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Operational issues encountered by Apollo astronauts relating to lunar dust were catalogued, including material abrasion that resulted in scratches and wear on spacesuit components, ultimately impacting visibility, joint mobility and pressure retention. Standard methods are being developed to measure abrasive wear on candidate construction materials to be used for spacesuits, spacecraft, and robotics. Calibration tests were conducted using a standard diamond stylus scratch tip on the common spacecraft structure aluminum, Al 6061-T6. Custom tips were fabricated from terrestrial counterparts of lunar minerals for scratching Al 6061-T6 and comparing to standard diamond scratches. Considerations are offered for how to apply standards when selecting materials and developing dust mitigation strategies for lunar architecture elements.

Smoke detection experiments were conducted in the Microgravity Science Glovebox (MSG) on the International Space Station (ISS) during Expedition 15 in an experiment entitled Smoke Aerosol Measurement Experiment (SAME). The preliminary results from these experiments are presented. In order to simulate detection of a prefire overheated-material event, samples of five different materials were heated to temperatures below the ignition point. The smoke generation conditions were controlled to provide repeatable sample surface temperatures and air flow conditions. The smoke properties were measured using particulate aerosol diagnostics that measure different moments of the size distribution. These statistics were combined to determine the count mean diameter which can be used to describe the overall smoke distribution.

The Vapor Phase Catalytic Ammonia Removal (VPCAR) system is being developed to recycle water for future NASA Exploration Missions [1,2,3,4,5]. Reduced gravity testing of the VPCAR System has been initiated to identify any potential problems with microgravity operation. Two microgravity testing campaigns have been conducted on NASA's C-9B Reduced Gravity Aircraft. These tests focused on the fluid dynamics of the unit's Wiped-Film Rotating Disk (WFRD) evaporator. The experiments used a simplified system to study the process of forming a thin film on a rotating disk. The configuration simulates the application of feed in the VPCAR's WFRD evaporator. The first round of aircraft testing, which was completed in early 2006, indicated that a problem with microgravity operation of the WFRD existed. It was shown that in reduced gravity the VPCAR wiper did not produce a uniform thin film [6]. The film was thicker near the axis of rotation where centrifugal forces are small.

Spacecraft radiators reject heat to their surroundings. Radiators can be deployable or mounted on the body of the spacecraft. NASA's Crew Exploration Vehicle is to use body mounted radiators. Coatings play an important role in heat rejection. The coatings provide the radiator surface with the desired optical properties of low solar absorptance and high infrared emittance. These specialized surfaces are applied to the radiator panel in a number of ways, including conventional spraying, plasma spraying, or as an appliqué. Not specifically designed for a weathering environment, little is known about the durability of conventional paints, coatings, and appliqués upon exposure to weathering and subsequent exposure to solar wind and ultraviolet radiation exposure. In addition to maintaining their desired optical properties, the coatings must also continue to adhere to the underlying radiator panel.

This paper will discuss the status of the Compact Flash Evaporator System (CFES) development at NASA Glenn. Three alternative heat sink technologies are being developed under Thermal Control for Advanced Capabilities within the Exploration Technology Development Program. One of them is CFES, a spray cooling concept related to the current Space Shuttle Orbiter Flash Evaporator System (FES). In the CFES concept, water is sprayed on the outside of a flat plate heat exchanger, through which flows the vehicle's primary vehicle heat transfer fluid. The steam is then exhausted to space in an open-loop system. Design, fabrication and testing of the CFES at NASA's Glenn Research Center will be reported.

Innovative multi-environment multimode thermal management architecture has been described that is capable of meeting widely varying thermal control requirements of various exploration mission scenarios currently under consideration. The proposed system is capable of operating in a single-phase or two-phase mode rejecting heat to the colder environment, operating in a two-phase mode with heat pump for rejecting heat to a warm environment, as well as using evaporative phase-change cooling for the mission phases where the radiator is incapable of rejecting the required heat. A single fluid loop can be used internal and external to the spacecraft for the acquisition, transport and rejection of heat by the selection of a working fluid that meets NASA safety requirements. Such a system may not be optimal for each individual mode of operation but its ability to function in multiple modes may permit global optimization of the thermal control system.

