of these are suited to measurement by the extremely sensitive methods of accelerator mass spectrometry.

Gregory, John C.↗
NASA NTRS1993

Selected results for LDEF thermal control coatings

Several different thermal control coatings were analyzed as part of the Long Duration Exposure Facility (LDEF) Materials Special Investigation Group activity and as part of the Space Environment Effects on Spacecraft Materials Experiment M0003. A brief discussion of the results obtained for these materials is presented.

Golden, Johnny L.↗
NASA NTRS2019

Spectroscopic Behavior of Composite, Black Thermal Paint, Solar Cell, and Multi-layered Insulation Materials in a GEO Simulated Environment

Materials currently populating Earth orbital regimes can be distinguished by comparing remote observational data to that of optical material measurements obtained in the laboratory. Experimentation for this research primarily involved the acquisition of spectroscopic measurements on materials of interest to the telescopic observational community for enhanced space situational awareness. Common spacecraft materials worthy of preeminent analysis for this investigation include a carbon-carbon (c-c) matrix composite, various black thermal paints, a GPS solar cell and three different cover glass components. These materials were subjected to a simulated geosynchronous Earth orbit (GEO) space environment with the intent of observing material optical property behavior over quantitative exposure time. The aforementioned materials have been measured in their pristine and GEO simulated exposed conditions. A reflectance spectrometer and a bi-directional reflectance distribution function (BRDF) optical system have been operated to perform material characterization, optical property analysis, and to further compare such data to telescopic observational data acquired on equal materials.

Reyes, Jacqueline A.↗
NASA NTRS1979

Materials testing for spacecraft use

A technique has been developed for determining the outgassing characteristics of materials intended for spacecraft use. A measurement of the total mass loss (TML) and of the collected volatile condensable materials (CVCM) is used as a guide for selecting those materials low in outgassing. The data were used to compare the outgassing from two conductive black paints. In addition, the apparatus was also used in an attempt to simulate expected launch temperatures by collecting the outgassed molecules on a glass plate, which simulates a nearby thermal control surface. The glass plate was then irradiated in vacuum by simulated sunlight. Transmittance patterns of the clean glass plate and also with the condensed material on it were compared with the transmittance pattern after irradiation of the condensate. The transmittance patterns showed a drastic change. The resultant discoloration indicated the detrimental effects of the contamination and aided in selecting the better conductive paint. Additional tests utilizing an aluminized mirror collector plate have been run to show the effect of irradiation upon the CVCM from certain different classes of materials often used on spacecraft.

Park, J. J.↗
NASA NTRS2013

Unmanned Vehicle Material Flammability Test

Microgravity combustion phenomena have been an active area of research for the past 3 decades however, there have been very few experiments directly studying spacecraft fire safety under low-gravity conditions. Furthermore, none of these experiments have studied sample and environment sizes typical of those expected in a spacecraft fire. All previous experiments have been limited to samples of the order of 10 cm in length and width or smaller. Terrestrial fire safety standards for all other habitable volumes on earth, e.g. mines, buildings, airplanes, ships, etc., are based upon testing conducted with full-scale fires. Given the large differences between fire behavior in normal and reduced gravity, this lack of an experimental data base at relevant length scales forces spacecraft designers to base their designs using 1-g understanding. To address this question a large scale spacecraft fire experiment has been proposed by an international team of investigators. This poster presents the objectives, status and concept of this collaborative international project to examine spacecraft material flammability at realistic scales. The concept behind this project is to utilize an unmanned spacecraft such as Orbital Cygnus vehicle after it has completed its delivery of cargo to the ISS and it has begun its return journey to earth. This experiment will consist of a flame spread test involving a meter scale sample ignited in the pressurized volume of the spacecraft and allowed to burn to completion while measurements are made. A computer modeling effort will complement the experimental effort. Although the experiment will need to meet rigorous safety requirements to ensure the carrier vehicle does not sustain damage, the absence of a crew removes the need for strict containment of combustion products. This will facilitate the examination of fire behavior on a scale that is relevant to spacecraft fire safety and will provide unique data for fire model validation. This will be the first opportunity to examine microgravity flame behavior at scales approximating a spacecraft fire.

Combustion↗
NASA NTRS1990

Laser supported detonation wave source of atomic oxygen for aerospace material testing

A pulsed high-flux source of nearly monoenergetic atomic oxygen was developed to perform accelerated erosion testing of spacecraft materials in a simulated low-earth orbit (LEO) environment. Molecular oxygen is introduced into an evacuated conical expansion nozzle at several atmospheres pressure through a pulsed molecular beam valve. A laser-induced breakdown is generated in the nozzle throat by a pulsed CO2 TEA laser. The resulting plasma is heated by the ensuing laser-supported detonation wave, and then it rapidly expands and cools. An atomic oxygen beam is generated with fluxes above 10 to the 18th atoms per pulse at 8 + or - 1.6 km/s with an ion content below 1 percent for LEO testing. Materials testing yielded the same surface oxygen enrichment in polyethylene samples as observed on the STS mission, and scanning electron micrographs of the irradiated polymer surfaces showed an erosion morphology similar to that obtained on low earth orbit.

Krech, Robert H.↗
NASA NTRS1987

Spectroscopic investigations and measurements of certain ARC jet parameters

The NASA/JSC Atmospheric Reentry Materials and Structures Evaluation Facility (ARMSEF) was intensively and extensively involved in ground testing of spacecraft materials and thermal protection systems (TPS) in simulated reentry conditions. Ground experiments on surface catalytic efficiency of such TPS requires a knowledge of the flow system in the arc jet. In the work described, spectroscopic diagnostic techniques are used to determine the free stream constituents. Specifically, the emission spectrum of the free stream constituents was obtained and the species therein identified. A laser system was added on, which will give the added capacity of studying the arc jet flow using Laser Raman Spectroscopy (LRS). The LRS technique will complement information obtained from the emission spectra. A short list of further work that can be done in the area of spectroscopic investigations on the arc jet is mentioned.

