ALBERT

Description: The designers of the Orion Crew Exploration Vehicle (CEV) utilize an intensive simulation program in order to predict the launch and landing characteristics of the Crew Impact Attenuation System (CIAS). The CIAS is the energy absorbing strut concept that dampens loads to levels sustainable by the crew during landing and consists of the crew module seat pallet that accommodates four to six seated astronauts. An important parameter required for proper dynamic modeling of the CIAS is knowledge of the suited center of mass (COM) variations within the crew population. Significant center of mass variations across suited crew configurations would amplify the inertial effects of the pallet and potentially create unacceptable crew loading during launch and landing. Established suited, whole-body, and posture-based mass properties were not available due to the uncertainty of the final CEV seat posture and suit hardware configurations. While unsuited segmental center of mass values can be obtained via regression equations from previous studies, building them into a model that was posture dependent with custom anthropometry and integrated suit components proved cumbersome and time consuming. Therefore, the objective of this study was to quantify the effects of posture, suit components, and the expected range of anthropometry on the center of mass of a seated individual. Several elements are required for the COM calculation of a suited human in a seated position: anthropometry; body segment mass; suit component mass; suit component location relative to the body; and joint angles defining the seated posture. Anthropometry and body segment masses used in this study were taken from a selection of three-dimensional human body models, called boundary manikins, which were developed in a previous project. These boundary manikins represent the critical anthropometric dimension extremes for the anticipated astronaut population. Six male manikins and 6 female manikins, representing a subset of the possible maximum and minimum sized crewmembers, were segmented using point-cloud software to create 17 major body segments. The general approach used to calculate the human mass properties was to utilize center of volume outputs from the software for each body segment and apply a homogeneous density function to determine segment mass 3-D coordinates. Suit components, based on the current consensus regarding predicted suit configuration values, were treated as point masses and were positioned using vector mathematics along the body segments based on anthropometry and COM position. A custom MATLAB script then articulates the body segment and suit positions into a selected seated configuration, using joint angles that characterize a standard seated position and a CEV specific seated position. Additional MATLAB(r) scripts are finally used to calculate the composite COM positions in 3-D space for all 12 manikins in both suited and unsuited conditions for both seated configurations. The analysis focused on two aspects: (1) to quantify how much the whole body COM varied from the smallest to largest subject and (2) the impacts of the suit components on the overall COM in each seat configuration. The location across all boundary manikins of the anterior- posterior COM varied by approximately 7cm, the vertical COM varied by approximately 9-10cm, and the mediolateral COM varied by approximately 1.2 cm from the midline sagittal plane for both seat configurations. This variation was surprisingly large given the relative proportionality of the mass distribution of the human body. The suit components caused an anterior shift of the total COM by approximately 2 cm and a shift to the right along the mediolateral axis of 0.4 cm for both seat configurations. When the seat configuration is in the standard posture, the suited vertical COM shifts inferiorly by up to 1 cm whereas in the CEV posture the vertical COM has no appreciable change. These general differences were due the high proportion of suit mass located in the boots and lower legs and their corresponding distance from the body COM as well as the prevalence of suit components on the right side of the body.

Description: This document constitutes the publication of work performed by the Space Human Factors Laboratory (mail code SF5 at the time) at the Johnson Space Center (JSC) in the months of June and July of 2000. At that time, the Space Human Factors Laboratory was part of the Space Human Factors Branch in the Flight Projects Division of the Space and Life Directorate. This report was originally to be a document for internal consumption only at JSC as it was seen to be only preliminary work for the further development of solid state illumination for general lighting on future space vehicles and the International Space Station (ISS). Due to funding constraints, immediate follow-on efforts were delayed and the need for publication of this document was overcome by other events. However, in recent years and with the development and deployment of a solid state light luminaire prototype on ISS, the time was overdue for publishing this information for general distribution and reference. Solid state lights (SSLs) are being developed to potentially replace the general luminaire assemblies (GLAs) currently in service in the International Space Station (ISS) and included in designs of modules for the ISS. The SSLs consist of arrays of light emitting diodes (LEDs), small solid state electronic devices that produce visible light in proportion to the electrical current flowing through them. Recent progressive advances in electrical power-to-light conversion efficiency in LED technology have allowed the consideration of LEDs as replacements for incandescent and fluorescent light sources in many circumstances, and their inherent advantages in ruggedness, reliability, and life expectancy make them attractive for applications in spacecraft. One potential area of application for the SSLs in the U.S. Laboratory Module of the ISS. This study addresses the suitability of the SSLs as replacements for the GLAs in this application.

Description: This document contains some of the descriptions of payload and experiment related to life support and habitation. These describe experiments that have or are scheduled to fly on the International Space Station. There are instructions, and descriptions of the fields that make up the database. The document is arranged in alphabetical order by the Payload

Description: Aircraft induced contrails have been found to have a net warming influence on the climate system, with strong regional dependence. Persistent linear contrails are detectable in 1 Km thermal imagery and, using an automated Contrail Detection Algorithm (CDA), can be identified on the basis of their different properties at the 11 and 12 m w av.el enTgthshe algorithm s ability to distinguish contrails from other linear features depends on the sensitivity of its tuning parameters. In order to keep the number of false identifications low, the algorithm imposes strict limits on contrail size, linearity and intensity. This paper investigates whether including additional information (i.e. meteorological data) within the CDA may allow for these criteria to be less rigorous, thus increasing the contrail-detection rate, without increasing the false alarm rate.

Description: NASA's exploration mission is to send humans to the Moon and Mars, in which the purpose is to learn how to live and work safely in those harsh environments. A critical aspect of living in an extreme environment is habitation, and within that habitation element there are key systems which monitor the habitation environment to provide a safe and comfortable living and working space for humans. Expandable habitats are one of the options currently being considered due to their potential mass and volume efficiencies. This paper discusses a joint project between the National Science Foundation (NSF), ILC Dover, and NASA in which an expandable habitat was deployed in the extreme environment of Antarctica to better understand the performance and operations over a one-year period. This project was conducted through the Innovative Partnership Program (IPP) where the NSF provided the location at McMurdo Station in Antarctica and support at the location, ILC Dover provided the inflatable habitat, and NASA provided the instrumentation and data system for monitoring the habitat. The outcome of this project provided lessons learned in the implementation of an inflatable habitat and the systems that support that habitat. These lessons learned will be used to improve current habitation capabilities and systems to meet the objectives of exploration missions to the moon and Mars.

Description: This paper presents a flutter analysis technique for the transonic flight regime. The technique uses an iterative approach to determine the critical dynamic pressure for a given mach number. Unlike other CFD-based flutter analysis methods, each iteration solves for the critical dynamic pressure and uses this value in subsequent iterations until the value converges. This process reduces the iterations required to determine the critical dynamic pressure. To improve the accuracy of the analysis, the technique employs a known structural model, leaving only the aerodynamic model as the unknown. The aerodynamic model is estimated using unsteady aeroelastic CFD analysis combined with a parameter estimation routine. The technique executes as follows. The known structural model is represented as a finite element model. Modal analysis determines the frequencies and mode shapes for the structural model. At a given mach number and dynamic pressure, the unsteady CFD analysis is performed. The output time history of the surface pressure is converted to a nodal aerodynamic force vector. The forces are then normalized by the given dynamic pressure. A multi-input multi-output parameter estimation software, ERA, estimates the aerodynamic model through the use of time histories of nodal aerodynamic forces and structural deformations. The critical dynamic pressure is then calculated using the known structural model and the estimated aerodynamic model. This output is used as the dynamic pressure in subsequent iterations until the critical dynamic pressure is determined. This technique is demonstrated on the Aerostructures Test Wing-2 model at NASA's Dryden Flight Research Center.

Description: We continue the development of our concept of a biotechnological loop for in-situ resource extraction along with propellant and food production at a future lunar outpost, based on the cultivation of litholytic cyanobacteria (LCB) with lunar regolith (LR) in a geobioreactor energized by sunlight. Our preliminary studies have shown that phototropic cultivation of LCB with simulants of LR in a low-mineralized medium supplemented with CO2 leads to rock dissolution (bioweathering) with the resulting accumulation of Fe, Mg and Al in cyanobacterial cells and in the medium. LCB cultivated with LR simulants produces more O2 than the same organisms cultivated in a high-mineralized medium. The loss of rock mass after bioweathering with LCB suggests the release of O from regolith. Further studies of chemical pathways of released O are required. The bioweathering process is limited by the availability of CO2, N, and P. Since lunar regolith is mainly composed of O, Si, Ca, Al and Mg, we propose to use organic waste to supply a geobioreactor with C, N and P. The recycling of organic waste, including urine, through a geobioreactor will allow for efficient element extraction as well as oxygen and biomass production. The most critical conclusion is that a biological life support system tied to a geobioreactor might be more efficient for supporting an extraterrestrial outpost than a closed environmental system.

Description: This paper describes the use of the molten oxide electrolysis (MOE) process for the extraction of oxygen for life support and propellant, and silicon and metallic elements for use in fabrication on the Moon. The Moon is rich in mineral resources, but it is almost devoid of chemical reducing agents, therefore, molten oxide electrolysis is ideal for extraction, since the electron is the only practical reducing agent. MOE has several advantages over other extraction methods. First, electrolytic processing offers uncommon versatility in its insensitivity to feedstock composition. Secondly, oxide melts boast the twin key attributes of highest solubilizing capacity for regolith and lowest volatility of any candidate electrolytes. The former is critical in ensuring high productivity since cell current is limited by reactant solubility, while the latter simplifies cell design by obviating the need for a gas-tight reactor to contain evaporation losses as would be the case with a gas or liquid phase fluoride reagent operating at such high temperatures. Alternatively, MOE requires no import of consumable reagents (e.g. fluorine and carbon) as other processes do, and does not rely on interfacing multiple processes to obtain refined products. Electrolytic processing has the advantage of selectivity of reaction in the presence of a multi-component feed. Products from lunar regolith can be extracted in sequence according to the stabilities of their oxides as expressed by the values of the free energy of oxide formation (e.g. chromium, manganese, Fe, Si, Ti, Al, magnesium, and calcium). Previous work has demonstrated the viability of producing Fe and oxygen from oxide mixtures similar in composition to lunar regolith by molten oxide electrolysis (electrowinning), also called magma electrolysis having shown electrolytic extraction of Si from regolith simulant. This paper describes recent advances in demonstrating the MOE process by a joint project with participation by NASA KSC and MSFC, and Ohio State University and MIT. Progress in measuring cell efficiency for oxygen production, development of non reacting electrodes, and cell feeding and withdrawal will be discussed.

Description: In 2009, the Cascade Distillation Subsystem (CDS) wastewater processor (Honeywell International, Torrance, CA) was assessed in the National Aeronautics and Space Administration (NASA) Exploration Life Support (ELS) distillation comparison test. The purpose of the test was to collect data to support down-selection and development of a primary distillation technology for application in a lunar outpost water recovery system. The CDS portion of the comparison test was conducted between May 6 and August 19, 2009. The system was challenged with two pretreated test solutions, each intended to represent a feasible wastewater generated in a surface habitat. The 30-day equivalent wastewater loading volume for a crew of four was processed for each wastewater solution. Test Solution 1 consisted of a mixed stream containing human-generated urine and humidity condensate. Test Solution 2 contained the addition of human-generated hygiene wastewater to the solution 1 waste stream components. Approximately 1500 kg of total wastewater was processed through the CDS during testing. Respective recoveries per solution were 93.4 +/- 0.7 and 90.3 +/- 0.5%. The average specific energy of the system was calculated to be less than 130 W-hr/kg. The following paper provides detailed information and data on the performance of the CDS as challenged per the ELS distillation comparison test.

Description: This paper summarizes investigations conducted on different orthostatic intolerance protection garments. This paper emphasizes on the engineering and operational aspects of the project. The current Shuttle pneumatic Anti-G Suit or AGS at 25 mmHg (0.5 psi) and customized medical mechanical compressive garments (20-30 mmHg) were tested on human subjects. The test process is presented. The preliminary results conclude that mechanical compressive garments can ameliorate orthostatic hypotension in hypovolemic subjects. A mechanical compressive garment is light, small and works without external pressure gas source; however the current garment design does not provide an adjustment to compensate for the loss of mass and size in the lower torso during long term space missions. It is also difficult to don. Compression garments that do not include an abdominal component are less effective countermeasures than garments which do. An early investigation conducted by the Human Adaptation and Countermeasures Division at Johnson Space Center (JSC) has shown there is no significant difference between the protection function of the AGS (at 77 mmHg or 1.5 psi) and the Russian anti-g suit, Kentavr (at 25 mmHg or 0.5 psi). Although both garments successfully countered hypovolemia-induced orthostatic intolerance, the Kentavr provided protection by using lower levels of compression pressure. This more recent study with a lower AGS pressure shows that pressures at 20-30 mmHg is acceptable but protection function is not as effective as higher pressure. In addition, a questionnaire survey with flight crewmembers who used both AGS and Kentavr during different missions was also performed.

Description: Reactive distillation designs are considered to reduce the presence of volatile organic compounds in the purified water. Reactive distillation integrates a reactor with a distillation column. A review of the literature in this field has revealed a variety of functional reactive columns in industry. Wastewater may be purified by a combination of a reactor and a distiller (e.g., the EWRS or VPCAR concepts) or, in principle, through a design which integrates the reactor with the distiller. A review of the literature in reactive distillation has identified some different designs in such combinations of reactor and distiller. An evaluation of reactive distillation and reactive air stripping is presented with regards to the reduction of volatile organic compounds in the contaminated water and air. Among the methods presented, an architecture is presented for the evaluation of the simultaneous oxidation of organics in air and water. These and other designs are presented in light of potential improvements in power consumptions and air and water purities for architectures which include catalytic activity integrated into the water processor. In particular, catalytic oxidation of organics may be useful as a tool to remove contaminants that more traditional distillation and/or air stripping columns may not remove. A review of the current leading edge at the commercial level and at the research frontier in catalytically active materials is presented. Themes and directions from the engineering developments in catalyst design are presented conceptually in light of developments in the nanoscale chemistry of a variety of catalyst materials.

