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NASA's proposed lunar lander, Altair, will be exposed to vastly different external temperatures following launch till its final destination on the moon. In addition, the heat rejection is lowest at the lowest environmental temperatures (0.5 kW @ 4K) and highest at the highest environmental temperature (4.5 kW @ 215K). This places a severe demand on the radiator design to handle these extreme turn-down requirements. A radiator with digital turn-down capability is currently under study at JPL as a robust means to meet the heat rejection demands and provide freeze protection while minimizing mass and power consumption. Turndown is achieved by independent control of flow branches with isolating latch valves and a gear pump to evacuate the isolated branches. A bench-top test was conducted to characterize the digital radiator concept. Testing focused on the demonstration of proper valve sequencing to achieve turn-down and recharge of flow legs.

Providing water necessary to maintain life support has been accomplished in spacecraft vehicles for over forty years. This paper will investigate how previous U.S. space vehicles provided potable water. The water source for the spacecraft, biocide used to preserve the water on-orbit, water stowage methodology, materials, pumping mechanisms, on-orbit water requirements, and water temperature requirements will be discussed. Where available, the hardware used to provide the water and the general function of that hardware will also be detailed. The Crew Exploration Vehicle (CEV or Orion) water systems will be generically discussed to provide a glimpse of how similar they are to water systems in previous vehicles. Conclusions, questions, and recommendations on strategies that could be applied to CEV based on previous spacecraft water system lessons learned will be made.

When a permanent human outpost is established on the Moon, various methods may be used to reject the heat generated by the base. One proposed concept is the use of a heat pump operating with a vertical, flow-through thermal radiator mounted on a Space Station type habitation module [1]. Since the temperature of the lunar surface varies over the day, the vertical radiator sink temperatures can reach much higher levels than the comfort and even survivability requirements of a habitation module. A high temperature lift heat pump will not only maintain a comfortable habitation module temperature, but will also decrease the size of the radiators needed to reject the waste heat. Thus, the heat pump will also decrease the mass of the entire thermal system. Engineers at the Johnson Space Center (JSC) have tested a High Lift Heat Pump design and are developing the next generation heat pump based on information and experience gained from this testing.

An evaporative heat sink has been designed and built by AlliedSignal for NASA's Johnson Space Center. The unit is a demonstrator of a primary heat exchanger for NASA's prototype Crew Return Vehicle (CRV), designated the X-38. The primary heat exchanger is responsible for rejecting the heat produced by both the flight crew and the avionics. Spacecraft evaporative heat sinks utilize space vacuum as a resource to control the vapor pressure of a liquid. For the X-38, water has been chosen as the heat transport fluid. A portion of this coolant flow is bled off for use as the evaporant. At sufficiently low pressures, the water can be made to boil at temperatures approaching its freezing point. Heat transferred to liquid water in this state will cause the liquid to evaporate, thus creating a heat sink for the spacecraft's coolant loop. The CRV mission requires the heat exchanger to be compact and low in mass.

We have completed preliminary tests that show the feasibility of an innovative concept for a spacesuit thermal control system using a lightweight, flexible heat pump/radiator. The heat pump/radiator is part of a regenerable LiCI/water absorption cooling device that absorbs an astronaut's metabolic heat and rejects it to the environment via thermal radiation at a relatively high temperature. We identified key design specifications for the system, demonstrated that it is feasible to fabricate the flexible radiator, measured the heat rejection capability of the radiator, and assessed the effects on overall mass of the PLSS. We specified system design features that will enable the flexible absorber/radiator to operate in a wide range of space exploration environments. The materials used to fabricate the flexible absorber/radiator samples were all found to be low off-gassing and many have already been qualified for use in space.

The Spacesuit Water Membrane Evaporator (SWME) is the baseline heat rejection technology selected for development for the Constellation lunar suit. The first SWME prototype, designed, built, and tested at Johnson Space Center in 1999 used a Teflon hydrophobic porous membrane sheet shaped into an annulus to provide cooling to the coolant loop through water evaporation to the vacuum of space. This present study describes the test methodology and planning to compare the test performance of three commercially available hollow fiber materials as alternatives to the sheet membrane prototype for SWME, in particular, a porous hydrophobic polypropylene, and two variants that employ ion exchange through non-porous hydrophilic modified Nafion. Contamination tests will be performed to probe for sensitivities of the candidate SWME elements to ordinary constituents that are expected to be found in the potable water provided by the vehicle, the target feedwater source.

In a crewed spacecraft environment, atmospheric carbon dioxide (CO2) and moisture control are crucial. Hamilton Sundstrand has developed a stable and efficient amine-based CO2 and water vapor sorbent, SA9T, that is well suited for use in a spacecraft environment. The sorbent is efficiently packaged in pressure-swing regenerable beds that are thermally linked to improve removal efficiency and minimize vehicle thermal loads. Flows are controlled with a single spool valve. This technology has been baselined for the new Orion spacecraft, but additional data was needed on the operational characteristics of the package in a simulated spacecraft environment. One unit was tested with simulated metabolic loads in a closed chamber at Johnson Space Center during the latter part of 2006. Those test results were reported in a 2007 ICES paper.

