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Search indexed NASA NTRS and DOE OSTI research on propulsion, heat transfer, battery materials and energy systems. Follow report and document links to the original sources.

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NASA NTRS2005

Utilizing Radioisotope Power System Waste Heat for Spacecraft Thermal Management

An advantage of using a Radioisotope Power System (RPS) for deep space or planetary surface missions is the readily available waste heat, which can be used for a number of beneficial purposes including: maintaining electronic components within a controlled temperature range, warming propulsion tanks and mobility actuators, and maintaining liquid propellants above their freezing temperature. Previous missions using Radioisotope Thermoelectric Generators (RTGs) dissipated large quantities of waste heat due to the low efficiency of the thermoelectric conversion technology. The next generation RPSs, such as the 110-Watt Stirling Radioisotope Generator (SRG110) will have higher conversion efficiencies, thereby rejecting less waste heat at a lower temperature and may require alternate approaches to transferring waste heat to the spacecraft. RTGs, with efficiencies of ~6 to 7 percent, reject their waste heat at the relatively high heat rejection temperature of 200 C. This is an advantage when rejecting heat to space; however, transferring heat to the internal spacecraft components requires a large and heavy radiator heat exchanger. At the same time, sensitive spacecraft instruments must be shielded from the thermal radiation of the RTG. The SRG110, with an efficiency around 22 percent and 50 C nominal housing surface temperature, can readily transfer the available waste heat directly via heat pipes, thermal straps, or fluid loops. The lower temperatures associated with the SRG110 avoid the chances of overheating other scientific components, eliminating the need for thermal shields. This provides the spacecraft designers more flexibility when locating the generator for a specific mission. A common misconception with high-efficiency systems is that there is not enough waste heat for spacecraft thermal management. This paper will dispel this misconception and investigate the use of a high-efficiency SRG110 for spacecraft thermal management and outline potential methods of waste heat utilization in several conceptual missions (Lunar Rover, Mars Rover, and Titan Lander). The advantages associated with the SRG110 as they relate to ease of assembly, less complex interfaces, and overall mass savings for a spacecraft will be highlighted.

Pantano, David R.↗
NASA NTRS2005

Utilizing Radioisotope Power System Waste Heat for Spacecraft Thermal Management

One of the advantages of using a Radioisotope Power System (RPS) for deep space or planetary surface missions is the readily available waste heat, which can be used to maintain electronic components within a controlled temperature range, to warm propulsion tanks and mobility actuators, and to gasify liquid propellants. Previous missions using Radioisotope Thermoelectric Generators (RTGs) dissipated a very large quantity of waste heat due to the relatively low efficiency of the thermoelectric conversion technology. The next generation RPSs, such as the 110-watt Stirling Radioisotope Generator (SRG110) will have much higher conversion efficiencies than their predecessors and therefore may require alternate approaches to transferring waste heat to the spacecraft. RTGs, with efficiencies of approx. 6 to 7% and 200 C housing surface temperatures, would need to use large and heavy radiator heat exchangers to transfer the waste heat to the internal spacecraft components. At the same time, sensitive spacecraft instruments must be shielded from the thermal radiation by using the heat exchangers or additional shields. The SRG110, with an efficiency around 22% and 50 C nominal housing surface temperature, can use the available waste heat more efficiently by more direct heat transfer methods such as heat pipes, thermal straps, or fluid loops. The lower temperatures allow the SRG110 much more flexibility to the spacecraft designers in configuring the generator without concern of overheating nearby scientific instruments, thereby eliminating the need for thermal shields. This paper will investigate using a high efficiency SRG110 for spacecraft thermal management and outline potential methods in several conceptual missions (Lunar Rover, Mars Rover, and Titan Lander) to illustrate the advantages with regard to ease of assembly, less complex interfaces, and overall mass savings.

