Search NASA⌕ Search

SEARCH · Search NASA

Results for “LUNAR PHASE”

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.

Quote a phrase for an exact phrase match. Source license links do not imply unrestricted reuse.

At least 685 records · Page 38
All sourcesNASA NTRSDOE OSTI
NASA NTRS2012

Experimental Investigation of Transient Sublimator Performance

Sublimators have been used as heat rejection devices for a variety of space applications including the Apollo Lunar Module and the Extravehicular Mobility Unit (EMU). Sublimators typically operate with steady-state feedwater utilization at or near 100%. However, sublimators are currently being considered for operations in a cyclical topping mode, which represents a new mode of operation for sublimators. Sublimators can be used as a supplemental heat rejection device during mission phases where the environmental temperature or heat rejection requirement changes rapidly. This scenario may occur during low lunar orbit, low earth orbit, or other planetary orbits. In these mission phases, the need for supplemental heat rejection will vary between zero and some fraction of the overall heat load. In particular, supplemental heat rejection is required for the portion of the orbit where the radiative sink temperature exceeds the system setpoint temperature. This paper will describe the effects of these transient starts and stops on the feedwater utilization during various feedwater timing scenarios. Experimental data from various scenarios is analyzed to investigate feedwater consumption efficiency under the cyclical conditions. Start up utilization tests were conducted to better understand the transient performance. This paper also provides recommendations for future sublimator design and transient operation.

Sheth, Rubik B.↗
NASA NTRS2017

Lunar COTS Mission Concept: Using the Moon's Resources to Enable an Economical and Sustainable Pathway to Mars and Beyond

To support and enable future human missions to the surface of the Moon, a new plan was formulated to gradually develop an economical, evolvable and sustainable lunar infrastructure using a public/private partnerships between NASA and industry to share cost and risk in the development phase and ultimately transfer operation of these infrastructure services to its industry owners. These infrastructure services may include but are not limited to the following: lunar cargo transportation, power stations, energy storage devices, communication towers and relay satellites, and resource extraction operations. The plan is called the Commercial Orbital Transfer Services (COTS) Program, using NASA's well-proven Commercial Orbital Transportation Services (COTS) Program acquisition model.

Zuniga, A. F.↗
NASA NTRS2024

EVA Operations for the Artemis III Science Team

The purpose of this presentation is to give the newly selected Artemis III Science Team some basic information on EVA Operations and potential interactions with the science team in both the planning and real-time phases of the mission.

EVA↗
NASA NTRSDate not supplied

EVA Operations for Artemis III External Teams

The purpose of this presentation is to give external Artemis III Teams (payloads, industry) some basic information on EVA Operations and hardware impacts in the design, the planning, and the real-time phases of the mission.

EVA↗
NASA NTRS1990

Lunar Observer - Scouting for a moon base

The proposed 1996 Lunar Observer mission to the moon is described. Lunar Observer is intended to acquire the definitive data set of the moon from which an enduring and flexible lunar base site selection can be made. A mission profile is summarized and mission phases defined. The fourteen proposed experiments are described and their measurement capabilities outlined. The spacecraft, a derivation of the planned Mars Observer spacecraft design, is described. A description is included of a subsatellite, which is used to obtain the first measurements of the lunar far side gravity field.

