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NASA NTRS1999

Seal Materials Compatible with the Electroplating Solvent Used in Constellation-X Mirrors

The existing gasket seals used in electroplating of the Constellation-X mirrors are difficult to assemble, and the current seal material is hydrophobic and too thick. The combination of the above problems result in: 1) non-uniform plating; 2) defect sites such as pits on the mirror edges; 3) "bear claws" on the edges of the mandrels and mirrors causing difficulties in shell-mirror separations; and 4) leakage of the plating solution past the seals into the mandrel causing chemical etching of the mandrel interior. This paper reports the results of this summer study in searching for alternate seal materials chemically compatible with the electroplating solvent. Fifteen common elastomeric rubber seal materials made-by Parker Seals were investigated including butyl, ethylene propylene, fluorosilicone, nitrile, Viton fluorocarbon, and silicone. Test results showed that Viton fluorocarbon compounds as a group were superior to the other tested compounds for chemical compatibility with the plating bath.

Pei, Xiong-Skiba↗

NASA NTRS1999

Design and Development of Lightweight Composite Tanks for the Mars Ascent Propulsion Technology

The investigation presented here focuses on the design and development of lightweight composite tanks for the Mars ascent propulsion technology. The proposed tanks are fabricated using the filament winding technique. The tanks will be used in the experimental permeability characterization of composite pressure vessels pressurized using cryogenic and kerosene fluids. We considered the geometry and composite material tailorability in the preliminary design formulation to obtain an isotensoid tank. The design formulation is based on membrane shell analysis. The tanks also include circular openings at the apex of the end caps for the installation of polar bosses. The development of a polar boss system was also investigated, and led to an innovative polar boss system that applies a uniform pressure on the o-ring gaskets. The permeability of these tanks was also considered and recommendations for improvement are presented.

Estrada, Hector↗

NASA NTRS2001

Atomic Oxygen Interactions With Silicone Contamination on Spacecraft in Low Earth Orbit Studied

Silicones have been widely used on spacecraft as potting compounds, adhesives, seals, gaskets, hydrophobic surfaces, and atomic oxygen protective coatings. Contamination of optical and thermal control surfaces on spacecraft in low Earth orbit (LEO) has been an ever-present problem as a result of the interaction of atomic oxygen with volatile species from silicones and hydrocarbons onboard spacecraft. These interactions can deposit a contaminant that is a risk to spacecraft performance because it can form an optically absorbing film on the surfaces of Sun sensors, star trackers, or optical components or can increase the solar absorptance of thermal control surfaces. The transmittance, absorptance, and reflectance of such contaminant films seem to vary widely from very transparent SiOx films to much more absorbing SiOx-based films that contain hydrocarbons. At the NASA Glenn Research Center, silicone contamination that was oxidized by atomic oxygen has been examined from LEO spacecraft (including the Long Duration Exposure Facility and the Mir space station solar arrays) and from ground laboratory LEO simulations. The findings resulted in the development of predictive models that may help explain the underlying issues and effects. Atomic oxygen interactions with silicone volatiles and mixtures of silicone and hydrocarbon volatiles produce glassy SiOx-based contaminant coatings. The addition of hydrocarbon volatiles in the presence of silicone volatiles appears to cause much more absorbing (and consequently less transmitting) contaminant films than when no hydrocarbon volatiles are present. On the basis of the LDEF and Mir results, conditions of high atomic oxygen flux relative to low contaminant flux appear to result in more transparent contaminant films than do conditions of low atomic oxygen flux with high contaminant flux. Modeling predictions indicate that the deposition of contaminant films early in a LEO flight should depend much more on atomic oxygen flux than it does later in a mission.