Pantano, David R.↗
NASA NTRS2005

Brain-Computer Interfaces for 1-D and 2-D Cursor Control: Designs Using Volitional Control of the EEG Spectrum or Steady-State Visual Evoked Potentials

We have developed and tested two EEG-based brain-computer interfaces (BCI) for users to control a cursor on a computer display. Our system uses an adaptive algorithm, based on kernel partial least squares classification (KPLS), to associate patterns in multichannel EEG frequency spectra with cursor controls. Our first BCI, Target Practice, is a system for one-dimensional device control, in which participants use biofeedback to learn voluntary control of their EEG spectra. Target Practice uses a KF LS classifier to map power spectra of 30-electrode EEG signals to rightward or leftward position of a moving cursor on a computer display. Three subjects learned to control motion of a cursor on a video display in multiple blocks of 60 trials over periods of up to six weeks. The best subject s average skill in correct selection of the cursor direction grew from 58% to 88% after 13 training sessions. Target Practice also implements online control of two artifact sources: a) removal of ocular artifact by linear subtraction of wavelet-smoothed vertical and horizontal EOG signals, b) control of muscle artifact by inhibition of BCI training during periods of relatively high power in the 40-64 Hz band. The second BCI, Think Pointer, is a system for two-dimensional cursor control. Steady-state visual evoked potentials (SSVEP) are triggered by four flickering checkerboard stimuli located in narrow strips at each edge of the display. The user attends to one of the four beacons to initiate motion in the desired direction. The SSVEP signals are recorded from eight electrodes located over the occipital region. A KPLS classifier is individually calibrated to map multichannel frequency bands of the SSVEP signals to right-left or up-down motion of a cursor on a computer display. The display stops moving when the user attends to a central fixation point. As for Target Practice, Think Pointer also implements wavelet-based online removal of ocular artifact; however, in Think Pointer muscle artifact is controlled via adaptive normalization of the SSVEP. Training of the classifier requires about three minutes. We have tested our system in real-time operation in three human subjects. Across subjects and sessions, control accuracy ranged from 80% to 100% correct with lags of 1-5 seconds for movement initiation and turning.

Trejo, Leonard J.↗
NASA NTRS2007

Cardiac and Vascular Function in Bedrested Volunteers: Effects of Artificial Gravity Training

Cardiovascular effects of an artificial gravity (AG) countermeasure on deconditioned male volunteers were studied. In two groups of men we measured cardiovascular parameters at rest and in response to 30 minutes of 80 deg. head up tilt (HUT) before, at the end of, and four days following 21 days of 6 deg. head down bed rest (HDBR). One group (N=7) underwent no countermeasure while the other (N=8) received a daily, one hour, dose (2.5 gz at the foot, decreasing to 1.0 gz at the heart) of AG training on the Johnson Space Center short radius centrifuge. Cardiovascular parameters measured included heart rate, blood pressure, stroke volume, cardiac output, peripheral vascular resistance, plasma volume shifts, and vasoactive hormones. Untrained subjects exhibited shorter tilt survival (on average 8 minutes shorter) compared to trained subjects. By the end of bed rest, mean heart rate (MHR) was elevated in both groups (both supine and during tilt). In addition, treated subjects demonstrated lower, tilt-induced, increases in MHR four days following HDBR, indicating a more rapid return to pre bed rest conditions. Results from an index of autonomic balance (percentage of MHR spectral power in the respiratory frequency range) in control of heart rate are consistent with the interpretation that parasympathetic nervous system withdrawal was responsible for both tilt- and bed rest-induced increases in MHR. Our data support our pre-study hypothesis that AG treatment would lessen cardiovascular effects of deconditioning in bed rested men and suggest that AG should be further pursued as a space flight countermeasure.

Meng, M.↗
NASA NTRS2007

Genesis Contingency Planning and Mishap Recovery: The Sample Curation View

Planning for sample preservation and curation was part of mission design from the beginning. One of the scientific objectives for Genesis included collecting samples of three regimes of the solar wind in addition to collecting bulk solar wind during the mission. Collectors were fabricated in different thicknesses for each regime of the solar wind and attached to separate frames exposed to the solar wind during specific periods of solar activity associated with each regime. The original plan to determine the solar regime sampled for specific collectors was to identify to which frame the collector was attached. However, the collectors were dislodged during the hard landing making identification by frame attachment impossible. Because regimes were also identified by thickness of the collector, the regime sampled is identified by measuring fragment thickness. A variety of collector materials and thin films applied to substrates were selected and qualified for flight. This diversity provided elemental measurement in more than one material, mitigating effects of diffusion rates and/or radiation damage. It also mitigated against different material and substrate strengths resulting in differing effects of the hard landing. For example, silicon crystal substrates broke into smaller fragments than sapphire-based substrates and diamond surfaces were more resilient to flying debris damage than gold. The primary responsibility of the curation team for recovery was process documentation. Contingency planning for the recovery phase expanded this responsibility to include not only equipment to document, but also gather, contain and identify samples from the landing area and the recovered spacecraft. The team developed contingency plans for various scenarios as part of mission planning that included topographic maps to aid in site recovery and identification of different modes of transport and purge capability depending on damage. A clean tent, set-up at Utah Test & Training Range to control the environment for processing the sample return capsule and cleanly installing a nitrogen purge to the canister, was used to control the environment for extracting collector fragments from the damaged canister and to document and package over 10,000 collector fragments.

