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NASA NTRS2015

Continued Advancement of Supported Liquid Membranes for Carbon Dioxide Control in Extravehicular Activity Applications

The development of a new, robust, portable life support system (PLSS) is a high priority for NASA in order to support longer and safer extravehicular activity (EVA) missions. One of the critical PLSS functions is maintaining the carbon dioxide (CO2) concentration in the suit at acceptable levels. Although the Metal Oxide (MetOx) canister has historically performed very well, it has a finite CO2 adsorption capacity. Therefore, the size and weight of the unit would have to be increased to extend EVA times. Consequently, new CO2 control technologies must be developed in order to meet mission objectives without increasing the size of the PLSS. Recent work has centered on sorbents that can be regenerated during the EVA; however, this strategy increases the system complexity and power consumption. A much simpler approach is to employ a membrane that vents CO2 to space and retains oxygen (O2). A membrane has many advantages over current technology: it is a continuous system with no limit on capacity, it requires no consumables, and it does not need any hardware to switch beds between absorption and regeneration. Unfortunately, conventional gas separation membranes do not have the needed selectivity for use in the PLSS. However, the required performance could be obtained with a supported liquid membrane (SLM), which consists of a microporous material filled with a liquid that selectively reacts with CO2 over O2. In a recently completed Phase II SBIR project, Reaction Systems, Inc. achieved the required CO2 permeance and selectivity with an SLM in a flat sheet configuration. This paper describes work to convert the SLM into a more compact form and to scale it up to handle more representative process flow rates.

Wickham, David T.↗
NASA NTRS2000

Analysis and Test of a Proton Exchange Membrane Fuel Cell Power System for Space Power Applications

An effort is underway to develop a prototype Proton Exchange Membrane (PEM) Fuel Cell breadboard system for fuhlre space applications. This prototype will be used to develop a comprehensive design basis for a space-rated PEM fuel cell powerplant. The prototype system includes reactant pressure regulators, ejector-based reactant pumps, a 4-kW fuel cell stack and cooling system, and a passive, membranebased oxygen / water separator. A computer model is being developed concurrently to analytically predict fluid flow in the oxidant reactant system. Fuel cells have historically played an important role in human-rated spacecraft. The Gemini and Apollo spacecraft used fuel cells for vehicle electrical power. The Space Shuttle currently uses three Alkaline Fuel Cell Powerplants (AFCP) to generate all of the vehicle's 15-20kW electrical power. Engineers at the Johnson Space Center have leveraged off the development effort ongoing in the commercial arena to develop PEM fuel cel ls for terrestrial uses. The prototype design originated from efforts to develop a PEM fuel cell replacement for the current Space Shuttle AFCP' s. In order to improve on the life and an already excellent hi storical record of reliability and safety, three subsystems were focused on. These were the fuel cell stack itself, the reactant circulation devices, and reactant / product water separator. PEM fuel cell stack performance is already demonstrating the potential for greater than four times the useful life of the current Shuttle's AFCP. Reactant pumping for product water removal has historically been accomplished with mechanical pumps. Ejectors offer an effective means of reactant pumping as well as the potential for weight reduction, control simplification, and long life. Centrifugal water separation is used on the current AFCP. A passive, membrane-based water separator offers compatibility with the micro-gravity environment of space, and the potential for control simplification, elimination of moving parts in an oxygen environment, and long life. The prototype system has been assembled from components that have previously been tested and evaluated at the component level. Preliminary data obtained from tests performed with the prototype system, as well as other published data, has been used to validate the analytical component models. These components have been incorporated into an integrated oxidant fluid system model. Results obtained from both the performance tests and the analytical model are presented.