Banks, Bruce A.↗

NASA NTRS2006

Compression Testing of Alumina Fiber Insulation

A series of tests were conducted to measure the response of alumina fiber insulation to compression loading. The alumina fiber insulation is a candidate gasket material for the Space Shuttle Government Furnished Equipment (GFE) Tile Overlay Repair. Tests were conducted at room temperature and 2300 F. The alumina fiber insulation is a fibrous insulation blanket which was supplied to Langley in two forms, a nominal 3 lb/ft3 version and a nominal 9 lb/ft3 version. The 3 lb/ft3 material was tested as sheets 0.15 and 0.25 inches thick and the 9 lb/ft3 material in sheets 1 inch thick. The material showed very non-linear compression behavior with the compressive resistance of the material increasing as the material was compressed. The 3 lb/ft3 0.15-inch thick material required 4.1 psi to reach the nominal installation thickness of 0.045 inches and retain a load of 2.1 lbs during unloading. Testing at 2300 F resulted in a stiffer more board-like material. The 3 lb/ft3 0.15-inch thick material retained 1 psi of compressive resistance after a 10 minute hold at 2300 F and 0.045 inches thickness.

Vaughn, Wallace L.↗

NASA NTRS2008

Final Report on Portable Laser Coating Removal Systems Field Demonstrations and Testing

Processes currently used throughout the National Aeronautics and Space Administration (NASA) to remove corrosion and coatings from structures, ground service equipment and small components results in waste streams consisting of toxic chemicals, spent media blast materials, and waste water. When chemicals are used in these processes they are typically high in volatile organic compounds (VOC) and are considered hazardous air pollutants (HAP). When blast media is used, the volume of hazardous waste generated is increased significantly. Many of the coatings historically used within NASA contain toxic metals such as hexavalent chromium, and lead. These materials are highly regulated and restrictions on worker exposure continue to increase. Most recently the EPA reduced the permissible exposure limit (PEL) for hexavalent chromium. The new standard lowers OSHA's PEL for hexavalent chromium from 52 to 5 micrograms of Cr(V1) per cubic meter of air as an 8-hour time-weighted average. Hexavalent chromium is found in the pretreatment and primer coatings used within the Shuttle Program. In response to the need to continue to protect assets within the agency and the growing concern over these new regulations, NASA is researching different ways to continue the required maintenance of both facility and flight equipment in a safe, efficient and environmentally preferable manner. The use of laser energy to remove prepare surfaces for a variety of processes, such as corrosion and coating removal, weld preparation and non destructive evaluation is a relatively new technology that has shown itself to be environmentally preferable and in many cases less labor intensive than currently used removal methods. The development of a Portable Laser Coating Removal System (PLCRS) started as the goal of a Joint Group on Pollution Prevention (JG-PP) project, led by the Air Force, where several types of lasers in several configurations were thoroughly evaluated. Following this project, NASA decided to evaluate the best performers on processes and coatings specific to the agency. Laser systems used during this project were all of a similar design, most of which had integrated vacuum systems in order to collect materials removed from substrate surfaces during operation. Due to the fact that the technology lends itself to a bide variety of processes, several site demonstrations were organized in order to allow for greater evaluation of the laser systems across NASA. The project consisted of an introductory demonstration and a more in-depth evaluation at Wright-Patterson Air Force Base. Additionally, field demonstrations occurred at Glenn Research Center and Kennedy Space Center. During these demonstrations several NASA specific applications were evaluated, including the removal of coatings within Orbiter tile cavities and Teflon from Space Shuttle Main Engine gaskets, removal of heavy grease from Solid Rocket Booster components and the removal of coatings on weld lines for Shuttle and general ground service equipment for non destructive evaluation (NDE). In addition, several general industry applications such as corrosion removal, structural coating removal, weld-line preparation and surface cleaning were evaluated. This included removal of coatings and corrosion from surfaces containing lead-based coatings and applications similar to launch-structure maintenance and Crawler maintenance. During the project lifecycle, an attempt was made to answer process specific concerns and questions as they arose. Some of these initially unexpected questions concerned the effects lasers might have on substrates used on flight equipment including strength, surface re-melting, substrate temperature and corrosion resistance effects. Additionally a concern was PPE required for operating such a system including eye, breathing and hearing protection. Most of these questions although not initially planned, were fully explored as a part of this project. Generally the results from tesng were very positive. Corrosion was effectively removed from steel, but less successfully from aluminum alloys. Coatings were able to be removed, with varying results, generally dark, matte and thin coatings were easier to remove. Steel and aluminum panels were able to be cleaned for welding, with no known deleterious effects and weld-lines were able to have coatings removed in critical areas for NDE while saving time as compared to other methods.

