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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 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

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.

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.

Smoke transport and detection were modeled numerically in the ISS Destiny module using the NIST, Fire Dynamics Simulator code. The airflows in Destiny were modeled using the existing flow conditions and the module geometry included obstructions that simulate the currently installed hardware on orbit. The smoke source was modeled as a 0.152 by 0.152 m region that emitted smoke particulate ranging from 1.46 to 8.47 mg/s. In the module domain, the smoke source was placed in the center of each Destiny rack location and the model was run to determine the time required for the two smoke detectors to alarm. Overall the detection times were dominated by the circumferential flow, the axial flow from the intermodule ventilation and the smoke source strength.

In carrying out NASA's Vision for Space Exploration, a number of different environments will be encountered that will require the crew to wear a protective space suit. Specifically, four suited mission phases are identified as Launch, Entry & Abort profiles, Contingency 0g (orbital) Extravehicular Activity (EVA), Lunar Surface EVA and Martian Surface EVA. This study presents conceptual design solutions based on a previous architecture assessment that defined space suit operational requirements for four proposed space suit configuration options. In addition, a subset of vehicle interface requirements are defined for enabling umbilical and physical connections between the suits and the various Constellation spacecraft in which they will be used. A summary of the resultant suit and component concepts and vehicle interface definitions is presented.

In this paper, we report on the development of an intelligent system for air quality monitoring and early detection and diagnosis of air contaminants. Optimal identification of contaminants is based upon the use of an Implicit Kalman Filter that uses both experimental data and a theoretical model to obtain optimal estimates. We have developed a three-dimensional unsteady-state model of contaminant transport, which uses a flow field generated numerically for the cabin using a finite element mesh. The optimal contaminant estimates are used as the basis for the detection of a contamination event. The algorithm is shown to distinguish between sensor faults and process faults.

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.

NASA and the FAA are funding the development of ground-based remote sensing systems specifically designed to detect and quantify the icing environment aloft. The goal of the NASA activity is to develop a relatively low cost stand-alone system that can provide practical icing information to the flight community. The goal of the FAA activity is to develop more advanced systems that can identify supercooled large drop (SLD) as well as general icing conditions and be integrated into the existing weather information infrastructure. Both activities utilize combinations of sensing technologies including radar, radiometry, and lidar, along with Internet-available external information such as numerical weather model output where it is found to be useful. In all cases the measured data of environment parameters will need to be converted into a measure of icing hazard before it will be of value to the flying community.

We consider the heat transfer characteristics of an ideal concentric disk used in the Wiped-Film Rotating-Disk (WFRD) evaporator for the Vapor Phase Catalytic Ammonia Removal (VPCAR) water recovery system. A mathematical model is derived to predict the radial temperature distribution and its average over the surface of the disk as a function of system parameters. The model shows self-similarity of the temperature distribution and the existence of a dimensionless parameter S (ratio of heat flux to convection) that can be used as a criterion to optimize the thermal characteristics of the disk in order to approach uniform surface temperature. Comparison of the model to experimental data using global (infrared imager) and local (resistive temperature devices) measurements shows that agreement with the model depends on the ambient condition denoted by the local heat transfer coefficient.

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.

This paper discusses technologies and software tools that are being implemented in a software toolkit currently under development at NASA Glenn Research Center. Its purpose is to help study the effects of icing on airfoil performance and assist with the aerodynamic simulation process which consists of characterization and modeling of ice geometry, application of block topology and grid generation, and flow simulation. Tools and technologies for each task have been carefully chosen based on their contribution to the overall process. For the geometry characterization and modeling, we have chosen an interactive rather than automatic process in order to handle numerous ice shapes. An Appendix presents features of a software toolkit developed to support the interactive process. Approaches taken for the generation of block topology and grids, and flow simulation, though not yet implemented in the software, are discussed with reasons for why particular methods are chosen.

Forty to seventy percent of astronauts and cosmonauts reportedly exhibit undesirable vestibular disturbances during the first few days of exposure to weightlessness, including Space Motion Sickness (SMS) and perceptual illusions. While SMS is the primary concern for short-duration missions, the effect of perceptual illusions during landing may be particularly problematic following long-duration missions such as returning from the International Space Station (ISS), or a Mars mission, where vestibular, perceptual and sensorimotor adaptation to 1g, to 0g, to 0.38g has occurred. The longer the mission, the more complete the adaptation is to hypogravity and the more severe the perceptual errors and sensorimotor control disturbances.

This paper describes the DC bus regulation control algorithm for the NASA flywheel energy storage system during charge, charge reduction and discharge modes of operation. The algorithm was experimentally verified in [1] and this paper presents the necessary models for simulation. Detailed block diagrams of the controller algorithm are given. It is shown that the flywheel system and the controller can be modeled in three levels of detail depending on the type of analysis required. The three models are explained and then compared using simulation results.

From November 2010 until May of 2011, NASA's Icing Remote Sensing System was positioned at Platteville, Colorado between the National Science Foundation's S-Pol radar and Colorado State University's CHILL radar (collectively known as FRONT, or ‘Front Range Observational Network Testbed’). This location was also underneath the flight-path of aircraft arriving and departing from Denver's International Airport, which allowed for comparison to pilot reports of in-flight icing. This work outlines how the NASA Icing Remote Sensing System's derived liquid water content and in-flight icing hazard profiles can be used to provide in-flight icing verification and validation during icing and non-icing scenarios with the purpose of comparing these times to profiles of polarized moment data from the two nearby research radars.

An experimental research effort was begun to develop a database of airplane aerodynamic characteristics with simulated ice accretion over a large range of incidence and sideslip angles. Wind-tunnel testing was performed at the NASA Langley 12-ft Low-Speed Wind Tunnel using a 3.5% scale model of the NASA Langley Generic Transport Model. Aerodynamic data were acquired from a six-component force and moment balance in static-model sweeps from α = -5 to 85 deg. and β = -45 to 45 deg. at a Reynolds number of 0.24x10⁶ and Mach number of 0.06. The 3.5% scale GTM was tested in both the clean configuration and with full-span artificial ice shapes attached to the leading edges of the wing, horizontal and vertical tail. Aerodynamic results for the clean airplane configuration compared favorably with similar experiments carried out on a 5.5% scale GTM.

Sending humans to the moon and Mars in support of NASA’s Vision for Space Exploration (VSE) presents a variety of operational environments in which astronauts will need to wear a space suit, both inside the vehicle and during Extravehicular Activity (EVA). Four feasible suit architectures were proposed by NASA in terms of the number and type of suits needed to enable task performance in scenarios ranging from launch and entry operations to conducting EVA’s in microgravity and on planetary surfaces. This study was aimed at defining space suit operational and functional needs across the spectrum of mission elements called out in the VSE, identifying temporal and technical design drivers, and establishing appropriate trade variables with associated weighting factors for analyzing the proposed architecture options. Recommendations from the analysis are offered for consideration in selecting from the four options.

Tests designed to quantify the gravitational effects on thermal mixing and reactant injection in a Supercritical Water Oxidation (SCWO) reactor have recently been performed in the Zero Gravity Facility (ZGF) at NASA's Glenn Research Center. An artificial waste stream, comprising aqueous mixtures of methanol, was pressurized to approximately 250 atm and then heated to 450°C. After uniform temperatures in the reactor were verified, a controlled injection of air was initiated through a specially designed injector to simulate diffusion limited reactions typical in most continuous flow reactors. Results from a thermal mapping of the reaction zone in both 1-g and 0-g environments are compared. Additionally, results of a numerical model of the test configuration are presented to illustrate first order effects on reactant mixing and thermal transport in the absence of gravity.