Vasquez, Arturo↗
NASA NTRS2015

Propulsion System Testing for the Iodine Satellite (iSAT) Demonstration Mission

CUBESATS are relatively new spacecraft platforms that are typically deployed from a launch vehicle as a secondary payload, providing low-cost access to space for a wide range of end-users. These satellites are comprised of building blocks having dimensions of 10x10x10 cm cu and a mass of 1.33 kg (a 1-U size). While providing low-cost access to space, a major operational limitation is the lack of a propulsion system that can fit within a CubeSat and is capable of executing high delta v maneuvers. This makes it difficult to use CubeSats on missions requiring certain types of maneuvers (i.e. formation flying, spacecraft rendezvous). Recently, work has been performed investigating the use of iodine as a propellant for Hall-effect thrusters (HETs) 2 that could subsequently be used to provide a high specific impulse path to CubeSat propulsion. 3, 4 Iodine stores as a dense solid at very low pressures, making it acceptable as a propellant on a secondary payload. It has exceptionally high ρIsp (density times specific impulse), making it an enabling technology for small satellite near-term applications and providing the potential for systems-level advantages over mid-term high power electric propulsion options. Iodine flow can also be thermally regulated, subliming at relatively low temperature (less than 100 C) to yield I2 vapor at or below 50 torr. At low power, the measured performance of an iodine-fed HET is very similar to that of a state-of-the-art xenon-fed thruster. Just as importantly, the current-voltage discharge characteristics of low power iodine-fed and xenon-fed thrusters are remarkably similar, potentially reducing development and qualifications costs by making it possible to use an already-qualified xenon-HET PPU in an iodine-fed system. Finally, a cold surface can be installed in a vacuum test chamber on which expended iodine propellant can deposit. In addition, the temperature doesn't have to be extremely cold to maintain a low vapor pressure in the vacuum chamber (it is under 10(exp -6) torr at −75 C), making it possible to 'cryopump' the propellant with lower-cost recirculating refrigerant-based systems as opposed to using liquid nitrogen or low temperature gaseous helium cryopanels. An iodine-based system is not without its challenges. The primary challenge is that the entire feed system must be maintained at an elevated temperature to prevent the iodine from depositing (transitioning from the gas phase directly back into the solid phase), which will block the propellant feed lines. Furthermore, deposition will occur unless the temperature in the lines is not greater than the temperature of the propellant reservoir. The flow rate can be controlled by adjusting the heating applied to the reservoir, but as with any thermal control there is a relatively slow response to changes in the heating rate. In the present paper, we describe the propulsion and propellant feed system for the iodine satellite (iSAT) flight demonstration mission. The system is based around the Busek BHT-200 Hall thruster, which has been modified for chemical compatibility with iodine vapor. While the gross propellant flow rate is maintained by the heated propellant reservoir, the flow to the anode and cathode are adjusted using two heated Vacco proportional flow control valves (PFCV), which provide very fast response on the flow rate adjustment. The flight mission design layout will be presented, showing how the system will be packaged into the overall 12-U spacecraft and the techniques being employed to protect the remaining spacecraft hardware from the propulsion system (e.g., plasma impingement, iodine deposition, thermal loads). In addition to the flight system design, results of testing the thruster and cathode with both operating on iodine propellant are presented. The tests are conducted on a thrust stand (see Fig. 1) in a large vacuum chamber containing a beam dump chilled to below -100 C to 'cryopump' the propellant. The thruster performance during these tests is presented, with these data used to evaluate the feed system and guide further refinements. Results of relatively long duration testing are presented to demonstrate the capability to operate for the length of the iSAT mission and to perform a number of re-starts as will be required by the mission concept of operations.

Polzin, Kurt A.↗
NASA NTRS1986

Amphibious Airplane

The airplane pictured is the new Air Shark I, a four-place amphibian that makes extensive use of composite materials and cruises at close to 200 miles per hour under power from a 200-horsepower engine. Air Shark I is a "homebuilt" airplane, assembled from a kit of parts and components furnished by Freedom Master Corporation, Satellite Beach, Florida. The airplane incorporates considerable NASA technology and its construction benefited from research assistance provided by Kennedy Space Center (KSC) In designing the Shark, company president Arthur M. Lueck was able to draw on NASA's aeronautical technology bank through KSC's computerized "recon" library. As a result of his work at KSC, the wing of the Air Shark I is a new airfoil developed by Langley Research Center for light aircraft. In addition, Lueck opted for NASA-developed "winglets," vertical extensions of the wing that reduce drag by smoothing air turbulence at the wingtips. The NASA technology bank also contributed to the hull design. Lueck is considering application of NASA laminar flow technology-means of smoothing the airflow over wing and fuselage-to later models for further improvement of the Shark's aerodynamic efficiency. A materials engineer, Lueck employed his own expertise in designing and selecting the materials for the composite segments, which include all structural members, exposed surfaces and many control components. The materials are fiber reinforced plastics, or FRP They offer a high strength-to-weight ratio, with a nominal strength rating about one and a half times that of structural steel. They provide other advantages: the materials can be easily molded into finished shapes without expensive tooling or machining, and they are highly corrosion resistant. The first homebuilt to be offered by Freedom Master, Air Shark I completed air and water testing in mid-1985 and the company launched production of kits.