Description: With the new vision of space travel aimed at traveling back to the Moon and eventually to Mars, NASA is designing a new spacesuit glove. The purpose of this study was to baseline hand strength while wearing the current Extravehicular Activity (EVA) glove, the Phase VI. By varying the pressure in the glove, hand strength could be characterized as a function of spacesuit pressure. This finding is of extreme importance when evaluating missions that require varying suit pressures associated with different operations within NASA's current human spaceflight program, Constellation. This characterization fed directly into the derivation of requirements for the next EVA glove. This study captured three types of maximum hand strength: grip, lateral pinch, and pulp-2 pinch. All three strengths were measured under varying pressures and compared to a bare-hand condition. The resulting standardized data was reported as a percentage of the bare-hand strength. The first wave of tests was performed while the subjects, four female and four male, were wearing an Extravehicular Mobility Unit (EMU) suit supported by a suit stand. This portion of the test collected data from the barehand, suited unpressurized, and suited pressurized (4.3 psi) conditions. In addition, the effects of the Thermal Micrometeoroid Garment (TMG) on hand strength were examined, with the suited unpressurized and pressurized cases tested with and without a TMG. It was found that, when pressurized and with the TMG, the Phase VI glove reduced applied grip strength to a little more than half of the subject s bare-hand strength. The lateral pinch strength remained relatively constant while the pulp-2 pinch strength actually increased with pressure. The TMG was found to decrease maximum applied grip strength by an additional 10% for both pressurized and unpressurized cases, while the pinch strengths saw little to no change. In developing requirements based on human subjects, it is important to attempt to derive results that encompass the variation within the entire population. The current EMU does not accommodate humans at the extremes of the anthropometric spectrum. To account for this and to ensure that these requirements cover the population, another phase of testing will be conducted in a differential pressure glove box. This phase will focus on smaller females and very large males that do not have a properly fitted EMU suit. Instead, they would wear smaller or larger gloves and be tested in the glove box as a means to compare and contrast their strength capabilities against the EMU accommodated hand size subjects. The glove box s ability to change pressures easily will also allow for a wider range of glove pressures to be tested. Compared to the data collected on the subjects wearing the EMU suit, it is expected that there will be similar ratios to bare-hand. It is recommended that this topic be sent to the Physical Ergonomics Board for review.

Description: The National Aeronautics and Space Administration (NASA) mission to expand space exploration will return humans to the Moon with the goal of maintaining a long-term presence. One challenge that NASA will face returning to the Moon is managing the lunar regolith found on the Moon's surface, which will collect on extravehicular activity (EVA) suits and other equipment. Based on the Apollo experience, the issues astronauts encountered with lunar regolith included eye/lung irritation, and various hardware failures (seals, screw threads, electrical connectors and fabric contamination), which were all related to inadequate lunar regolith mitigation. A vacuum cleaner capable of detaching, transferring, and efficiently capturing lunar regolith has been proposed as a method to mitigate the lunar regolith problem in the habitable environment on lunar surface. In order to develop this vacuum, a modified "off-the-shelf" vacuum cleaner has been used to determine detachment efficiency, vacuum requirements, and optimal cleaning techniques to ensure efficient dust removal in habitable lunar surfaces, EVA spacesuits, and air exchange volume. During the initial development of the Lunar Surface System vacuum cleaner, systematic testing was performed with varying flow rates on multiple surfaces (fabrics and metallics), atmospheric (14.7 psia) and reduced pressures (10.2 and 8.3 psia), different vacuum tool attachments, and several vacuum cleaning techniques to determine the performance requirements for the vacuum cleaner. The data recorded during testing was evaluated by calculating percent removal, relative to the retained simulant on the tested surface. In addition, Scanning Electron Microscopy (SEM) imaging was used to determine particle size distribution retained on the surface. The scope of this paper is to explain the initial phase of vacuum cleaner development, including historical Apollo mission data, current state-of-the-art vacuum cleaner technology, and vacuum cleaner testing that has focused on detachment capabilities varying pressure environments.

Description: The International Space Station (ISS) crew compliment has increased in size from 3 to 6 crew members . In order to support this increase in crew on ISS, the United States on-orbit Segment (USOS) has been outfitted with a suite of regenerative Environmental Control and Life Support (ECLS) hardware including an Oxygen Generation System(OGS), Waste and Hygiene Compartment (WHC), and a Water Recovery System (WRS). The WRS includes the Urine Processor Assembly (UPA) and the Water Processor Assembly (WPA). With this additional life support hardware, the ISS has achieved full redundancy in its on-orbit life support system between the USOS and Russian Segment (RS). The additional redundancy created by the Regenerative ECLS hardware creates the opportunity for independent support capabilities between segments, and for the first time since the start of ISS, the necessity to revise Life Support strategy agreements. Independent operating strategies coupled with the loss of the Space Shuttle supply and return capabilities in 2010 offer new and unique challenges. This paper will discuss the evolution of the ISS Life Support hardware strategy in support of 6-Crew on ISS, as well as the continued work that is necessary to ensure the support of crew and ISS Program objectives through the life of station.

Description: Two fluid life tests have been conducted to evaluate propylene glycol-based fluids for use in Constellation habitats and vehicles. The first test was conducted from November 2008 to January 2009 to help determine the compatibility of the propylene glycol-based fluid selected for Orion at the time. When the first test uncovered problems with the fluid selection, an investigation and selection of a new fluid were conducted. A second test was started in March 2010 to evaluate the new selection. For the first test, the fluid was subjected to a thermal fluid loop that had flight-like properties, as compared to Orion. The fluid loop had similar wetted materials, temperatures, flow rates, and aluminum wetted surface area to fluid volume ratio. The test was designed to last for 10 years, the life expectancy of the lunar habitat. However, the test lasted less than two months. System filters became clogged with precipitate, rendering the fluid system inoperable. Upon examination of the precipitate, it was determined that the precipitate composition contained aluminum, which could have only come from materials in the test stand, as aluminum is not part of the original fluid composition. Also, the fluid pH was determined to have increased from 10.1, at the first test sample, to 12.2, at the completion of the test. This high of a pH is corrosive to aluminum and was certainly a contributing factor to the development of precipitate. Due to the problems encountered during this test, the fluid was rejected as a coolant candidate for Orion. A new propylene glycol-based fluid was selected by the Orion project for use in the Orion vehicle. The Orion project has conducted a series of screening tests to help verify that there will be no problems with the new fluid selection. To compliment testing performed by the Orion project team, a new life test was developed to test the new fluid. The new test bed was similar to the original test bed, but with some improvements based on experience gained from the earlier test bed. The surface area of both aluminum and nickel in the test bed were designed to be similar to that of the Orion fluid loop, since the Orion fluid loop was expected to have high concentrations of both metals in the system. Also, additional sample materials were added to the test bed to match recent updates to materials selections for Orion. At the time of this paper publication, approximately five months of testing will have been completed. This paper gives a status of the testing completed to date.

Description: This paper discusses the effort in evaluating and selecting a light weight impact protection material for the Constellation Space Suit Element (CSSE) Portable Life Support Subsystem (PLSS) conceptual packaging study. A light weight material capable of holding and protecting the components inside the PLSS is required to demonstrate the viability of the flexible PLSS packaging concept. The material needs to distribute, dissipate, and absorb the impact energy of the PLSS falling on the lunar surface. It must also be very robust and function in the extreme lunar thermal vacuum environment for up to one hundred Extravehicular Activity (EVA) missions. This paper documents the performance requirements for selecting a foam protection material, and the methodologies for evaluating commercial off-the-shelf (COTS) foam protection materials. It also presents the materials properties test results and impact drop test results of the various foam materials evaluated in the study. The findings from this study suggest that a foam based flexible protection system is a viable solution for PLSS packaging. However, additional works are needed to optimize COTS foam properties or to develop a composite foam system that will meet all the performance requirements for the CSSE PLSS flexible packaging.

Description: The International Space Station (ISS) United States Operational Segment (USOS) has four permanent rack sized ISS Crew Quarters (CQ) providing a private crewmember space. The CQ uses Node 2 cabin air for ventilation/thermal cooling, as opposed to conditioned ducted air from the ISS Temperature Humidity Control System or the ISS fluid cooling loop connections. Consequently, CQ can only increase the air flow rate to reduce the temperature delta between the cabin and the CQ interior. However, increasing airflow causes increased acoustic noise so efficient airflow distribution is an important design parameter. The CQ utilized a two fan push-pull configuration to ensure fresh air at the crewmember s head position and reduce acoustic exposure. The CQ interior needs to be below Noise Curve 40 (NC-40). The CQ ventilation ducts are open to the significantly louder Node 2 cabin aisle way which required significantly acoustic mitigation controls. The design implementation of the CQ ventilation system and acoustic mitigation are very inter-related and require consideration of crew comfort balanced with use of interior habitable volume, accommodation of fan failures, and possible crew uses that impact ventilation and acoustic performance. This paper illustrates the types of model analysis, assumptions, vehicle interactions, and trade-offs required for CQ ventilation and acoustics. Additionally, on-orbit ventilation system performance and initial crew feedback is presented. This approach is applicable to any private enclosed space that the crew will occupy.

Description: In order to control system and component temperatures, many spacecraft thermal control systems use a radiator coupled with a pumped fluid loop to reject waste heat from the vehicle. Since heat loads and radiation environments can vary considerably according to mission phase, the thermal control system must be able to vary the heat rejection. The ability to "turn down" the heat rejected from the thermal control system is critically important when designing the system.. Electrochromic technology as a radiator coating is being investigated to vary the amount of heat being rejected by a radiator. Coupon level tests were performed to test the feasibility of the technology. Furthermore, thermal math models were developed to better understand the turndown ratios required by full scale radiator architectures to handle the various operation scenarios during a mission profile for Altair Lunar Lander. This paper summarizes results from coupon level tests as well as thermal math models developed to investigate how electrochromics can be used to provide the largest turn down ratio for a radiator. Data from the various design concepts of radiators and their architectures are outlined. Recommendations are made on which electrochromic radiator concept should be carried further for future thermal vacuum testing.

Description: The Trace Contaminant Control System (TCCS), located within the ventilation loop of the Portable Life Support System (PLSS) of the Constellation Space Suit Element (CSSE), is responsible for removing hazardous trace contaminants from the space suit ventilation flow. This paper summarizes the results of a trade study that evaluated if trace contaminant control could be accomplished without a TCCS, relying on suit leakage, ullage loss from the carbon dioxide and humidity control system, and other factors. Trace contaminant generation rates were revisited to verify that values reflect the latest designs for CSSE pressure garment materials and PLSS hardware. Additionally, TCCS sizing calculations were performed and a literature survey was conducted to review the latest developments in trace contaminant technologies.

Description: NASA s Constellation Program includes the Orion, Altair, and Lunar Surface Systems project offices. The first two elements, Orion and Altair, are manned space vehicles while the third element is broader and includes several subelements including Rovers and a Lunar Habitat. The upcoming planned missions involving these systems and vehicles include several risks and design challenges. Due to the unique thermal environment, many of these risks and challenges are associated with the vehicles thermal control system. NASA s Exploration Systems Mission Directorate (ESMD) includes the Exploration Technology Development Program (ETDP). ETDP consists of several technology development projects. The project chartered with mitigating the aforementioned risks and design challenges is the Thermal Control System Development for Exploration Project. The risks and design challenges are addressed through a rigorous technology development process that culminates with an integrated thermal control system test. The resulting hardware typically has a Technology Readiness Level (TRL) of six. This paper summarizes the development efforts being performed by the technology development project. The development efforts involve heat acquisition and heat rejection hardware including radiators, heat exchangers, and evaporators. The project has also been developing advanced phase change material heat sinks and performing assessments for thermal control system fluids. The current paper will provide an update to a similar overview paper published at last year s International Conference on Environmental Systems (ICES).

Description: A space suit s mobility is critical to an astronaut s ability to perform work efficiently. As mobility increases, the astronaut can perform tasks for longer durations with less fatigue. The term mobility, with respect to space suits, is defined in terms of two key components: joint range of motion and joint torque. Individually these measures describe the path which in which a joint travels and the force required to move it through that path. Previous space suits mobility requirements were defined as the collective result of these two measures and verified by the completion of discrete functional tasks. While a valid way to impose mobility requirements, such a method does necessitate a solid understanding of the operational scenarios in which the final suit will be performing. Because the Constellation space suit system requirements are being finalized with a relatively immature concept of operations, the Space Suit Element team elected to define mobility in terms of its constituent parts to increase the likelihood that the future pressure garment will be mobile enough to enable a broad scope of undefined exploration activities. The range of motion requirements were defined by measuring the ranges of motion test subjects achieved while performing a series of joint maximizing tasks in a variety of flight and prototype space suits. The definition of joint torque requirements has proved more elusive. NASA evaluated several different approaches to the problem before deciding to generate requirements based on unmanned joint torque evaluations of six different space suit configurations being articulated through 16 separate joint movements. This paper discusses the experiment design, data analysis and results, and the process used to determine the final values for the Constellation pressure garment joint torque requirements.

Description: The Sublimator Driven Coldplate (SDC) is a unique piece of thermal control hardware that has several advantages over a traditional thermal control scheme. The principal advantage is the possible elimination of a pumped fluid loop, potentially increasing reliability and reducing complexity while saving both mass and power. Furthermore, the Integrated Sublimator Driven Coldplate (ISDC) concept couples a coolant loop with the previously described SDC hardware. This combination allows the SDC to be used as a traditional coldplate during long mission phases. The previously developed SDC technology cannot be used for long mission phases due to the fact that it requires a consumable feedwater for heat rejection. Adding a coolant loop also provides for dissimilar redundancy on the Altair Lander ascent module thermal control system, which is the target application for this technology. Tests were performed on an Engineering Development Unit at NASA s Johnson Space Center to quantify and assess the performance of the SDC. Correlated thermal math models were developed to help explain the test data. The paper also outlines the preliminary results of an ISDC concept being developed.

Description: An EMU water processing kit (Airlock Coolant Loop Recovery A/L CLR) was developed as a corrective action to Extravehicular Mobility Unit (EMU) coolant flow disruptions experienced on the International Space Station (ISS) in May of 2004 and thereafter. Conservative schedules for A/L CLR use and component life were initially developed and implemented based on prior analysis results and analytical modeling. The examination of postflight samples and EMU hardware in November of 2006 indicated that the A/L CLR kits were functioning well and had excess capacity that would allow a relaxation of the initially conservative schedules of use and component life. A relaxed use schedule and list of component lives was implemented thereafter. Since the adoption of the relaxed A/L CLR schedules of use and component lives, several A/L CLR kit components, transport loop water samples and sensitive EMU transport loop components have been examined to gage the impact of the relaxed requirements. The intent of this paper is to summarize the findings of that evaluation, and to outline updated schedules for A/L CLR use and component life.