The NASA Glenn Research Center has been developing a spacecraft open loop spray evaporative heat sink for use in pressure environments near sea-level, where evaporative cooling of water is not effective. The working fluid is R134a, a common refrigerant used in household appliances, considered safe and non-toxic for humans. The concept uses an open loop spray of R134a impinging on a heated flat plate, through which a closed loop of hot coolant flows, having acquired the heat from spacecraft electronics boxes, the cabin heat exchanger, and other heat sources. The latent heat of evaporation cools the outside of the hot plate, and through heat conduction, reduces the temperature of the coolant. The testing at NASA Glenn has used an electrically heated cylindrical copper target to simulate the hot plate. This paper will discuss the R134a feed system, the test matrix, and test results.

Oxygen storage and delivery systems for advanced Lunar Exploration Missions are substantially different than those of the International Space Station (ISS) or Apollo missions. The oxygen must be stored without venting for durations of 180 to 210 days prior to use and then used to supply both the steady, low pressure oxygen for the crew, and the higher-pressure oxygen for the extra-vehicular mobility unit. The baseline design is a high pressure gaseous oxygen storage system. Alternate technologies that may offer substantial advantages in terms of the equivalent system mass over the baseline design are being currently evaluated. This study examines both the supercritical and subcritical liquid oxygen storage options, including one with active cooling using a cryocooler. It is found that an actively cooled sub-critical storage system offered the lowest mass system that could satisfy the requirements.

A fire in spacecraft or habitat supporting NASA's Exploration mission could jeopardize the system, mission, and/or crew. Given adequate measures for fire prevention, the hazard from a fire can be significantly reduced if fire detection is rapid and occurs in the early stages of fire development. The simultaneous detection of both particulate and gaseous products has been proven to rapidly detect fires and accurately distinguish between real fires and nuisance sources. This paper describes the development status of gaseous and particulate sensor elements, integrated sensor systems, and system testing. It is concluded that while development is still necessary, the fundamental approach of smart, miniaturized, multisensor technology has the potential to significantly improve the safety of NASA space exploration systems.

This paper presents results from a smoke detection experiment entitled Smoke Aerosol Measurement Experiment (SAME) which was conducted in the Microgravity Science Glovebox on the International Space Station (ISS) during Expedition 15. Five different materials representative of those found in spacecraft were pyrolyzed at temperatures below the ignition point with conditions controlled to provide repeatable sample surface temperatures and air flow conditions. The sample materials were Teflon®, Kapton®, cellulose, silicone rubber and dibutylphthalate. The transport time from the smoke source to the detector was simulated by holding the smoke in an aging chamber for times ranging from 10 to1800 seconds. Smoke particle samples were collected on Transmission Electron Microscope (TEM) grids for post-flight analysis.

The Vapor Phase Catalytic Ammonia Removal (VPCAR) system is being developed to recycle water for future NASA Exploration Missions. Testing was recently conducted on NASA's C-9B Reduced Gravity Aircraft to determine the microgravity performance of a key component of the VPCAR water recovery system. Six flights were conducted to evaluate the fluid dynamics of the Wiped-Film Rotating Disk (WFRD) distillation component of the VPCAR system in microgravity, focusing on the water delivery method. The experiments utilized a simplified system to study the process of forming a thin film on a disk similar to that in the evaporator section of VPCAR. Fluid issues are present with the current configuration, and the initial alternative configurations were only partial successful in microgravity operation. The underlying causes of these issues are understood, and new alternatives are being designed to rectify the problems.

The history and current status of spacecraft smoke detection is discussed including a review of the state of understanding of the effect of gravity on the resultant smoke particle size. The results from a spacecraft experiment (Comparative Soot Diagnostics (CSD)) which measured microgravity smoke particle sizes are presented. Five different materials were tested producing smokes with different properties including solid aerosol smokes and liquid droplets aerosol smokes. The particulate size distribution for the solid particulate smokes increased substantially in microgravity and the results suggested a corresponding increase for the smokes consisting of a liquid aerosol. A planned follow on experiment that will resolve the issues raised by CSD is presented. Early results from this effort have provided the first measurements of the ambient aerosol environment on the ISS (International Space Station) and suggest that the ISS has very low ambient particle levels.

Packed bed reactors are well known for their vast and diverse applications in the chemical industry; from gas absorption, to stripping, to catalytic conversion. Use of this type of reactor in terrestrial applications has been rather extensive because of their simplicity and relative ease of operation. Developing similar reactors for use in microgravity is critical to many space-based advanced life support systems. However, the hydrodynamics of two-phase flow packed bed reactors in this new environment and the effects of one physicochemical process on another has not been adequately assessed. Surface tension or capillary forces play a much greater role which results in a shifting in flow regime transitions and pressure drop. Results from low gravity experiments related to flow regimes and two-phase pressure drop models are presented in this paper along with a description of plans for a flight experiment on the International Space Station (ISS).