Rothgeb, Matthew J.↗

NASA NTRS1954

Preliminary Altitude Operational Characteristics of a J57-P1 Turbojet Engine

The operational characteristics of a J57-P1 turbojet engine have been investigated at altitudes between 15,000 and 66,000 feet in the Lewis altitude wind tunnel. Included in this study is a discussion of fuel nozzle coking, the altitude operating limits with and without the standard engine control, the compressor surge characteristics, and the engine starting and windmilling characteristics. Severe circumferential turbine outlet temperature gradients which occurred at high altitude as a result of fuel nozzle coking were alleviated by the manufacturer's change in the fuel flow divider schedule and in a nozzle gasket material. Compressor air bleed is required to prevent surge of the outboard compressor in the low engine speed region. The maximum altitude at which the engine was operated without the control was about 66,000 feet at 0.8 flight Mach number and at a reduced engine speed to avoid compressor surge; with the engine control in operation, the altitude operating limit is reduced to approximately 59,000 feet. The maximum altitude at which the engine was started was about 40,000 feet.

Wallner, Lewis E.↗

NASA NTRS1947

Cooling Characteristics of the V-1650-7 Engine: Coolant-Flow Distribution, Cylinder Temperatures, and Heat Rejections at Typical Operating Conditions - 1

An investigation was conducted to determine the coolant-flow distribu tion, the cylinder temperatures, and the heat rejections of the V-165 0-7 engine . The tests were run a t several power levels varying from minimum fuel consumption to war emergency power and at each power l evel the coolant flows corresponded to the extremes of those likely t o be encountered in typical airplane installations, A mixture of 30-p ercent ethylene glycol and 70-percent water was used as the coolant. The temperature of each cylinder was measured between the exhaust val ves, between the intake valves, in the center of the head, on the exh aust-valve guide, at the top of the barrel on the exhaust side, and o n each exhaust spark-plug gasket. For an increase in engine power fro m 628 to approximately 1700 brake horsepower the average temperature for the cylinder heads between the exhaust valves increased from 437 deg to 517 deg F, the engine coolant heat rejection increased from 12 ,600 to 22,700 Btu. per minute, the oil heat rejection increased from 1030 to 4600 Btu per minute, and the aftercooler-coolant heat reject ion increased from 450 to 3500 Btu -per minute.

Povolny, John H.↗

NASA NTRS2006

NASA Tech Briefs, June 2006

Topics covered include: Magnetic-Field-Response Measurement-Acquisition System; Platform for Testing Robotic Vehicles on Simulated Terrain; Interferometer for Low-Uncertainty Vector Metrology; Rayleigh Scattering for Measuring Flow in a Nozzle Testing Facility; "Virtual Feel" Capaciflectors; FETs Based on Doped Polyaniline/Polyethylene Oxide Fibers; Miniature Housings for Electronics With Standard Interfaces; Integrated Modeling Environment; Modified Recursive Hierarchical Segmentation of Data; Sizing Structures and Predicting Weight of a Spacecraft; Stress Testing of Data-Communication Networks; Framework for Flexible Security in Group Communications; Software for Collaborative Use of Large Interactive Displays; Microsphere Insulation Panels; Single-Wall Carbon Nanotube Anodes for Lithium Cells; Tantalum-Based Ceramics for Refractory Composites; Integral Flexure Mounts for Metal Mirrors for Cryogenic Use; Templates for Fabricating Nanowire/Nanoconduit- Based Devices; Measuring Vapors To Monitor the State of Cure of a Resin; Partial-Vacuum-Gasketed Electrochemical Corrosion Cell; Theodolite Ring Lights; Integrating Terrain Maps Into a Reactive Navigation Strategy; Reducing Centroid Error Through Model-Based Noise Reduction; Adaptive Modeling Language and Its Derivatives; Stable Satellite Orbits for Global Coverage of the Moon; and Low-Cost Propellant Launch From a Tethered Balloon

