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All sourcesNASA NTRSDOE OSTI
NASA NTRS2002

Vacuum Ultraviolet Radiation and Atomic Oxygen Durability Evaluation of HST Bi-Stem Thermal Shield Materials

Bellows-type thermal shields were used on the bi-stems of replacement solar arrays installed on the Hubble Space Telescope (HST) during the first HST servicing mission (SMI) in December 1993. These thermal shields helped reduce the problem of thermal gradient- induced jitter observed with the original HST solar arrays during orbital thermal cycling and have been in use on HST for eight years. This paper describes ground testing of the candidate solar array bi-stem thermal shield materials including backside aluminized Teflon(R)FEP (fluorinated ethylene propylene) with and without atomic oxygen (AO) and ultraviolet radiation protective surface coatings for durability to AO and combined AO and vacuum ultraviolet (VOV) radiation. NASA Glenn Research Center (GRC) conducted VUV and AO exposures of samples of candidate thermal shield materials at HST operational temperatures and pre- and post-exposure analyses as part of an overall program coordinated by NASA Goddard Space Flight Center (GSFC) to determine the on-orbit durability of these materials. Coating adhesion problems were observed for samples having the AO- and combined AO/UV-protective coatings. Coating lamination occurred with rapid thermal cycling testing which simulated orbital thermal cycling. This lack of adhesion caused production of coating flakes from the material that would have posed a serious risk to HST optics if the coated materials were used for the bi-stem thermal shields. No serious degradation was observed for the uncoated aluminized Teflon(R) as evaluated by optical microscopy, although atomic force microscopy (AFM) microhardness testing revealed that an embrittled surface layer formed on the uncoated Teflon(R) surface due to vacuum ultraviolet radiation exposure. This embrittled layer was not completely removed by AO erosion, No cracks or particle flakes were produced for the embrittled uncoated material upon exposure to VUV and AO at operational temperatures to an equivalent exposure of approximately five years in the HST environment. Uncoated aluminized FEP Teflon(R) was determined to be the most appropriate thermal shield material and was used on the bi-stems of replacement solar arrays installed on HST during SMI in December 1993. The SMI -installed solar arrays air scheduled to be replaced during MST's fourth servicing mission (SM3B) in early 2002.

Dever, Joyce↗
NASA NTRS2007

Lunar In Situ Materials-Based Surface Structure Technology Development Efforts at NASA/MSFC

For long-duration missions on other planetary bodies, the use of in situ materials will become increasingly critical. As man's presence on these bodies expands, so must the structures to accommodate them, including habitats, laboratories, berms, radiation shielding for surface reactors, garages, solar storm shelters, greenhouses, etc. The use of in situ materials will significantly offset required launch upmass and volume issues. Under the auspices of the In Situ Fabrication & Repair (ISFR) Program at NASA/Marshall Space Flight Center (MSFC), the Surface Structures project has been developing materials and construction technologies to support development of these in situ structures. This paper will report on the development of several of these technologies at MSFC's Prototype Development Laboratory (PDL). These technologies include, but are not limited to, development of extruded concrete and inflatable concrete dome technologies based on waterless and water-based concretes, development of regolith-based blocks with potential radiation shielding binders including polyurethane and polyethylene, pressure regulation systems for inflatable structures, production of glass fibers and rebar derived from molten lunar regolith simulant, development of regolithbag structures, and others, including automation design issues. Results to date and lessons learned will be presented, along with recommendations for future activities.