Parker, G. L.↗
NASA NTRS2025

South Pole Lunar Lighting Studies for Driving Exploration on the Lunar Surface

NASA’s Artemis lunar missions will face new exploration challenges due to inherently low sun angles in the lunar south pole region. Whereas the Apollo missions were afforded sunlight for approximately 8 to 50-degrees above the horizon, sunlight for the next lunar missions will only be 1 to 2-degrees about the horizon. With this low angle, long shadows and high contrasts of light and dark areas will be faced by the crew and remote teleop operators while exploring the lunar surface. With this in mind, NASA developed an integrated virtual Lighting and Navigation Simulation to understand these challenges. Two studies have been conducted in this Lunar South Pole environment to evaluate the effects of natural and artificial lighting on driving and navigating a lunar rover across the surface. Early NASA studies, such as this, are used to aid in developing techniques and explore concepts of operations to promote mission success and crew safety. In the Phase One development evaluation, six astronauts were teleported to ten different lighting conditions. For each location, drivers were to drive to an imaginary target approximately 200-meters straight ahead and provide subjective feedback on their ability to drive under the given lighting conditions. Phase Two, a more operational study, four astronauts and a remote operator tested five different Artemis lunar mission scenarios. Results indicated sun direction at such low angles, especially when driving a high speed, can severely impact the crew’s ability to safely drive the rover. In the up-sun situation, with the sun directly in the driver’s eyes and compounded long shadows, the consistent preference among drivers was to initiate a tacking strategy of approximately +/- 20 to 30-degrees to improve visibility. This maneuver does require more time and rover energetics. Conversely, driving down-sun required the crew to tack as well to avoid the shadow of rover blocking the terrain. Less appreciated is how the surrounding landscape is lit. Traveling into shadowed areas, especially while facing a lit terrain beyond the shadowed area, drivers enter at their own risk due to pupil contraction making artificial lights useless. Additionally, the constant transitioning between dark and light areas are mentally taxing to the crew and natural navigational references such as the stars are invisible and therefore unusable. Slope and depth magnitude are very difficult hazards to judge when approaching a shadowed crater. This naturally leads to slower driving speeds than originally anticipated. As for observing scientifically interesting features, assessing them accurately varies greatly by the lighting condition. In most cases, the task can be completed, but the strategy is to use the sun to one’s advantage. Crew workload distribution in the cockpit for driving operations was split amongst crew. The driver primarily focused attention on visual terrain (80%) but referenced displays approximately 20% of the time, while the co-pilot/navigator generally provided directional cues to the driver. These natural lighting conditions present significant challenges for safe rover operations; however, early studies having given investigators a “first-look” into understanding of how operating on the South Lunar Pole can be accomplished.

LTv↗
NASA NTRS2014

Lunar Cube Transfer Trajectory Options

Contingent upon the modification of an initial condition of the injected or deployed orbit. Additionally, these designs can be restricted by the selection of the Cubesat subsystem design such as propulsion or communication. Nonetheless, many trajectory options can be designed with have a wide range of transfer durations, fuel requirements, and final destinations. Our investigation of potential trajectories highlights several design options including deployment into low Earth orbit (LEO), geostationary transfer orbits (GTO), and higher energy direct lunar transfer orbits. In addition to direct transfer options from these initial orbits, we also investigate the use of longer duration Earth-Moon dynamical systems. For missions with an intended lunar orbit, much of the design process is spent optimizing a ballistic capture while other science locations such as Sun-Earth libration or heliocentric orbits may simply require a reduced Delta-V imparted at a convenient location along the trajectory. In this article we examine several design options that meet the above limited deployment and subsystem drivers. We study ways that both impulsive and low-thrust Solar Electric Propulsion (SEP) engines can be used to place the Cubesat first into a highly eccentric Earth orbit, enter the Moon's Sphere of Influence, and finally achieve a highly eccentric lunar orbit. We show that such low-thrust transfers are feasible with a realistic micro-thruster model, assuming that the Cubesat can generate sufficient power for the SEP. Two examples are shown here: (1) A Cubestat injected by Exploration Mission 1 (EM-1) then employing low thrust; and (2) a CubSat deployed in a GTO, then employing impulsive maneuvers. For the EM-1 injected initial design, we increase the EM-1 targeted lunar flyby distance to reduce the energy of the lunar flyby to match that of a typical lMoon system heteroclinic manifold. Figure 1 presents an option that encompasses the similar dynamics as that of the ARTEMIS mission design. Low-thrust maneuvers are used along the manifold trajectory to raise perigee to that of a lunar orbit, adjust the timing with respect to the Moon, rotate the line of apsides, and target a ballistic lunar encounter. In this design a second flyby decreases the orbital energy with respect to the Moon, so that C3 -0.1 km2s2. Another design, shown in Figure 2 emanates from a GTO then uses impulsive maneuvers to phase onto a local Earth-Moon manifold, which then transfers the CubeSat to a lunar encounter.

