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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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At least 109 records · Page 6

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NASA NTRS1973

Spacecraft microbial burden reduction due to atmospheric entry heating - Jupiter.

An effort has been underway at Jet Propulsion Laboratory to develop the tools necessary to predict temperature histories for a typical spacecraft during inadvertent entry. In order that the results have general applicability, parametric analyses are performed. The thermal response of the spacecraft components and debris resulting from disintegration is determined. The temperature histories of small particles and composite materials, such as thermal blankets and an antenna, are given special attention. Guidelines are given to indicate the types of components and debris most likely to contain viable organisms, which could contaminate the lower layers of the Jovian atmosphere.

Gonzalez, C. C.↗

NASA NTRS1981

Preflow stresses in Martian rampart ejecta blankets - A means of estimating the water content

Measurements of extents of rampart ejecta deposits as a function of the size of the parent craters support models which, for craters larger than about 6 km diameter, constrain ejecta blankets to all have a similar maximum thickness regardless of the crater size. These volatile-rich ejecta blankets may have failed under their own weights, then flowed radially outward. Assuming this to be so, some of the physicomechanical properties of the ejecta deposits at the time of their emplacement can then be determined. Finite-element studies of the stress magnitudes, distributions, and directions in hypothetical Martian rampart ejecta blankets reveal that the material most likely failed when the shear stresses were less than 500 kPa and the angle of internal friction was between 26 and 36 deg. These figures imply that the ejecta has a water content between 16 and 72%. Whether the upper limit or the lower limit is more appropriate depends on the mode of failure which one presumes: namely, viscous flow of plastic deformation.

Woronow, A.↗

NASA NTRS1984

Lunar magma ocean and its implication for origin of the Moon

A plausible accretional model of the Moon using as a constraint the formation of a magma ocean is discussed. Recently, it was shown that the surface of a planet growing by planetestimal impacts was heated over the melting temperature of surface materials due to the blanketing effect of an impact induced atmosphere. Using the same calculational scheme for the Earth, the early thermal history of the Moon growing by planetestimal impacts can be calculated for various accretional models. It is shown that a magma ocean covering the entire surface was formed in both models. Most important parameters related to surface temperature are safronov number and accretion time. Results show that a very small safronov number is needed for formation of the magma ocean. Safronov number is usually larger than 1 for accretion of planetestimals in heliocentric orbit. However, safronov number decreases when the Moon's growth is dominated by the proximity of the Earth. According to Harris, safronov number falls in the range of 0.02-0.1 for the binary accretion case. Therefore, it is suggested that the Moon was formed by accretion of planetestimals in geocentric orbits.

Matsui, T.↗

NASA NTRS1986

Retractable Sun Shade

Window-shade type spring roller contains blanket, taken up by rotating cylindrical frame and held by frame over area to be shaded. Blanket made of tough, opaque polyimide material. Readily unfurled by mechanism to protect space it encloses from Sun. Blanket forms arched canopy over space and allows full access to it from below. When shading not needed, retracted mechanism stores blanket compactly. Developed for protecting sensitive Space Shuttle payloads from direct sunlight while cargo-bay doors open. Adapted to shading of greenhouses, swimming pools, and boats.

Frank, A.↗

NASA NTRS1990

Spacecraft studies of Phobos and Mars

Utilizing the Termoskan data set of the Phobos '88 mission we have recognized a new feature on Mars: Ejecta blanket Distinct In the THermal infrared (EDITH). Virtually all of the more than one hundred of these features discovered in the Termoskan data are located on the plains near Valles Manneris. EDITH's have a startlingly clear dependence upon terrains of Hesperian age, implying a spatial or temporal dependence on Hesperian terrains. Almost no thermally distinct ejecta blankets are associated with any of the thousands of craters within the data set that occur on the older Noachian units. EDITH's also do not appear on the portions of the younger Tharsis Amazonian units seen in the data. The Hesperian terrain dependence cannot be explained by either atmospheric or impactor variations; Noachian and Hesperian terrains must have experienced identical atmospheric and impactor conditions during Hesperian times. Thermally distinct eject a blankets therefore reflect target material differences and/or secondary modification processes. A further discussion of EIDTH's is presented.