Description: For future lunar extravehicular activities (EVA), one method under consideration for rejecting crew and electronics heat involves evaporating water through a hydrophobic, porous Teflon(Registered Trademark) membrane. A Spacesuit Water Membrane Evaporator (SWME) prototype using this membrane was successfully tested by Ungar and Thomas (2001) with predicted performance matching test data well. The above referenced work laid the foundation for the design of a compact sheet membrane SWME development unit for use in the Constellation System Spacesuit Element Portable Life Support System (Vogel and et. al., ICES 2008). Major design objectives included minimizing mass, volume, and manufacturing complexity while rejecting a minimum of 810 watts of heat from water flowing through the SWME at 91 kg/hr with an inlet temperature of 291K. The design meeting these objectives consisted of three concentric cylindrical water channels interlaced with four water vapor channels. Two units were manufactured for the purpose of investigating manufacturing techniques and performing thermal testing. The extensive thermal test measured SWME heat rejection as a function of water inlet temperatures, water flow-rates, water absolute pressures, water impurities, and water vapor back-pressures. This paper presents the test results and subsequent analysis, which includes a comparison of SWME heat rejection measurements to pretest predictions. In addition, test measurements were taken such that an analysis of the commercial-off-the-shelf vapor pressure control valve could be performed.

Description: The Constellation Program requires the development of a space suit system to meet new requirements for launch, entry, and abort crew survival functions, microgravity intravehicular and extravehicular activities, and lunar surface exploration. This paper summarizes recent work and the current status of the NASA Constellation Space Suit Element Pressure Garment and Crew Survival Subsystem (PG/CS). The emphasis of the work by the PGS/CS team has been in the areas of feasibility studies toward PGS/CS architecture definition, risk mitigation, and requirements development. Included are results from component level engineering studies, testing in the Orion Vehicle and Orion seat mockups, microgravity testing on the Reduced Gravity Aircraft, occupant protection sled testing, analyses and studies, and their implications on Constellation PG/CS subsystem.

Description: The development of the Advanced Life Support (ALS) Sizing Analysis Tool (ALSSAT) using Microsoft(Registered TradeMark) Excel was initiated by the Crew and Thermal Systems Division (CTSD) of Johnson Space Center (JSC) in 1997 to support the ALS and Exploration Offices in Environmental Control and Life Support System (ECLSS) design and studies. It aids the user in performing detailed sizing of the ECLSS for different combinations of the Exploration Life support (ELS) regenerative system technologies. This analysis tool will assist the user in performing ECLSS preliminary design and trade studies as well as system optimization efficiently and economically. The latest ALSSAT related publication in ICES 2004 detailed ALSSAT s development status including the completion of all six ELS Subsystems (ELSS), namely, the Air Management Subsystem, the Biomass Subsystem, the Food Management Subsystem, the Solid Waste Management Subsystem, the Water Management Subsystem, and the Thermal Control Subsystem and two external interfaces, including the Extravehicular Activity and the Human Accommodations. Since 2004, many more regenerative technologies in the ELSS were implemented into ALSSAT. ALSSAT has also been used for the ELS Research and Technology Development Metric Calculation for FY02 thru FY06. It was also used to conduct the Lunar Outpost Metric calculation for FY08 and was integrated as part of a Habitat Model developed at Langley Research Center to support the Constellation program. This paper will give an update on the analysis tool s current development status as well as present the analytical results of one of the trade studies that was performed.

Description: NASA is planning to return to the moon and then explore Mars. A permanent base at the south pole of the moon will be the test bed for Mars. At the moon base, two crewmembers are expected to conduct Extravehicular Activity (EVA) six days every week. Current spacesuits are cooled by the sublimation of water ice into vacuum. A single 7 hour EVA near the lunar equator in daylight can expend up to 5 kilograms of water. Because of the high cost of transporting spacesuit cooling water to the moon, the water for one EVA could cost hundreds of thousands of dollars. The lunar south pole and Mars have low surface temperatures that make cooling much easier than at the lunar equator. Alternate cooling methods and keeping to cool environments can reduce or eliminate the loss of water for spacesuit cooling. If cooling water is not needed, a recycling life support system can provide all the required crew water and oxygen without transporting additional water from Earth.

Description: A habitable atmosphere is a fundamental requirement for human spaceflight. To meet such a requirement, the cabin atmosphere must be constantly scrubbed to maintain human life and system functionality. The primary system for atmospheric scrubbing of the US on-orbit segment (USOS) of the International Space Station (ISS) is the Trace Contaminant Control System (TCCS). As part of the Environmental Control and Life Support Systems (ECLSS) atmosphere revitalization rack in the US Lab, the TCCS operates continuously, scrubbing trace contaminants generated primarily by two sources: the metabolic offgassing of crew members and the offgassing of equipment in the ISS. It has been online for approximately 95% since activated in February 2001. The TCCS is comprised of a charcoal bed, a catalytic oxidizer, and a lithium hydroxide post-sorbent bed, all of which are designed to be replaced onorbit when necessary. In 2006, all three beds were replaced following an observed increase in the system resistance that occurred over a period several months. The beds were returned to ground and subjected to a test, teardown and evaluation to investigate the root cause(s) of the decrease in flow rate through the system. In addition, various chemical and physical analyses of the bed materials were performed to determine contaminant loading and any changes in performance. This paper will mainly focus on the results of these analyses and how this correlates with what has been observed from archival sampling and onorbit events. This may provide insight into the future performance of the TCCS and rate of change for orbital replacement units in the TCCS.

Description: NASA is currently engaged in an activity to facilitate effective operations on the International Space Station (ISS) after the Space Shuttle retires. Currently, the Space Shuttle delivers crew and cargo to and from ISS. The Space Shuttle provides the only large scale method of hardware return from ISS to the ground. Hardware that needs to be periodically repaired, refurbished, or recalibrated must come back from ISS on the Shuttle. One example of NASA flight hardware that is used on ISS and refurbished on the ground is the Compound Specific Analyzer for Oxygen (CSA-O2). The CSA-O2 is an electrochemical sensor that is used on orbit for about 12 months (depending on Shuttle launch schedules), then returned to the ground for sensor replacement. The shuttle is scheduled to retire in 2010, and the ISS is scheduled to operate until 2016. NASA needs a hand held sensor that measures oxygen in the ISS environment and has a 5-10 year service life. After conducting a survey of oxygen sensor systems, NASA selected a Tunable Diode Laser Absorption Spectrometer (TDLAS) as the method of measurement that best addresses the needs for ISS. These systems are compact, meet ISS accuracy requirements, and because they use spectroscopic techniques, the sensors are not consumed or altered after making a measurement. TDLAS systems have service life ratings of 5-10 years, based on the lifetime of the laser. NASA is engaged in modifying a commercially available sensor, the Vaisala OMT 355, for the ISS application. The Vaisala OMT 355 requires three significant modifications to meet ISS needs. The commercial sensor uses a wall mount power supply, and the ISS sensor needs to use a rechargeable battery as its source of power. The commercial sensor has a pressure correction setpoint: the sensor can be adjusted to operate at reduced pressure conditions, but the sensor does not self correct dynamically and automatically. The ISS sensor needs to operate in the airlock, and make accurate measurements in an environment that can change from 14.7 psia to 10.2 psia in 15 minutes. The commercial sensor needs to be repackaged into a configuration that is more compact, and better suited for ISS airlock operations. NASA has recently completed a prototype of the reconfigured system. The unit has been repackaged in a way that the optical path of the spectrometer is unchanged, but the electronics has been integrated into a case measuring 10.7 X 7.2 X 3.0 inches. Two flight qualified rechargeable batteries have been integrated into system. The batteries can power the sensor for 10 hours on a single charge. A pressure sensor has been added to the system. The modified unit automatically compensates for changes in pressure, and meets 0.2% accuracy requirements for oxygen measurements in an environment with 18 to 32% oxygen across a pressure range of 10.0 to 15.0 psia.

Description: The design and mass cost of a starship and its life support system are investigated. The mission plan for a multi generational interstellar voyage to colonize a new planet is used to describe the starship design, including the crew habitat, accommodations, and life support. Only current technology is assumed. Highly reliable life support systems can be provided with reasonably small additional mass, suggesting that they can support long duration missions. Bioregenerative life support, growing crop plants that provide food, water, and oxygen, has been thought to need less mass than providing stored food for long duration missions. The large initial mass of hydroponics systems is paid for over time by saving the mass of stored food. However, the yearly logistics mass required to support a bioregenerative system exceeds the mass of food solids it produces, so that supplying stored dehydrated food always requires less mass than bioregenerative food production. A mixed system that grows about half the food and supplies the other half dehydrated has advantages that allow it to breakeven with stored dehydrated food in about 66 years. However, moderate increases in the hydroponics system mass to achieve high reliability, such as adding spares that double the system mass and replacing the initial system every 100 years, increase the mass cost of bioregenerative life support. In this case, the high reliability half food growing, half food supplying system does not breakeven for 389 years. An even higher reliability half and half system, with three times original system mass and replacing the system every 50 years, never breaks even. Growing food for starship life support requires more mass than providing dehydrated food, even for multigeneration voyages of hundreds of years. The benefits of growing some food may justify the added mass cost. Much more efficient recycling food production is wanted but may not be possible. A single multigenerational interstellar voyage to colonize a new planet would have cost similar to that of the Apollo program. Cost is reduced if a small crew travels slowly and lands with minimal equipment. We can go to the stars!

Description: CO2 acquisition and utilization technologies will have a vital role in determining sustained and affordable life support and in-situ fuel production architectures for human and robotic exploration of Moon and Mars. For long-term human exploration to be practical, reliable technologies have to be implemented to capture and chemically reduce the metabolic CO2 from the cabin air to restitute oxygen consumption. Technologies that facilitate the in-situ capture and conversion of atmospheric CO2 to fuel are essential for a viable human mission to Mars and their demonstration on the moon is critical as well. This paper describes the concept and experimental investigation of a CO2 capture and reduction system that comprises an adsorption compressor and a CO2 and steam co-electrolysis unit. The process products include oxygen for life support and Syngas (CO and H2) for synthetic fuel production. Electrochemical performance in terms of CO2 conversion, oxygen production, and power consumption of a system with a capacity to process 1kg CO2 per day (1-person equivalent) will be discussed.

Description: Three NASA centers are working together to address the challenge of operating robotic assets in support of human exploration of the Moon. This paper describes the combined work to date of the Ames Research Center (ARC), Jet Propulsion Laboratory (JPL) and Johnson Space Center (JSC) on a common support framework to control and monitor lunar robotic assets. We discuss how we have addressed specific challenges including time-delayed operations, and geographically distributed collaborative monitoring and control, to build an effective architecture for integrating a heterogeneous collection of robotic assets into a common work. We describe the design of the Robot Application Programming Interface Delegate (RAPID) architecture that effectively addresses the problem of interfacing a family of robots including the JSC Chariot, ARC K-10 and JPL ATHLETE rovers. We report on lessons learned from the June 2008 field test in which RAPID was used to monitor and control all of these assets. We conclude by discussing some future directions to extend the RAPID architecture to add further support for NASA's lunar exploration program.

Description: An assessment of APNASA was conducted at NASA Glenn Research Center under the Fundamental Aeronautics Program to determine their predictive capabilities. The geometry selected for this study was Stage 35 which is a single stage transonic compressor. A speedline at 100% speed was generated and compared to experimental data at 100% speed for two turbulence models. Performance of the stage at 100% speed and profiles of several key aerodynamic parameters are compared to the survey data downstream of the stator in this report. In addition, hub leakage was modeled and compared to solutions without leakage and the available experimental data.

Description: The Shear History Extensional Rheology Experiment (SHERE) is an International Space Station (ISS) glovebox experiment designed to study the effect of preshear on the transient evolution of the microstructure and viscoelastic tensile stresses for monodisperse dilute polymer solutions. The SHERE experiment hardware was launched on Shuttle Mission STS-120 (ISS Flight 10A) on October 22, 2007, and 20 fluid samples were launched on Shuttle Mission STS-123 (ISS Flight 10/A) on March 11, 2008. Astronaut Gregory Chamitoff performed experiments during Increment 17 on the ISS between June and September 2008. A summary of the ten year history of the hardware development, the experiment's science objectives, and Increment 17's flight operations are discussed in the paper. A brief summary of the preliminary science results is also discussed.

Description: An ideal pulse tube cryocooler using an ideal gas can operate at any temperature. This is not true for real gases. The enthalpy flow resulting from the real gas effects of 3He, 4He, and their mixtures in ideal pulse tube cryocoolers puts limits on the operating temperature of pulse tube cryocoolers. The discussion of these effects follows a previous description of the real gas effects in ideal pulse tube cryocoolers and makes use of models of the thermophysical properties of 3He and 4He. Published data is used to extend the analysis to mixtures of 3He and 4He. The analysis was done for pressures below 2 MPa and temperatures below 2.5 K. Both gases and their mixtures show low temperature limits for pulse tube cryocoolers. These limits are in the 0.5-2.2 K range and depend on pressure and mixture. In some circumstances, even lower temperatures may be possible. Pulse tube cryocoolers using the ha-fluid properties of dilute 3He in superfluid 4He appear to have no limit.

Description: Resonant effects and energy dissipation due to sloshing fuel inside propellant tanks are problems that arise in the initial design of any spacecraft or launch vehicle. A faster and more reliable method for calculating these effects during the design stages is needed. Using Computational Fluid Dynamics (CFD) techniques, a model of these fuel tanks can be created and used to predict important parameters such as resonant slosh frequency and damping rate. This initial study addresses the case of free surface slosh. Future studies will focus on creating models for tanks fitted with propellant management devices (PMD) such as diaphragms and baffles.