Source record↗
NASA NTRS1994

Multiple Hollow Cathode Wear Testing

A hollow cathode-based plasma contactor has been baselined for use on the Space Station to reduce station charging. The plasma contactor provides a low impedance connection to space plasma via a plasma produced by an arc discharge. The hollow cathode of the plasma contactor is a refractory metal tube, through which xenon gas flows, which has a disk-shaped plate with a centered orifice at the downstream end of the tube. Within the cathode, arc attachment occurs primarily on a Type S low work function insert that is next to the orifice plate. This low work function insert is used to reduce cathode operating temperatures and energy requirements and, therefore, achieve increased efficiency and longevity. The operating characteristics and lifetime capabilities of this hollow cathode, however, are greatly reduced by oxygen bearing contaminants in the xenon gas. Furthermore, an optimized activation process, where the cathode is heated prior to ignition by an external heater to drive contaminants such as oxygen and moisture from the insert absorbed during exposure to ambient air, is necessary both for cathode longevity and a simplified power processor. In order to achieve the two year (approximately 17,500 hours) continuous operating lifetime requirement for the plasma contactor, a test program was initiated at NASA Lewis Research Center to demonstrate the extended lifetime capabilities of the hollow cathode. To date, xenon hollow cathodes have demonstrated extended lifetimes with one test having operated in excess of 8000 hours in an ongoing test utilizing contamination control protocols developed by Sarver-Verhey. The objectives of this study were to verify the transportability of the contamination control protocols developed by Sarver-Verhey and to evaluate cathode contamination control procedures, activation processes, and cathode-to-cathode dispersions in operating characteristics with time. These were accomplished by conducting a 2000 hour wear test of four hollow cathodes with different xenon gas purities and activation processes. The following will be presented: a description of the facility and test hardware, testing procedures and operating conditions, a discussion of test results, and conclusions.

Soulas, George C.↗
NASA NTRS2008

Rescue Shuttle Flight Re-Entry: Controlling Astronaut Thermal Exposure

A rescue mission for the STS-125 Hubble Telescope Repair Mission requires reentry from space with 11 crew members aboard, exceeding past cabin thermal load experience and risking crew thermal stress potentially causing cognitive performance and physiological decrements. The space shuttle crew cabin air revitalization system (ARS) was designed to support a nominal crew complement of 4 to 7 crew and 10 persons in emergencies, all in a shirt-sleeve environment. Subsequent to the addition of full pressure suits with individual cooling units, the ARS cannot maintain a stable temperature in the crew cabin during reentry thermal loads. Bulk cabin thermal models, used for rescue mission planning and analysis of crew cabin air, were unable to accurately represent crew workstation values of air flow, carbon dioxide, and heat content for the middeck. Crew temperature models suggested significantly elevated core temperatures. Planning for an STS-400 potential rescue of seven stranded crew utilized computational fluid dynamics (CFD) models to demonstrate inhomogeneous cabin thermal properties and improve analysis compared to bulk models. In the absence of monitoring of crew temperature, heart rate, metabolic rate and incomplete engineering data on the performance of the integrated cooling garment/cooling unit (ICG/CU) at cabin temperatures above 75 degrees F, related systems & models were reevaluated and tests conducted with humans in the loop. Changes to the cabin ventilation, ICU placement, crew reentry suit-donning procedures, Orbiter Program wave-off policy and post-landing power down and crew extraction were adopted. A second CFD and core temperature model incorporated the proposed changes and confirmed satisfactory cabin temperature, improved air distribution, and estimated core temperatures within safe limits. CONCLUSIONS: These changes in equipment, in-flight and post-landing procedures, and policy were implemented for the STS-400 rescue shuttle & will be implemented in any future rescue flights from the International Space Station of stranded shuttle crews.

