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Future space exploration missions will require a lightweight spacesuit that expends no consumables. This paper describes the design and performance of a prototype heat rejection system that weighs less than current systems and vents zero water. The system uses regenerable LiCl/water absorption cooling. Absorption cooling boosts the heat absorbed from the crew member to a high temperature for rejection to space from a compact, non-venting radiator. The system is regenerated by heating to 100°C for two hours. The system provides refrigeration at 17°C and rejects heat at temperatures greater than 50°C. The overall cooling capacity is over 100 W-hr/kg.

In order to provide effective cooling for astronauts during extravehicular activities (EVAs), a liquid cooling and ventilation garment (LCVG) is used to remove heat by a series of tubes through which cooling water is circulated. To better predict the effectiveness of the LCVG and determine possible modifications to improve performance, computer simulations dealing with the interaction of the cooling garment with the human body have been run using the Wissler Human Thermal Model. Simulations have been conducted to predict the heat removal rate for various liquid cooled garment configurations. The current LCVG uses 48 cooling tubes woven into a fabric with cooling water flowing through the tubes. The purpose of the current project is to decrease the overall weight of the LCVG system. In order to achieve this weight reduction, advances in the garment heat removal rates need to be obtained.

Lunar and martian materials can be processed and used at planetary outposts to reduce the need (and thus the cost) of transporting supplies from Earth. A variety of uses for indigenous, on-site materials have been suggested, including uses as rocket propellants, construction materials, and life support materials. Utilization of on-site resources will supplement Regenerative Life Support Systems (RLSS) that will be needed to regenerate air, water, and wastes, and to produce food (e.g., plants) for human consumption during long-duration space missions.

The International Space Station (ISS) Active Thermal Control System (ATCS) (Ref. 1,2) has changed over the past several years to address problems and to improve its assembly and operation on-orbit. This paper captures the ways in which the Internal (I) ATCS and External (E) ATCS have changed design characteristics and operations both for the system currently operating on-orbit and the new elements of the system that are about to be added and/or activated. The rationale for changes in ATCS design, assembly and operation will provide insights into the lessons learned during ATCS development. The state of the assembly of the integrated ATCS will be presented to provide a status of the build-up of the system. The capabilities of the on-orbit system will be presented with a summary of the elements of the ISS ATCS that are functional on-orbit plus the plans for launch of remaining parts of the integrated ISS ATCS.

A porous plate sublimator (based on an existing Lunar Module LM-209 design) is baselined as a heat rejection device for the X-38 vehicle due to its simplicity, reliability, and flight readiness. The sublimator is a passive device used for rejecting heat to the vacuum of space by sublimating water to obtain efficient heat rejection in excess of 1,000 Btu/lb of water. It is ideally suited for the X-38/CRV mission as it requires no active control, has no moving parts, has 100% water usage efficiency, and is a well-proven technology. Two sublimators have been built and tested for the X-38 program, one of which will fly on the NASA V-201 space flight demonstrator vehicle in 2001. The units satisfied all X-38 requirements with margin and have demonstrated excellent performance. Minor design changes were made to the LM-209 design for improved manufacturability and parts obsolescence.

The NASA-wide CEV Smart Buyer Team (SBT) was assembled in January 2006 and was tasked with the development of a NASA in-house design for the CEV Crew Module (CM), Service Module (SM), and Launch Abort System (LAS). This effort drew upon over 250 engineers from all of the 10 NASA Centers. In 6 weeks, this in-house design was developed. The Thermal Systems Team was responsible for the definition of the active and passive design architecture. The SBT effort for Thermal Systems can be best characterized as a design architecting activity. Proof-of-concepts were assessed through system-level trade studies and analyses using simplified modeling. This nimble design approach permitted definition of a point design and assessing its design robustness in a timely fashion. This paper will describe the architecting process and present trade studies and proposed thermal designs

As part of its integrated system test bed capability, NASA's Advanced Life Support Program has undertaken the development of a large-scale advanced life support facility capable of supporting long-duration testing of integrated, regenerative biological and physicochemical life support systems. This facility--the Advanced Life Support Human-Rated Test Facility (HRTF) is currently being built at the Johnson Space Center. The HRTF is comprised of a series of interconnected chambers with a sealed internal environment capable of supporting a test crew of four for periods exceeding one year. The life support system will consist of both biological and physicochemical components and will perform air revitalization, water recovery, food production, solid waste processing, thermal management, and integrated command and control functions. Currently, a portion of this multichamber facility has been constructed and is being outfitted with basic utilities and infrastructure.

Hamilton Sundstrand has developed a scalable evaporative heat rejection system called the Multi-Fluid Evaporator (MFE). It was designed to support the Orion Crew Module and to support future Constellation missions. The MFE would be used from Earth sea level conditions to the vacuum of space. This system combines the functions of the Space Shuttle flash evaporator and ammonia boiler into a single compact package with improved freeze-up protection. The heat exchanger core is designed so that radial flow of the evaporant provides increasing surface area to keep the back pressure low. The multiple layer construction of the core allows for efficient scale up to the desired heat rejection rate. A full-scale unit uses multiple core sections that, combined with a novel control scheme, manage the risk of freezing the heat exchanger cores. A four-core MFE prototype was built in 2007.

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 all controlled with a single spool valve. This technology has been baselined for the new Orion spacecraft. However, more data was needed on the operational characteristics of the package in a simulated spacecraft environment. A unit was therefore tested with simulated metabolic loads in a closed chamber at Johnson Space Center during the last third of 2006. Tests were run at a variety of cabin temperatures and with a range of operating conditions varying cycle time, vacuum pressure, air flow rate, and crew activity levels.