Description: Launch operations engineers at the Kennedy Space Center have identified an Integrated Refrigeration and Storage system as a promising technology to reduce launch costs and enable advanced cryogenic operations. This system uses a close cycle Brayton refrigerator to remove energy from the stored cryogenic propellant. This allows for the potential of a zero loss storage and transfer system, as well and control of the state of the propellant through densification or re-liquefaction. However, the behavior of the fluid in this type of system is different than typical cryogenic behavior, and there will be a learning curve associated with its use. A 400 liter research cryostat has been designed, fabricated and delivered to KSC to test the thermo fluid behavior of liquid oxygen as energy is removed from the cryogen by a simulated DC cycle cryocooler. Results of the initial testing phase focusing on heat exchanger characterization and zero loss storage operations using liquid oxygen are presented in this paper. Future plans for testing of oxygen densification tests and oxygen liquefaction tests will also be discussed. KEYWORDS: Liquid Oxygen, Refrigeration, Storage

Description: A computational fluid dynamics (CFD) method is adapted, validated and applied to spinning gear systems with emphasis on predicting windage losses. Several spur gears and a disc are studied. The CFD simulations return good agreement with measured windage power loss. Turbulence modeling choices, the relative importance of viscous and pressure torques with gear speed and the physics of the complex 3-D unsteady flow field in the vicinity of the gear teeth are studied.

Description: No abstract available

Description: As the provision of potable water is critical for successful habitation of the International Space Station (ISS), life support systems were installed in December 2008 to recycle both humidity from the atmosphere and urine to conserve available water in the vehicle. Pre-consumption testing from the dispensing needle at the Potable Water Dispenser (PWD) indicated that bacterial concentrations exceeded the current ISS specifications of 50 colony forming units (CFU) per ml. Subsequent investigations revealed that a corrugated stainless steel flex hose upstream of the dispensing needle in the PWD was filled with non-sterile water and left at room temperature for over one month before launch. To simulate biofilm formation that was suspected in the flight system, sterile flex hoses were seeded with a consortium of bacterial isolates previously recovered from other ISS water systems, which included Ralstonia pickettii, Burkholderia multivorans, Caulobacter vibrioides., and Cupriavidus pauculus. After 5 days of incubation, these hoses were challenged with various chemical disinfectants including hydrogen peroxide, colloidal silver, and buffered pH solutions to determine the ability of the disinfectants to decrease and maintain bacterial concentrations below ISS specifications. Disinfection efficacy over time was measured by collecting daily heterotrophic plate counts following exposure to the disinfectants. A single flush with either 6% hydrogen peroxide solution or a mixture of 3% hydrogen peroxide and 400 ppb colloidal silver effectively reduced the bacterial concentrations to less than 1 CFU/ml for a period of up to 2 months. Testing results indicated that hydrogen peroxide and mixtures of hydrogen peroxide and colloidal silver have tremendous potential as alternative disinfectants for ISS water systems.

Description: The International Space Station (ISS) Environmental Control and Life Support (ECLS) system includes regenerative and non-regenerative technologies that provide the basic life support functions to support the crew, while maintaining a safe and habitable shirtsleeve environment. This paper provides a summary of the U.S. ECLS system activities over the past year, covering the period of time between March 2009 and February 2010. The ISS continued permanent crew operations, with the start of Phase 3 of the ISS Assembly Sequence and an increase of the ISS crew size from three to six. Work continues on the last of the Phase 3 pressurized elements.

Description: Accidents and incidents show that high workload-induced stress and poor teamwork skills lead to performance decrements and errors. Research on teamwork shows that effective teams are able to adapt to stressful situations, and to reduce workload by using successful strategies for communication and decision making, and through dynamic redistribution of tasks among team members. Furthermore, superior teams are able to recognize signs and symptoms of workload-induced stress early, and to adapt their coordination and communication strategies to the high workload, or stress conditions. Mission Control Center (MCC) teams often face demanding situations in which they must operate as an effective team to solve problems with crew and vehicle during onorbit operations. To be successful as a team, flight controllers (FCers) must learn effective teamwork strategies. Such strategies are the focus of Space Flight Resource Management (SFRM) training. SFRM training in MOD has been structured to include some classroom presentations of basic concepts and case studies, with the assumption that skill development happens in mission simulation. Integrated mission simulations do provide excellent opportunities for FCers to practice teamwork, but also require extensive technical knowledge of vehicle systems, mission operations, and crew actions. Such technical knowledge requires lengthy training. When SFRM training is relegated to integrated simulations, FCers can only practice SFRM after they have already mastered the technical knowledge necessary for these simulations. Given the centrality of teamwork to the success of MCC, holding SFRM training till late in the flow is inefficient. But to be able to train SFRM earlier in the flow, the training cannot rely on extensive mission-specific technical knowledge. Hence, the need for a generic SFRM training framework that would allow FCers to develop basic teamwork skills which are mission relevant, but without the required mission knowledge. Work on SFRM training has been conducted in collaboration with the Expedition Vehicle Division at the Mission Operations Directorate (MOD) and with United Space Alliance (USA) which provides training to Flight Controllers. The space flight resource management training work is part of the Human Factors in Training Directed Research Project (DRP) of the Space Human Factors Engineering (SHFE) Project under the Space Human Factors and Habitability (SHFH) Element of the Human Research Program (HRP). Human factors researchers at the Ames Research Center have been investigating team work and distributed decision making processes to develop a generic SFRM training framework for flight controllers. The work proposed for FY10 continues to build on this strong collaboration with MOD and the USA Training Group as well as previous research in relevant domains such as aviation. In FY10, the work focuses on documenting and analyzing problem solving strategies and decision making processes used in MCC by experienced FCers.

Description: In support of the Constellation Program, which will return humans to the moon and establish an Outpost, NASA has conducted an analysis of crew clothing and laundry options. Single-use or "disposable" clothing has been used from Apollo until International Space Station (ISS) missions, meaning that clothes were worn for the whole mission or thrown away when they became too dirty to wear any longer. This is justified for short duration missions; however, as the Constellation mission will last much longer and each individual Outpost mission is expected to last up to 180 days, mission goals and launch penalties for mass and volume may lead to a different conclusion. Furthermore, the habitat atmosphere pressure and therefore oxygen volume percentage will be different from ISS or Shuttle. Almost daily EVA sorties will be a norm during Outpost exploration missions. All of these factors will have impacts on selection of crew clothing and laundry options for Outpost missions. Mass and volume estimates for disposable crew clothing have been shown as a major penalty in long-duration manned space exploration missions in previous analyses. Assuming disposable clothing like ISS, Equivalent System Mass (ESM) of crew clothing and hygiene towels was estimated to be 11,000 kg or about 11% of total life support system ESM for a 10-year Lunar Outpost mission with 4 crew members. Ways to reduce this clothing penalty, which are discussed in this paper, include: a) Reduce clothing supply rate through using clothes made of advanced fabrics; b) Reduce daily usage rate by extending its use duration before disposing; and c) Use laundry and reusable clothing. The report summarizes recent research efforts in advanced clothing, proposed clothing supply rates for Exploration missions, results of a trade-off study between disposable clothing and laundry, and conclusions and suggestions for Constellation Program clothing.

Description: As the International Space Station (ISS) nears completion, and full international utilization is achieved, we are at a scientific crossroads. ISS is the premier location for research aimed at understanding the effects of microgravity on the human body. For applications to future human exploration, it is key for validation, quantification, and mitigation of a wide variety of spaceflight risks to health and human performance. Understanding and mitigating these risks is the focus of NASA s Human Research Program. However, NASA s approach to defining human research objectives is only one of many approaches within the ISS international partnership (including Roscosmos, the European Space Agency, the Canadian Space Agency, and the Japan Aerospace Exploration Agency). Each of these agencies selects and implements their own ISS research, with independent but related objectives for human and life sciences research. Because the science itself is also international and collaborative, investigations that are led by one ISS partner also often include cooperative scientists from around the world. The operation of the ISS generates significant additional data that is not directly linked to specific investigations. Such data comes from medical monitoring of crew members, life support and radiation monitoring, and from the systems that have been implemented to protect the health of the crew (such as exercise hardware). We provide examples of these international synergies in human research on ISS and highlight key early accomplishments that derive from these broad interfaces. Taken as a whole, the combination of diverse research objectives, operational data, international sharing of research resources on ISS, and scientific collaboration provide a robust research approach and capability that no one partner could achieve alone.

Description: During post-flight processing of STS-116, damage to crewmember Robert Curbeam's Phase VI Glove Thermal Micrometeoroid Garment was discovered. This damage consisted of: loss of RTV-157 palm pads on the thumb area on the right glove, a 0.75 inch cut in the Vectran adjacent to the seam and thumb pad (single event cut), constituting the worst glove damage ever recorded for the U.S. space program. The underlying bladder and restraint were found not be damaged by this event. Evaluation of glove damage found that the outer Vectran fibers were sliced as a result of contact with a sharp edge or pinch point rather than general wear or abrasion (commonly observed on the RTV pads). Damage to gloves was also noted on STS-118 and STS-120. One potential source of EMU glove damages are sharp crater lips on external handrails, generated by micrometeoroid and orbital debris (MMOD) impacts. In this paper, the results of a hypervelocity impact (HVI) test program on representative and actual ISS handrails are presented. These tests were performed in order to characterize impact damage profiles on ISS handrails and evaluate alternatives for limiting risk to future missions. It was determined that both penetrating and non-penetrating MMOD impacts on aluminum and steel ISS handrails are capable of generating protruding crater profiles which exceed the heights required for EMU glove abrasion risk by an order of magnitude. Testing demonstrated that flexible overwraps attached to the outside of existing handrails are capable of limiting contact between hazardous crater formations and crewmember gloves during extravehicular activity (EVA). Additionally, replacing metallic handrails with high strength, low ductility, fiber reinforced composite materials would limit the formation of protruding crater lips on new ISS modules.

Description: Although fitting a garment is often considered more of an art than a science, experts suggest that a subjectively poor fit is a symptom of inappropriate ease, the space between the wearer and the garment. The condition of poor suit fit is a unique problem for the space program and it can be attributed primarily to: a) NASA s policy to accommodate a wide variety of people (males and females from 1st to 99th percentile range and with various shapes and sizes) and b) its requirement to deploy a minimum number of suit sizes for logistical reasons. These factors make the space suit fit difficult to assess, where a wide range of people must be fit by the minimum possible number of suits, and yet, fit is crucial for operability and safety. Existing simplistic sizing scheme do not account for wide variations in shape within a diverse population with very limited sizing options. The complex issue of fit has been addressed by a variety of methods, many of which have been developed by the military, which has always had a keen interest in fitting its diverse population but with a multitude of sizing options. The space program has significantly less sizing options, so a combination of these advanced methods should be used to optimize space suit size and assess space suit fit. Multivariate methods can be used to develop sizing schemes that better reflect the wearer population, and integrated sizing systems can form a compromise between fitting men and women. Range of motion and operability testing can be combined with subjective feedback to provide a comprehensive evaluation of fit. The amount of ease can be tailored using these methods, to provide enough extra room where it is needed, without compromising mobility and comfort. This paper discusses the problem of fit in one of its most challenging applications: providing a safe and comfortable spacesuit that will protect its wearer from the extreme environment of space. It will discuss the challenges and necessity of closely fitting its potential wearers, a group of people from a broad spectrum of the population, and will detail some of the methods that can be employed to ensure and validate a good fit.

Description: An amine-based carbon dioxide (CO2) and water vapor sorbent in pressure-swing regenerable beds has been developed by Hamilton Sundstrand and baselined for the Orion Atmosphere Revitalization System (ARS). In three previous years at this conference, reports were presented on extensive Johnson Space Center (JSC) testing of this technology in a sea-level pressure environment with simulated and real human metabolic loads in both open and closed-loop configurations. The test article design was iterated a third time before the latest series of such tests, which was performed in the first half of 2009. The new design incorporates a canister configuration modification for overall unit compactness and reduced pressure drop, as well as a new process flow control valve that incorporates both compressed gas purge and dual-end vacuum desorption capabilities. This newest test article is very similar to the flight article designs. Baseline tests of the new unit were performed to compare its performance to that of the previous test articles. Testing of compressed gas purge operations helped refine launchpad operating condition recommendations developed in earlier testing. Operating conditions used in flight program computer models were tested to validate the model projections. Specific operating conditions that were recommended by the JSC test team based on past test results were also tested for validation. The effects of vacuum regeneration line pressure on resulting cabin conditions was studied for high metabolic load periods, and a maximum pressure is recommended.

Description: Every spacecraft atmosphere contains trace contaminants resulting from offgassing by cabin materials and human passengers. An amine-based carbon dioxide (CO2) and water vapor sorbent in pressure-swing regenerable beds has been developed by Hamilton Sundstrand and baselined for the Orion Atmosphere Revitalization System (ARS). Part of the risk mitigation effort for this new technology is the study of how atmospheric trace contaminants will affect and be affected by the technology. One particular area of concern is ammonia, which, in addition to the normal spacecraft sources, can also be off-gassed by the amine-based sorbent. In the first half of 2009, tests were performed with typical cabin atmosphere levels of five of the most common trace gases, most of which had not yet been tested with this technology. A subscale sample of the sorbent was exposed to each of the chemicals mixed into a stream of moist, CO2-laden air, and the CO2 adsorption capacity of the sorbent was compared before and after the exposure. After these typical-concentration chemicals were proven to have negligible effect on the subscale sample, tests proceeded on a full-scale test article in a sealed chamber with a suite of eleven contaminants. To isolate the effects of various test rig components, several extended-duration tests were run: without injection or scrubbing, with injection and without scrubbing, with injection and scrubbing by both the test article and dedicated trace contaminant filters, and with injection and scrubbing by only the test article. The high-level results of both the subscale and full-scale tests are examined in this paper.