The Sublimator Driven Coldplate (SDC) is a unique piece of thermal control hardware that has several advantages over a more traditional thermal control system. The principal advantage is the possible elimination of a pumped fluid loop, potentially saving mass, power, and complexity. Because this concept relies on evaporative heat rejection techniques, it is primarily useful for short mission durations. Additionally, the concept requires a conductive path between the heat-generating component and the heat rejection device. Therefore, it is mostly a relevant solution for a vehicle with a relatively low heat rejection requirement and/or short transport distances. Tests were performed on coupons and an Engineering Development Unit (EDU) at NASA's Johnson Space Center to better understand the basic operational principles and to validate the analytical methods being used for the SDC development.

The aim of this study was to explore if fingernail delamination injury following EMU glove use may be caused by compression-induced blood flow occlusion in the finger. During compression tests, finger blood flow decreased more than 60%, however this occurred more rapidly for finger pad compression (4 N) than for fingertips (10 N). A pressure bulb compression test resulted in 50% and 45% decreased blood flow at 100 mmHg and 200 mmHg, respectively. These results indicate that the finger pad pressure required to articulate stiff gloves is more likely to contribute to injury than the fingertip pressure associated with tight fitting gloves.

In the design of a new space suit it is necessary to have requirements that define what mobility space suit joints should be capable of achieving in both a system and at the component level. NASA elected to divide mobility into its constituent parts -- range of motion (ROM) and torque -- in an effort to develop clean design requirements that limit subject performance bias and are easily verified. Unfortunately, the measurement of mobility can be difficult to obtain. Current technologies, such as the Vicon motion capture system, allow for the relatively easy benchmarking of range of motion (ROM) for a wide array of space suit systems. The ROM evaluations require subjects in the suit to accurately evaluate the ranges humans can achieve in the suit. However, when it comes to torque, there are significant challenges for both benchmarking current performance and writing requirements for future suits.

A generalized computer program for analysis of pressure and composition in multiple volume systems has been under development by the National Aeronautics and Space Administration (NASA) since 1976. This paper describes the most recent developments in the program. These improvements include the expansion of the program to nine volumes, improvements to the model of the International Space Station (ISS) carbon dioxide removal system, and addition of a detailed Sabatier carbon dioxide reduction mode. An evaluation of the feasibility of adding of trace contaminant tracking was also performed. This paper will also present the results of an analysis that compares model predictions with ISS flight data for carbon dioxide (CO2) maintenance.

During the period July 2001 to March 2002, the performance of a water-jacketed high intensity discharge lamp of advanced design was evaluated within a lamp test stand at The University of Arizona (UA), Controlled Environment Agriculture Center (CEAC) in Tucson, Arizona. The lamps and test stand system were developed by Mr. Phil Sadler of Sadler Machine Company, Tempe, Arizona, and supported by a Space Act Agreement between NASA-Johnson Space Center (JSC) and UA. The purpose was for long term testing of the prototype lamp and demonstration of an improved procedure for use of water-jacketed lamps for plant production within the close confines of controlled environment facilities envisioned by NASA within Bioregenerative Life Support Systems. The lamp test stand consisted of six, 400 watt water-cooled, high pressure sodium HID lamps, mounted within a framework.

Research into increased capacity solid amine sorbents has found a candidate (SA9T) that will provide enough increase in cyclic carbon dioxide removal capacity to produce a fully redundant Regenerative Carbon Dioxide Removal System (RCRS). This system will eliminate the need for large quantities of backup LiOH, thus gaining critical storage space on board the shuttle orbiter. This new sorbent has shown an ability to package two fully redundant (four) sorbent beds together with their respective valves, fans and plumbing to create two operationally independent systems. The increase in CO2 removal capacity of the new sorbent will allow these two systems to fit within the envelope presently used by the RCRS. This paper reports on the sub-scale amine testing performed in support of the development effort. In addition, this paper will provide a preliminary design schematic of a fully redundant RCRS.

This paper provides an overview of the IMMWPS Integrated Ground Test Program, completed at the NASA Johnson Space Center (JSC) during October and November 2001. The JSC Crew and Thermal Systems Division (CTSD) has developed the IMMWPS orbital flight experiment to test the feasibility of a microbe-based water purifier for use in zero-gravity conditions. The IMMWPS design utilizes a Microbial Processor Assembly (MPA) inoculated with facultative anaerobes to convert organic contaminants in wastewater to carbon dioxide and biomass. The primary purpose of the ground test program was to verify functional operations and procedures. A secondary objective was to provide initial ground data for later comparison to on-orbit performance. This paper provides a description of the overall test program, including the test article hardware and the test sequence performed to simulate the anticipated space flight test program. In addition, a summary of significant results from the testing is provided.