Gillis, David B.↗
NASA NTRS1953

Low-Speed Investigation of the Static Lateral Stability and Control Characteristics of a 1/6-Scale Model of the Republic XF-84H Airplane with the Propeller Operating

An investigation was made to determine the static lateral stability and control characteristics of a l/6-scale model of the Republic XF-84H airplane with the propeller operating. The model had a 40deg swept wing of aspect ratio 3.45 and had a thin 3-blade supersonic-type propeller. Many modifications to the basic configuration were investigated in attempts to alleviate lateral and directional trim problems which appeared to be associated with propeller slipstream rotation. Although significant benefits were realized with several modifications, none of those tested would be expected to afford satisfactory behavior for all normal flight conditions. A marked left-wing roll-off tendency was indicated at high angles of attack for the basic model configuration. Projection of only the left slat was the most effective remedy found for this problem with the propeller operating. The use of differential wing-flap deflection also appeared to offer a promising means for reducing the roll-off tendency with power on. The large sidewash over the vertical tail, associated with slip- stream rotation, severely restricted the conditions for which directional , trim could be maintained. A small triangular dorsal fin, oriented opposite to the slipstream rotation, was found very effective in reducing the adverse sidewash flow at the tail.

Sleeman, William C., Jr.↗
NASA NTRS2011

NEXT Propellant Management System Integration With Multiple Ion Thrusters

As a critical part of the NEXT test validation process, a multiple-string integration test was performed on the NEXT propellant management system and ion thrusters. The objectives of this test were to verify that the PMS is capable of providing stable flow control to multiple thrusters operating over the NEXT system throttling range and to demonstrate to potential users that the NEXT PMS is ready for transition to flight. A test plan was developed for the sub-system integration test for verification of PMS and thruster system performance and functionality requirements. Propellant management system calibrations were checked during the single and multi-thruster testing. The low pressure assembly total flow rates to the thruster(s) were within 1.4 percent of the calibrated support equipment flow rates. The inlet pressures to the main, cathode, and neutralizer ports of Thruster PM1R were measured as the PMS operated in 1-thruster, 2-thruster, and 3-thruster configurations. It was found that the inlet pressures to Thruster PM1R for 2-thruster and 3-thruster operation as well as single thruster operation with the PMS compare very favorably indicating that flow rates to Thruster PM1R were similar in all cases. Characterizations of discharge losses, accelerator grid current, and neutralizer performance were performed as more operating thrusters were added to the PMS. There were no variations in these parameters as thrusters were throttled and single and multiple thruster operations were conducted. The propellant management system power consumption was at a fixed voltage to the DCIU and a fixed thermal throttle temperature of 75 C. The total power consumed by the PMS was 10.0, 17.9, and 25.2 W, respectively, for single, 2-thruster, and 3-thruster operation with the PMS. These sub-system integration tests of the PMS, the DCIU Simulator, and multiple thrusters addressed, in part, the NEXT PMS and propulsion system performance and functionality requirements.

Sovey, James S.↗
NASA NTRS2001

Quantitative Species Measurements in Microgravity Combustion Flames

Flame-vortex interactions are canonical configurations that can be used to study the underlying processes occurring in complicated turbulent reacting flows. The elegant simplicity of the flame-vortex interaction permits the study of these complex interactions under relatively controllable experimental configurations, in contrast to direct measurements in turbulent flames. The ability to measure and model the fundamental phenomena that occur in a turbulent flame, but with time and spatial scales which are amenable to our diagnostics, permits significant improvements in the understanding of turbulent combustion under both normal and reduced gravity conditions. In this paper, we report absolute mole fraction measurements of methane in a reacting vortex ring. These microgravity experiments are performed in the 2.2-sec drop tower at NASA Glenn Research Center. In collaboration with Drs. Chen and Dahm at the University of Michigan, measured methane absorbances are incorporated into a new model from which the temperature and concentrations of all major gases in the flame can be determined at all positions and times in the development of the vortex ring. This is the first demonstration of the ITAC (Iterative Temperature with Assumed Chemistry) approach, and the results of these computations and analyses are presented in a companion paper by Dahm and Chen at this Workshop. We believe that the ITAC approach will become a powerful tool in understanding a wide variety of combustion flames under both equilibrium and non-equilibrium conditions.