Murray, Bruce C.↗

NASA NTRS1999

Compaction as the Origin of the Unusual Craters on the Asteroid Mathilde

Asteroid Mathilde has been pummeled by at least five giant impacts (Figure 1). Previous experience with cratering suggests Mathilde's giant craters should each be surrounded by kilometer-deep blankets of ejecta, i.e. material excavated during the impact events. Curiously, there appears to be very little ejecta around Mathilde's craters; they show no evidence of filling by ejecta from adjacent large craters. A previous explanation for the missing ejecta, based on computer simulations, is that Mathilde's unusually high porosity (50 +/- 20%) confines the deposited impact kinetic energy to a localized volume, and produces excavation velocities so high (greater than approximately 20m/s) that nearly all ejecta escape Mathilde's gravitational field. Here we report on laboratory experiments in a highly porous material that give a different explanation. The crater is formed primarily by compaction, not excavation. The small amount of material that is lofted has velocities and ranges so small that nearly all of it is re-deposited within the crater bowl, thereby sparing neighboring craters from ejecta in-filling. This peculiar style of cratering implies that highly porous asteroids are minor contributors of meteorites, because essentially no ejecta escape these asteroids.

Housen, Kevin R.↗

NASA NTRS2001

Further Investigations of the Passive Optical Sample Assembly (POSA) - I Flight Experiment

The Passive Optical Sample Assembly-I (POSA-I), part of the Mir Environmental Effects Payload (MEEP), was designed to study the combined effects of contamination, atomic oxygen, ultraviolet radiation, vacuum, then-nal cycling, and other constituents of the space environment on spacecraft materials. The MEEP program is a Phase I International Space Station Risk Mitigation Experiment. Candidate materials for the International Space Station (ISS) were exposed in a specially designed "suitcase" carrier, with identical specimens facing either Mir or space. The payload was attached by EVA to the exterior of the Mir docking module during the STS-76 mission (f'ig. 1). It was removed during the STS-86 mission after an 18-month exposure. During the mission, it received approximately 7 x 1019 atoMS/CM2 atomic oxygen, as calculated by polymer mass loss, and 413 ESH of solar ultraviolet radiation on the Mir-facing side. The side facing away from Mir received significant contaminant deposition, so atomic oxygen fluence has not been reliably determined. The side facing away from Mir received 571 ESH of solar UV. Contamination was observed on both the Mir-facing and space-facing sides of the POSA-I experiment , with a greater amount of deposition on the space facing side than the Mir side. The contamination has been determined to be outgassed silicone photofixed by ultraviolet radiation and converted to silicate by atomic oxygen interaction. Electron spectroscopy for chemical analysis (ESCA) with depth profiling indicated the presence of 26 - 31 nm silicate on the Mir-facing side and 500 - 1000 nm silicate on the space-facing side. The depth profiling also showed that the contaminant layer was uniform, with a small amount of carbon present on the surface and trace amounts of nitrogen, phosphorus, sulfur, and tin. The surface carbon layer is likely due to post-flight exposure in the laboratory and is similar to carbonaceous deposits on control samples. EDAX and FTIR analysis concurred with ESCA for the presence of silicon, oxygen, and carbon. Nearly 400 samples were exposed on POSA-I, which included materials such as thermal control coatings polymeric films, optical materials, and multi-layer insulation blankets. A previous paper discussed the effects of the space environment exposure and contaminant deposition on candidate materials for ISS, including Z93P inorganic thermal control coating, various anodizes, and multi-layer insulation blankets. This paper details the investigation of environmental effects on the remainder of POSA-I samples, particularly the innovative conductive thermal control coatings developed by AZ Technology of Huntsville, AL and HT Research Institute of Chicago, IL. The silicone/silicate contamination had a significant impact on the solar absorptance of white inorganic thermal control coatings on the space-facing side of POSA-I. The effect of contamination on electrical conductivity is discussed. Samples of conductive anodized aluminum developed by Boundary Technologies of Buffalo Grove, IL were also exposed on POSA-I. The effects of the space environment and contaminant deposition on the optical and electrical properties of the conductive anodized aluminum are discussed.

Finckenor, Miria M.↗

NASA NTRS2007

Thermal Performance Testing of Cryogenic Insulation Systems

Efficient methods for characterizing thermal performance of materials under cryogenic and vacuum conditions have been developed. These methods provide thermal conductivity data on materials under actual-use conditions and are complementary to established methods. The actual-use environment of full temperature difference in combination with vacuum-pressure is essential for understanding insulation system performance. Test articles include solids, foams, powders, layered blankets, composite panels, and other materials. Test methodology and apparatus design for several insulation test cryostats are discussed. The measurement principle is liquid nitrogen boil-off calorimetry. Heat flux capability ranges from approximately 0.5 to 500 watts per square meter; corresponding apparent thermal conductivity values range from below 0.01 up to about 60 mW/m- K. Example data for different insulation materials are also presented. Upon further standardization work, these patented insulation test cryostats can be available to industry for a wide range of practical applications.