Description: The gloved hand is an astronaut s primary means of interacting with the environment, so performance on an EVA is strongly impacted by any restrictions imposed by the glove. As a result, these restrictions have been the subject of study for decades. However, previous studies have generally been unsuccessful in quantifying glove mobility and tactility. Instead, studies have tended to focus on the dexterity, strength and functional performance of the gloved hand. Therefore, it has been difficult to judge the impact of each type of restriction on the glove s overall capability. The lack of basic information on glove mobility in particular, is related to the difficulty in instrumenting a gloved hand to allow an accurate evaluation. However, the current study aims at developing novel technological capabilities to provide metrics for mobility and tactility that can be used to assess the performance of a glove in a way that could enable designers and engineers to improve upon their current designs. A series of evaluations were performed in ungloved, unpressurized and pressurized (4.3 psi) conditions, to allow a comparison across pressures and to the baseline barehanded condition. In addition, a subset of the testing was also performed with the Thermal Micrometeoroid Garment (TMG) removed. This test case in particular provided some interesting insight into how much of an impact the TMG has on gloved mobility -- in some cases, as much as pressurization of the glove. Previous rule-of-thumb estimates had assumed that the TMG would have a much lower impact on mobility, while these results suggest that an improvement in the TMG could actually have a significant impact on glove performance. Similarly, tactility testing illustrated the impact of glove pressurization on tactility and provided insight on the design of interfaces to the glove. The metrics described in this paper have been used to benchmark the Phase VI EVA glove and to develop requirements for the next generation glove for the Constellation program.

Description: Computations are performed to investigate the effect of rocket control motors on flush air-data sensor systems. Such sensors are critical for the control of space vehicles during launch and re-entry, but are prone to interference from rocket motors, hypersonic-flow effects, etc. Computational analyses provide a means for studying these interference effects and exploring opportunities for mitigating them, either through design techniques or through appropriate processing of the sensor outputs. In the present work, the influence of rocket control motors on the nosecone flush air-data sensors of a launch-abort vehicle is studied. Particular attention is paid to the differential effect of various control-jet combinations on surface pressures. The relative effectiveness of inviscid, viscous, turbulent and two-phase-flow approximations in addressing this problem is also investigated.

Description: A method has been developed which integrates a fluid flow analyzer and a thermal analyzer to produce both steady state and transient results of 1-D, 2-D, and 3-D analysis models. The Generalized Fluid System Simulation Program (GFSSP) is a one dimensional, general purpose fluid analysis code which computes pressures and flow distributions in complex fluid networks. The MSC Systems Improved Numerical Differencing Analyzer (MSC.SINDA) is a one dimensional general purpose thermal analyzer that solves network representations of thermal systems. Both GFSSP and MSC.SINDA have graphical user interfaces which are used to build the respective model and prepare it for analysis. The SINDA/GFSSP Conjugate Integrator (SGCI) is a formbase graphical integration program used to set input parameters for the conjugate analyses and run the models. The contents of this paper describes SGCI and its thermo-fluids conjugate analysis techniques and capabilities by presenting results from some example models including the cryogenic chill down of a copper pipe, a bar between two walls in a fluid stream, and a solid plate creating a phase change in a flowing fluid.

Description: The Constellation Program (CxP) consists of spacecrafts, launch vehicles, and support systems to execute the Exploration Architecture. The Program is currently divided into three distinct phases. The first phase is to develop a vehicle to provide limited cargo resupply capability and allow crew member rotation to the International Space Station (ISS). The second phase is to support the return of humans to the moon. The final phase is currently envisioned to allow the delivery of humans and cargo to Mars for an extended time. To implement this phased approach the CxP is currently working on the first vehicle and support systems to replace the Space Shuttle and allow continued access to space. This paper provides a summary of the CxP Thermal and Environmental Control and Life Support (ECLS) work that that has occurred across the different parts of the Program in support of these three phases over the past year.

Description: Systems engineering could also be called the art of compromise. At its heart, systems engineering seeks to find that solution which maximizes the utility of the system, usually compromising the performance of each individual subsystem. While seemingly straightforward, systems engineering methodology is complicated when the utility to be maximized is unclear and the costs to each individual subsystem are not - or not easily - quantifiable. In this paper, we explore one such systems engineering problem within the Constellation Program as a case study in applied systems engineering. During suited operations, astronauts within Orion will be connected to an umbilical to receive and return breathing gas. The pressure drop associated with this umbilical must be overcome by the Orion vehicle. A smaller umbilical, which is desirable for crew operations, means a higher pressure drop, resulting in additional mass and power for the vehicle. We outline the technical considerations in the development of this integrated system and discuss the method by which we reached the ultimate solution. This paper, while just one example of the kind of problem solving that happens every day, offers insight into what happens when the theories of systems engineering are put into practice.

Description: Node 1 (Unity) flew to International Space Station (ISS) on Flight 2A. Node 1 was the first module of the United States On-Orbit Segment (USOS) launched to ISS. The Node 1 ISS Environmental Control and Life Support (ECLS) design featured limited ECLS capability. The main purpose of Node 1 was to provide internal storage by providing four stowage rack locations within the module and to allow docking of multiple modules and a truss segment to it. The ECLS subsystems inside Node 1 were routed through the element prior to launch to allow for easy integration of the attached future elements, particularly the Habitation Module which was planned to be located at the nadir docking port of Node 1. After Node 1 was on-orbit, the Program decided not to launch the Habitation Module and instead, to replace it with Node 3 (Tranquility). In 2007, the Program became concerned with a potential Russian docking port approach issue for the Russian FGB nadir docking port after Node 3 is attached to Node 1. To solve this concern the Program decided to relocate Node 3 from Node 1 nadir to Node 1 port. To support the movement of Node 3 the Program decided to build a modification kit for Node 1, an on-orbit feedthrough leak test device, and new vestibule jumpers to support the ECLS part of the relocation. This paper provides a design overview of the modification kit, a summary of the Node 1 ECLS re-verification to support the Node 3 relocation from Node 1 nadir to Node 1 port, and a status of the ECLS modification kit installation into Node 1.

Description: The Orion Crew Exploration Vehicle (CEV) is the first crew transport vehicle to be developed by the National Aeronautics and Space Administration (NASA) in the last thirty years. The CEV is being developed to transport the crew safely from the Earth to the Moon and back again. This year, the vehicle continued to go through design refinements to reduce weight, meet requirements, and operate reliably. Preliminary Design Review was performed and long lead procurement items were started. The design of the Orion Environmental Control and Life Support (ECLS) system, which includes the life support and active thermal control systems, is progressing through the design stage into manufacturing. This paper covers the Orion ECLS development from April 2009 to April 2010.

Description: NASA will build an outpost on the lunar surface for long-duration human habitation and research. The surface of the Moon is covered by a layer of fine, reactive dust, and the living quarters in the lunar outpost are expected to be contaminated by lunar dust. Because the toxicity of lunar dust is not known, NASA has tasked its toxicology laboratory to evaluate the risk of exposure to the dust and to establish safe exposure limits for astronauts working in the lunar habitat. Studies of the pulmonary toxicity of a dust are generally done first in rodents by intratracheal/intrapharyngeal instillation. This toxicity screening test is then followed by an inhalation study, which requires much more of the test dust and is labor intensive. Preliminary results obtained by examining lung lavage fluid from dust-treated mice show that lunar dust was somewhat toxic (more toxic than TiO2, but less than quartz dust). More extensive studies are in progress to further examine lung lavage fluid for biomarkers of toxicity and lung tissues for histopathological lesions in rodents exposed to aged and activated (ground) lunar dust samples. In these studies, reference dusts (TiO2 and quartz) of known toxicities and have industrial exposure limits will be studied in parallel so the relative toxicity of lunar dust can be determined. The results from the instillation studies will be useful for choosing exposure concentrations for the animal inhalation study. The animal inhalation exposure will be conducted with lunar dust simulant prior to the study with the lunar dust. The experiment with the simulate will ensure that the study techniques used with actual lunar dust will be successful. The results of instillation and inhalation studies will reveal the toxicological risk of exposures and are essential for setting exposure limits on lunar dust for astronauts living in the lunar habitat.

Description: A Metabolic heat-regenerated Temperature Swing Adsorption (MTSA) system is being developed for carbon dioxide, water and thermal control in a lunar and martian portable life support system (PLSS). A previous system analysis was performed to evaluate the impact of MTSA on PLSS design. That effort was Mars specific and assumed liquid carbon dioxide (LCO2) coolant made from martian resources. Transient effects were not considered but rather average conditions were used throughout the analysis. This effort takes into further consideration the transient effects inherent in the cycling MTSA system as well as assesses the use of water as coolant. Standard heat transfer, thermodynamic, and heat exchanger methods are presented to conduct the analysis. Assumptions and model verification are discussed. The tool was used to perform various system studies. Coolant selection was explored and takes into account different operational scenarios as the minimum bed temperature is driven by the sublimation temperature of the coolant (water being significantly higher than LCO2). From this, coolant mass is sized coupled with sorbent bed mass because MTSA adsorption performance decreases with increasing sublimation temperature. Reduction in heat exchanger performance and even removal of certain heat exchangers, like a recuperative one between the two sorbent beds, is also investigated. Finally, the coolant flow rate is varied over the cycle to determine if there is a more optimal means of cooling the bed from a mass perspective. Results of these studies and subsequent recommendations for system design are presented.

Description: This paper describes modeling methods for the three core components of a Metabolic heat regenerated Temperature Swing Adsorption (MTSA) subassembly: the sorbent bed, a sublimation (cooling) heat exchanger (SHX), and a condensing icing (warming) heat exchanger (CIHX). The primary function of the MTSA, removing carbon dioxide from a ventilation loop, is performed via the sorbent bed. The CIHX is used to heat the sorbent bed for desorption and to remove moisture from the ventilation loop while the SHX is alternately employed to cool the sorbent bed via sublimation of a spray of water at low pressure to prepare the reconditioned bed for the next cycle. This paper describes a system level model of the MTSA as developed in Thermal Desktop and SINDA/FLUINT including assumptions on geometry and physical phenomena, modeling methodology and relevant pa ra mete rizatio ns. Several areas of particular modeling interest are discussed. In the sorbent bed, capture of the translating CO2 saturation front and associated local energy and mass balance in both adsorbing and desorbing modes is covered. The CIHX poses particular challenges for modeling in SINDA/FLUINT as accounting for solids states in fluid submodels are not a native capability. Methods for capturing phase change and latent heat of ice as well as the transport properties across a layer of low density accreted frost are developed. This extended modeling capacity is applicable to temperatures greater than 258 K. To extend applicability to the minimum device temperature of 235 K, a method for a mapped transformation of temperatures from below the limit temperatures to some value above is given along with descriptions for associated material property transformations and the resulting impacts to total heat and mass transfer. Similar considerations are shown for the SHX along with assumptions for flow mechanics and resulting model methods for sublimation in a flow.

Description: Metabolic heat regenerated Temperature Swing Adsorption (MTSA) technology is being developed for thermal and carbon dioxide (CO 2) control for a future Portable Life Support System (PLSS), as well as water recycling. CO 2 removal and rejection is accomplished by driving a sorbent through a temperature swing of approximately 210 K to 280 K . The sorbent is cooled to these sub-freezing temperatures by a Sublimating Heat Exchanger (SHX) with liquid coolant expanded to sublimation temperatures. Water is the baseline coolant available on the moon, and if used, provides a competitive solution to the current baseline PLSS schematic. Liquid CO2 (LCO2) is another non-cryogenic coolant readily available from Martian resources which can be produced and stored using relatively low power and minimal infrastructure. LCO 2 expands from high pressure liquid (~5800 kPa) to Mars ambient (0.8 kPa) to produce a gas / solid mixture at temperatures as low as 156 K. Analysis and experimental work are presented to investigate factors that drive the design of a heat exchanger to effectively use this sink. Emphasis is given to enabling efficient use of the CO 2 cooling potential and mitigation of heat exchanger clogging due to solid formation. Minimizing mass and size as well as coolant delivery are also considered. The analysis and experimental work is specifically performed in an MTSA-like application to enable higher fidelity modeling for future optimization of a SHX design. In doing so, the work also demonstrates principles and concepts so that the design can be further optimized later in integrated applications (including Lunar application where water might be a choice of coolant).

Description: Freezable radiators offer an attractive solution to the issue of thermal control system scalability. As thermal environments change, a freezable radiator will effectively scale the total heat rejection it is capable of as a function of the thermal environment and flow rate through the radiator. Scalable thermal control systems are a critical technology for spacecraft that will endure missions with widely varying thermal requirements. These changing requirements are a result of the space craft s surroundings and because of different thermal loads during different mission phases. However, freezing and thawing (recovering) a radiator is a process that has historically proven very difficult to predict through modeling, resulting in highly inaccurate predictions of recovery time. This paper summarizes tests on three test articles that were performed to further empirically quantify the behavior of a simple freezable radiator, and the culmination of those tests into a full scale design. Each test article explored the bounds of freezing and recovery behavior, as well as providing thermo-physical data of the working fluid, a 50-50 mixture of DowFrost HD and water. These results were then used as a tool for developing correlated thermal model in Thermal Desktop which could be used for modeling the behavior of a full scale thermal control system for a lunar mission. The final design of a thermal control system for a lunar mission is also documented in this paper.

Description: The need to resolve new patterns of sensory feedback in altered gravitoinertial environments requires cognitive processes to develop appropriate reference frames for spatial orientation awareness. The purpose of this study was to examine deficits in spatial cognitive performance during adaptation to conflicting tilt-translation stimuli. Fourteen subjects were tilted within a lighted enclosure that simultaneously translated at one of 3 frequencies. Tilt and translation motion was synchronized to maintain the resultant gravitoinertial force aligned with the longitudinal body axis, resulting in a mismatch analogous to spaceflight in which the canals and vision signal tilt while the otoliths do not. Changes in performance on different spatial cognitive tasks were compared 1) without motion, 2) with tilt motion alone (pitch at 0.15, 0.3 and 0.6 Hz or roll at 0.3 Hz), and 3) with conflicting tilt-translation motion. The adaptation paradigm was continued for up to 30 min or until the onset of nausea. The order of the adaptation conditions were counter-balanced across 4 different test sessions. There was a significant effect of stimulus frequency on both motion sickness and spatial cognitive performance. Only 3 of 14 were able to complete the full 30 min protocol at 0.15 Hz, while 7 of 14 completed 0.3 Hz and 13 of 14 completed 0.6 Hz. There were no changes in simple visual-spatial cognitive tests, e.g., mental rotation or match-to-sample. There were significant deficits during 0.15 Hz adaptation in both accuracy and reaction time during a spatial reference task in which subjects are asked to identify a match of a 3D reoriented cube assemblage. Our results are consistent with antidotal reports of cognitive impairment that are common during sensorimotor adaptation with G-transitions. We conclude that these cognitive deficits stem from the ambiguity of spatial reference frames for central processing of inertial motion cues.