Stansbery, E. K.↗
NASA NTRS2008

Space Technology 5 Observations of Auroral Field-Aligned Currents

During its three month long technology validation mission, Space Technology 5 (ST-5) returned high quality multi-point measurements of the near-Earth magnetic field. Its three micro-satellites were launched into a 300 x 4500 km, dawn - dusk, sun synchronous orbit (inclination = 105.60) orbit with a period of 138 min by a Pegasus launch vehicle on March 22, 2006. The spacecraft were maintained in a "pearls on a sting" constellation with controlled spacings ranging from just over 5000 km down to under 50 km. The individual micro-satellites were 48 cm tall octagons with diameters of 50 cm. They were spin-stabilized at approximately 20 rpm at deployment and slowly spun-down to about 15 rpm by the end of the mission. Each spacecraft carried a miniature tri-axial fluxgate magnetometer (MAG) provided by the University of California at Los Angeles mounted at the end of a ultra-low mass 72 cm boom. These data allow, for the first time, the separation of temporal and spatial variations in field-aligned current (FAC) perturbations measured in low-Earth orbit on time scales of 10 sec to 10 min. The constellation measurements are used to directly determine field-aligned current sheet motion, thickness. and current density. Two multi-point methods for the inference of FAC current density that have not previously been possible in low-Earth orbit are demonstrated: 1) the -standard method." based upon s/c velocity, but corrected for FAC current sheet motion. and 2) the "gradiometer method" which uses simultaneous magnetic field measurements at two points with known separation. Future studies will apply these methods to the entire ST-5 data sct and expand to include horizontal ionospheric currents. ULF waves and geomagnetic field gradient analyses.

Slavin, James↗
NASA NTRS2009

Spacecraft Thermal Management

In the 21st century, the National Aeronautics and Space Administration (NASA), the Russian Federal Space Agency, the National Space Agency of Ukraine, the China National Space Administration, and many other organizations representing spacefaring nations shall continue or newly implement robust space programs. Additionally, business corporations are pursuing commercialization of space for enabling space tourism and capital business ventures. Future space missions are likely to include orbiting satellites, orbiting platforms, space stations, interplanetary vehicles, planetary surface missions, and planetary research probes. Many of these missions will include humans to conduct research for scientific and terrestrial benefits and for space tourism, and this century will therefore establish a permanent human presence beyond Earth s confines. Other missions will not include humans, but will be autonomous (e.g., satellites, robotic exploration), and will also serve to support the goals of exploring space and providing benefits to Earth s populace. This section focuses on thermal management systems for human space exploration, although the guiding principles can be applied to unmanned space vehicles as well. All spacecraft require a thermal management system to maintain a tolerable thermal environment for the spacecraft crew and/or equipment. The requirements for human rating and the specified controlled temperature range (approximately 275 K - 310 K) for crewed spacecraft are unique, and key design criteria stem from overall vehicle and operational/programatic considerations. These criteria include high reliability, low mass, minimal power requirements, low development and operational costs, and high confidence for mission success and safety. This section describes the four major subsystems for crewed spacecraft thermal management systems, and design considerations for each. Additionally, some examples of specialized or advanced thermal system technologies are presented, which may be enabling to future space missions never before attempted like a crewed mission to Mars.

Hurlbert, Kathryn Miller↗
NASA NTRS2010

STS 132 Return Samples: Assessment of Air Quality Aboard the Shuttle (STS-132) and International Space Station (ULF4)

The toxicological assessments of 2 grab sample canisters (GSCs) from the Shuttle are reported. Analytical methods have not changed from earlier reports. The recoveries of the 3 surrogates (13C-acetone, fluorobenzene, and chlorobenzene) from the 2 Shuttle GSCs averaged 93, 85%, and 88%, respectively. Based on the end-of-mission sample, the Shuttle atmosphere was acceptable for human respiration. The toxicological assessment of 7 GSCs from the ISS is also shown. The recoveries of the 3 standards (as listed above) from the GSCs averaged 78, 96 and 90%, respectively. Recovery from formaldehyde control badges ranged from 90 to 112%.