Fiske, M. R.↗
NASA NTRS2007

Mars Radiation Risk Assessment and Shielding Design for Long-term Exposure to Ionizing Space Radiation

NASA is now focused on the agency's vision for space exploration encompassing a broad range of human and robotic missions including missions to Moon, Mars and beyond. As a result, there is a focus on long duration space missions. NASA is committed to the safety of the missions and the crew, and there is an overwhelming emphasis on the reliability issues for space missions and the habitat. The cost-effective design of the spacecraft demands a very stringent requirement on the optimization process. Exposure from the hazards of severe space radiation in deep space and/or long duration missions is a critical design constraint and a potential 'show stopper'. Thus, protection from the hazards of severe space radiation is of paramount importance to the agency's vision. It is envisioned to have long duration human presence on the Moon for deep space exploration. The exposures from ionizing radiation - galactic cosmic radiation and solar particle events - and optimized shield design for a swing-by and a long duration Mars mission have been investigated. It is found that the technology of today is inadequate for safe human missions to Mars, and revolutionary technologies need to be developed for long duration and/or deep space missions. The study will provide a guideline for radiation exposure and protection for long duration missions and career astronauts and their safety.

Tripathi, Ram K.↗
NASA NTRS1980

A solar encounter mission - Starprobe

Accurate measurements of the internal mass distribution and dynamics of the sun as well as in situ observations of its environs are proposed through a JPL-designed satellite, STARPROBE. Scheduled for launch in 1988 under NASA's five-year plan, the probe will conduct radiometric experiments to determine the solar quadrupole moment (J2), fields-and-particles experiments (evaluating such phenomena as heating, acceleration, and peak random velocity of solar wind), and optical experiments potentially superior to those of the Earth-orbiting Solar Optical Telescope (SOT). With flight times ranging from 2.7 to 5.9 years, present delivery options comprise planetary gravity-assist trajectories that will give the spacecraft a perihelion of four solar radii at a 90 deg-inclination. Owing to science- and mission requirements, the thermal shield of the craft should tolerate a solar flux 3000 times greater than that at earth, radioisotope thermoelectric generators supplying 700 watts of power, and a three-axis stabilization system with reaction wheels providing control torque. Implementation concepts now under study may change some of the details of the design.

Randolph, J. E.↗
NASA NTRS2012

Effect of Solar Exposure on the Atomic Oxygen Erosion of Hubble Space Telescope Aluminized-Teflon Thermal Shields

When exposed to low Earth orbital (LEO) environment, external spacecraft materials degrade due to radiation, thermal cycling, micrometeoroid and debris impacts, and atomic oxygen (AO) interaction. Collisions between AO and spacecraft can result in oxidation of external spacecraft surface materials, which can lead to erosion and severe structural and/or optical property deterioration. It is therefore essential to understand the AO erosion yield (Ey), the volume loss per incident oxygen atom (cu cm/atom), of polymers to assure durability of spacecraft materials. The objective of this study was to determine whether solar radiation exposure can increase the rate of AO erosion of polymers in LEO. The material studied was a section of aluminized-Teflon (DuPont) fluorinated ethylene propylene (Al-FEP) thermal shield exposed to space on the Hubble Space Telescope (HST) for 8.25 years. Retrieved samples were sectioned from the circular thermal shield and exposed to ground laboratory thermal energy AO. The results indicate that the average Ey of the solar facing HST Al-FEP was 1.9 10(exp -24)cu cm/atom, while the average Ey of the anti-solar HST Al-FEP was 1.5 10(exp -24)cu cm/atom. The Ey of the pristine samples was 1.6- 1.7 10(exp -24)cu cm/atom. These results indicate that solar exposure affects the post-flight erosion rate of FEP in a plasma asher. Therefore, it likely affects the erosion rate while in LEO.

