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The Lunar Electric Rover (LER), which was formerly called the Small Pressurized Rover (SPR), is currently being carried as an integral part of the lunar surface architectures that are under consideration in the Constellation Program. One element of the LER is the suit port, which is the means by which crew members perform Extravehicular Activities (EVAs). Two suit port deliverables were produced in fiscal year 2008: a 1-g suit port concept evaluator for functional integrated testing with the LER 1-g concept vehicle and a functional and pressurizable Engineering Unit (EU). This paper focuses on the 1-g suit port concept evaluator test results from the Desert Research and Technology Studies (D-RATS) October 2008 testing at Black Point Lava Flow (BPLF), Arizona. The 1-g suit port concept evaluator was integrated with the 1-g LER cabin and chassis concepts.

A preliminary assessment of the reach envelope and field of vision (FOV) for a subject wearing a Mark III space suit was requested for use in human-machine interface design of the Science Crew Operations and Utility Testbed (SCOUT) vehicle. The reach and view of two suited and unsuited subjects were evaluated while seated in the vehicle using 3-dimensional position data collected during a series of reaching motions. Data was interpolated and displayed in orthogonal views and cross-sections. Compared with unsuited conditions, medio-lateral reach was not strongly affected by the Mark III suit, whereas vertical and antero-posterior reach were inhibited by the suit. Lateral FOV was reduced by approximately 40° in the suit. The techniques used in this case study may prove useful in human-machine interface design by providing a new means of developing and displaying reach envelopes.

Spacecraft hardware trade studies compare options primarily on mass while considering impacts to cost, risk, and schedule. Historically, other factors have been considered in these studies, such as reliability, technology readiness level (TRL), volume and crew time. In most cases, past trades compared two or more technologies across functional and TRL boundaries, which is an uneven comparison of the technologies. For example, low TRL technologies with low mass were traded directly against flight-proven hardware without consideration for requirements and the derived architecture. To provide for even comparisons of spacecraft hardware, trades need to consider functionality, mission constraints, integer vs. real number of flight hardware units, and mass growth allowances by TRL.

Carbon dioxide (CO2), a natural product of human metabolism, accumulates quickly in sealed environments when humans are present, and can induce headaches, among other symptoms. Major resources are expended to control CO2 levels to concentrations that are tolerable to the crews of spacecraft and submersible craft. It is not practical to control CO2 levels to those found in the ambient environment on earth. As NASA looks ahead to long-duration missions conducted far from earth, difficult issues arise related to the management and effects of human exposure to CO2. One is the problem of “pockets” of CO2 in the habitat caused by excess generation of the gas in one location without a mechanism to purge the area with fresh air. This results in the crew rebreathing CO2 from their exhaled breath, exposing them to a much higher concentration of CO2 than whole-module measurements would suggest. Another issue is the potential increased sensitivity to CO2 in microgravity.

Deneb's Interactive Graphic Robot Instruction Progam (IGRIP) and Envision software packages with the Ergonomic analysis option enabled were used for manufacturing process analysis and maintainability / human factors design evaluation in the Lockheed Martin Tactical Aircraft Systems - Fort Worth facility. The initial objective of both the manufacturing and maintainability engineering community was to validate the use of ergonomic modeling and simulation tools in an effort to gain acceptance of this new technology. Each discipline selected an existing operation to baseline the validation. Manufacturing selected the F-16 vertical fin as it is assembled from detail parts into a complete assembly, ready to be mated to the aircraft. Maintainability selected the removal of the Expanded Data Entry Electronics Unit (EXDEEU) located behind the ejection seat of the F-16 aircraft.

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.

Space suits for advanced missions have baselined radiators as the primary means of heat rejection in order to minimize consumables and logistics requirements. While radiators have been used in the active thermal control system for spacecraft since Gemini, the use of radiators in spacesuits introduces many unique requirements. These include the ability to reduce the amount of heat rejection when overcooling or overheating of the crew member is a concern. Overcooling can occur with low metabolic rates, cold environments or a combination of the two, and overheating can occur with high metabolic rates in a warm environment. The main goal of the Radiator Advanced Demonstrator (RAD) program is to build and fly a radiator on the current Extravehicular Mobility Unit (EMU) in order to verify thermal performance capabilities in actual flight conditions. The RAD incorporates an aluminum plate separated from the primary water panel with a silicone gasket.