Silver, Joel A.↗
NASA NTRS2003

Flame Synthesis Of Single-Walled Carbon Nanotubes And Nanofibers

Carbon nanotubes are widely sought for a variety of applications including gas storage, intercalation media, catalyst support and composite reinforcing material [1]. Each of these applications will require large scale quantities of CNTs. A second consideration is that some of these applications may require redispersal of the collected CNTs and attachment to a support structure. If the CNTs could be synthesized directly upon the support to be used in the end application, a tremendous savings in post-synthesis processing could be realized. Therein we have pursued both aerosol and supported catalyst synthesis of CNTs. Given space limitations, only the aerosol portion of the work is outlined here though results from both thrusts will be presented during the talk. Aerosol methods of SWNT, MWNT or nanofiber synthesis hold promise of large-scale production to supply the tonnage quantities these applications will require. Aerosol methods may potentially permit control of the catalyst particle size, offer continuous processing, provide highest product purity and most importantly, are scaleable. Only via economy of scale will the cost of CNTs be sufficient to realize the large-scale structural and power applications on both earth and in space. Present aerosol methods for SWNT synthesis include laser ablation of composite metalgraphite targets or thermal decomposition/pyrolysis of a sublimed or vaporized organometallic [2]. Both approaches, conducted within a high temperature furnace, have produced single-walled nanotubes (SWNTs). The former method requires sophisticated hardware and is inherently limited by the energy deposition that can be realized using pulsed laser light. The latter method, using expensive organometallics is difficult to control for SWNT synthesis given a range of gasparticle mixing conditions along variable temperature gradients; multi-walled nanotubes (MWNTs) are a far more likely end products. Both approaches require large energy expenditures and produce CNTs at prohibitive costs, around $500 per gram. Moreover these approaches do not possess demonstrated scalability. In contrast to these approaches, flame synthesis can be a very energy efficient, low-cost process [3]; a portion of the fuel serves as the heating source while the remainder serves as reactant. Moreover, flame systems are geometrically versatile as illustrated by innumerable boiler and furnace designs. Addressing scalability, flame systems are commercially used for producing megatonnage quantities of carbon black [4]. Although it presents a complex chemically reacting flow, a flame also offers many variables for control, e.g. temperature, chemical environment and residence times [5]. Despite these advantages, there are challenges to scaling flame synthesis as well.

Wal, Randy L. Vander↗
NASA NTRS2012

NASA Tech Briefs, August 2012

Topics covered include: Mars Science Laboratory Drill; Ultra-Compact Motor Controller; A Reversible Thermally Driven Pump for Use in a Sub-Kelvin Magnetic Refrigerator; Shape Memory Composite Hybrid Hinge; Binding Causes of Printed Wiring Assemblies with Card-Loks; Coring Sample Acquisition Tool; Joining and Assembly of Bulk Metallic Glass Composites Through Capacitive Discharge; 670-GHz Schottky Diode-Based Subharmonic Mixer with CPW Circuits and 70-GHz IF; Self-Nulling Lock-in Detection Electronics for Capacitance Probe Electrometer; Discontinuous Mode Power Supply; Optimal Dynamic Sub-Threshold Technique for Extreme Low Power Consumption for VLSI; Hardware for Accelerating N-Modular Redundant Systems for High-Reliability Computing; Blocking Filters with Enhanced Throughput for X-Ray Microcalorimetry; High-Thermal-Conductivity Fabrics; Imidazolium-Based Polymeric Materials as Alkaline Anion-Exchange Fuel Cell Membranes; Electrospun Nanofiber Coating of Fiber Materials: A Composite Toughening Approach; Experimental Modeling of Sterilization Effects for Atmospheric Entry Heating on Microorganisms; Saliva Preservative for Diagnostic Purposes; Hands-Free Transcranial Color Doppler Probe; Aerosol and Surface Parameter Retrievals for a Multi-Angle, Multiband Spectrometer LogScope; TraceContract; AIRS Maps from Space Processing Software; POSTMAN: Point of Sail Tacking for Maritime Autonomous Navigation; Space Operations Learning Center; OVERSMART Reporting Tool for Flow Computations Over Large Grid Systems; Large Eddy Simulation (LES) of Particle-Laden Temporal Mixing Layers; Projection of Stabilized Aerial Imagery Onto Digital Elevation Maps for Geo-Rectified and Jitter-Free Viewing; Iterative Transform Phase Diversity: An Image-Based Object and Wavefront Recovery; 3D Drop Size Distribution Extrapolation Algorithm Using a Single Disdrometer; Social Networking Adapted for Distributed Scientific Collaboration; General Methodology for Designing Spacecraft Trajectories; Hemispherical Field-of-View Above-Water Surface Imager for Submarines; and Quantum-Well Infrared Photodetector (QWIP) Focal Plane Assembly.