Research committed by the Langley Research Center through 1995 resulting in the HZETRN code provides the current basis for shield design methods according to NASA STD-3000 (2005). With this new prominence, the database, basic numerical procedures, and algorithms are being re-examined with new methods of verification and validation being implemented to capture a well defined algorithm for engineering design processes to be used in this early development phase of the Bush initiative. This process provides the methodology to transform the 1995 HZETRN research code into the 2005 HZETRN engineering code to be available for these early design processes. In this paper, we will review the basic derivations including new corrections to the codes to insure improved numerical stability and provide benchmarks for code verification.

The NASA Johnson Space Center has plans to integrate a Solid Amine Water Desorbed (SAWD II) carbon dioxide removal subsystem into the Variable Pressure Growth Chamber (VPGC). The SAWD II subsystem will be used to remove any excess carbon dioxide (CO2) input into the VPGC which is not assimilated by the plants growing in the chamber. An analysis of the integrated VPGC-SAWD II system was performed using a mathematical model of the system implemented in the Computer-Aided System Engineering and Analysis (CASE/A) package. The analysis consisted of an evaluation of the SAWD II subsystem configuration within the VPGC, the planned operations for the subsystem, and the overall performance of the subsystem and other VPGC subsystems. Based on the model runs, recommendations were made concerning the SAWD II subsystem configuration and operations, and the chambers' automatic CO2 injection control subsystem.

Poly (vinyl chloride) (PVC) matrix membranes which incorporate the ionophore nonactin have been evaluated as cation exchange membranes for ammonium ion transport in an electrolytic cell configuration. Interest exists for the development of cation selective membranes for removal of low levels (<200ppm) of ammonium ions commonly found in recycled effluent streams in such diverse applications as expected in a Space Station and commercial fisheries. Ammonium ions are generated as a decomposition product of urea and over time build up in concentration, thus rendering the water unsuitable for human consumption. Nonactin is commonly used in a PVC matrix for ion-selective electrodes.

The Orion crew module (CM) is being designed to perform survivable land and water landings. There are many issues associated with post-landing crew survival. In general, the most challenging of the realistic Orion landing scenarios from an environmental control standpoint is the off-nominal water landing. Available power and other consumables will be very limited after landing, and it may not be possible to provide full environmental control within the crew cabin for very long after splashdown. Given the bulk and thermal insulation characteristics of the crew-worn pressure suits, landing in a warm tropical ocean area would pose a risk to crew survival from elevated core body temperatures, if for some reason the crewmembers were not able to remove their suits and/or exit the vehicle. This paper summarizes the analyses performed and conclusions reached regarding post-landing crew survival following a water landing, from the standpoint of the crew's core body temperatures.

To date, 59 man-EVA's have been conducted in the Shuttle Program with minimum physiological problems or limitations. The physiological requirements for life support in the Shuttle EVA include pressure, gas composition, inspired CO2 pressure, heat- removal capability, in-suit water replacement, and caloric replacement. These requirements and their basis in verification testing or analysis are reviewed. The operational measures are identified. The suit pressure in combination with a gas composition of at least 92 percent assures that sufficient O2 pressure is available to the crewmember. The nominal suit pressure of 4.3 psi±0.1 psi was maintained during all 59 man-EVA's. The contingency suit pressure was never required to be used. The suit pressure in combination with the cabin pressure and pre-EVA denitrogenation procedures minimize the risk of altitude decompression sickness. There has been no incidence of decompression sickness during Shuttle EVA.

Since the early 1980’s, the Shuttle Extra Vehicular Activity (EVA) glove design has evolved to meet the challenge of space based tasks. These tasks have typically been satellite retrieval and repair or EVA based flight experiments. With the start of the International Space Station (ISS) assembly, the number of EVA based missions is increasing far beyond what has been required in the past; this has commonly been referred to as the “Wall of EVA’s”. To meet this challenge, it was determined that the evolution of the current glove design would not meet future mission objectives. Instead, a revolution in glove design was needed to create a high performance tool that would effectively increase crewmember mission efficiency. The results of this effort have led to the design, certification and implementation of the Phase VI EVA glove into the Shuttle flight program.

The recovery of potable water from waste water produced by humans in regenerative life support systems is essential for success of long-duration space missions. The Lunar-Mars Life Support Test Project (LMLSTP) Phase II test was performed to validate candidate technologies to support these missions. The test was conducted in the Crew and Thermal Systems Division (CTSD) Life Support Systems Integration Facility (LSSIF) at Johnson Space Center (JSC). Discussed in this paper are the water recovery system (WRS) results of this test. A crew of 4-persons participated in the test and lived in the LSSIF chamber for a duration of 30-days from June 12 to July 12, 1996. The crew had accommodations for personal hygiene, the air was regenerated for reuse, and the waste water was processed to potable and hygiene quality for reuse by the crew during this period. The waste water consisted of shower, laundry, handwash, urine and humidity condensate.

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.

Processes to remove and recover carbon dioxide (CO2) and water vapor from air are essential for successful long-duration space missions. This paper presents results of a developmental program focused on the use of a liquid- sorbent/membrane-contactor (LSMC) system for removal of CO2 and water vapor from air. In this system, air from the spacecraft cabin atmosphere is circulated through one side of a hollow-fiber membrane contactor. On the other side of the membrane contactor is flowed a liquid sorbent, which absorbs the CO2 and water vapor from the feed air. The liquid sorbent is then heated to desorb the CO2 and water vapor. The CO2 is subsequently removed from the system as a concentrated gas stream, whereas the water vapor is condensed, producing a water stream. A breadboard system based on this technology was designed and constructed. Tests showed that the LSMC breadboard system can produce a CO2 stream and a liquid- water stream.

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.

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.