An experimental study has been made of compressible jet mixing in an axisymmetric ejector of converging-diverging geometry. Three different jet sizes, 0.01, 0.0235, and 0.045 in. diameter were tested with three different mixer sizes, 0.25, 0.286, and 0.36 in. diameter. Jet and mixer combination were tested along with varying jet to mixer distances. The jet pressure varied from 20 to 200 psig, jet mass varied from 0.3 lbm/hr to 10 lbm/hr., and jet temperature varied from 21 to 24 deg. F. The secondary loop pressure varied from 3.7 to 25 psia, secondary mass flow varied from 1 to 70 lbm/hr, secondary loop pressure drop varied from 4 inH20 to 10 inH20, and secondary loop temperature varied same as jet temperature. The mass flow ratio was in the range of 2 to 14. The results were analyzed and compared with the Hickman and Nuckols and Sexton prediction models. The loss factor in Nuckols and Sexton model was adjusted to match the test results.

In this investigation, analytical models were developed to predict the performance characteristics of axisymmetric single jet ejector. The ejector is divided into four parts, jet, mixer, nozzle, and diffuser. Basic flow equations were combined to calculate end to end flow characteristics for each of the four ejector components. Different jets and mixer combination were tested using three jet and three mixers. Characteristics curves have been drawn to predict flow characteristics of the ejector. Different configuration of jet and mixer incorporated different loss coefficient. Hence to get correct flow characteristics of the ejector right loss coefficient should be used.

The Lunar Mars Life Support Test Project (LMLSTP) Phase III test was the final test in a series of tests conducted to evaluate regenerative life support systems performance over increasingly longer durations. The Phase III test broke new ground for the U.S. Space Program by being the first test to look at integration of biological and physical-chemical systems for air, water and solid waste recovery for a crew of four for 91 days. Microbial bioreactors were used as the first step in the water recovery system (WRS). This biologically based WRS continuously recovered 100% of the water used by the crew consistent with NASA's strict potable standards. The air revitalization system was a combination of physical-chemical hardware and wheat plants which worked together to remove and reduce the crew's metabolically produced carbon dioxide and provide oxygen.

The Lunar-Mars Life Support Systems Test Project's Phase iii Test utilized the Variable Pressure Growth Chamber to contribute to the air revitalization and food requirements of a crew of four for a period of 91 days. USU-Apogee wheat was planted and harvested using a staged approach to provide more uniform levels of air revitalization and a staggered production of grain. The wheat crop provided an average of 1 .1 person-equivalents per day of carbon dioxide removal for air revitalization over the 91 -day human test. Over 34 kg of grain was harvested. it was found that staged cropping required more intensive management of the nutrient solution than single batch cropping. it was also found that salts which were biologically recovered from the plant biomass were as effective as conventional reagent-grade salts for use in the hydroponic nutrient solution.

A technology project to produce a new space suit for planetary applications started in January of 1997, with a thermal vacuum test of the system, including a suited crew member, expected in the year 2000. This will be a progress report on the activities that occurred during the project's first year. The four year project is funded out of Code M at NASA Headquarters and is an effort to integrate the latest EVA technology into a maintainable modular design. The project will use as much off-the-shelf hardware as practical in an effort to lower development cost and decrease development time. Three pressurized garment configurations will be evaluated and two different portable life support systems will be built. The first year was primarily spent developing laboratories, bench-top working laboratory subsystems, analytical models, and the overall requirements and architecture of the system.

The optimum radiator configuration in hot lunar thermal environments is one in which the radiator is parallel to the ground and has no view to the hot lunar surface. However, typical spacecraft configurations have limited real estate available for top-mounted radiators, resulting in a desire to use the spacecraft's vertically oriented sides. Vertically oriented, flat panel radiators will have a large view factor to the lunar surface, and thus will be subjected to significant incident lunar infrared heat. Consequently, radiator fluid temperatures will need to exceed ~325 K (assuming standard spacecraft radiator optical properties) in order to provide positive heat rejection at lunar noon. Such temperatures are too high for crewed spacecraft applications in which a heat pump is to be avoided.