James. John T.↗
NASA NTRS2011

Wide-Field Optic for Autonomous Acquisition of Laser Link

An innovation reported in Two-Camera Acquisition and Tracking of a Flying Target, NASA Tech Briefs, Vol. 32, No. 8 (August 2008), p. 20, used a commercial fish-eye lens and an electronic imaging camera for initially locating objects with subsequent handover to an actuated narrow-field camera. But this operated against a dark-sky background. An improved solution involves an optical design based on custom optical components for the wide-field optical system that directly addresses the key limitations in acquiring a laser signal from a moving source such as an aircraft or a spacecraft. The first challenge was to increase the light collection entrance aperture diameter, which was approximately 1 mm in the first prototype. The new design presented here increases this entrance aperture diameter to 4.2 mm, which is equivalent to a more than 16 times larger collection area. One of the trades made in realizing this improvement was to restrict the field-of-view to +80 deg. elevation and 360 azimuth. This trade stems from practical considerations where laser beam propagation over the excessively high air mass, which is in the line of sight (LOS) at low elevation angles, results in vulnerability to severe atmospheric turbulence and attenuation. An additional benefit of the new design is that the large entrance aperture is maintained even at large off-axis angles when the optic is pointed at zenith. The second critical limitation for implementing spectral filtering in the design was tackled by collimating the light prior to focusing it onto the focal plane. This allows the placement of the narrow spectral filter in the collimated portion of the beam. For the narrow band spectral filter to function properly, it is necessary to adequately control the range of incident angles at which received light intercepts the filter. When this angle is restricted via collimation, narrower spectral filtering can be implemented. The collimated beam (and the filter) must be relatively large to reduce the incident angle down to only a few degrees. In the presented embodiment, the filter diameter is more than ten times larger than the entrance aperture. Specifically, the filter has a clear aperture of about 51 mm. The optical design is refractive, and is comprised of nine custom refractive elements and an interference filter. The restricted maximum angle through the narrow-band filter ensures the efficient use of a 2-nm noise equivalent bandwidth spectral width optical filter at low elevation angles (where the range is longest), at the expense of less efficiency for high elevations, which can be tolerated because the range at high elevation angles is shorter. The image circle is 12 mm in diameter, mapped to 80 x 360 of sky, centered on the zenith.

Page, Norman A.↗
NASA NTRS2011

An Experimental Study of Upward Burning Over Long Solid Fuels: Facility Development and Comparison

As NASA's mission evolves, new spacecraft and habitat environments necessitate expanded study of materials flammability. Most of the upward burning tests to date, including the NASA standard material screening method NASA-STD-6001, have been conducted in small chambers where the flame often terminates before a steady state flame is established. In real environments, the same limitations may not be present. The use of long fuel samples would allow the flames to proceed in an unhindered manner. In order to explore sample size and chamber size effects, two large chambers were developed at NASA GRC under the Flame Prevention, Detection and Suppression (FPDS) project. The first was an existing vacuum facility, VF-13, located at NASA John Glenn Research Center. This 6350 liter chamber could accommodate fuels sample lengths up to 2 m. However, operational costs and restricted accessibility limited the test program, so a second laboratory scale facility was developed in parallel. By stacking additional two chambers on top of an existing combustion chamber facility, this 81 liter Stacked-chamber facility could accommodate a 1.5 m sample length. The larger volume, more ideal environment of VF-13 was used to obtain baseline data for comparison with the stacked chamber facility. In this way, the stacked chamber facility was intended for long term testing, with VF-13 as the proving ground. Four different solid fuels (adding machine paper, poster paper, PMMA plates, and Nomex fabric) were tested with fuel sample lengths up to 2 m. For thin samples (papers) with widths up to 5 cm, the flame reached a steady state length, which demonstrates that flame length may be stabilized even when the edge effects are reduced. For the thick PMMA plates, flames reached lengths up to 70 cm but were highly energetic and restricted by oxygen depletion. Tests with the Nomex fabric confirmed that the cyclic flame phenomena, observed in small facility tests, continued over longer sample. New features were also observed at the higher oxygen/pressure conditions available in the large chamber. Comparison of flame behavior between the two facilities under identical conditions revealed disparities, both qualitative and quantitative. This suggests that, in certain ranges of controlling parameters, chamber size and shape could be one of the parameters that affect the material flammability. If this proves to be true, it may limit the applicability of existing flammability data.