Source record↗

NASA NTRS2006

Self-Passivating Lithium/Solid Electrolyte/Iodine Cells

Robust lithium/solid electrolyte/iodine electrochemical cells that offer significant advantages over commercial lithium/ iodine cells have been developed. At room temperature, these cells can be discharged at current densities 10 to 30 times those of commercial lithium/iodine cells. Moreover, from room temperature up to 80 C, the maximum discharge-current densities of these cells exceed those of all other solid-electrolyte-based cells. A cell of this type includes a metallic lithium anode in contact with a commercial flexible solid electrolyte film that, in turn, is in contact with an iodine/ graphite cathode. The solid electrolyte (the chemical composition of which has not been reported) offers the high ionic conductivity needed for high cell performance. However, the solid electrolyte exhibits an undesirable chemical reactivity to lithium that, if not mitigated, would render the solid electrolyte unsuitable for use in a lithium cell. In this cell, such mitigation is affected by the formation of a thin passivating layer of lithium iodide at the anode/electrolyte interface. Test cells of this type were fabricated from iodine/graphite cathode pellets, free-standing solid-electrolyte films, and lithium-foil anodes. The cathode mixtures were made by grinding together blends of nominally 10 weight percent graphite and 90 weight percent iodine. The cathode mixtures were then pressed into pellets at 36 kpsi (248 MPa) and inserted into coin-shaped stainless-steel cell cases that were coated with graphite paste to minimize corrosion. The solid-electrolyte film material was stamped to form circular pieces to fit in the coin cell cases, inserted in the cases, and pressed against the cathode pellets with polyethylene gaskets. Lithium-foil anodes were placed directly onto the electrolyte films. The layers described thus far were pressed and held together by stainless- steel shims, wave springs, and coin cell caps. The assembled cells were then crimped to form hermetic seals. It was found that the solid electrolyte films became discolored within seconds after they were placed in contact with the cathodes - a result of facile diffusion of iodine through the solid electrolyte material (see figure).

Bugga, Ratnakumar↗

NASA NTRS2005

Hexagonal and Pentagonal Fractal Multiband Antennas

Multiband dipole antennas based on hexagonal and pentagonal fractals have been analyzed by computational simulations and functionally demonstrated in experiments on prototypes. These antennas are capable of multiband or wide-band operation because they are subdivided into progressively smaller substructures that resonate at progressively higher frequencies by virtue of their smaller dimensions. The novelty of the present antennas lies in their specific hexagonal and pentagonal fractal configurations and the resonant frequencies associated with them. These antennas are potentially applicable to a variety of multiband and wide-band commercial wireless-communication products operating at different frequencies, including personal digital assistants, cellular telephones, pagers, satellite radios, Global Positioning System receivers, and products that combine two or more of the aforementioned functions. Perhaps the best-known prior multiband antenna based on fractal geometry is the Sierpinski triangle antenna (also known as the Sierpinski gasket), shown in the top part of the figure. In this antenna, the scale length at each iteration of the fractal is half the scale length of the preceding iteration, yielding successive resonant frequencies related by a ratio of about 2. The middle and bottom parts of the figure depict the first three iterations of the hexagonal and pentagonal fractals along with typical dipole-antenna configuration based on the second iteration. Successive resonant frequencies of the hexagonal fractal antenna have been found to be related by a ratio of about 3, and those of the pentagonal fractal antenna by a ratio of about 2.59.