Description: The Small Pressurized Rover (SPR) is currently being carried as an integral part of the current Lunar Surface Architectures under consideration in the Constellation program. One element of the SPR is the suit port, the means by which the crew performs Extravehicular Activities (EVAs). Two suit port deliverables were produced in fiscal year 2008: an aft bulkhead mockup for functional integrated testing with the 1-G SPR mockup and a functional and pressurizable engineering unit. This paper focuses on the test results and lessons learned on the aft bulkhead mockup. The suit port aft bulkhead mockup was integrated with the mockup of the SPR cabin and chassis. It is located on the aft bulkhead of the SPR cabin structure and includes hatches, a locking mechanism, seals, interior and exterior suit don/doff aids, and exterior platforms to accommodate different crewmember heights. A lightweight mockup of the Mark III suit was tested with the suit port aft bulkhead mockup. There are several limitations to the suit port and mockup suits, and results of the suit port evaluation are presented and interpreted within the context of the limitations.

Description: The Constellation Space Suit Element (CSSE) is an integrated assembly made up of primarily a Pressure Garment System (PGS) and a Portable Life Support System (PLSS). The PLSS is further composed of an oxygen (O2) subsystem, a ventilation subsystem, and a thermal subsystem. This paper baselines a detailed schematic of the CSSE PLSS to provide a basis for current and future CSSE PLSS development efforts. Both context diagrams and detailed schematics describe the hardware components and overall functions for all three of the PLSS subsystems. Additionally, PLSS functions are presented for multiple operational scenarios as follows: 1) Nominal Extravehicular Activity (EVA) Mode; 2) Umbilical Modes; a) No Recharge, b) With Recharge; 3) Decompression Sickness (DCS) Treatment Mode; 4) Buddy Mode; 5) Secondary O2 Modes; a) Helmet Purge; b) Suit Purge; c) Operational; and 5) PLSS Removed Umbilical Mode. A performance modeling effort is being performed to provide a preliminary confirmation of this layout and the current state of the thermal hydraulic modeling efforts being conducted for the PLSS is presented. The goal of these efforts is to provide realistic simulations of the PLSS under various modes of operation. Modeling approaches and assumptions are discussed as well as component model descriptions. Results from the models are included that show PLSS operations at steady-state and transient conditions. Finally, conclusions and recommendations are offered that summarize results, identify PLSS design weaknesses uncovered during review of the analysis results, and propose areas for improvement to increase model fidelity and accuracy.

Description: The Orbiter delivers water to the International Space Station (ISS) in Contingency Water Containers (CWCs) on each flight to the ISS. These CWCs are routinely sampled during each mission to verify the quality of the delivered water. Of the 5 samples returned on STS118/ 13A.1 in August 2007, two exhibited microbial growth exceeding potable water acceptability limits and historical data by orders of magnitude . The microbe was identified as Wautersia species and an investigation was launched to find the source of the contamination. Since then, samples collected on subsequent flights indicated additional CWCs had the same bacteria, as well as several on-orbit systems. An investigation was launched to try to find and address the source of the bacterial contamination. This paper will discuss how Wautersia was found, what Wautersia is, the investigation, and resolution.

Description: In a vehicle constrained by mass and power, it is necessary to ensure that during the process of reducing hardware mass and power that the health and well being of the crew is not compromised in the design process. To that end, it is necessary to ensure that in the final phase of flight - recovery, that the crew core body temperature remains below the crew cognitive deficit set by the Constellation program. This paper will describe the models used to calculate the thermal environment of the spacecraft after splashdown as well as the human thermal model used to calculate core body temperature. Then the results of these models will be examined to understand the key drivers for core body temperature. Finally, the analysis results will be used to show that additional cooling capability must be added to the vehicle to ensure crew member health post landing.

Description: An amine-based carbon dioxide (CO2) and water vapor sorbent in pressure-swing regenerable beds has been developed by Hamilton Sundstrand and baselined for the Orion Atmosphere Revitalization System (ARS). In two previous years at this conference, reports were presented on extensive Johnson Space Center (JSC) testing of this technology in a sea-level pressure environment with simulated human metabolic loads. Another paper at this year s conference discusses similar testing with real human metabolic loads, including some closed-loop testing with emergency breathing masks. The Orion ARS is designed to also support extravehicular activity operations from a depressurized cabin. The next step in developmental testing at JSC was, therefore, to test this ARS technology in a typical closed space suit loop environment with low-pressure pure oxygen inside the process loop and vacuum outside the loop. This was the first instance of low-pressure oxygen loop testing of a new Orion ARS technology, and was conducted with simulated human metabolic loads in December 2008. The test investigated pressure drops through two different styles of prototype suit umbilical connectors and general swing-bed performance with both umbilical configurations as well as with a short jumper line installed in place of the umbilicals. Other interesting results include observations on the thermal effects of swing-bed operation in a vacuum environment and a recommendation of cycle time to maintain acceptable atmospheric CO2 and moisture levels.

Description: Dynamic simulation of the lunar outpost habitat life support was undertaken to investigate the impact of life support failures and to investigate responses. Some preparatory static analysis for the Lunar Outpost life support model, an earlier version of the model, and an investigation into the impact of Extravehicular Activity (EVA) were reported previously. (Jones, 2008-01-2184, 2008-01-2017) The earlier model was modified to include possible resupply delays, power failures, recycling system failures, and atmosphere and other material storage failures. Most failures impact the lunar outpost water balance and can be mitigated by reducing water usage. Food solids, nitrogen can be obtained only by resupply from Earth. The most time urgent failure is a lass of carbon dioxide removal capability. Life support failures might be survivable if effective operational solutions are provided in the system design.

Description: In this paper we describe the web services, processes, communication protocols and ad-hoc service chains utilized in the late summer and early fall 2007 Ikhana UAS response to the wildfires burning in southern California. Additionally, we describe the lessons learned that will be applied to the upcoming Global Hawk UAS Aura Satellite Validation Experiment planned for early 2009.

Description: Recycling life support systems can achieve ultra reliability by using spares to replace failed components. The added mass for spares is approximately equal to the original system mass, provided the original system reliability is not very low. Acceptable reliability can be achieved for the space shuttle and space station by preventive maintenance and by replacing failed units, However, this maintenance and repair depends on a logistics supply chain that provides the needed spares. The Mars mission must take all the needed spares at launch. The Mars mission also must achieve ultra reliability, a very low failure rate per hour, since it requires years rather than weeks and cannot be cut short if a failure occurs. Also, the Mars mission has a much higher mass launch cost per kilogram than shuttle or station. Achieving ultra reliable space life support with acceptable mass will require a well-planned and extensive development effort. Analysis must define the reliability requirement and allocate it to subsystems and components. Technologies, components, and materials must be designed and selected for high reliability. Extensive testing is needed to ascertain very low failure rates. Systems design should segregate the failure causes in the smallest, most easily replaceable parts. The systems must be designed, produced, integrated, and tested without impairing system reliability. Maintenance and failed unit replacement should not introduce any additional probability of failure. The overall system must be tested sufficiently to identify any design errors. A program to develop ultra reliable space life support systems with acceptable mass must start soon if it is to produce timely results for the moon and Mars.

Description: The International Space Station (ISS) crew compliment will be increasing in size from 3 to 6 crew members in the summer of 2009. In order to support this increase in crew on ISS, the United States on-orbit Segment (USOS) has been outfitted with a suite of regenerative Environmental Control and Life Support (ECLS) hardware including an Oxygen Generation System(OGS), Waste and Hygiene Compartment (WHC), and a Water Recovery System (WRS). The WRS includes the Urine Processor Assembly (UPA) and the Water Processor Assembly (WPA). A critical step in advancing to a 6Crew support capability on ISS is a full checkedout and verification of the Regenerative ECLS hardware. With a successful checkout, the ISS will achieve full redundancy in its onorbit life support system between the USOS and Russian Segment (RS). The additional redundancy created by the Regenerative ECLS hardware creates the opportunity for independent support capabilities between segments, and for the first time since the start of ISS, the necessity to revise Life Support strategy agreements. Independent operating strategies coupled with the loss of the Space Shuttle supply and return capabilities in 2010 offers additional challenges. These challenges create the need for a higher level of onorbit consumables reserve to ensure crewmember life support during a system failure. This paper will discuss the evolution of the ISS Life Support hardware strategy in support of 6Crew on ISS, as well as the continued work which will be necessary to ensure the support of crew and ISS Program objectives through the life of station.

Description: Projection photolithography has been used as a fabrication method for radial grove gratings. Use of photolithographic method for diffraction grating fabrication represents the most significant breakthrough in grating technology in the last 60 years, since the introduction of holographic written gratings. Unlike traditional methods utilized for grating fabrication, this method has the advantage of producing complex diffractive groove contours that can be designed at pixel-by-pixel level, with pixel size currently at the level of 45 45 nm. Typical placement accuracy of the grating pixels is 10 nm over 30 nm. It is far superior to holographic, mechanically ruled or direct e-beam written gratings and results in high spatial coherence and low spectral cross-talk. Due to the smooth surface produced by reactive ion etch, such gratings have a low level of randomly scattered light. Also, due to high fidelity and good surface roughness, this method is ideally suited for fabrication of radial groove gratings. The projection mask is created using a laser writer. A single crystal silicon wafer is coated with photoresist, and then the projection mask, with its layer of photoresist, is exposed for patterning in a stepper or scanner. To develop the photoresist, the fabricator either removes the exposed areas (positive resist) of the unexposed areas (negative resist). Next, the patterned and developed photoresist silicon substrate is subjected to reactive ion etching. After this step, the substrate is cleaned. The projection mask is fabricated according to electronic design files that may be generated in GDS file format using any suitable CAD (computer-aided design) or other software program. Radial groove gratings in off-axis grazing angle of incidence mount are of special interest for x-ray spectroscopy, as they allow achieving higher spectral resolution for the same grating area and have lower alignment tolerances than traditional in-plane grating scheme. This is especially critical for NASA Constellation- X project that will utilize hundreds of gratings all of which need to be precisely aligned for x-ray observation of space.

Description: A method has been developed for fabricating gratings on flat substrates, and then reproducing the groove pattern on a curved (concave or convex) substrate and a corresponding grating device. First, surface relief diffraction grating grooves are formed on flat substrates. For example, they may be fabricated using photolithography and reactive ion etching, maskless lithography, holography, or mechanical ruling. Then, an imprint of the grating is made on a deformable substrate, such as plastic, polymer, or other materials using thermoforming, hot or cold embossing, or other methods. Interim stamps using electroforming, or other methods, may be produced for the imprinting process or if the same polarity of the grating image is required. The imprinted, deformable substrate is then attached to a curved, rigid substrate using epoxy or other suitable adhesives. The imprinted surface is facing away from the curved rigid substrate. As an alternative fabrication method, after grating is imprinted on the deformable substrate as described above, the grating may be coated with thin conformal conductive layer (for example, using vacuum deposition of gold). Then the membrane may be mounted over an opening in a pressured vessel in a manner of a membrane on a drum, grating side out. The pressure inside of the vessel may be changed with respect to the ambient pressure to produce concave or convex membrane surface. The shape of the opening may control the type of the surface curvature (for example, a circular opening would create spherical surface, oval opening would create toroidal surface, etc.). After that, well-known electroforming methods may be used to create a replica of the grating on the concave or convex membrane. For example, the pressure vessel assembly may be submerged into an electro-forming solution and negative electric potential applied to the metal coated membrane using an insulated wire. Positive electric potential may be then applied to a nickel or other metal plate submerged into the same solution. Metal ions would transfer from the plate through the solution into the membrane, producing high fidelity metal replica of the grating on the membrane. In one variation, an adhesive may be deposited on the deformable substrate, and then cured without touching the rigid, curved substrate. Edges of the deformable substrate may be attached to the rigid substrate to ensure uniform deformation of the deformable substrate. The assembly may be performed in vacuum, and then taken out to atmospheric pressure conditions to ensure that no air is trapped between the deformable and rigid substrates. Alternatively, a rigid surface with complementary curvature to the rigid substrate may be used to ensure uniform adhesion of the deformable substrate to the rigid substrate. Liquid may be applied to the surface of the deformable substrate to uniformly distribute pressure across its surface during the curing or hardening of the adhesive, or the film may be pressed into the surface using a deformable object or surface. After the attachment is complete, the grooves may be coated with reflective or dielectric layers to improve diffraction efficiency.

Description: A promising method of detecting imminent tsunamis and estimating their destructive potential involves the use of Global Positioning System (GPS) data in addition to seismic data. Application of the method is expected to increase the reliability of global tsunami-warning systems, making it possible to save lives while reducing the incidence of false alarms. Tsunamis kill people every year. The 2004 Indian Ocean tsunami killed about 230,000 people. The magnitude of an earthquake is not always a reliable indication of the destructive potential of a tsunami. The 2004 Indian Ocean quake generated a huge tsunami, while the 2005 Nias (Indonesia) quake did not, even though both were initially estimated to be of the similar magnitude. Between 2005 and 2007, five false tsunami alarms were issued worldwide. Such alarms result in negative societal and economic effects. GPS stations can detect ground motions of earthquakes in real time, as frequently as every few seconds. In the present method, the epicenter of an earthquake is located by use of data from seismometers, then data from coastal GPS stations near the epicenter are used to infer sea-floor displacements that precede a tsunami. The displacement data are used in conjunction with local topographical data and an advanced theory to quantify the destructive potential of a tsunami on a new tsunami scale, based on the GPS-derived tsunami energy, much like the Richter Scale used for earthquakes. An important element of the derivation of the advanced theory was recognition that horizontal sea-floor motions contribute much more to generation of tsunamis than previously believed. The method produces a reliable estimate of the destructive potential of a tsunami within minutes typically, well before the tsunami reaches coastal areas. The viability of the method was demonstrated in computational tests in which the method yielded accurate representations of three historical tsunamis for which well-documented ground-motion measurements were available. Development of a global tsunami-warning system utilizing an expanded network of coastal GPS stations was under consideration at the time of reporting the information for this article.