Guo, Aobo↗
NASA NTRS2000

Radiation measurements in low Earth orbit: U.S. and Russian results

The radiation environment in low-Earth orbital flights is complex. It is strongly influenced by altitude, orbital inclination, time within a given solar cycle, flight duration, and shielding configuration. At any specified shielded location, both primary and secondary particles generated by nuclear interactions of primary particles with spacecraft structure are present. In addition, there are atmospheric secondary albedo protons and neutrons. No single detector can adequately measure this complex radiation field, and measurements of very high linear energy transfer target fragmentation products are particularly difficult. Crew radiation exposure have exclusively been measured using passive thermoluminescent detectors (TLDs). The cosmonaut exposures on the Mir station, uncorrected for the TLD inefficiency and neutron contribution, have varied from a low of 2.43 cGy to a high of 8.70 cGy. These correspond to dose rates of 144 microGy d(-1) to 468 microGy d(-1). These are consistent with rates observed by the D2 ion-chamber. Using the rates measured by the D1 chamber, dose rates under 4 cm of water vary from about 60 microGy d-1 to about 350 microGy d(-1). There is variation of about a factor of two between the dose rates at various locations in the same module. There is also a variation of dose rates of about a factor two between various modules. The highest astronaut dose for a Shuttle flight (STS-82) was 3.205 cGy with a dose rate of 3,221 microGy d(-1). Neutron contribution could be 36 +/- 15% of the astronaut charged particle dose equivalent. East-West asymmetry of dose rate is significant for spacecrafts that fly in an fixed altitude, such as the International Space Station.

NASA Center JSC↗
NASA NTRS2013

Radiation Shielding Optimization on Mars

Future space missions to Mars will require radiation shielding to be optimized for deep space transit and an extended stay on the surface. In deep space, increased shielding levels and material optimization will reduce the exposure from most solar particle events (SPE) but are less effective at shielding against galactic cosmic rays (GCR). On the surface, the shielding provided by the Martian atmosphere greatly reduces the exposure from most SPE, and long-term GCR exposure is a primary concern. Previous work has shown that in deep space, additional shielding of common materials such as aluminum or polyethylene does not significantly reduce the GCR exposure. In this work, it is shown that on the Martian surface, almost any amount of aluminum shielding increases exposure levels for humans. The increased exposure levels are attributed to neutron production in the shield and Martian regolith as well as the electromagnetic cascade induced in the Martian atmosphere. This result is significant for optimization of vehicle and shield designs intended for the surface of Mars.

Slaba, Tony C.↗
NASA NTRS1962

Analysis of Thermal-Protection Systems for Space-Vehicle Cryogenic-Propellant Tanks

Analytical techniques are presented that permit the calculation of heat-transfer rates with various thermal-protection systems for liquid-cryogenic-propellant tanks subjected to on-board, solar, and planetary heat fluxes. The thermal-protection systems considered include using closely spaced reflective surfaces (foils) and widely spaced reflective surfaces (shadow shields), insulation, arrangement of vehicle components, orientation with respect to radiant heating sources, and coatings for the control of solar absorptivity. The effectiveness of these thermal-protection systems in reducing propellant heating is shown both for ideal heat-transfer models and for a simplified hydrogen-oxygen terminal stage on a Mars mission. The proper orientation of a space-vehicle cryogenic tank with respect to the Sun is one of the more beneficial methods of reducing the heating effect of solar flux. Shadow shields can be extremely effective in reducing the propellant heating due to both solar and on-board fluxes. However, low-altitude planet orbits can result in high propellant heating rates due to planetary radiation reflected from the shields. For low-altitude orbits of more than a few days, foils appear to be desirable for all cryogenic-tank surfaces. Foils are also effective in reducing the on-board heating. A choice of shadow shields or foils cannot be made until a particular vehicle and a particular mission are chosen. The thermal conductivity of insulation materials would have to be lower by about two orders of magnitude with no increase in density before insulation could compete with reflective surfaces for use in long-duration thermal protection of cryogenic tanks in space. To demonstrate the application of the methods devised, thermal-protection systems are developed for a hydrogen-oxygen terminal stage for typical Mars missions.

SPACE VEHICLE↗
NASA NTRS1986

Summary of radiation dosimetry results on U.S. and Soviet manned spacecraft

Measurements of the radiation environment aboard U.S. and Soviet manned spacecraft are reviewed and summarized. Data obtained mostly from passive and some active radiation detectors now exist for the case of low-earth-orbit missions. Major uncertainties still exist for space exposure in high-altitude, high-inclination geostationary orbits, in connection with solar effects and that of shielding. Data from active detectors flown in Spacelabs 1 and 2 suggest that a variety of phenomena must be understood before the effects of long-term exposure at the Space Station type of orbit and shielding can be properly assessed.