Tang, Philip W.↗

NASA NTRS2005

Variable-Pressure Washer

The variable-pressure washer (VPW) is a proposed device that is so named because (1) it would play the role similar to that played by an ordinary washer, except that (2) the clamping pressure applied by it would vary with either circumferential or radial position. In a typical contemplated application, the radially varying clamping pressure would be used to obtain more nearly uniform compression on a pair of concentric seals (e.g., an O-ring or a gasket) in an assembly that experiences larger deformations normal to the sealing surface for locations around the outer diameter of the attachment flange when compared to locations around the inner diameter. The VPW (see figure) would include two interlocking channel rings pushed axially away from each other by compression spring-like components located at two or more radial positions. Each spring would have a different stiffness based on the radial location. Overlapping splits in each interlocking channel ring would allow for the non-uniform deformation in the rings. Each spring would be held in place by retaining cups attached to the inner flat surfaces of the channel rings. A plunger attached to one channel ring on the central axis would be captured in a plunger housing attached to the other channel ring: The capture of the plunger would hold the VPW together. When the VPW was clamped between two flat surfaces, the clamping force would be distributed unevenly across the face of the washer in the radial direction. The different stiffnesses of the springs would be chosen, in conjunction with other design parameters, to obtain a specified radial variation of clamping pressure in the presence of a specified clamping force.

Smeltzer, Stanley S., III↗

NASA NTRS2011

International Space Station (ISS) Emergency Mask (EM) Development

The Emergency Mask (EM) is considered a secondary response emergency Personal Protective Equipment (PPE) designed to provide respiratory protection to the International Space Station (ISS) crewmembers in response to a post-fire event or ammonia leak. The EM is planned to be delivered to ISS in 2012 to replace the current air purifying respirator (APR) onboard ISS called the Ammonia Respirator (AR). The EM is a one ]size ]fits ]all model designed to fit any size crewmember, unlike the APR on ISS, and uses either two Fire Cartridges (FCs) or two Commercial Off-the-Shelf (COTS) 3M(Trademark). Ammonia Cartridges (ACs) to provide the crew with a minimum of 8 hours of respiratory protection with appropriate cartridge swap ]out. The EM is designed for a single exposure event, for either post ]fire or ammonia, and is a passive device that cannot help crewmembers who cannot breathe on their own. The EM fs primary and only seal is around the wearer fs neck to prevent a crewmember from inhaling contaminants. During the development of the ISS Emergency Mask, several design challenges were faced that focused around manufacturing a leak free mask. The description of those challenges are broadly discussed but focuses on one key design challenge area: bonding EPDM gasket material to Gore(Registered Trademark) fabric hood.

Toon, Katherine P.↗

NASA NTRS2004

Development of a Lunar Consumables Storage and Distribution Depot

NASA is in the preliminary planning stages for a future lunar base as a response to President George W. Bush's recent announcement of a new sustained exploration program beyond low earth orbit. Kennedy Space Center engineers are supporting this program by utilizing experience in Spaceport system design and operations to help develop a Lunar Consumables Depot. This depot will store propellants, life support fluids, and other consumables either transported from Earth or manufactured from In Situ resources. The depot will distribute these consumables in an energy efficient manner to end users including spacecraft, habitation modules, and rovers. This paper addresses some of the changes to lunar base architecture design as a result of advances in knowledge of lunar resources over the past 35 years, as well as technology advances in the area of In Situ Resource Utilization and consumable storage and distribution. A general system level description of the depot will be presented, including overall design philosophy and high level requirements. Finally, specific subsystem technologies that have been or will be developed by KSC will be addressed. Examples of these technologies are automated umbilicals, cryogenic refrigerators, novel storage vessels, advanced heat switches and heat exchangers, and self healing gaskets and wires.