Description: The Space Human Factors Engineering (SHFE) Standing Review Panel (SRP) evaluated 22 gaps and 39 tasks in the three risk areas assigned to the SHFE Project. The area where tasks were best designed to close the gaps and the fewest gaps were left out was the Risk of Reduced Safety and Efficiency dire to Inadequate Design of Vehicle, Environment, Tools or Equipment. The areas where there were more issues with gaps and tasks, including poor or inadequate fit of tasks to gaps and missing gaps, were Risk of Errors due to Poor Task Design and Risk of Error due to Inadequate Information. One risk, the Risk of Errors due to Inappropriate Levels of Trust in Automation, should be added. If astronauts trust automation too much in areas where it should not be trusted, but rather tempered with human judgment and decision making, they will incur errors. Conversely, if they do not trust automation when it should be trusted, as in cases where it can sense aspects of the environment such as radiation levels or distances in space, they will also incur errors. This will be a larger risk when astronauts are less able to rely on human mission control experts and are out of touch, far away, and on their own. The SRP also identified 11 new gaps and five new tasks. Although the SRP had an extremely large quantity of reading material prior to and during the meeting, we still did not feel we had an overview of the activities and tasks the astronauts would be performing in exploration missions. Without a detailed task analysis and taxonomy of activities the humans would be engaged in, we felt it was impossible to know whether the gaps and tasks were really sufficient to insure human safety, performance, and comfort in the exploration missions. The SRP had difficulty evaluating many of the gaps and tasks that were not as quantitative as those related to concrete physical danger such as excessive noise and vibration. Often the research tasks for cognitive risks that accompany poor task or information design addressed only part, but not all, of the gaps they were programmed to fill. In fact the tasks outlined will not close the gap but only scratch the surface in many cases. In other cases, the gap was written too broadly, and really should be restated in a more constrained way that can be addressed by a well-organized and complementary set of tasks. In many cases, the research results should be turned into guidelines for design. However, it was not clear whether the researchers or another group would construct and deliver these guidelines.

Description: Through-silicon via (TSV) is the latest in a progression of technologies for stacking silicon devices in three dimensions (3D). Driven by the need for improved performance, methods to use short vertical interconnects to replace the long interconnects found in 2D structures have been developed. The industry is moving past the feasibility (research and development [R and D]) phase for TSV technology into the commercialization phase where economic realities will determine which technologies are adopted. Low-cost fine via hole formation and highly reliable via filling technologies have been demonstrated; process equipment and materials are available. Even though design, thermal, and test issues remain, much progress has been made.

Description: In a Stirling radioisotope system, heat must continually be removed from the GPHS modules, to maintain the GPHS modules and surrounding insulation at acceptable temperatures. Normally, the Stirling convertor provides this cooling. If the Stirling convertor stops in the current system, the insulation is designed to spoil, preventing damage to the GPHS, but also ending the mission. An alkali-metal Variable Conductance Heat Pipe (VCHP) is under development to allow multiple stops and restarts of the Stirling convertor. The status of the ongoing effort in developing this technology is presented in this paper. An earlier, preliminary design had a radiator outside the Advanced Stirling Radioisotope Generator (ASRG) casing, used NaK as the working fluid, and had the reservoir located on the cold side adapter flange. The revised design has an internal radiator inside the casing, with the reservoir embedded inside the insulation. A large set of advantages are offered by this new design. In addition to reducing the overall size and mass of the VCHP, simplicity, compactness and easiness in assembling the VCHP with the ASRG are significantly enhanced. Also, the permanently elevated temperatures of the entire VCHP allows the change of the working fluid from a binary compound (NaK) to single compound (Na). The latter, by its properties, allows higher performance and further mass reduction of the system. Preliminary design and analysis shows an acceptable peak temperature of the ASRG case of 140 C while the heat losses caused by the addition of the VCHP are 1.8 W.

Description: An integrated waveguide-to-MMIC (monolithic microwave integrated circuit) chip operating in the 300-GHz range is designed to operate well on high-permittivity semiconductor substrates typical for an MMIC amplifier, and allows a wider MMIC substrate to be used, enabling integration with larger MMICs (power amplifiers). The waveguide-to- CBCPW (conductor-backed coplanar waveguide) transition topology is based on an integrated dipole placed in the E-plane of the waveguide module. It demonstrates low loss and good impedance matching. Measurement and simulation demonstrate that the loss of the transition and waveguide loss is less than 1-dB over a 340-to-380-GHz bandwidth. A transition is inserted along the propagation direction of the waveguide. This transition uses a planar dipole aligned with the maximum E-field of the TE10 waveguide mode as an inter face between the waveguide and the MMIC. Mode conversion between the coplanar striplines (CPS) that feed the dipole and the CBCPW transmission line is accomplished using a simple air-bridge structure. The bottom side ground plane is truncated at the same reference as the top-side ground plane, leaving the end of the MMIC suspended in air.

Description: The Supply-Chain Optimization Template (SCOT) is an instructional guide for identifying, evaluating, and optimizing (including re-engineering) aerospace- oriented supply chains. The SCOT was derived from the Supply Chain Council s Supply-Chain Operations Reference (SCC SCOR) Model, which is more generic and more oriented toward achieving a competitive advantage in business.

Description: A system of electronic hardware and software synthesizes, in nearly real time, an image of a portion of a scene surveyed by as many as eight video cameras aimed, in different directions, at portions of the scene. This is a prototype of systems that would enable a pilot to view the scene outside a remote or windowless cockpit. The outputs of the cameras are digitized. Direct memory addressing is used to store the data of a few captured images in sequence, and the sequence is repeated in cycles. Cylindrical warping is used in merging adjacent images at their borders to construct a mosaic image of the scene. The mosaic-image data are written to a memory block from which they can be rendered on a head-mounted display (HMD) device. A subsystem in the HMD device tracks the direction of gaze of the wearer, providing data that are used to select, for display, the portion of the mosaic image corresponding to the direction of gaze. The basic functionality of the system has been demonstrated by mounting the cameras on the roof of a van and steering the van by use of the images presented on the HMD device.

Description: The volume scattering function (VSF) of seawater affects visibility, remote sensing properties, in-water light propagation, lidar performance, and the like. Currently, it s possible to measure only small forward angles of VSF, or to use cumbersome, large, and non-autonomous systems. This innovation is a method of measuring the full range of VSF using a portable instrument. A single rapid-sensing photosensor is used to scan a green laser beam, which delivers the desired measurement. By using a single sensor, inter-calibration is avoided. A compact design is achieved by using drift-free detector electronics, fiber optics, and a new type of photomultiplier. This provides a high angular resolution of 1 or better, as well as the ability to focus in on a VSF region of particular interest. Currently, the total scattering of light is measured as a difference from the other two parts of the light budget equation. This innovation will allow the direct calculation of the total scattering of light by taking an integral of the VSF over all angles. This directly provides one of the three components of the light budget equation, allowing greater versatility in its calculation.

Description: A method of heating a foam-based catalyst bed has been developed using silicon carbide as the catalyst support due to its readily accessible, high surface area that is oxidation-resistant and is electrically conductive. The foam support may be resistively heated by passing an electric current through it. This allows the catalyst bed to be heated directly, requiring less power to reach the desired temperature more quickly. Designed for heterogeneous catalysis, the method can be used by the petrochemical, chemical processing, and power-generating industries, as well as automotive catalytic converters. Catalyst beds must be heated to a light-off temperature before they catalyze the desired reactions. This typically is done by heating the assembly that contains the catalyst bed, which results in much of the power being wasted and/or lost to the surrounding environment. The catalyst bed is heated indirectly, thus requiring excessive power. With the electrically heated catalyst bed, virtually all of the power is used to heat the support, and only a small fraction is lost to the surroundings. Although the light-off temperature of most catalysts is only a few hundred degrees Celsius, the electrically heated foam is able to achieve temperatures of 1,200 C. Lower temperatures are achievable by supplying less electrical power to the foam. Furthermore, because of the foam s open-cell structure, the catalyst can be applied either directly to the foam ligaments or in the form of a catalyst- containing washcoat. This innovation would be very useful for heterogeneous catalysis where elevated temperatures are needed to drive the reaction.

Description: An eye-safe LADAR system weighs under 500 grams and has range resolution of 1 mm at 10 m. This laser uses an adjustable, tiny microelectromechanical system (MEMS) mirror that was made in SiWave to sweep laser frequency. The size of the laser device is small (70x50x13 mm). The LADAR uses all the mature fiber-optic telecommunication technologies in the system, making this innovation an efficient performer. The tiny size and light weight makes the system useful for commercial and industrial applications including surface damage inspections, range measurements, and 3D imaging.

Description: This work extends the lossless data compression technique described in Fast Lossless Compression of Multispectral- Image Data, (NPO-42517) NASA Tech Briefs, Vol. 30, No. 8 (August 2006), page 26. The original technique was extended to include a near-lossless compression option, allowing substantially smaller compressed file sizes when a small amount of distortion can be tolerated. Near-lossless compression is obtained by including a quantization step prior to encoding of prediction residuals. The original technique uses lossless predictive compression and is designed for use on multispectral imagery. A lossless predictive data compression algorithm compresses a digitized signal one sample at a time as follows: First, a sample value is predicted from previously encoded samples. The difference between the actual sample value and the prediction is called the prediction residual. The prediction residual is encoded into the compressed file. The decompressor can form the same predicted sample and can decode the prediction residual from the compressed file, and so can reconstruct the original sample. A lossless predictive compression algorithm can generally be converted to a near-lossless compression algorithm by quantizing the prediction residuals prior to encoding them. In this case, since the reconstructed sample values will not be identical to the original sample values, the encoder must determine the values that will be reconstructed and use these values for predicting later sample values. The technique described here uses this method, starting with the original technique, to allow near-lossless compression. The extension to allow near-lossless compression adds the ability to achieve much more compression when small amounts of distortion are tolerable, while retaining the low complexity and good overall compression effectiveness of the original algorithm.

Description: This compilation of trade studies performed from 2005 to 2006 addressed a number of power system design issues for the Constellation Program Extravehicular Activity Spacesuit. Spacesuits were required for spacewalks and in-space activities as well as lunar and Mars surface operations. The trades documented here considered whether solar power was feasible for spacesuits, whether spacesuit power generation should be a distributed or a centralized function, whether self-powered in-space spacesuits were better than umbilically powered ones, and whether the suit power system should be recharged in place or replaced.

Description: A new approach has been proposed for increasing astronaut comfort and speech capture. Currently, the special design of a spacesuit forms an extreme acoustic environment making it difficult to capture clear speech without compromising comfort. The proposed Integrated Spacesuit Audio (ISA) system is to incorporate the microphones into the helmet and use software to extract voice signals from background noise.

Description: Wind-US is a computational platform which may be used to numerically solve various sets of equations governing physical phenomena. Currently, the code supports the solution of the Euler and Navier-Stokes equations of fluid mechanics, along with supporting equation sets governing turbulent and chemically reacting flows. Wind-US is a product of the NPARC Alliance, a partnership between the NASA Glenn Research Center (GRC) and the Arnold Engineering Development Center (AEDC) dedicated to the establishment of a national, applications-oriented flow simulation capability. The Boeing Company has also been closely associated with the Alliance since its inception, and represents the interests of the NPARC User's Association. The "Wind-US User's Guide" describes the operation and use of Wind-US, including: a basic tutorial; the physical and numerical models that are used; the boundary conditions; monitoring convergence; the files that are read and/or written; parallel execution; and a complete list of input keywords and test options.

Description: This paper considers the streamline-upwind Petrov/Galerkin (SUPG) method applied to the compressible Euler and Navier-Stokes equations in conservation-variable form. The spatial discretization, including a modified approach for interpolating the inviscid flux terms in the SUPG finite element formulation, is briefly reviewed. Of particular interest is the behavior of the shock capturing operator, which is required to regularize the scheme in the presence of strong, shock-induced gradients. A standard shock capturing operator which has been widely used in previous studies by several authors is presented and discussed. Specific modifications are then made to this standard operator which are designed to produce a more physically consistent discretization in the presence of strong shock waves. The actual implementation of the term in a finite dimensional approximation is also discussed. The behavior of the standard and modified scheme is then compared for several supersonic/hypersonic flows. The modified shock capturing operator is found to preserve enthalpy in the inviscid portion of the flowfield substantially better than the standard operator.

Description: Capillary flow in containers or conduits with interior corners are common place in nature and industry. The majority of investigations addressing such flows solve the problem numerically in terms of a friction factor for flows along corners with contact angles below the Concus-Finn critical wetting condition for the particular conduit geometry of interest. This research effort provides missing numerical data for the flow resistance function F(sub i) for partially and nonwetting systems above the Concus-Finn condition. In such cases the fluid spontaneously de-wets the interior corner and often retracts into corner-bound drops. A banded numerical coefficient is desirable for further analysis and is achieved by careful selection of length scales x(sub s) and y(sub s) to nondimensionalize the problem. The optimal scaling is found to be identical to the wetting scaling, namely x(sub s) = H and y(sub s) = Htan (alpha), where H is the height from the corner to the free surface and a is the corner half-angle. Employing this scaling produces a relatively weakly varying flow resistance F(sub i) and for subsequent analyses is treated as a constant. Example solutions to steady and transient flow problems are provided that illustrate applications of this result.

Description: Hydrologic models use relatively simple mathematical equations to conceptualize and aggregate the complex, spatially distributed, and highly interrelated water, energy, and vegetation processes in a watershed. A consequence of process aggregation is that the model parameters often do not represent directly measurable entities and must, therefore, be estimated using measurements of the system inputs and outputs. During this process, known as model calibration, the parameters are adjusted so that the behavior of the model approximates, as closely and consistently as possible, the observed response of the hydrologic system over some historical period of time. This Chapter reviews the current state-of-the-art of model calibration in watershed hydrology with special emphasis on our own contributions in the last few decades. We discuss the historical background that has led to current perspectives, and review different approaches for manual and automatic single- and multi-objective parameter estimation. In particular, we highlight the recent developments in the calibration of distributed hydrologic models using parameter dimensionality reduction sampling, parameter regularization and parallel computing.

