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NASA NTRS2010

A General Closed-Form Solution for the Lunar Reconnaissance Orbiter (LRO) Antenna Pointing System

The National Aeronautics and Space Administration s (NASA) Lunar Reconnaissance Orbiter (LRO) launched on June 18, 2009 from the Cape Canaveral Air Force Station aboard an Atlas V launch vehicle into a direct insertion trajectory to the Moon LRO, designed, built, and operated by the NASA Goddard Space Flight Center in Greenbelt, MD, is gathering crucial data on the lunar environment that will help astronauts prepare for long-duration lunar expeditions. During the mission s nominal life of one year its six instruments and one technology demonstrator will find safe landing site, locate potential resources, characterize the radiation environment and test new technology. To date, LRO has been operating well within the bounds of its requirements and has been collecting excellent science data images taken from the LRO Camera Narrow Angle Camera (LROC NAC) of the Apollo landing sites have appeared on cable news networks. A significant amount of information on LRO s science instruments is provided at the LRO mission webpage. LRO s Attitude Control System (ACS), in addition to controlling the orientation of the spacecraft is also responsible for pointing the High Gain Antenna (HGA). A dual-axis (or double-gimbaled) antenna, deployed on a meter-long boom, is required to point at a selected Earth ground station. Due to signal loss over the distance from the Moon to Earth, pointing precision for the antenna system is very tight. Since the HGA has to be deployed in spaceflight, its exact geometry relative to the spacecraft body is uncertain. In addition, thermal distortions and mechanical errors/tolerances must be characterized and removed to realize the greatest gain from the antenna system. These reasons necessitate the need for an in-flight calibration. Once in orbit around the moon, a series of attitude maneuvers was conducted to provide data needed to determine optimal parameters to load onboard, which would account for the environmental and mechanical errors at any antenna orientation. The nominal geometry for the HGA involves an outer gimbal axis that is exactly perpendicular to the inner gimbal axis, and a target direction that is exactly perpendicular to the outer gimbal axis. For this nominal geometry, closed-form solutions of the desired gimbal angles are simple to get for a desired target direction specified in the spacecraft body fame. If the gimbal axes and the antenna boresight are slightly misaligned, the nominal closed-form solution is not sufficiently accurate for computing the gimbal angles needed to point at a target. In this situation, either a general closed-form solution has to be developed for a mechanism with general geometries, or a correction scheme has to be applied to the nominal closed-form solutions. The latter has been adopted for Solar Dynamics Observatory (SDO) as can be seen in Reference 1, and the former has been used for LRO. The advantage of the general closed-form solution is the use of a small number of parameters for the correction of nominal solutions, especially in the regions near singularities. Singularities here refer to cases when the nominal closed-form solutions have two or more solutions. Algorithm complexity, however, is the disadvantage of the general closed-form solution.

Shah, Neerav↗

NASA NTRS1991

Radiation protection for human missions to the Moon and Mars

Radiation protection assessments are performed for advanced Lunar and Mars manned missions. The Langley cosmic ray transport code and the nucleon transport code are used to quantify the transport and attenuation of galactic cosmic rays and solar proton flares through various shielding media. Galactic cosmic radiation at solar maximum and minimum, as well as various flare scenarios are considered. Propagation data for water, aluminum, liquid hydrogen, lithium hydride, lead, and lunar and Martian regolith (soil) are included. Shield thickness and shield mass estimates required to maintain incurred doses below 30 day and annual limits (as set for Space Station Freedom and used as a guide for space exploration) are determined for simple geometry transfer vehicles. On the surface of Mars, dose estimates are presented for crews with their only protection being the carbon dioxide atmosphere and for crews protected by shielding provided by Martian regolith for a candidate habitat.