Mueller, Robert P.↗

NASA NTRS2012

Sealing Materials for Use in Vacuum at High Temperatures

Sealing materials that can be applied and left in place in vacuum over a wide range of temperatures (especially temperatures of a few thousand degrees Celsius) have been conceived and investigated for potential utility in repairing thermal-protection tiles on the space shuttles in orbit before returning to Earth. These materials are also adaptable to numerous terrestrial applications that involve vacuum processing and/or repair of structures that must withstand high temperatures. These materials can be formulated to have mechanical handling characteristics ranging from almost freely flowing liquid-like consistency through paste-like consistency to stiff puttylike consistency, and to retain these characteristics in vacuum until heated to high curing temperatures. A sealing material of this type can be formulated to be used in any of several different ways for example, to be impregnated into a high-temperature-fabric patch, impregnated into a high-temperature-fabric gasket for sealing a patch, applied under a patch, or applied alone in the manner of putty or wallboard compound. The sealing material must be formulated to be compatible with, and adhere to, the structural material(s) to be repaired. In general, the material consists of a vacuum-compatible liquid containing one or more dissolved compound(s) and/or mixed with suspended solid particles. Depending on the intended application, the liquid can be chosen to be of a compound that can remain in place in vacuum for a time long enough to be useful, and/or to evaporate or decompose in a controlled way to leave a useful solid residue behind. The evaporation rate is determined by proper choice of vapor pressure, application of heat, and/or application of ultraviolet light or other optical radiation. The liquid chosen for the original space shuttle application is a commercial silicone vacuum-pump oil.

Pettit, Donald R.↗

NASA NTRS2010

Mechanics of Re-Torquing in Bolted Flange Connections

It has been widely accepted that the phenomenon of time-dependent loosening of flange connections is a strong consequence of the viscous nature of the compression seal material. Characterizing the coupled interaction between gasket creep and elastic bolt stiffness has been useful in predicting conditions that facilitate leakage. Prior advances on this sub-class of bolted joints has lead to the development of (1) constitutive models for elastomerics, (2) initial tightening strategies, (3) etc. The effect of re-torque, which is a major consideration for typical bolted flange seals used on the Space Shuttle fleet, has not been fully characterized, however. The current study presents a systematic approach to characterizing bolted joint behavior as the consequence of sequentially applied torques. Based on exprimenta1 and numerical results, the optimal re-torquing parameters have been identified that allow for the negligible load loss after pre-load application

Gordon, Ali P.↗

NASA NTRS2014

Parametric Mass Reliability Study

The International Space Station (ISS) systems are designed based upon having redundant systems with replaceable orbital replacement units (ORUs). These ORUs are designed to be swapped out fairly quickly, but some are very large, and some are made up of many components. When an ORU fails, it is replaced on orbit with a spare; the failed unit is sometimes returned to Earth to be serviced and re-launched. Such a system is not feasible for a 500+ day long-duration mission beyond low Earth orbit. The components that make up these ORUs have mixed reliabilities. Components that make up the most mass-such as computer housings, pump casings, and the silicon board of PCBs-typically are the most reliable. Meanwhile components that tend to fail the earliest-such as seals or gaskets-typically have a small mass. To better understand the problem, my project is to create a parametric model that relates both the mass of ORUs to reliability, as well as the mass of ORU subcomponents to reliability.