Murty, Vedula S.↗
NASA NTRS2003

"Smoke": Characterization Of Smoke Particulate For Spacecraft Fire Detection

The "Smoke" experiment is a flight definition investigation that seeks to increase our understanding of spacecraft fire detection through measurements of particulate size distributions of preignition smokes from typical spacecraft materials. Owing to the catastrophic risk posed by even a very small fire in a spacecraft, the design goal for spacecraft fire detection is to detect the fire as quickly as possible, preferably in the preignition phase before a real flaming fire has developed. Consequently the target smoke for detection is typically not soot (typical of established hydrocarbon fires) but instead, pyrolysis products, and recondensed polymer particles. At the same time, false alarms are extremely costly as the crew and the ground team must respond quickly to every alarm. The U.S. Space Shuttle (STS: Space Transportation System) and the International Space Station (ISS) both use smoke detection as the primary means of fire detection. These two systems were designed in the absence of any data concerning low-gravity smoke particle (and background dust) size distributions. The STS system uses an ionization detector coupled with a sampling pump and the ISS system is a forward light scattering detector operating in the near IR. These two systems have significantly different sensitivities with the ionization detector being most sensitive (on a mass concentration basis) to smaller particulate and the light scattering detector being most sensitive to particulate that is larger than 1 micron. Since any smoke detection system has inherent size sensitivity characteristics, proper design of future smoke detection systems will require an understanding of the background and alarm particle size distributions that can be expected in a space environment.

Urban, David L.↗
NASA NTRS2017

Micrometeoroid and Orbital Debris Risk Assessment With Bumper 3

The Bumper 3 computer code is the primary tool used by NASA for micrometeoroid and orbital debris (MMOD) risk analysis. Bumper 3 (and its predecessors) have been used to analyze a variety of manned and unmanned spacecraft. The code uses NASA's latest micrometeoroid (MEM-R2) and orbital debris (ORDEM 3.0) environment definition models and is updated frequently with ballistic limit equations that describe the hypervelocity impact performance of spacecraft materials. The Bumper 3 program uses these inputs along with a finite element representation of spacecraft geometry to provide a deterministic calculation of the expected number of failures. The Bumper 3 software is configuration controlled by the NASA/JSC Hypervelocity Impact Technology (HVIT) Group. This paper will demonstrate MMOD risk assessment techniques with Bumper 3 used by NASA's HVIT Group. The Permanent Multipurpose Module (PMM) was added to the International Space Station in 2011. A Bumper 3 MMOD risk assessment of this module will show techniques used to create the input model and assign the property IDs. The methodology used to optimize the MMOD shielding for minimum mass while still meeting structural penetration requirements will also be demonstrated.

Hyde, J.↗
NASA NTRS2012

Space Weather Impacts on Spacecraft Design and Operations in Auroral Charging Environments

Spacecraft in low altitude, high inclination (including sun-synchronous) orbits are widely used for remote sensing of the Earth s land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems are episodically exposed to environments characterized by a high flux of energetic (approx.1 to 10 s kilovolt) electrons in regions of very low background plasma density which is similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. While it is well established that charging conditions in geostationary orbit are responsible for many anomalies and even spacecraft failures, to date there have been relatively few such reports due to charging in auroral environments. This presentation first reviews the physics of the space environment and its interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments and discuss how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Minow, Joseph I.↗
NASA NTRS2012

Extreme Spacecraft Charging in Polar Low Earth Orbit

Spacecraft in low altitude, high inclination (including sun-synchronous) orbits are widely used for remote sensing of the Earth's land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems episodically charge to frame potentials in the kilovolt range when exposed to space weather environments characterized by a high flux of energetic (~10 s kilovolt) electrons in regions of low background plasma density which is similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. We first review the physics of space environment interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments. Finally, a set of extreme DMSP charging events are described varying in maximum negative frame potential from ~0.6 kV to ~2 kV, focusing on the characteristics of the charging events that are of importance both to the space system designer and to spacecraft operators. The goal of the presentation is to bridge the gap between scientific studies of auroral charging and the need for engineering teams to understand how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Colson, Andrew D.↗
NASA NTRS2012

Extreme Spacecraft Charging in Polar Low Earth Orbit

Spacecraft in low altitude, high inclination (including sun -synchronous) orbits are widely used for remote sensing of the Earth fs land surface and oceans, monitoring weather and climate, communications, scientific studies of the upper atmosphere and ionosphere, and a variety of other scientific, commercial, and military applications. These systems episodically charge to frame potentials in the kilovolt range when exposed to space weather environments characterized by a high flux of energetic (approx.10 fs kilovolt) electrons in regions of low background plasma density. Auroral charging conditions are similar in some ways to the space weather conditions in geostationary orbit responsible for spacecraft charging to kilovolt levels. We first review the physics of space environment interactions with spacecraft materials that control auroral charging rates and the anticipated maximum potentials that should be observed on spacecraft surfaces during disturbed space weather conditions. We then describe how the theoretical values compare to the observational history of extreme charging in auroral environments. Finally, a set of extreme DMSP charging events are described varying in maximum negative frame potential from approx.0.6 kV to approx.2 kV, focusing on the characteristics of the charging events that are of importance both to the space system designer and to spacecraft operators. The goal of the presentation is to bridge the gap between scientific studies of auroral charging and the need for engineering teams to understand how space weather impacts both spacecraft design and operations for vehicles on orbital trajectories that traverse auroral charging environments.

Colson, Andrew D.↗