Simonsen, Lisa C.↗

NASA NTRS1990

Equilibrium radiative heating tables for Earth entry

The recent resurgence of interest in blunt-body atmospheric entry for applications such as aeroassisted orbital transfer and planetary return has engendered a corresponding revival of interest in radiative heating. Radiative heating may be of importance in these blunt-body flows because of the highly energetic shock layer around the blunt nose. Sutton developed an inviscid, stagnation point, radiation coupled flow field code for investigating blunt-body atmospheric entry. The method has been compared with ground-based and flight data, and reasonable agreement has been found. To provide information for entry body studies in support of lunar and Mars return scenarios of interest in the 1970's, the code was exercised over a matrix of Earth entry conditions. Recently, this matrix was extended slightly to reflect entry vehicle designs of current interest. Complete results are presented.

Sutton, Kenneth↗

NASA NTRS2010

Lunar Reconnaissance Orbiter (LRO): Observations for Lunar Exploration and Science

The Lunar Reconnaissance Orbiter (LRO) was implemented to facilitate scientific and engineering-driven mapping of the lunar surface at new spatial scales and with new remote sensing methods, identify safe landing sites, search for in situ resources, and measure the space radiation environment. After its successful launch on June 18,2009, the LRO spacecraft and instruments were activated and calibrated in an eccentric polar lunar orbit until September 15, when LRO was moved to a circular polar orbit with a mean altitude of 50 km. LRO will operate for at least one year to support the goals of NASA's Exploration Systems Mission Directorate (ESMD), and for at least two years of extended operations for additional lunar science measurements supported by NASA's Science Mission Directorate (SMD). LRO carries six instruments with associated science and exploration investigations, and a telecommunications/radar technology demonstration. The LRO instruments are: Cosmic Ray Telescope for the Effects of Radiation (CRaTER), Diviner Lunar Radiometer Experiment (DLRE), Lyman-Alpha Mapping Project (LAMP), Lunar Exploration Neutron Detector (LEND), Lunar Orbiter Laser Altimeter (LOLA), and Lunar Reconnaissance Orbiter Camera (LROC). The technology demonstration is a compact, dual-frequency, hybrid polarity synthetic aperture radar instrument (Mini-RF). LRO observations also support the Lunar Crater Observation and Sensing Satellite (LCROSS), the lunar impact mission that was co-manifested with LRO on the Atlas V (401) launch vehicle. This paper describes the LRO objectives and measurements that support exploration of the Moon and that address the science objectives outlined by the National Academy of Science's report on the Scientific Context for Exploration of the Moon (SCEM). We also describe data accessibility by the science and exploration community.

Vondrak, Richard↗

NASA NTRS2007

Solid Lubricants and Coatings for Extreme Environments: State-of-the-Art Survey

An investigation was conducted to survey anticipated requirements for solid lubricants in lunar and Martian environments, as well as the effects of these environments on lubricants and their performance and durability. The success of habitats and vehicles on the Moon and Mars, and ultimately, of the human exploration of and permanent human presence on the Moon and Mars, are critically dependent on the correct and reliable operation of many moving mechanical assemblies and tribological components. The coefficient of friction and lifetime of any lubricant generally vary with the environment, and lubricants have very different characteristics under different conditions. It is essential, therefore, to select the right lubrication technique and lubricant for each mechanical and tribological application. Several environmental factors are hazardous to performance integrity on the Moon and Mars. Potential threats common to both the Moon and Mars are low ambient temperatures, wide daily temperature swings (thermal cycling), solar flux, cosmic radiation, and large quantities of dust. The surface of Mars has the additional challenges of dust storms, winds, and a carbon dioxide atmosphere. Solid lubricants and coatings are needed for lunar and Martian applications, where liquid lubricants are ineffective and undesirable, and these lubricants must perform well in the extreme environments of the Moon, Mars, and space, as well as on Earth, where they will be assembled and tested. No solid lubricants and coatings and their systems currently exist or have been validated that meet these requirements, so new solid lubricants must be designed and validated for these applications.