Source record↗
NASA NTRS2006

The Exploration Water Recovery System

The Exploration Water Recovery System is designed towards fulfillment of NASA s Vision for Space Exploration, which will require elevation of existing technologies to higher levels of optimization. This new system, designed for application to the Exploration infrastructure, presents a novel combination of proven air and water purification technologies. The integration of unit operations is modified from that of the current state-of-the-art water recovery system so as to optimize treatment of the various waste water streams, contaminant loads, and flow rates. Optimization is achieved primarily through the removal of volatile organic contaminants from the vapor phase prior to their absorption into the liquid phase. In the current state-of-the-art system, the water vapor in the cabin atmosphere is condensed, and the volatile organic contaminants present in that atmosphere are absorbed into the aqueous phase. Removal of contaminants the5 occurs via catalytic oxidation in the liquid phase. Oxidation kinetics, however, dictate that removal of volatile organic contaminants from the vapor phase can inherently be more efficient than their removal from the aqueous phase. Taking advantage of this efficiency reduces the complexity of the water recovery system. This reduction in system complexity is accompanied by reductions in the weight, volume, power, and resupply requirements of the system. Vapor compression distillation technology is used to treat the urine, condensate, and hygiene waste streams. This contributes to the reduction in resupply, as incorporation of vapor compression distillation technology at this point in the process reduces reliance on the expendable ion exchange and adsorption media used in the current state-of-the-art water recovery system. Other proven technologies that are incorporated into the Exploration Water Recovery System include the Trace Contaminant Control System and the Volatile Removal Assembly.

ORourke, Mary Jane E.↗
NASA NTRS2012

Control Algorithms Charge Batteries Faster

On March 29, 2011, NASA s Mercury Surface, Space Environment, Geochemistry and Ranging (MESSENGER) spacecraft beamed a milestone image to Earth: the first photo of Mercury taken from orbit around the solar system s innermost planet. (MESSENGER is also the first spacecraft to orbit Mercury.) Like most of NASA s deep space probes, MESSENGER is enabled by a complex power system that allows its science instruments and communications to function continuously as it travels millions of miles from Earth. "Typically, there isn't one particular power source that can support the entire mission," says Linda Taylor, electrical engineer in Glenn Research Center s Power Systems Analysis Branch. "If you have solar arrays and you are in orbit, at some point you re going to be in eclipse." Because of this, Taylor explains, spacecraft like MESSENGER feature hybrid power systems. MESSENGER is powered by a two-panel solar array coupled with a nickel hydrogen battery. The solar arrays provide energy to the probe and charge the battery; when the spacecraft s orbit carries it behind Mercury and out of the Sun s light, the spacecraft switches to battery power to continue operations. Typically, hybrid systems with multiple power inputs and a battery acting alternately as storage and a power source require multiple converters to handle the power flow between the devices, Taylor says. (Power converters change the qualities of electrical energy, such as from alternating current to direct current, or between different levels of voltage or frequency.) This contributes to a pair of major concerns for spacecraft design. "Weight and size are big drivers for any space application," Taylor says, noting that every pound added to a space vehicle incurs significant costs. For an innovative solution to managing power flows in a lightweight, cost-effective manner, NASA turned to a private industry partner.