Holt, James P.↗

NASA NTRS2015

Testing of an Arcjet Thruster with Capability of Direct-Drive Operation

Electric thrusters typically require a power processing unit (PPU) to convert the spacecraft provided power to the voltage-current that a thruster needs for operation. Testing has been initiated to study whether an arcjet thruster can be operated directly with the power produced by solar arrays without any additional conversion. Elimination of the PPU significantly reduces system-level complexity of the propulsion system, and lowers developmental cost and risk. The work aims to identify and address technical questions related to power conditioning and noise suppression in the system and heating of the thruster in long-duration operation. The apparatus under investigation has a target power level from 400-1,000 W. However, the proposed direct-drive arcjet is potentially a highly scalable concept, applicable to solar-electric spacecraft with up to 100's of kW and beyond. A direct-drive electric propulsion system would be comprised of a thruster that operates with the power supplied directly from the power source (typically solar arrays) with no further power conditioning needed between those two components. Arcjet thrusters are electric propulsion devices, with the power supplied as a high current at low voltage; of all the different types of electric thruster, they are best suited for direct drive from solar arrays. One advantage of an arcjet over Hall or gridded ion thrusters is that for comparable power the arcjet is a much smaller device and can provide more thrust and orders of magnitude higher thrust density (approximately 1-10 N/sq m), albeit at lower I(sub sp) (approximately 800-1000 s). In addition, arcjets are capable of operating on a wide range of propellant options, having been demonstrated on H2, ammonia, N2, Ar, Kr, Xe, while present SOA Hall and ion thrusters are primarily limited to Xe propellant. Direct-drive is often discussed in terms of Hall thrusters, but they require 250-300 V for operation, which is difficult even with high-voltage solar arrays. The arcjet requires under 100 V, which is more in-line with what is easily possible with a solar array. Direct-drive of an electric propulsion system confers the advantage of reducing or eliminating the power processing unit (PPU) that is typically needed to convert the spacecraft-provided power to the voltage and current needed for thruster operation. Since the PPU is typically the most expensive piece of an electric thruster system, from both a fabrication and qualification standpoint, its elimination offers the potential for major reductions in system cost and risk. The design of the arcjet built for this effort was based on previous low power (1 kW class) arcjets. It has a precision machined 99.95% pure tungsten anode which also serves as the nozzle. The anode constrictor region is 1 mm (0.040-in) diameter and 1 mm (0.040-in) long. The cathode is a tungsten welding electrode doped with LaO2; its tip was precision ground to a 30◦ angle ending in a blunt end. The two electrodes are separated by a boron-nitride insulator which also serves as the propellant injection manifold; it ends in six small holes which introduce the propellant gas in the diverging section of the nozzle, directly adjacent to the cathode. The electrodes and insulator are housed in a stainless-steel outer-body, with a Macor insulator at the mid-plane to provide thermal isolation between the front and back halves of the device. The gas seals were made using Grafoil gaskets. Figure 1A shows the assembled thruster in the vacuum chamber; figure 1B shows the thruster in operation on argon at a flow rate of 676 sccm (20 mg/s). Initial testing was conducted in a 3.5-ft diameter vacuum chamber; the ultimate pressure reached during quasi-steady operation of the thruster was about 330 millitorr. The thruster was powered with a high-current, 0-100A, 15 kW power supply. The discharge was initiated with a high-voltage (approximately 10 kV) spark initiator that was isolated from the supply by a stack of diodes. The testing indicated that an operating point exists within the I-V characteristics that is compatible with direct-drive solar-electric operation; for a flow rate of 20 mg/s (argon) the arc could be sustained at a voltage of about 20 V and a current of 25 A (500W).

Martin, Adam K.↗

NASA NTRS2016

Hexavalent Chrome Free Coatings for Electronics: Electromagnetic Interference (EMI) Shielding Effectiveness (SE)

Determine the suitability of trivalent chromium conversion coatings that meet the requirements of MIL-DTL-5541, Type II, for use in applications where high-frequency electrical performance is important. Evaluate the ability of hexavalent chrome free pretreated aluminum to form adequate EMI seals, and maintain that seal while being subjected to harsh environmental conditions. Assess the performance of trivalent chromium pretreatments against a known control hexavalent chrome pretreatment before and after they have been exposed to a set of environmental conditions. It is known that environmental testing causes a decrease in shielding effectiveness when hexavalent chrome pretreatments are used (Alodine 1200s). Need to determine how shielding effectiveness will be affected with the use of hexavalent chrome free pretreatments. Performance will be assessed by evaluating shielding effectiveness (SE) test data from a variety of test samples comprised of different aluminum types and/or conversion coatings. The formation of corrosion will be evaluated between the mating surfaces and gasket to assess the corrosion resistant properties of the pretreatments, comparing the hexavalent control to the hexavalent chrome free pretreatments.

hexavalent chrome free↗