Description: A physical principle (more precisely, a principle that incorporates mathematical models used in physics) has been conceived as the basis of a method of generating randomness in Monte Carlo simulations. The principle eliminates the need for conventional random-number generators. The Monte Carlo simulation method is among the most powerful computational methods for solving high-dimensional problems in physics, chemistry, economics, and information processing. The Monte Carlo simulation method is especially effective for solving problems in which computational complexity increases exponentially with dimensionality. The main advantage of the Monte Carlo simulation method over other methods is that the demand on computational resources becomes independent of dimensionality. As augmented by the present principle, the Monte Carlo simulation method becomes an even more powerful computational method that is especially useful for solving problems associated with dynamics of fluids, planning, scheduling, and combinatorial optimization. The present principle is based on coupling of dynamical equations with the corresponding Liouville equation. The randomness is generated by non-Lipschitz instability of dynamics triggered and controlled by feedback from the Liouville equation. (In non-Lipschitz dynamics, the derivatives of solutions of the dynamical equations are not required to be bounded.)

Description: A report presents the concept of a machine aboard the space shuttle that would cut oversized thermal-tile blanks to precise sizes and shapes needed to replace tiles that were damaged or lost during ascent to orbit. The machine would include a computer-controlled jigsaw enclosed in a clear acrylic shell that would prevent escape of cutting debris. A vacuum motor would collect the debris into a reservoir and would hold a tile blank securely in place. A database stored in the computer would contain the unique shape and dimensions of every tile. Once a broken or missing tile was identified, its identification number would be entered into the computer, wherein the cutting pattern associated with that number would be retrieved from the database. A tile blank would be locked into a crib in the machine, the shell would be closed (proximity sensors would prevent activation of the machine while the shell was open), and a "cut" command would be sent from the computer. A blade would be moved around the crib like a plotter, cutting the tile to the required size and shape. Once the tile was cut, an astronaut would take a space walk for installation.

Description: A computer program performs level- 1B processing (the term 1B is explained below) of data from observations of the limb of the Earth by the Earth Observing System (EOS) Microwave Limb Sounder (MLS), which is an instrument aboard the Aura spacecraft. This software accepts, as input, the raw EOS MLS scientific and engineering data and the Aura spacecraft ephemeris and attitude data. Its output consists of calibrated instrument radiances and associated engineering and diagnostic data. [This software is one of several computer programs, denoted product generation executives (PGEs), for processing EOS MLS data. Starting from level 0 (representing the aforementioned raw data, the PGEs and their data products are denoted by alphanumeric labels (e.g., 1B and 2) that signify the successive stages of processing.] At the time of this reporting, this software is at version 2.2 and incorporates improvements over a prior version that make the code more robust, improve calibration, provide more diagnostic outputs, improve the interface with the Level 2 PGE, and effect a 15-percent reduction in file sizes by use of data compression.

Description: A system of electronic hardware and software for synchronizing recordings from multiple, physically separated video cameras is being developed, primarily for use in multiple-look-angle video production. The system, the time code used in the system, and the underlying method of synchronization upon which the design of the system is based are denoted generally by the term "Geo-TimeCode(TradeMark)." The system is embodied mostly in compact, lightweight, portable units (see figure) denoted video time-code units (VTUs) - one VTU for each video camera. The system is scalable in that any number of camera recordings can be synchronized. The estimated retail price per unit would be about $350 (in 2006 dollars). The need for this or another synchronization system external to video cameras arises because most video cameras do not include internal means for maintaining synchronization with other video cameras. Unlike prior video-camera-synchronization systems, this system does not depend on continuous cable or radio links between cameras (however, it does depend on occasional cable links lasting a few seconds). Also, whereas the time codes used in prior video-camera-synchronization systems typically repeat after 24 hours, the time code used in this system does not repeat for slightly more than 136 years; hence, this system is much better suited for long-term deployment of multiple cameras.

Description: Three documents provide information on a digital processing unit (DPU) for the planned Aquarius mission, in which a radiometer aboard a spacecraft orbiting Earth is to measure radiometric temperatures from which data on sea-surface salinity are to be deduced. The DPU is the interface between the radiometer and an instrument-command-and-data system aboard the spacecraft. The DPU cycles the radiometer through a programmable sequence of states, collects and processes all radiometric data, and collects all housekeeping data pertaining to operation of the radiometer. The documents summarize the DPU design, with emphasis on innovative aspects that include mainly the following: a) In the radiometer and the DPU, conversion from analog voltages to digital data is effected by means of asynchronous voltage-to-frequency converters in combination with a frequency-measurement scheme implemented in field-programmable gate arrays (FPGAs). b) A scheme to compensate for aging and changes in the temperature of the DPU in order to provide an overall temperature-measurement accuracy within 0.01 K includes a high-precision, inexpensive DC temperature measurement scheme and a drift-compensation scheme that was used on the Cassini radar system. c) An interface among multiple FPGAs in the DPU guarantees setup and hold times.

Description: The term precision linear actuator, direct drive ( PLADD ) refers to a robust linear actuator designed to be capable of repeatedly performing, over a lifetime of the order of 5 to 10 years, positioning maneuvers that include, variously, submicron increments or slews of the order of a centimeter. The PLADD is capable of both long stroke (120 mm) and high resolution (repeatable increments of 20 nm). Unlike precise linear actuators of prior design, the PLADD contains no gears, levers, or hydraulic converters. The PLADD, now at the prototype stage of development, is intended for original use as a coarse-positioning actuator in a spaceborne interferometer. The PLADD could also be adapted to terrestrial applications in which there are requirements for long stroke and high resolution: potential applications include medical imaging and fabrication of semiconductor devices. The PLADD (see figure) includes a commercially available ball-screw actuator driven directly by a commercially available three-phase brushless DC motor. The ball-screw actuator comprises a spring-preloaded ball nut on a ball screw that is restrained against rotation as described below. The motor is coupled directly (that is, without an intervening gear train) to a drive link that, in turn, is coupled to the ball nut. By eliminating the gear train, the direct-drive design eliminates the complexity, backlash, and potential for misalignment associated with a gear train. To prevent inadvertent movement, there is a brake that includes flexured levers compressed against the drive link by preload springs. This is a power-off brake: There are also piezoelectric stacks that can be activated to oppose the springs and push the levers away from the drive link. Hence, power must be applied to the piezoelectric stacks to release the drive link from braking. To help ensure long operational life, all of the mechanical drive components are immersed in an oil bath within hermetically sealed bellows. The outer end of the bellows holds the outer end of the ball screw, thereby preventing rotation of the ball screw. Positioning is controlled by an electronic control system that includes digital and analog subsystems that interact with the motor and brake and with two sensor/encoder units: a Hall-effect-sensor rotation encoder and a linear glass-scale encoder. This system implements a proportional + integral + derivative control algorithm that results in variation of voltage commands to each of the three pairs of windings of the brushless DC motor. In one of two alternative control modes, the voltages are applied to the windings in a trapezoidal commutation scheme on the basis of timing signals obtained from the Hall-effect sensors; this scheme yields relatively coarse positioning - 24 steps per motor revolution. The second control mode involves a sinusoidal commutation scheme in which the output of the linear glass-scale encoder is transposed to rotational increments to yield much finer position feedback - more than 400,000 steps per revolution.

Description: Auto-Generated Semantic Processing (AGSP) Services is a suite of software tools for automated generation of other computer programs, denoted cross-platform semantic adapters, that support interoperability of computer-based communication systems that utilize a variety of both new and legacy communication software running in a variety of operating- system/computer-hardware combinations. AGSP has numerous potential uses in military, space-exploration, and other government applications as well as in commercial telecommunications. The cross-platform semantic adapters take advantage of common features of computer- based communication systems to enforce semantics, messaging protocols, and standards of processing of streams of binary data to ensure integrity of data and consistency of meaning among interoperating systems. The auto-generation aspect of AGSP Services reduces development time and effort by emphasizing specification and minimizing implementation: In effect, the design, building, and debugging of software for effecting conversions among complex communication protocols, custom device mappings, and unique data-manipulation algorithms is replaced with metadata specifications that map to an abstract platform-independent communications model. AGSP Services is modular and has been shown to be easily integrable into new and legacy NASA flight and ground communication systems.

Description: Several metal alloys have shown promise as improved catalysts for catalytic thermal decomposition of hydrocarbon gases to produce carbon nanotubes (CNTs). Heretofore almost every experiment on the production of carbon nanotubes by this method has involved the use of iron, nickel, or cobalt as the catalyst. However, the catalytic-conversion efficiencies of these metals have been observed to be limited. The identification of better catalysts is part of a continuing program to develop means of mass production of high-quality carbon nanotubes at costs lower than those achieved thus far (as much as $100/g for purified multi-wall CNTs or $1,000/g for single-wall CNTs in year 2002). The main effort thus far in this program has been the design and implementation of a process tailored specifically for high-throughput screening of alloys for catalyzing the growth of CNTs. The process includes an integral combination of (1) formulation of libraries of catalysts, (2) synthesis of CNTs from decomposition of ethylene on powders of the alloys in a pyrolytic chemical-vapor-decomposition reactor, and (3) scanning- electron-microscope screening of the CNTs thus synthesized to evaluate the catalytic efficiencies of the alloys. Information gained in this process is put into a database and analyzed to identify promising alloy compositions, which are to be subjected to further evaluation in a subsequent round of testing. Some of these alloys have been found to catalyze the formation of carbon nano tubes from ethylene at temperatures as low as 350 to 400 C. In contrast, the temperatures typically required for prior catalysts range from 550 to 750 C.

Description: Seats to prevent or limit crash injuries to astronauts aboard the crew vehicle of the Orion spacecraft are undergoing development. The design of these seats incorporates and goes beyond crash-protection concepts embodied in prior spacecraft and racing-car seats to afford superior protection against impacts. Although the seats are designed to support astronauts in a recumbent, quasi-fetal posture that would likely not be suitable for non-spacecraft applications, parts of the design could be adapted to military and some civilian aircraft seats and to racing car seats to increase levels of protection. The main problem in designing any crashworthy seat is to provide full support of the occupant against anticipated crash and emergency-landing loads so as to safely limit motion, along any axis, of any part of the occupant s body relative to (1) any other part of the occupant s body, (2) the spacecraft or other vehicle, and (3) the seat itself. In the original Orion spacecraft application and in other applications that could easily be envisioned, the problem is complicated by severe limits on space available for the seat, a requirement to enable rapid egress by the occupant after a crash, and a requirement to provide for fitting of the seat to a wide range of sizes and shapes of a human body covered by a crash suit, space suit, or other protective garment. The problem is further complicated by other Orion-application-specific requirements that must be omitted here for the sake of brevity. To accommodate the wide range of crewmember body lengths within the limits on available space in the original Orion application, the design provides for taller crewmembers to pull their legs back closer toward their chests, while shorter crewmembers can allow their legs to stretch out further. The range of hip-support seat adjustments needed to effect this accommodation, as derived from NASA s Human Systems Integration Standard, was found to define a parabolic path along which the knees must be positioned. For a given occupant, the specific position along the path depends on the distance from the heel to the back of the knee. The application of the concept of parabolic adjustment of the hip-support structure caused the seat pan to also take on a parabolic shape, yielding the unanticipated additional benefit that the seat pan fits the occupant s buttocks and thighs more nearly conformally than do seat pans of prior design. This more nearly conformal fit effectively eliminates a void between the occupant s body and the seat pan, thereby helping to prevent what, in prior seat designs, was shifting of the occupant s body into that void during an impact.

Description: The purpose of this study was to determine the effect of atomic oxygen (AO) exposure on the hydrophilicity of nine different polymers for biomedical applications. Atomic oxygen treatment can alter the chemistry and morphology of polymer surfaces, which may increase the adhesion and spreading of cells on Petri dishes and enhance implant growth. Therefore, nine different polymers were exposed to atomic oxygen and water-contact angle, or hydrophilicity, was measured after exposure. To determine whether hydrophilicity remains static after initial atomic oxygen exposure, or changes with higher fluence exposures, the contact angles between the polymer and water droplet placed on the polymer s surface were measured versus AO fluence. The polymers were exposed to atomic oxygen in a 100-W, 13.56-MHz radio frequency (RF) plasma asher, and the treatment was found to significantly alter the hydrophilicity of non-fluorinated polymers. Pristine samples were compared with samples that had been exposed to AO at various fluence levels. Minimum and maximum fluences for the ashing trials were set based on the effective AO erosion of a Kapton witness coupon in the asher. The time intervals for ashing were determined by finding the logarithmic values of the minimum and maximum fluences. The difference of these two values was divided by the desired number of intervals (ideally 10). The initial desired fluence was then multiplied by this result (2.37), as was each subsequent desired fluence. The flux in the asher was determined to be approximately 3.0 x 10(exp 15) atoms/sq cm/sec, and each polymer was exposed to a maximum fluence of 5.16 x 10(exp 20) atoms/sq cm.

Description: The figure schematically depicts some alternative designs of proposed compact, lightweight optoelectronic microscopes that would contain no lenses and would generate magnified video images of specimens. Microscopes of this type were described previously in Miniature Microscope Without Lenses (NPO - 20218), NASA Tech Briefs, Vol. 22, No. 8 (August 1998), page 43 and Reflective Variants of Miniature Microscope Without Lenses (NPO 20610), NASA Tech Briefs, Vol. 26, No. 9 (September 1999), page 6a. To recapitulate: In the design and construction of a microscope of this type, the focusing optics of a conventional microscope are replaced by a combination of a microchannel filter and a charge-coupled-device (CCD) image detector. Elimination of focusing optics reduces the size and weight of the instrument and eliminates the need for the time-consuming focusing operation. The microscopes described in the cited prior articles contained two-dimensional CCDs registered with two-dimensional arrays of microchannels and, as such, were designed to produce full two-dimensional images, without need for scanning. The microscopes of the present proposal would contain one-dimensional (line image) CCDs registered with linear arrays of microchannels. In the operation of such a microscope, one would scan a specimen along a line perpendicular to the array axis (in other words, one would scan in pushbroom fashion). One could then synthesize a full two-dimensional image of the specimen from the line-image data acquired at one-pixel increments of position along the scan. In one of the proposed microscopes, a beam of unpolarized light for illuminating the specimen would enter from the side. This light would be reflected down onto the specimen by a nonpolarizing beam splitter attached to the microchannels at their lower ends. A portion of the light incident on the specimen would be reflected upward, through the beam splitter and along the microchannels, to form an image on the CCD. If the nonpolarizing beam splitter were replaced by a polarizing one, then the specimen would be illuminated by s-polarized light. Upon reflection from the specimen, some of the s-polarized light would become p-polarized. Only the p-polarized light would contribute to the image on the CCD; in other words, the image would contain information on the polarization rotating characteristic of the specimen.