Fesmire, James E.↗

NASA NTRS1993

Beta cloth durability assessment for Space Station Freedom (SSF) Multi-Layer Insulation (MLI) blanket covers

MLI blankets for the Space Station Freedom (SSF) must comply with general program requirements and recommendations for long life and durability in the low-Earth orbit (LEO) environment. Atomic oxygen and solar ultraviolet/vacuum ultraviolet are the most important factors in the SSF natural environment which affect materials life. Two types of Beta cloth (Teflon coated woven glass fabric), which had been proposed as MLI blanket covers, were tested for long-term durability in the LEO environment. General resistance to atomic oxygen attack and permeation were evaluated in the high velocity atomic oxygen beam system at Los Alamos National Laboratories. Long-term exposure to the LEO environment was simulated in the laboratory using a radio frequency oxygen plasma asher. The plasma asher treated Beta cloth specimens were tested for thermo-optical properties and mechanical durability. Space exposure data from the Long Duration Exposure Facility and the Intelsat Solar Array Coupon were also used in the durability assessment. Beta cloth fabricated to Rockwell specification MBO 135-027 (Chemglas 250) was shown to have acceptable durability for general use as an MLI blanket cover material in the LEO environment while Sheldahl G414500 should be used only in locations which are protected from direct Ram atomic oxygen.

Koontz, Steven L.↗

NASA NTRS1991

The status of lightweight photovoltaic space array technology based on amorphous silicon solar cells

Ultralight, flexible photovoltaic (PV) array of amorphous silicon (a-Si) was identified as a potential low cost power source for small satellites. A survey was conducted of the status of the a-Si PV array technology with respect to present and future performance, availability, cost, and risks. For existing, experimental array blankets made of commercial cell material, utilizing metal foil substrates, the Beginning of Life (BOL) performance at Air Mass Zero (AM0) and 35 C includes total power up to 200 W, power per area of 64 W/sq m and power per weight of 258 W/kg. Doubling of power per weight occurs when polyimide substrates are used. Estimated End of Life (EOL) power output after 10 years in a nominal low earth orbit would be 80 pct. of BOL, the degradation being due to largely light induced effects (-10 to -15 pct.) and in part (-5 pct.) to space radiation. Predictions for the year 1995 for flexible PV arrays, made on the basis of published results for rigid a-Si modules, indicate EOL power output per area and per weight of 105 W/sq m and 400 W/kg, respectively, while predictions for the late 1990s based on existing U.S. national PV program goals indicate EOL values of 157 W/sq m and 600 W/kg. Cost estimates by vendors for 200 W ultralight arrays in volume of over 1000 units range from $100/watt to $125/watt. Identified risks include the lack of flexible, space compatible encapsulant, the lack of space qualification effort, recent partial or full acquisitions of US manufacturers of a-Si cells by foreign firms, and the absence of a national commitment for a long range development program toward developing of this important power source for space.

Hanak, Joseph J.↗

NASA NTRS1991

Performance analysis of advanced spacecraft TPS

Spacecraft entering a planetary atmosphere require a very sophisticated thermal protection system. The materials used must be tailored to each specific vehicle based on its planned mission profiles. Starting with the Space Shuttle, many types of ceramic insulation with various combinations of thermal properties have been developed by others. The development of two new materials is described: A Composite Flexible Blanket Insulation which has a significantly lower effective thermal conductivity than other ceramic blankets; and a Silicon Matrix Composite which has applications at high temperature locations such as wing leading edges. Also, a systematic study is described that considers the application of these materials for a proposed Personnel Launch System. The study shows how most of these available ceramic materials would perform during atmospheric entry of this vehicle. Other specific applications of these thermal protection materials are discussed.