Kleinhenz, Julie↗
NASA NTRS2011

Design and Analysis of Mirror Modules for IXO and Beyond

Advancements in X-ray astronomy demand thin, light, and closely packed thin optics which lend themselves to segmentation of the annular mirrors and, in turn, a modular approach to the mirror design. The functionality requirements of such a mirror module are well understood. A baseline modular concept for the proposed International X-Ray Observatory (IXO) Flight Mirror Assembly (FMA) consisting of 14,000 glass mirror segments divided into 60 modules was developed and extensively analyzed. Through this development, our understanding of module loads, mirror stress, thermal performance, and gravity distortion have greatly progressed. The latest progress in each of these areas is discussed herein. Gravity distortion during horizontal X-ray testing and on-orbit thermal performance have proved especially difficult design challenges. In light of these challenges, fundamental trades in modular X-ray mirror design have been performed. Future directions in module X-ray mirror design are explored including the development of a 1.8 m diameter FMA utilizing smaller mirror modules. The effect of module size on mirror stress, module self-weight distortion, thermal control, and range of segment sizes required is explored with advantages demonstrated from smaller module size in most cases.

McClelland, Ryan S.↗
NASA NTRS2012

Imidazolium-Based Polymeric Materials as Alkaline Anion-Exchange Fuel Cell Membranes

Polymer electrolyte membranes that conduct hydroxide ions have potential use in fuel cells. A variety of polystyrene-based quaternary ammonium hydroxides have been reported as anion exchange fuel cell membranes. However, the hydrolytic stability and conductivity of the commercially available membranes are not adequate to meet the requirements of fuel cell applications. When compared with commercially available membranes, polystyrene-imidazolium alkaline membrane electrolytes are more stable and more highly conducting. At the time of this reporting, this has been the first such usage for imidazolium-based polymeric materials for fuel cells. Imidazolium salts are known to be electrochemically stable over wide potential ranges. By controlling the relative ratio of imidazolium groups in polystyrene-imidazolium salts, their physiochemical properties could be modulated. Alkaline anion exchange membranes based on polystyrene-imidazolium hydroxide materials have been developed. The first step was to synthesize the poly(styrene-co-(1-((4-vinyl)methyl)-3- methylimidazolium) chloride through a free-radical polymerization. Casting of this material followed by in situ treatment of the membranes with sodium hydroxide solutions provided the corresponding hydroxide salts. Various ratios of the monomers 4-chloromoethylvinylbenzine (CMVB) and vinylbenzine (VB) provided various compositions of the polymer. The preferred material, due to the relative ease of casting the film, and its relatively low hygroscopic nature, was a 2:1 ratio of CMVB to VB. Testing confirmed that at room temperature, the new membranes outperformed commercially available membranes by a large margin. With fuel cells now in use at NASA and in transportation, and with defense potential, any improvement to fuel cell efficiency is a significant development.

Narayan, Sri R.↗
NASA NTRS2013

Post-Flight EDL Entry Guidance Performance of the 2011 Mars Science Laboratory Mission

The 2011 Mars Science Laboratory was the first Mars guided entry which safely delivered the rover to a landing within a touchdown ellipse of 19.1 km x 6.9 km. The Entry Terminal Point Controller guidance algorithm is derived from the final phase Apollo Command Module guidance and, like Apollo, modulates the bank angle to control the range flown. The guided entry performed as designed without any significant exceptions. The Curiosity rover was delivered about 2.2 km from the expected touchdown. This miss distance is attributed to little time to correct the downrange drift from the final bank reversal and a suspected tailwind during heading alignment. The successful guided entry for the Mars Science Laboratory lays the foundation for future Mars missions to improve upon.