Dynamical System↗
NASA NTRSDate not supplied

Autonomous Navigation of a Lunar Relay Using GNSS and Other Measurements

Many of the highest priority destinations at the Moon lack a continuous view of Earth, such as the lunar poles or lunar far side. Exploration of these sites will require spacecraft in cislunar space to relay communications and provide position, navigation, and timing (PNT) services. Accurate knowledge of relay position, velocity, and time is essential to these services. This paper describes a concept for a PNT Instrument being developed for the Lunar Communications Relay and Navigation Systems (LCRNS) Project. The instrument is intended as a payload that would enable autonomous, on-board, real-time navigation and timing using Global Navigation Satellite System (GNSS), optical navigation, and one-way measurements from Earth-based ground stations. Hardware-in-the-loop simulations using flight software are used to realistically characterize performance on hardware platforms with a path to flight. These results provide preliminary validation of the proposed PNT Instrument, demonstrate the benefits of augmenting GNSS with other measurements, and serve as an insightful reference for the design of future lunar missions, including those that will operate within the LunaNet framework of standards. This instrument concept relies on several technologies developed at NASA Goddard Space Flight Center (GSFC). For GNSS observables, the instrument relies on the high-altitude NavCube 3 mini (NC3m) GNSS receiver specifically designed for cislunar applications. The autoNGC system, which consists of flight software and a hardware platform, is responsible for fusing the observables using its extended Kalman filter, the Goddard Enhanced Onboard Navigation System (GEONS). Optical navigation observables are processed within autoNGC (“autonomous Navigation, Guidance, and Control”) using the Goddard Image Analysis & Navigation Tool (GIANT) which is also responsible for simulating high-fidelity images for test and analysis. In addition to describing the PNT Instrument and its components, the paper will present predicted performance based on simulation results. As a baseline, it will present GNSS-only hardware-in-the loop results using a NC3m test unit to process Spirent-simulated GPS signals in a potential lunar relay trajectory: a 12-hour elliptical frozen lunar orbit (ELFO). GEONS then processes the GPS pseudorange and time differenced carrier phase measurements to estimate and propagate the relay state (position, velocity, and time). These results extend previously published work that showed preliminary ELFO performance. Previous work has shown the importance of other measurement types, so additional simulations are performed which augment GNSS with ground station observables and several methods of optical navigation, including celestial navigation, limb-finding (e.g., observations of the lunar horizon), and terrain relative navigation (TRN). TRN involves correlating simulated predicted images of the lunar surface with actual imagery; misalignments of landmarks identified in each image are translated into relay state updates. TRN is valuable as a measurement of the relay’s state relative to the Moon, especially during GNSS outages or after maneuvers. One-way Pseudorange and Doppler measurements from Earth-based ground stations are also simulated. The full set of observables is processed using autoNGC. These simulations make use of autoNGC and NC3m test units, a lab atomic clock, and a pulse-per-second (PPS) generation and distribution system. This combination of subsystems, and the hardware platforms used in this analysis, represents a PNT Instrument that could be flown on a lunar relay. Results from the hardware-in-the-loop simulations presented in this paper provide a preliminary assessment of the achievable navigation performance of this instrument concept. PNT Instrument performance is compared to the GPS-only performance, and a discussion is provided on the apparent merits and challenges of each measurement type.

Ben Ashman↗
NASA NTRS1985

Launching rockets and small satellites from the lunar surface

Scientific payloads and their propulsion systems optimized for launch from the lunar surface differ considerably from their counterparts for use on earth. For spin-stabilized payloads, the preferred shape is a large diameter-to-length ratio to provide stability during the thrust phase. The rocket motor required for a 50-kg payload to reach an altitude of one lunar radius would have a mass of about 41 kg. To place spin-stabilized vehicles into low altitude circular orbits, they are first launched into an elliptical orbit with altitude about 840 km at aposelene. When the spacecraft crosses the desired circular orbit, small retro-rockets are fired to attain the appropriate direction and speed. Values of the launch angle, velocity increments, and other parameters for circular orbits of several altitudes are tabulated. To boost a 50-kg payload into a 100-km altitude circular orbit requires a total rocket motor mass of about 90 kg.