Pitts, William C.↗

NASA NTRS1994

Atomic oxygen durability evaluation of the flexible batten for the photovoltaic array mast on Space Station

A test program was conducted at the National Aeronautics and Space Administration's Lewis Research Center (LeRC) to evaluate the long term low Earth orbital (LEO) atomic oxygen (AO) durability of a flexible (fiberglass-epoxy composite) batten. The flexible batten is a component used to provide structural rigidity in the photovoltaic array mast on Space Station. The mast is used to support and articulate the photovoltaic array, therefore, the flexible batten must be preloaded for the 15 year lifetime of an array blanket. Development hardware and composite materials were evaluated in ground testing facilities for AO durability and dynamic retraction-deployment cyclic loading representative of expected full life in-space application. The CV1144 silicone (AO protective) coating was determined to provide adequate protection against AO degradation of the composite material and provided fiber containment, thus the structural integrity of the flexible batten was maintained. Both silicone coated and uncoated flexible battens maintained load carrying capabilities. Results of the testing did indicate that the CV1144 silicone protective coating was oxidized by AO reactions to form a brittle glassy (SiO2) skin that formed cracking patterns on all sides of the coated samples. The cracking was observed in samples that were mechanically stressed as well as samples in non-stressed conditions. The oxidized silicon was observed to randomly spall in small localized areas, on the flexible battens that underwent retraction-deployment cycling. Some darkening of the silicon, attributed to vacuum ultraviolet (VUV) radiation, was observed.

Stidham, Curtis R.↗

NASA NTRS1997

STS-82 Flight Day 07 Highlights

The seventh day of the STS-82 mission begins with the crew, Commander Kenneth D. Bowersox, Pilot Scott J. Horowitz, Payload Commander Mark C. Lee, and Mission Specialists Gregory J. Harbaugh, Steven L. Smith, Joseph R. Tanner, and Steven A. Hawley performing their third spacewalk of the mission by emerging from Discovery's airlock. Their first task is the replacement of a Solar Array Drive Electronics package which is used to control the positioning of Hubble's solar arrays. Harbaugh and Tanner next venture to the top of the telescope where they replaced covers over Hubble's magnetometers, which are used to sense the telescope's position in relation to the Earth through data acquired from the Earth's magnetic field. The spacewalking astronauts then place thermal blankets of multi-layer material over two areas of degraded insulation around the light shield portion of the telescope just below the top of the astronomical observatory.

Source record↗

NASA NTRS2004

Hot Structure Control Surface Progress for X-37 Technology Development Program

The NASA Marshall Space Flight Center (MSFC) has been leading the development of technologies that will enable the development, fabrication, and flight of the automated X-37 Orbital Vehicle (OV). With the Administration s recent announcement of the Vision for Space Exploration, NASA placed the X-37 OV design on hold while developing detailed requirements for a Crew Exploration Vehicle, but has continued funding the development of high-risk, critical technologies for potential future space exploration vehicle applications. Hot Structure Control Surfaces (HSCS) technology development is one of the high-priority areas being funded at this time. The goal of HSCS research is to mitigate risk by qualifying the lightest possible components that meet the stringent X-37 OV weight and performance requirements, including Shuttle-type reen- try environments with peak temperatures of 2800 OF. The small size of the X-37 OV (25.7-feet long and 14.9-foot wingspan) drives the need for advanced HSCS because the vehicle's two primary aerodynamic surfaces, the flaperons and ruddervators, have thicknesses ranging from approximately 5 in. down to 1 in. Traditional metallic or polymer-matrix composites covered with tile or blanket thermal protection system (TPS) materials cannot be used as there is insufficient volume to fabricate such multi-component structures. Therefore, carbon-carbon (C-C) and carbodsilicon-carbide (C-SiC) composite HSCS structures are being developed in parallel by two teams supporting the X-37 prime contractor (The Boeing Company). The Science Applications International Coy. (SAIC) and Carbon-Carbon Advanced Technologies, Inc. (C-CAT) team is developing the C-C HSCS, while the General Electric Energy Power Systems Composites (GE-PSC) and Materials Research and Design (MRD) team is developing the C-SiC HSCS. These two teams were selected to reduce the high level of risk associated with developing advanced control surface components. They have continued HSCS development work as part of the X-37 critical technology development contract. The SAIC/C-CAT team is using Advanced Carbon-Carbon (ACC) because its fabrication is very similar to the process used for Space Shuttle Reinforced Carbon-Carbon fabrication, including the Sic-based pack cementation conversion coating systems using with both materials. ACC was selected over RCC because it has much higher tension and compressions strengths, and because T-300 fiber is readily available, whereas RCC rayon fiber is no longer manufactured. The GE-PSC/MRD team is using a T-300 fiber-reinforced Sic matrix composite material densified by chemical vapor infiltration. The C-Sic material has an Sic-based environmental barrier coating. Major accomplishments have been made over the past year by both HSCS teams. C-C and C- SiC flaperon subcomponents, which are truncated full-scale versions of flight hardware, have been fabricated and are undergoing testing at the NASA Dryden Flight Research Center, NASA Langley Research Center, and U.S. Air Force Research Laboratory. By the end of 2004, ruddervator subcomponents also will be delivered and tested. As NASA moves forward in realizing the Vision for Space Exploration, it will continue to invest in advanced research and development aimed at making new generations of spacecraft safer, more reliable, and more affordable. The X-37 HSCS effort ultimately will benefit the Agency's vision and mission.