Miyoshi, Kazuhisa↗

NASA NTRS2005

Exploration Rover Concepts and Development Challenges

This paper presents an overview of exploration rover concepts and the various development challenges associated with each as they are applied to exploration objectives and requirements for missions on the Moon and Mars. A variety of concepts for surface exploration vehicles have been proposed since the initial development of the Apollo-era lunar rover. This paper provides a brief description of the rover concepts, along with a comparison of their relative benefits and limitations. In addition, this paper outlines, and investigates a number of critical development challenges that surface exploration vehicles must address in order to successfully meet the exploration mission vision. These include: mission and environmental challenges, design challenges, and production and delivery challenges. Mission and environmental challenges include effects of terrain, extreme temperature differentials, dust issues, and radiation protection. Design methods are discussed that focus on optimum methods for developing highly reliable, long-life and efficient systems. In addition, challenges associated with delivering a surface exploration system is explored and discussed. Based on all the information presented, modularity will be the single most important factor in the development of a truly viable surface mobility vehicle. To meet mission, reliability, and affordability requirements, surface exploration vehicles, especially pressurized rovers, will need to be modularly designed and deployed across all projected Moon and Mars exploration missions.

Zakrajsek, James J.↗

NASA NTRS1992

Space crew radiation exposure analysis system based on a commercial stand-alone CAD system

Major improvements have recently been completed in the approach to spacecraft shielding analysis. A Computer-Aided Design (CAD)-based system has been developed for determining the shielding provided to any point within or external to the spacecraft. Shielding analysis is performed using a commercially available stand-alone CAD system and a customized ray-tracing subroutine contained within a standard engineering modeling software package. This improved shielding analysis technique has been used in several vehicle design projects such as a Mars transfer habitat, pressurized lunar rover, and the redesigned Space Station. Results of these analyses are provided to demonstrate the applicability and versatility of the system.

Appleby, Matthew H.↗

NASA NTRS2017

A First Look at the Navigation Design and Analysis for the Orion Exploration Mission 2

This paper will detail the navigation and dispersion design and analysis of the first Orion crewed mission. The optical navigation measurement model will be described. The vehicle noise includes the residual acceleration from attitude deadbanding, attitude maneuvers, CO2 venting, wastewater venting, ammonia sublimator venting and solar radiation pressure. The maneuver execution errors account for the contribution of accelerometer scale-factor on the accuracy of the maneuver execution. Linear covariance techniques are used to obtain the navigation errors and the trajectory dispersions as well as the DV performance. Particular attention will be paid to the accuracy of the delivery at Earth Entry Interface and at the Lunar Flyby.