Anderson, K. A.↗
NASA NTRS2008

Solar Concentrator Demonstrator for Lunar Regolith Processing

NASA at the Marshall Space Flight Center (MSFC) is building a portable inflatable solar concentrator ground demonstrator for use in testing in-situ resource utilization (ISRU) lunar regolith processing methods. Of primary interest is the production of oxygen as a propellant oxidizer and for life support. There are various processes being proposed for the in-situ reduction of the lunar regolith, the leading processes are hydrogen reduction, carbothermal reduction and vapor phase pyrolysis. The concentrator system being built at MSFC could support demonstrations of all of these processes. The system consists of a light inflatable concentrator that will capture sunlight and focus it onto a receiver inside a vacuum chamber. Inflatable concentrators are good for space based applications due to their low weight and dense packaging at launch. The hexapod design allows the spot size to be increased to reduce the power density if needed for the process being demonstrated. In addition to the hardware development, a comprehensive simulation model is being developed and will be verified and validated using the system hardware. The model will allow for the evaluation of different lunar locations and operational scenarios for the lunar regolith processing with a high confidence in the predicted results.

Fikes, John C.↗
NASA NTRS2012

Global Maps of Lunar Neutron Fluxes from the LEND Instrument

The latest neutron spectrometer measurements with the Lunar Exploration Neutron Detector (LEND) onboard the Lunar Reconnaissance Orbiter (LRO) are presented. It covers more than 1 year of mapping phase starting on 15 September 2009. In our analyses we have created global maps showing regional variations in the flux of thermal (energy range < 0.015 eV) and fast neutrons (>0.5 MeV), and compared these fluxes to variances in soil elemental composition, and with previous results obtained by the Lunar Prospector Neutron Spectrometer (LPNS). We also processed data from LEND collimated detectors and derived a value for the collimated signal of epithermal neutrons based on the comparative analysis with the LEND omnidirectional detectors. Finally, we have compared our final (after the data reduction) global epithermal neutron map with LPNS data.

Epithermal neutrons↗
NASA NTRS1979

Lunar physical properties from analysis of magnetometer data

The electromagnetic properties of the lunar interior are discussed with emphasis on (1) bulk, crustal, and local anomalous conductivity; (2) bulk magnetic permeability measurements, iron abundance estimates, and core size limits; (3) lunar ionosphere and atmosphere; and (4) crustal magnetic remanence: scale size measurements and constraints on remanence origin. Appendices treat the phase relationship between the energetic particle flux modulation and current disc penetrations in the Jovian magnetosphere (Pioneer 10 inbound) theories for the origin of lunar magnetism; electrical conductivity anomalies associated with circular lunar maria; electromagnetic properties of the Moon; Mare Serenitatis conductivity anomaly detected by Apollo 16 and Lunokhod 2 magnetometers; and lunar properties from magnetometer data: effects of data errors.

Daily, W. D.↗
NASA NTRS1979

New determinations of far-side lunar radii from Apollo photography

Determinations of far side lunar radii based on the reduction of photogrammetric measurements derived from Apollo 16 photographs obtained during the trans-earth phase of the spacecraft trajectory are presented. Reduction is accomplished by determining the Apollo 16 metric camera orientation relative to control points on the lunar surface whose coordinates are known exactly, than charting the surface on the basis of data coverage. The positions of 66 surface features in the area between longitudes 90 and 130 deg E and 10 and 60 deg N relative to the center of mass of the moon, with a relative accuracy of 500 km, are presented. A topographical map which can resolve basin-sized features has been derived from radius determinations. It is found that the craters Fabry, Riemann and Szilard comprise a topographically depressed region about 500 km in diameter with a floor 2.4 to 3.4 km below the 1730.0 km reference sphere and 4.8 to 5.8 km below the northern rim of the unfilled basin.