Spacecraft that must operate in cold environments at reduced heat load are at risk of radiator freezing. For a vehicle that lands at the Lunar South Pole, the design thermal environment is 215 K, but the radiator working fluid must also be kept from freezing during the 0 K sink of transit. A radiator bypass flow setpoint control design such as those used on the Space Shuttle Orbiter and ISS would require more than 30% of the design heat load to avoid radiator freezing during transit - even with a very low freezing point working fluid. By changing the traditional active thermal control system (ATCS) architecture to include a regenerating heat exchanger inboard of the radiator and using a regenerator bypass flow control valve to maintain system setpoint, the required minimum system heat load can be reduced by more than half. This gives the spacecraft much more flexibility in design and operation. The present work describes the regenerator bypass ATCS setpoint control methodology.

NASA has unique requirements for the development and application of air quality standards for human space flight. Such standards must take into account the continuous nature of exposures, the possibility of increased susceptibility of crewmembers to the adverse effects of air pollutants because of the stresses of space flight, and the recognition that rescue options may be severely limited in remote habitats. NASA has worked with the National Research Council Committee on Toxicology (NRCCOT) since the early 1990s to set and document appropriate standards. The process has evolved through 2 rounds. The first was to set standards for the space station era, and the second was to set standards for longer stays in space and update the original space station standards. The update was to be driven by new toxicological data and by new methods of risk assessment for predicting safe levels from available data. The last phase of this effort has been completed.

Designers of future space vehicles envision simplifying the Atmosphere Revitalization (AR) system by combining the functions of trace contaminant (TC) control and carbon dioxide removal into one swing-bed system. Flow rates and bed sizes of the TC and CO2 systems have historically been very different. There is uncertainty about the ability of trace contaminant sorbents to adsorb adequately in a high-flow or short bed length configurations, and to desorb adequately during short vacuum exposures. This paper describes preliminary results of a comparative experimental investigation into adsorbents for trace contaminant control. Ammonia sorbents and low temperature catalysts for CO oxidation are the foci. The data will be useful to designers of AR systems for Constellation. Plans for extended and repeated vacuum exposure of ammonia sorbents are also presented.

Orion is the next vehicle for human space travel. Humans will be sustained in space by the Orion subystem, environmental control and life support (ECLS). The ECLS concept at the subsystem level is outlined by function and technology. In the past two years, the interface definition with other subsystems has increased through different integrated studies. The paper presents the key requirements and discusses three recent studies (e.g., unpressurized cargo) along with the respective impacts on the ECLS design moving forward.

Advancements in electrical, mechanical, and structural design onboard modern more electric aircraft have added significant stress to the thermal management systems (TMS). A thermal management system level analysis tool has been created in MATLAB/Simulink to facilitate rapid system analysis and optimization to meet the growing demands of modern aircraft. It is anticipated that the tracking of thermal energy through numerical integration will lead to more accurate predictions of worst case TMS sizing conditions. In addition, the non-proprietary nature of the tool affords users the ability to modify component models and integrate advanced conceptual designs that can be evaluated over multiple missions to determine the impact at a system level.

The On-line Project Information System (OPIS) is the Exploration Life Support (ELS) mechanism for task data sharing and annual reporting. Fiscal year 2008 (FY08) was the first year in which ELS Principal Investigators (PI's) were required to complete an OPIS annual report. The reporting process consists of downloading a template that is customized to the task deliverable type(s), completing the report, and uploading the document to OPIS for review and approval. In addition to providing a general status and overview of OPIS features, this paper describes the user critiques and resulting system modifications of the first year of OPIS reporting efforts. Specifically, this paper discusses process communication and logistics issues, user interface ambiguity, report completion challenges, and the resultant or pending system improvements designed to circumvent such issues for the fiscal year 2009 reporting effort.

To work outside a space craft, humans must wear a protective suit. The required suit pressurization creates additional resistance for the wearer while performing work. How much does the suit effect work and fatigue? To answer these questions, dynamic torque was collected for the shoulder, elbow and wrist for six subjects in an Extra-vehicular Mobility Unit (EMU). In order to quantify fatigue, the subjects were to exert maximum voluntary torque for five minutes or until their maximum fell below 50% of their initial maximum for three consecutive repetitions. Using the collected torque and time data, logarithmic based functions were derived to estimate torque decay to within an absolute error of 20%. These results will be used in the development of a generalized tool for prediction of maximum available torque over time for humans using the current EMU.