The problem of determining equilibrium temperatures for re-radiating surfaces in space vacuum was analyzed and the resulting mathematical relationships were incorporated in a code to determine space sink temperatures in the solar system. A brief treatment of planetary atmospheres is also included. Temperature values obtained with the code are in good agreement with available spacecraft telemetry and meteorological measurements for Venus and Earth. The code has been used in the design of space power system radiators for future interplanetary missions.

Using a new fiber electrode technology, currently developed and produced by Bekaert Corporation (Bekaert), Electro Energy, Inc., (EEI) Mobile Energy Products Group (formerly, Eagle-Picher Technologies, LLC, Power Systems Department) in Colorado Springs, CO has demonstrated that it is feasible to manufacture flight weight nickel hydrogen cells having about twice the specific energy (80 vs. 40 watt-hr./kg) as state-of-the-art nickel hydrogen cells that are currently flown on geosynchronous communications satellites. Although lithium-ion battery technology has made large in-roads to replace the nickel alkaline technology (nickel-cadmium, nickel-metal hydride), the technology offered here competes with lithium-ion weight and offers alternatives not present in the lithium-ion chemistry such as: ability to undergo a continuous overcharge, reversal on discharge, and sustain rate capability sufficient to start automotive and aircraft engines at subzero temperatures.

The International Space Station Carbon Dioxide Removal Assembly (CDRA) uses regenerable adsorption technology to remove carbon dioxide (CO2) from cabin air. CO2 product water vapor measurements from a CDRA test bed unit at the NASA Marshall Space Flight Center were made using a tunable infrared diode laser differential absorption spectrometer (TILDAS) provided by NASA Glenn Research Center. The TILDAS instrument exceeded all the test specifications, including sensitivity, dynamic range, time response, and unattended operation. During the CO2 desorption phase, water vapor concentrations as low as 5 ppmv were observed near the peak of CO2 evolution, rising to levels of ∼40 ppmv at the end of a cycle. Periods of high water concentration (>100 ppmv) were detected and shown to be caused by an experimental artifact.

Research in microgravity (low-gravity) combustion promises innovations and improvements in fire prevention and response for human-crew spacecraft. Findings indicate that material flammability and fire spread in microgravity are significantly affected by atmospheric flow rate, oxygen concentration, and diluent composition. This information can lead to modifications and correlations to standard material-assessment tests for prediction of fire resistance in space. Research on smoke-particle changes in microgravity promises future improvements and increased sensitivity of smoke detectors in spacecraft. Research on fire suppression by extinguishing agents and venting can yield new information on effective control of the rare, but serious fire events in spacecraft.

The International Space Station (ISS) is a complex spacecraft that will take several years to assemble in orbit. During many of the assembly and maintenance procedures, the space station’s large solar arrays must be locked, which can significantly reduce power generation. To date, power generation analyses have not included power generation from the backside of the solar cells in a desire to produce a conservative analysis. This paper describes the testing of ISS solar cell backside power generation, analytical modeling, and analysis results on an ISS assembly mission.

The Interim Control Module (ICM) is being built by the US Naval Research Laboratory (NRL) for NASA as a propulsion module for the International Space Station (ISS). Originally developed as a spinning spacecraft used to move payloads to their final orbit, for ISS, the ICM will be in a fixed orientation and location for long periods resulting in substantial solar panel shadowing. This paper describes the methods used to determine the incident energy on the ICM solar panels and the power capability of the electric power system (EPS). Applying this methodology has resulted in analyses and assessments used to identify ICM early design changes/options, placement and orientations that enable successful operation of the EPS under a wide variety of anticipated conditions.

NASA Glenn Research Center and the Wallops Flight Facility jointly conducted a PEM fuel cell power system development effort for high altitude balloon applications. This was the first phase of NASA efforts to offer higher balloon payload power levels with extended duration mission capabilities for atmospheric science missions. At present, lead-acid batteries typically supply about 100 watts of power to the balloon payload for approximately 8 hours duration. The H2-O2 PEM fuel cell demonstration system developed for this effort can supply at least 200 watts for 48 hours duration. The system was designed and fabricated, then tested in ambient ground environments as well as in a thermal vacuum chamber to simulate operation at 75 kft. altitude. Initially, this program was planned to culminate with a demonstration flight test but no flight has been scheduled, thus far.