Valentine, P. G.↗

NASA NTRS2002

Multilayer composite material and method for evaporative cooling

A multilayer composite material and method for evaporative cooling of a person employs an evaporative cooling liquid that changes phase from a liquid to a gaseous state to absorb thermal energy. The evaporative cooling liquid is absorbed into a superabsorbent material enclosed within the multilayer composite material. The multilayer composite material has a high percentage of the evaporative cooling liquid in the matrix. The cooling effect can be sustained for an extended period of time because of the high percentage of phase change liquid that can be absorbed into the superabsorbent. Such a composite can be used for cooling febrile patients by evaporative cooling as the evaporative cooling liquid in the matrix changes from a liquid to a gaseous state to absorb thermal energy. The composite can be made with a perforated barrier material around the outside to regulate the evaporation rate of the phase change liquid. Alternatively, the composite can be made with an imperveous barrier material or semipermeable membrane on one side to prevent the liquid from contacting the person's skin. The evaporative cooling liquid in the matrix can be recharged by soaking the material in the liquid. The multilayer composite material can be fashioned into blankets, garments and other articles.

Buckley, Theresa M.↗

NASA NTRS1993

A photovoltaic catenary-tent array for the Martian surface

To provide electrical power during an exploration mission to Mars, a deployable tent-shaped structure with a flexible photovoltaic (PV) blanket is proposed. The array is designed with a self-deploying mechanism utilizing pressurized gas expansion. The structural design for the array uses a combination of cables, beams, and columns to support and deploy the PV blanket. Under the force of gravity, a cable carrying a uniform load will take the shape of a catenary curve. A catenary-tent collector is self shading which must be taken into account in the solar radiation calculation. The shape and the area of the shadow on the array was calculated and used in the determination of the global irradiance on the array. The PV blanket shape and structure dimensions were optimized to achieve a configuration which maximizes the specific power (W/kg). The optimization was performed for three types of PV blankets (silicon, gallium arsenide over germanium, and amorphous silicon) and two types of structural materials (carbon composite and arimid fiber composite). The results show that the catenary shape of the PV blanket corresponding to zero end angle at the base with respect to the horizontal results in the highest specific power. The tent angle is determined by optimizing the specific mass and the output power for maximum specific power. The combination of carbon fiber structural material and amorphous silicon blanket produces the highest specific power.

Colozza, Anthony J.↗

NASA NTRS1995

Automated Laser Cutting In Three Dimensions

Computer-controlled machine-tool system uses laser beam assisted by directed flow of air to cut refractory materials into complex three-dimensional shapes. Velocity, position, and angle of cut varied. In original application, materials in question were thermally insulating thick blankets and tiles used on space shuttle. System shapes tile to concave or convex contours and cuts beveled edges on blanket, without cutting through outer layer of quartz fabric part of blanket. For safety, system entirely enclosed to prevent escape of laser energy. No dust generated during cutting operation - all material vaporized; larger solid chips dislodged from workpiece easily removed later.

Bird, Lisa T.↗

NASA NTRS1985

Perforating Thin Metal Sheets

Sheets only few mils thick bonded together, punched, then debonded. Three-step process yields perforated sheets of metal. (1): Individual sheets bonded together to form laminate. (2): laminate perforated in desired geometric pattern. (3): After baking, laminate separates into individual sheets. Developed for fabricating conductive layer on blankets that collect and remove ions; however, perforated foils have other applications - as conductive surfaces on insulating materials; stiffeners and conductors in plastic laminates; reflectors in antenna dishes; supports for thermal blankets; lightweight grille cover materials; and material for mockup of components.

Davidson, M. E.↗