D'Souza, Chris D.↗

NASA NTRS2007

Biomedical Support of U.S. Extravehicular Activity

The world's first extravehicular activity (EVA) was performed by A. A. Leonov on March 18, 1965 during the Russian Voskhod-2 mission. The first US EVA was executed by Gemini IV astronaut Ed White on June 3, 1965, with an umbilical tether that included communications and an oxygen supply. A hand-held maneuvering unit (HHMU) also was used to test maneuverability during the brief EVA; however the somewhat stiff umbilical limited controlled movement. That constraint, plus difficulty returning through the vehicle hatch, highlighted the need for increased thermal control and improved EVA ergonomics. Clearly, requirements for a useful EVA were interrelated with the vehicle design. The early Gemini EVAs generated requirements for suits providing micro-meteor protection, adequate visual field and eye protection from solar visual and infrared radiation, gloves optimized for dexterity while pressurized, and thermal systems capable of protecting the astronaut while rejecting metabolic heat during high workloads. Subsequent Gemini EVAs built upon this early experience and included development of a portable environmental control and life support systems (ECLSS) and an astronaut maneuvering unit. The ECLSS provided a pressure vessel and controller with functional control over suit pressure, oxygen flow, carbon dioxide removal, humidity, and temperature control. Gemini EVA experience also identified the usefulness of underwater neutral buoyancy and altitude chamber task training, and the importance of developing reliable task timelines. Improved thermal management and carbon dioxide control also were required for high workload tasks. With the Apollo project, EVA activity was primarily on the lunar surface; and suit durability, integrated liquid cooling garments, and low suit operating pressures (3.75 pounds per square inch absolute [psia] or 25.8 kilopascal [kPa],) were required to facilitate longer EVAs with ambulation and significant physical workloads with average metabolic rates of 1000 BTU/hr and peaks of up to 2200 BTU/hr. Mobility was further augmented with the Lunar Roving Vehicle. The Apollo extravehicular mobility unit (EMU) was made up of over 15 components, ranging from a biomedical belt for capturing and transmitting biomedical data, urine and fecal containment systems, a liquid cooling garment, communications cap, a modular portable life support system (PLSS), a boot system, thermal overgloves, and a bubble helmet with eye protection. Apollo lunar astronauts performed successful EVAs on the lunar surface from a 5 psia (34.4 kPa) 100% oxygen environment in the Lunar Lander. A maximum of three EVAs were performed on any mission. For Skylab a modified A7LB suit, used for Apollo 15, was selected. The Skylab astronaut life support assembly (ALSA) provided umbilical support through the life support umbilical (LSU) and used open loop oxygen flow, rather than closed-loop as in Apollo missions. Thermal control was provided by liquid water circulated by spacecraft pumps and electrical power also was provided from the spacecraft via the umbilical. The cabin atmosphere of 5 psia (34.4 kPa), 70% oxygen, provided a normoxic atmosphere and because of the very low nitrogen partial pressures, no special protocols were required to protect against decompression sickness (DCS) as was the case with the Apollo spacecraft with a 5 psi, 100% oxygen environment.

Gernhardt, Michael L.↗

NASA NTRS2008

Ground Plane and Near-Surface Thermal Analysis for NASA's Constellation Program

Most spacecraft thermal analysis tools assume that the spacecraft is in orbit around a planet and are designed to calculate solar and planetary fluxes, as well as radiation to space. On NASA Constellation projects, thermal analysts are also building models of vehicles in their pre-launch condition on the surface of a planet. This process entails making some modifications in the building and execution of a thermal model such that the radiation from the planet, both reflected albedo and infrared, is calculated correctly. Also important in the calculation of pre-launch vehicle temperatures are the natural environments at the vehicle site, including air and ground temperatures, sky radiative background temperature, solar flux, and optical properties of the ground around the vehicle. A group of Constellation projects have collaborated on developing a cohesive, integrated set of natural environments that accurately capture worst-case thermal scenarios for the pre-launch and launch phases of these vehicles. The paper will discuss the standardization of methods for local planet modeling across Constellation projects, as well as the collection and consolidation of natural environments for launch sites. Methods for Earth as well as lunar sites will be discussed.

Gasbarre, Joseph F.↗

NASA NTRS2012

Evaluating Three Active Thermal Architectures for Exploration Missions

Mass and cost are typically the two biggest challenges facing space craft designers. Active thermal control systems for crewed space-craft are typically among the more massive and costly systems on the vehicle. A study was completed evaluating three different thermal control system architectures to evaluate overall performance, mass and cost for a typical exploration mission profile. The architectures that were evaluated were 1 - a two-loop system using an internal liquid loop interfacing with an external liquid loop and flow loop with flow through radiators; 2 - a-single loop architecture with flow through radiators utilizing a regenerative heat exchanger and heater; and 3 - a single-loop architecture with heat pipe radiators. Environmental conditions, calculated for a given lunar exploration mission, and mission heat load profiles, generated based on previous Orion time lines, were evalauated through the phases of the on orbit mission. Performance for each of the architectures was evaluated along with the resultant mass of each system. Recommendations include adding a thermal topping system to lunar missions due to the extreme hot environments encountered in near-lunar approaches.