Mendeck, Gavin F.↗
NASA NTRS2015

The Age of Lunar South Circumpolar Craters Haworth, Shoemaker, Faustini, and Shackleton: Implications for Regional Geology, Surface Processes, and Volatile Sequestration

The interiors of the lunar south circumpolar craters Haworth, Shoemaker, Faustini, and Shackleton contain permanently shadowed regions (PSRs) and have been interpreted to contain sequestered volatiles including water ice. Altimetry data from the Lunar Orbiter Laser Altimeter (LOLA) onboard the Lunar Reconnaissance Orbiter provide a new means of examining the permanently shadowed interiors of these craters in unprecedented detail. In this study, we used extremely high-resolution gridded LOLA data of Haworth, Shoemaker, Faustini, and Shackleton to determine the size-frequency distributions and the spatial density of craters superposing their rims, inner slopes, and floors. Based on their population of superposed D greater than or equal to 2 km craters, Haworth, Shoemaker, and Faustini have pre-Nectarian formation ages. Shackleton is interpreted as having a Late Imbrian age on the basis of craters with diameter D greater than or equal to 0.5 km superposed on its rim. The local density of craters with sub-km diameters across our study area is strongly dependent on slope; because of its steep interior slopes, the lifetime of craters on the interior of Shackleton is limited. The slope-dependence of the small crater population implies that the population in this size range is controlled primarily by the rate at which craters are destroyed. This is consistent with the hypothesis that crater removal and resurfacing is a result of slopedependent processes such as diffusive mass wasting and seismic shaking, linked to micrometeorite and meteorite bombardment. Epithermal neutron flux data and UV albedo data show that these circumpolar PSRs, particularly Shoemaker, may have approximately 1-2% water ice by mass in their highly porous surface regolith, and that Shoemaker may have approximately 5% or more water ice by mass in the near subsurface. The ancient formation ages of Shoemaker, Faustini and Haworth, and the Late Imbrian (approximately 3.5 Ga) crater retention ages of their floors suggests that any water ice that might have been deposited in their permanently shadowed areas was insufficient to modify the superposed crater population since that time.

Circumpolar↗
NASA NTRS2016

Concept of Operations for a Prospective "Proving Ground" in the Lunar Vicinity

NASA is studying conceptual architectures for a "Proving Ground" near the Moon or in high lunar orbit to conduct human space exploration missions that bridge the gap between today's operations with the International Space Station (ISS) and future human exploration of Mars beginning in the 2030s. This paper describes the framework of a concept of operations ("Conops") for candidate activities in the Proving Ground. The Conops discusses broad goals that the Proving Ground might address, such as participation from commercial entities, support for human landings on the Moon, use of mature technologies, and growth of capability through a steady cadence of increasingly ambitious piloted missions. Additional Proving Ground objectives are outlined in a companion paper. Key elements in the Conops include the Orion spacecraft (with mission kits for docking and other specialized operations) and the Space Launch System (SLS) heavy-lift rocket. Potential additions include a new space suit, commercial launch vehicles and logistics carriers, Solar Electric Propulsion (SEP) stages to move elements between different orbits and eventually take them on excursions to deep space, a core module with multiple docking ports, a habitation block, and robotic and piloted lunar landers. The landers might include reusable ascent modules which could remain docked to in-space elements between lunar sorties. A module providing advanced regenerative life support functions could launch to the ISS, and later move to the Proving Ground. The architecture will include infrastructure for launch preparation, communication, mission control, and range safety. The Conops describes notional missions chosen to guide the design of the architecture and its elements. One such mission might be the delivery of a approximately 10-t Transit Habitat element, comanifested with Orion on a Block 1B SLS launcher, to the Proving Ground. In another mission, the architecture might participate in direct human exploration of an asteroidal boulder brought to high lunar orbit by the Asteroid Redirect Mission. The Proving Ground stack could serve as a staging point and tele-operation center for robotic and piloted Moon landings. With the addition of a SEP stage, the architecture could support months-long excursions within and beyond the Earth's sphere of influence, possibly culminating in a year-long mission to land humans on a near-Earth asteroid. In the last case, after returning to near-lunar space, two of the asteroid explorers could join two crewmembers freshly arrived from Earth for a Moon landing, helping to quantify the risk of landing deconditioned crews on Mars. In a conceptual mission particularly stressing to system design, Proving Ground elements could transit to Mars orbit. Other possible design-driving operations include relocation of the stack with no crew on board, the unpiloted journey of the advanced life support module from ISS to the lunar vicinity, excursions to other destinations in near-Earth space, and additional support for Mars exploration in conjunction with the Evolvable Mars Campaign. The Proving Ground Conops concludes with a discussion of aborts and contingency operations