Benton, E. V.↗
NASA NTRS1991

Cellular track model of biological damage to mammalian cell cultures from galactic cosmic rays

The assessment of biological damage from the galactic cosmic rays (GCR) is a current interest for exploratory class space missions where the highly ionizing, high-energy, high-charge ions (HZE) particles are the major concern. The relative biological effectiveness (RBE) values determined by ground-based experiments with HZE particles are well described by a parametric track theory of cell inactivation. Using the track model and a deterministic GCR transport code, the biological damage to mammalian cell cultures is considered for 1 year in free space at solar minimum for typical spacecraft shielding. Included are the effects of projectile and target fragmentation. The RBE values for the GCR spectrum which are fluence-dependent in the track model are found to be more severe than the quality factors identified by the International Commission on Radiological Protection publication 26 and seem to obey a simple scaling law with the duration period in free space.

Cucinotta, Francis A.↗
NASA NTRS1992

Environmental effects on an optical-UV-IR synthesis array

The Moon offers a stable platform with excellent seeing conditions for the Lunar Optical-UV-IR Synthesis Array (LOUISA). Some troublesome aspects of the lunar environment will need to be overcome to realize the full potential of the Moon as an observatory site. Mitigation of negative effects of vacuum, thermal radiation, dust, and micrometeorite impact is feasible with careful engineering and operational planning. Shields against impact, dust, and solar radiation need to be developed. Means of restoring degraded surfaces are probably essential for optical and thermal control surfaces deployed in long-lifetime lunar facilities. Precursor missions should be planned to validate and enhance the understanding of the lunar environment (e.g., dust behavior without and with human presence and to determine environmental effects on surfaces and components. Precursor missions should generate data useful in establishing keepout zones around observatory facilities while rocket launches and landings, mining, and vehicular traffic could be detrimental to observatory operation.

Johnson, Stewart W.↗
NASA NTRS1992

Environmental effects on lunar astronomical observatories

The Moon offers a stable platform with excellent seeing conditions for astronomical observations. Some troublesome aspects of the lunar environment will need to be overcome to realize the full potential of the Moon as an observatory site. Mitigation of negative effects of vacuum, thermal radiation, dust, and micrometeorite impact is feasible with careful engineering and operational planning. Shields against impact, dust, and solar radiation need to be developed. Means of restoring degraded surfaces are probably essential for optical and thermal control surfaces deployed in long-lifetime lunar facilities. Precursor missions should be planned to validate and enhance the understanding of the lunar environment (e.g., dust behavior without and with human presence) and to determine environmental effects on surfaces and components. Precursor missions should generate data useful in establishing keepout zones around observatory facilities where rocket launches and landings, mining, and vehicular traffic could be detrimental to observatory operation.

Johnson, Stewart W.↗
NASA NTRS2003

Videogrammetry Using Projected Circular Targets: Proof-of-Concept Test

Videogrammetry is the science of calculating 3D object coordinates as a function of time from image sequences. It expands the method of photogrammetry to multiple time steps enabling the object to be characterized dynamically. Photogrammetry achieves the greatest accuracy with high contrast, solid-colored, circular targets. The high contrast is most often effected using retro-reflective targets attached to the measurement article. Knowledge of the location of each target allows those points to be tracked in a sequence of images, thus yielding dynamic characterization of the overall object. For ultra-lightweight and inflatable gossamer structures (e.g. solar sails, inflatable antennae, sun shields, etc.) where it may be desirable to avoid physically attaching retro-targets, a high-density grid of projected circular targets - called dot projection - is a viable alternative. Over time the object changes shape or position independently of the dots. Dynamic behavior, such as deployment or vibration, can be characterized by tracking the overall 3D shape of the object instead of tracking specific object points. To develop this method, an oscillating rigid object was measured using both retroreflective targets and dot projection. This paper details these tests, compares the results, and discusses the overall accuracy of dot projection videogrammetry.