Cross, Cynthia D.↗

NASA NTRSDate not supplied

Space Radiation Technologies for Human Missions beyond Low-Earth-Orbit

Ionizing radiation exposure is a pervasive occupational risk to astronauts and remains a major health risk associated with human exploration beyond low-Earth-orbit. Sustained lunar and Mars exploration crew will be at increased risk of radiogenic cancers and cardiovascular disease, as well as, the potential for in-mission performance decrements, immune system decrements, and late neurodegenerative conditions impacting crew health. Optimizing astronaut space radiation protection will incorporate both physical and biological radiological aspects and will implement protection strategies that span pre-mission, in-flight, and post mission timelines. Major facets of physical risk mitigation include accuracy in predicting the space radiation solar environment, monitoring environmental and human exposures, and minimizing exposure through vehicle shielding and design. Biological mitigation facets include understanding and predicting biological response to the space environment, developing advanced medical countermeasures, and identifying appropriate surveillance strategies to preserve long-term health. Coordinated investments in these major capability areas will support development of an integrated radiation protection solution to minimize astronaut risk for moon to Mars exploration missions.

Lisa C Simonsen↗

NASA NTRS2001

Application of Monte-Carlo Analyses for the Microwave Anisotropy Probe (MAP) Mission

The Microwave Anisotropy Probe (MAP) is the third launch in the National Aeronautics and Space Administration's (NASA's) a Medium Class Explorers (MIDEX) program. MAP will measure, in greater detail, the cosmic microwave background radiation from an orbit about the Sun-Earth-Moon L2 Lagrangian point. Maneuvers will be required to transition MAP from it's initial highly elliptical orbit to a lunar encounter which will provide the remaining energy to send MAP out to a lissajous orbit about L2. Monte-Carlo analysis methods were used to evaluate the potential maneuver error sources and determine their effect of the fixed MAP propellant budget. This paper will discuss the results of the analyses on three separate phases of the MAP mission - recovering from launch vehicle errors, responding to phasing loop maneuver errors, and evaluating the effect of maneuver execution errors and orbit determination errors on stationkeeping maneuvers at L2.

Mesarch, Michael A.↗

NASA NTRS2003

Orbit Determination Support for the Microwave Anisotropy Probe (MAP)

THe microwave Anisotropy Probe (MAP) ia the second Medium Class Explorer (MIDEX) mission of the National Aeronautics and Space Administration (NASA). The main goal of the MAP observatory is to measure the temperature fluctuations, known as anisotropy, of the cosmic microwave background (CBG) radiation over the entire sky and to produce a map of the CMB anisotropies with an angular resolution of approximately 3 degrees. MAP was launched from the Cape Canaveral Air Force Station Complex 17 aboard a Delta II 7425-10 expendable launch vehicle at exactly 19:46:46.183 UTC on June 30, 20001. The spacecraft receiver a nominal direct insertion by the Delta expendable launch vehicle into a 185-km circular orbit with a 28.7 deg. inclination. MAP was than maneuvered into a sequence of phasing loops designed to set up a lunar gravity-assisted acceleration of the spacecraft onto a transfer trajectory to a lissajous orbit about the Earth-Sun L2 Lagrange point, about 1.5 million km from Earth. The science mission minimum lifetime is two years of observations at L2 with a desired lifetime of 4 years. The MAP transfer orbit consisted of 3.5 phasing loops.The MAP trajectory schematic all the way through L2 is shown. The first loop had a period of 7 days, the second and third loops were 10 days long, and the last half loop was 5 days. The periselene (i.e., lunar encounter or swingby) took place approximately 30 days after launch. After the periselene, the spacecraft cruised for approximately 60 days before it arrived in the vicinity of the L2 libration point. Two mid-course correction (MCC) maneuvers were performed to refine MAP's post-launch trajectory-one after periselene and one prior to arrival at vicinity of L2. Now that MAP is at its operational L2 lissajous orbit, the MAP satellite is commanded to perform occasional station-keeping (SK) maneuvers in order to maintain its orbit around L2. Because of its complex orbital characteristics, the mission provided a unique challenge to orbit determination (OD) support in many orbital regimes.