A thermal analysis of the compressible carbon dioxide (CO2) flow for the Portable Fire Extinguisher (PFE) system has been performed. A SINDA/FLUINT model has been developed for this analysis. The model includes the PFE tank and the Temporary Sleep Station (TeSS) nozzle, and both have an initial temperature of 72 °F. In order to investigate the thermal effect on the nozzle due to discharging CO2, the PFE TeSS nozzle pipe has been divided into three segments. This model also includes heat transfer predictions for PFE tank inner and outer wall surfaces. The simulation results show that the CO2 discharge rates and component wall temperatures fall within the requirements for the PFE system. The simulation results also indicate that after 50 seconds, the remaining CO2 in the tank may be near the triple point (gas, liquid and solid) state and, therefore, restricts the flow.

The Major Constituent Analyzer (MCA) was activated on-orbit on 2/13/01 and provided essentially continuous readings of partial pressures for oxygen, nitrogen, carbon dioxide, methane, hydrogen and water in the ISS atmosphere. The MCA plays a crucial role in the operation of the Laboratory ECLSS and EVA operations from the airlock. This paper discusses the performance of the MCA as compared to specified accuracy requirements. The MCA has an on-board self-calibration capability and the frequency of this calibration could be relaxed with the level of instrument stability observed on-orbit. This paper also discusses anomalies the MCA experienced during the first year of on-orbit operation. Extensive Built In Test (BIT) and fault isolation capabilities proved to be invaluable in isolating the causes of anomalies. The process of fault isolation is discussed along with development of workaround solutions and implementation of permanent on-orbit corrections.

The International Space Station Waste Collector Subsystem Risk Mitigation Experiment (ISS WCS RME) was flown as the primary (Shuttle) WCS on Space Shuttle flight STS-104 (ISS-7A) in July 2001, to validate new design enhancements. In general, the WCS is utilized for collecting, storing, and compacting fecal & associated personal hygiene waste, in a zero gravity environment. In addition, the WCS collects and transfers urine to the Shuttle waste storage tank. All functions are executed while controlling odors and providing crew comfort. The ISS WCS previously flew on three Shuttle flights as the Extended Duration Orbiter (EDO) WCS, as it was originally designed to support extended duration Space Shuttle flights up to 30 days in length. Soon after its third flight, the Space Shuttle Program decided to no longer require 30 day extended mission duration capability and provided the EDO WCS to the ISS Program.

Gas-separation and reverse-osmosis membrane models are being developed in conjunction with membrane testing at NASA JSC. The completed gas-separation membrane model extracts effective component permeabilities from multicomponent test data, and predicts the effects of flow configuration, operating conditions, and membrane dimensions on module performance. Variable feed- and permeate-side pressures are considered. The model has been applied to test data for hollow-fiber membrane modules with simulated cabin-air feeds. Results are presented for a membrane designed for air drying applications. Extracted permeabilities are used to predict the effect of operating conditions on water enrichment in the permeate. A first-order reverse-osmosis model has been applied to test data for spiral wound membrane modules with a simulated hygiene water feed. The model estimates an effective local component rejection coefficient under pseudo-steady-state conditions.

For over 40 years, NASA has been putting humans safely into space in part by minimizing microbial risks to crew members. Success of the program to minimize such risks has resulted from a combination of engineering and design controls as well as active monitoring of the crew, food, water, hardware, and spacecraft interior. The evolution of engineering and design controls is exemplified by the implementation of HEPA filters for air treatment, antimicrobial surface materials, and the disinfection regimen currently used on board the International Space Station. Data from spaceflight missions confirm the effectiveness of current measures; however, fluctuations in microbial concentrations and trends in contamination events suggest the need for continued diligence in monitoring and evaluation as well as further improvements in engineering systems. The knowledge of microbial controls and monitoring from assessments of past missions will be critical in driving the design of future spacecraft.

This ground-based study compared the risk of microbubbles during decompression under simulated space extravehicular activities (EVA) after three days of six-degree head-down bed rest with three days of ambulatory control. Test subjects were exposed to a pressure of 44.8 kPa (6.5 psi), breathed 100% oxygen, and exercised at reduced pressure either in the supine (during experimental) or upright (control) position. Circulating microbubbles were monitored by a precordial Doppler ultrasound device, and were found in 52% (12/23) of control and 26% (6/23) of experimental exposures. Survival analysis using Cox proportional hazards regression showed that there was 0.22 times (95% confidence interval=0.07-0.68) reduction in the risk of high grade microbubbles after bed rest, compared to controls (p=0.004). This finding is of importance in evaluating the risk of DCS during EVA.