Love, Stanley G.↗
NASA NTRS2016

Concept of Operations for a Prospective "Proving Ground" in the Lunar Vicinity

NASA is studying a "Proving Ground" near the Moon to conduct human space exploration missions in preparation for future flights to Mars. This paper describes a concept of operations ("conops") for activities in the Proving Ground, focusing on the construction and use of a mobile Cislunar Transit Habitat capable of months-long excursions within and beyond the Earth-Moon system. Key elements in the conops include the Orion spacecraft (with mission kits for docking and other specialized operations) and the Space Launch System heavy-lift rocket. Potential additions include commercial launch vehicles and logistics carriers, solar electric propulsion stages to move elements between different orbits and eventually take them on excursions to deep space, a node module with multiple docking ports, habitation and life support blocks, and international robotic and piloted lunar landers. The landers might include reusable ascent modules which could remain docked to in-space elements between lunar sorties. The architecture will include infrastructure for launch preparation, communication, mission control, and range safety. The conops describes "case studies" of notional missions chosen to guide the design of the architecture and its elements. One such mission is the delivery of a ~10-ton pressurized element, co-manifested with an Orion on a Block 1B Space Launch System rocket, to the Proving Ground. With a large solar electric propulsion stage, the architecture could enable a year-long mission to land humans on a near-Earth asteroid. In the last case, after returning to near-lunar space, two of the asteroid explorers could join two crewmembers freshly arrived from Earth for a Moon landing, helping to safely quantify the risk of landing deconditioned crews on Mars. The conops also discusses aborts and contingency operations. Early return to Earth may be difficult, especially during later Proving Ground missions. While adding risk, limited-abort conditions provide needed practice for Mars, from which early return is likely to be impossible.

Love, Stanley G.↗
NASA NTRS2022

Dynamic Path Planning Automation Concept for Advanced Air Mobility

Advanced Air Mobility (AAM) aims to develop an air transportation system for novel air vehicles between local, regional, intraregional, and urban places. Safety and efficiency of increasingly complex AAM operations are expected to require extensive use of automation, ranging from controlling the revolutionary new aircraft to managing flights dynamically in the high tempo airspace and aerodrome operations. Both onboard and ground automation will play central roles in assisting AAM operators with managing the flight paths of their fleet. This document presents a concept for dynamic path planning (DPP) automation applicable to AAM and other flight operations. The role of the DPP automation system is fivefold: (1) it creates a flight path with desired qualities of being feasible, deconflicted, harmonized, flexible, and optimal; (2) it monitors the progress of flight in a dynamic operating environment; (3) it supports the user in evaluating continued acceptability of the flight path in changing conditions; (4) it revises the flight path as needed to maintain the desired flight path qualities; and (5) it coordinates the flight path with airspace users and service providers. Key users of the DPP automation system include flight planners, pilots, and airspace service providers The concept allows for the system to be installed onboard the aircraft as well as on the ground. The system responds automatically to the dynamic operating environment to ensure that a safe and operationally acceptable flight path is available throughout the flight.

Vivek Sharma↗
DOE OSTI2026

Beam-water Interaction System for PFAS Treatment

Per- and polyfluoroalkyl substances (PFAS) are persistent environmental contaminants that resist conventional water treatment due to the strength of their carbon fluorine bonds. Electron-beam irradiation has shown potential for PFAS degradation; however, its effectiveness is limited by shallow energy penetration, restricting treatment to the upper region of the water surface. This project addresses this limitation through the design of a compact mixing chamber and recirculation system to enhance surface renewal and improve treatment efficiency. The system, constrained to approximately 10 cm 20 cm 20 cm, was developed using computer-aided design, computational fluid dynamics (CFD), and MATLAB-based hydraulic analysis to evaluate flow behavior, head loss, and pump compatibility. A perforated-baffle geometry and controlled flow range of 1 3 GPM were implemented to promote uniform mixing while maintaining a stable free surface for electron-beam exposure. Modeling results indicate improved flow distribution and reduced dead zones, supporting more consistent interaction between contaminated water and the beam. The final design enhances hydrodynamic performance within system constraints, contributing to scalable electron-beam water treatment technologies for improved PFAS degradation.

Ramos, Gerardo [Northern Illinois U.]↗