Source record↗
NASA NTRS2009

Efficient Ionization Investigation for Flow Control and Energy Extraction

Nonequilibrium ionization of air by nonthermal means is explored for hypersonic vehicle applications. The method selected for evaluation generates a weakly ionized plasma using pulsed nanosecond, high-voltage discharges sustained by a lower dc voltage. These discharges promise to provide a means of energizing and sustaining electrons in the air while maintaining a nearly constant ion/neutral molecule temperature. This paper explores the use of short approx.5 nsec, high-voltage approx.12 to 22 kV, repetitive (40 to 100 kHz) discharges in generating a weakly ionized gas sustained by a 1 kV dc voltage in dry air at pressures from 10 to 80 torr. Demonstrated lifetimes of the sustainer discharge current approx.10 to 25 msec are over three orders of magnitude longer than the 5 nsec pulse that generates the electrons. This life is adequate for many high speed flows, enabling the possibility of exploiting weakly ionized plasma phenomena in flow-fields such as those in hypersonic inlets, combustors, and nozzles. Results to date are obtained in a volume of plasma between electrodes in a bell jar. The buildup and decay of the visible emission from the pulser excited air is photographed on an ICCD camera with nanosecond resolution and the time constants for visible emission decay are observed to be between 10 to 15 nsec decreasing as pressure increases. The application of the sustainer voltage does not change the visible emission decay time constant. Energy consumption as indicated by power output from the power supplies is 194 to 669 W depending on pulse repetition rate.

Schneider, Steven J.↗
NASA NTRS2009

The Evolution and Development of the Lunar Regolith and Implications for Lunar Surface Operations and Construction

The lunar regolith consists of about 90% submillimeter particles traditionally termed lunar soil. The remainder consists of larger particles ranging up to boulder size rocks. At the lower size end, soil particles in the 10s of nanometer sizes are present in all soil samples. Lunar regolith overlies bedrock which consists of either lava flows in mare regions or impact-produced megaregolith in highland regions. Lunar regolith has been produced over billions of years by a combination of breaking and communition of bedrock by meteorite bombardment coupled with a variety of complex space weathering processes including solar wind implantation, solar flare and cosmic ray bombardment with attendant radiation damage, melting, vaporization, and vapor condensation driven by impact, and gardening and turnover of the resultant soil. Lunar regolith is poorly sorted compared to most terrestrial soils, and has interesting engineering properties including strong grain adhesion, over-compacted soil density, an abundance of agglutinates with sharp corners, and a variety of properties related to soil maturity. The NASA program has supported a variety of engineering test research projects, the production of bricks by solar or microwave sintering, the production of concrete, the in situ sintering and glazing of regolith by microwave, and the extraction of useful resources such as oxygen, hydrogen, iron, aluminum, silicon and other products. Future requirements for a lunar surface base or outpost will include construction of protective berms, construction of paved roadways, construction of shelters, movement and emplacement of regolith for radiation shielding and thermal control, and extraction of useful products. One early need is for light weight but powerful digging, trenching, and regolith-moving equipment.

McKay, David↗
NASA NTRS2001

A microwave-powered sterilizable interface for aseptic access to bioreactors that are vulnerable to microbial contamination

Novel methods and apparatus that employ the rapid heating characteristics of microwave irradiation to facilitate the aseptic transfer of nutrients, products, and other materials between microbially sensitive systems and the external environment are described. The microwave-sterilizable access port (MSAP) consists of a 600-W magnetron emitting at a frequency of 2.45 GHz, a sterilization chamber with inlet and outlet flow lines, and a specimen transfer interface. Energy is routed to the sterilization chamber via a coaxial transmission line where small quantities of water couple strongly with the incident radiation to produce a superheated vapor phase. The efficiency of energy transfer is enhanced through the use of microwave susceptors within the sterilization chamber. Mating surfaces are thermally sterilized through direct contact with the hot gas. Efficacy has been demonstrated using the thermophile Bacillus stearothermophilus.