Pappa, Richard S.↗
NASA NTRS2003

Videogrammetry Using Projected Circular Targets: Proof-of-Concept Test

Videogrammetry is the science of calculating 3D object coordinates as a function of time from image sequences. It expands the method of photogrammetry to multiple time steps enabling the object to be characterized dynamically. Photogrammetry achieves the greatest accuracy with high contrast, solid-colored circular targets. The high contrast is most often effected using retro-reflective targets attached to the measurement article. Knowledge of the location of each target allows those points to be tracked in a sequence of images, thus yielding dynamic characterization of the overall object. For ultra-lightweight and inflatable gossamer structures (e.g. solar sails, inflatable antennae, sun shields, etc.) where it may be desirable to avoid physically attaching retro-targets, a high-density grid of projected circular targets - called dot projection - is a viable alternative. Over time the object changes shape or position independently of the dots. Dynamic behavior, such as deployment or vibration, can be characterized by tracking the overall 3D shape of the object instead of tracking specific object points. To develop this method, an oscillating rigid object was measured using both retro- reflective targets and dot projection. This paper details these tests, compares the results, and discusses the overall accuracy of dot projection videogrammetry.

Black, Jonathan T.↗
NASA NTRS2005

Single Particle Damage Events in Candidate Star Camera Sensors

This viewgraph presentation presents information on the following topics: 1) Solar and trapped protons and shielding; 2) Proton interactions in Si; 3) Displacement damage effects in detectors; 4) Hot pixel mechanisms, introduction rates, and annealing.

Marshall, Paul↗
NASA NTRS2017

Design and Technical Study of Neutrino Detector Spacecraft

A neutrino detector is proposed to be developed for use on a space probe in close orbit of the Sun. The detector will also be protected from radiation by a tungsten shield Sun shade, active veto array and passive cosmic shielding. With the intensity of solar neutrinos substantially greater in a close solar orbit than on the Earth only a small 250 kg detector is needed. It is expected that this detector and space probe studying the core of the Sun, its nuclear furnace and particle physics basic properties will bring new knowledge beyond what is currently possible for Earth bound solar neutrino detectors.

Design and Technical Study↗
NASA NTRS2015

Performance and Facility Background Pressure Characterization Tests of NASAs 12.5-kW Hall Effect Rocket with Magnetic Shielding Thruster

NASA's Space Technology Mission Directorate (STMD) Solar Electric Propulsion Technology Demonstration Mission (SEP/TDM) project is funding the development of a 12.5-kW Hall thruster system to support future NASA missions. The thruster designated Hall Effect Rocket with Magnetic Shielding (HERMeS) is a 12.5-kW Hall thruster with magnetic shielding incorporating a centrally mounted cathode. HERMeS was designed and modeled by a NASA GRC and JPL team and was fabricated and tested in vacuum facility 5 (VF5) at NASA GRC. Tests at NASA GRC were performed with the Technology Development Unit 1 (TDU1) thruster. TDU1's magnetic shielding topology was confirmed by measurement of anode potential and low electron temperature along the discharge chamber walls. Thermal characterization tests indicated that during full power thruster operation at peak magnetic field strength, the various thruster component temperatures were below prescribed maximum allowable limits. Performance characterization tests demonstrated the thruster's wide throttling range and found that the thruster can achieve a peak thruster efficiency of 63% at 12.5 kW 500 V and can attain a specific impulse of 3,000 s at 12.5 kW and a discharge voltage of 800 V. Facility background pressure variation tests revealed that the performance, operational characteristics, and magnetic shielding effectiveness of the TDU1 design were mostly insensitive to increases in background pressure.

Performance↗