Mohan, S. N.↗
NASA NTRSDate not supplied

3D Stereo Immersive Environment Utilizing Mobile Devices / Peripherals

This project expands upon some of the FY23 IRAD funded "Full Lunar Surface Simulation Platform" (FLS-SP) derived technologies. The results have demonstrated a low-cost, compact, immersive technology that supports training, interaction, mission planning, and other man-in-loop activities in simulated environments called “3D Stereo Immersion Environment Utilizing Mobile Devices” (3D SIEUMD). These activities will become crucial for safety and mission, planning, and assurance in the Artemis program, as well as future commercial crew programs. The FLS-SP enabled the visualization of simulated lunar surface operations for rovers, landers, habitation, and EVA. The 3D SIEUMD technology allows for NASA S&MA experts, tasked with defining safety protocols and procedures, to interact in an immersive environment without expensive immersion gear. Reduction of the overall safety risk by immersing safety concepts into the design and engineering phase for very small cost is accomplished even before hardware is deployed to the lunar surface.

Scott Johnson↗
NASA NTRS1967

Study of life support systems for space missions exceeding one year in duration, phase 1A. Volume 2 - Mission studies

The study covered by this report is a continuation of the work conducted under ' contract NAS 2-3012, Study of Life Support Systems For Space Missions Exceeding One Year In Duration Phase I. The previous work compared biological and chemical food processing systems with a stored food system for hypothetical missions of from 10 to 100 men for one to three years. The follow-on work Phase IA, the subject of this report, reviewed new concepts for waste processing and food manufacture in the closed system such as wastes for radiation shielding, wastes for propulsion, glycerol synthesis using chromatographic reactors, glycerol synthesis using acetylene as an intermediate, animal links, the use of radiophosphors for algae illumination, and new biological system approaches. The most promising of these new concepts were combined with the food producing systems evolved in Phase I and compared for three specific model missions; Earth orbit, lunar and mars. The effects of spacecraft power penalty, leakage rate, radiation shielding requirements and extravehicular expendable requirements were assessed and a cost effectiveness analysis was completed for each mission. The cost analysis compared the launch cost savings of the advanced systems with the R&D costs required to produce a flight ready system. The study results indicate that systems using wastes for radiation shielding, leakage makeup, or EVA expendables are very competitive with the biological and chemical food systems for all missions. The stored food baseline system and wastes for propulsion ranked poorly in the comparison studies.

R. B. Jagow↗
NASA NTRS1972

Progress in regenerative life support systems for a lunar laboratory.

Research and development work for application of Environmental and Thermal Control/Life Support System (ETC/LSS) on a lunar base mission is reviewed, covering lunar mission requirements and constraints, a Lunar Base ETC/LSS reliability assessment, food regeneration, the water and waste system, the atmosphere regeneration subsystem, and atmosphere contaminant control. The establishment of detailed system design criteria for the Lunar Surface Base LSS is considered to be premature at this phase of the project. Some recommendations are given instead for guidance in further R & D efforts.

Pecoraro, J. N.↗
NASA NTRS1973

Green spherules from Apollo 15 - Inferences about their origin from inert gas measurements.

Green spherules from the 'clod' 15426 and from fines 15421 contain about 100 times less trapped inert gases than normal bulk fines from Apollo 15. These spherules have apparently never been directly exposed to the solar wind. Spherules from other fines contain about 10 times more trapped gas than those from the 'clod.' The gas in the former is surface correlated. However, spherules from fines 15401 are exceptionally gas-poor. The trapped gases can be of solar-wind origin, but this origin requires a two-stage model for the spherules from the clods. Another possibility is that the gases were absorbed from an ambient gas phase. The trapped gases may also be assumed to represent primordial lunar gas. The composition of this gas is then similar to the 'solar' or 'unfractionated' component of gas-rich meteorites, but unlike that in most of the carbonaceous chondrites.

Lakatos, S.↗