Validation Studies↗
NASA NTRS1992

Potential pressurized payloads: Fluid and thermal experiments

Space Station Freedom (SSF) presents the opportunity to perform long term fluid and thermal experiments in a microgravity environment. This presentation provides perspective on the need for fluids/thermal experimentation in a microgravity environment, addresses previous efforts, identifies possible experiments, and discusses the capabilities of a proposed fluid physics/dynamics test facility. Numerous spacecraft systems use fluids for their operation. Thermal control, propulsion, waste management, and various operational processes are examples of such systems. However, effective ground testing is very difficult. This is because the effect of gravity induced phenomena, such as hydrostatic pressure, buoyant convection, and stratification, overcome such forces as surface tension, diffusion, electric potential, etc., which normally dominate in a microgravity environment. Hence, space experimentation is necessary to develop and validate a new fluid based technology. Two broad types of experiments may be performed on SSF: basic research and applied research. Basic research might include experiments focusing on capillary phenomena (with or without thermal and/or solutal gradients), thermal/solutal convection, phase transitions, and multiphase flow. Representative examples of applied research might include two-phase pressure drop, two-phase flow instabilities, heat transfer coefficients, fluid tank fill/drain, tank slosh dynamics, condensate removal enhancement, and void formation within thermal energy storage materials. In order to better support such fluid/thermal experiments on board SSF, OSSA has developed a conceptual design for a proposed Fluid Physics/Dynamics Facility (FP/DF). The proposed facility consists of one facility rack permanently located on SSF and one experimenter rack which is changed out as needed to support specific experiments. This approach will minimize the on-board integration/deintegration required for specific experiments. The FP/DF will have acceleration/vibration compensation, power and thermal interfaces, computer command/data collection, a video imaging system, and a portable glove box for operations. This facility will allow real-time astronaut interaction with the testing.

Swanson, Theodore D.↗
NASA NTRS2008

System Engineering on the Use for Ares I,V - the Simpler, the Better

The Ares I and Ares V Vehicles will utilize the J-2X rocket engine developed for NASA by the Pratt & Whitney Rocketdyne Company. The J-2X is an improved higher power version of the original J-2 engine used during the Apollo program. With higher power and updated requirements for safety and performance, the J-2X becomes a new engine using state-of-the-art design methodology, materials and manufacturing processes. The implementation of Systems Engineering (SE) principles enables the rapid J-2X development program to remain aligned with the ARES I and V vehicle programs, Meeting the aggressive development schedule is a challenge. Coordinating the best expertise thai NASA and PWR have to offer requires effectively utilizing resources at multiple sites. This presents formidable communication challenges. SE allows honest and open discussions of issues and problems. This simple idea is often overlooked in large and complex SE programs. Regular and effective meetings linking SE objectives to component designs are used to voice differences of opinions with customer and contractor in attendance so that the best mutual decisions can be made on the shortest possible schedule. Regular technical interchange meetings on secure program wide computer networks and CM processes are effective,in the "Controlled Change" process that exemplifies good SE. Good communication is a key effective SE implementation. The System of Systems approach is the vision of the Orion program which facilitates the establishment of dynamic SE processes at all levels including the engine. SE enables requirements evolution by facilitating organizational and process agility. Flow down and distribution of requirements is controlled by Allocation Reports which breakdown numerical design objectives (weight, reliability, etc.) into quanta goals for each component area. Linked databases of design and verification requirements helps eliminate redundancy and potential mistakes inherent m separated systems. Another tool, the Architecture Design Description, is being used to control J-2X system architecture and effectively communicate configuration changes to those involved in the design process. But the proof is in successful program accomplishment. The SE is the methodology being used to meet the challenge of completing J-2X engine certification 2 years ahead of any engine program ever developed at PWR. The Ares I SE system of systems has delivered according to expectations thus far. All major design reviews (SRR. PDR, CDR) have been successfully conducted to satisfy overall program objectives using SE as the basis for accomplishment. The paper describes SE tools and techniques utilized to achieve this success.

Kelly, William↗