Truong, Son H.↗

NASA NTRS1989

Lunar construction utility vehicle

The lunar construction utility vehicle (LCUV) is an all-purpose construction vehicle which will aid in the robotic assembly of a lunar outpost. The LCUV will have the following capabilities: (1) must be self supporting including repairs; (2) must offload itself from a lunar lander; (3) must be telerobotic and semi-autonomous; (4) must be able to transport one space station common module; (5) must allow for man-rated operation; and (6) must be able to move lunar regolith for site preparation. This study recommends the use of an elastic tracked vehicle. Detailed material analyses of most of the LCUV components were accomplished. The body frame, made of pinned truss elements, was stress analyzed using NASTRAN. A track connection system was developed; however, kinematic and stress analyses are still required. This design recommends the use of hydrogen-oxygen fuel cells for power. Thermal control has proven to be a problem which may be the most challenging technically. A tentative solution has been proposed which utilizes an onboard and towable radiator. Detailed study of the heat dissipation requirements is needed to finalize radiator sizing. Preliminary work on a man-rated cabin has begun; however, this is not required during the first mission phase of the LCUV. Finally, still in the conceptual phases, are the communication, navigation and mechanical arm systems.

Source record↗

NASA NTRS2015

Mars Conjunction Crewed Missions With a Reusable Hybrid Architecture

A new crew Mars architecture has been developed that provides many potential benefits for NASA-led human Mars moons and surface missions beginning in the 2030s or 2040s. By using both chemical and electric propulsion systems where they are most beneficial and maintaining as much orbital energy as possible, the Hybrid spaceship that carries crew round trip to Mars is pre-integrated before launch and can be delivered to orbit by a single launch. After check-out on the way to cis-lunar space, it is refueled and can travel round trip to Mars in less than 1100 days, with a minimum of 300 days in Mars vicinity (opportunity dependent). The entire spaceship is recaptured into cis-lunar space and can be reused. The spaceship consists of a habitat for 4 crew attached to the Hybrid propulsion stage which uses long duration electric and chemical in-space propulsion technologies that are in use today. The hybrid architecture's con-ops has no in-space assembly of the crew transfer vehicle and requires only rendezvous of crew in a highly elliptical Earth orbit for arrival at and departure from the spaceship. The crew transfer vehicle does not travel to Mars so it only needs be able to last in space for weeks and re-enter at lunar velocities. The spaceship can be refueled and resupplied for multiple trips to Mars (every other opportunity). The hybrid propulsion stage for crewed transits can also be utilized for cargo delivery to Mars every other opportunity in a reusable manner to pre-deploy infrastructure required for Mars vicinity operations. Finally, the Hybrid architecture provides evolution options for mitigating key long-duration space exploration risks, including crew microgravity and radiation exposure.

Merrill, Raymond G.↗

NASA NTRS1991

Selenia: A habitability study for the development of a third generation lunar base

When Apollo astronauts landed on the Moon, the first generation of lunar bases was established. They consisted essentially of a lunar module and related hardware capable of housing two astronauts for not more than several days. Second generation lunar bases are being developed, and further infrastructure, such as space station, orbital transfer, and reusable lander vehicles will be necessary, as prolonged stay on the Moon is required for exploration, research, and construction for the establishment of a permanent human settlement there. Human life in these habitats could be sustained for months, dependent on a continual flow of life-support supplies from Earth. Third-generation lunar bases will come into being as self sufficiency of human settlements becomes feasible. Regeneration of water, oxygen production, and development of indigenous construction materials from lunar resources will be necessary. Greenhouses will grow food supplies in engineered biospheres. Assured protection from solar flares and cosmic radiation must be provided, as well as provision for survival under meteor showers, or the threat of meteorite impact. All these seem to be possible within the second decade of the next century. Thus, the builders of Selenia, the first of the third-generation lunar bases are born today. During the last two years students from the School of Architecture of the University of Puerto Rico have studied the problems that relate to habitability for prolonged stay in extraterrestrial space. An orbital personnel transport to Mars developed originally by the Aerospace Engineering Department of the University of Michigan was investigated and habitability criteria for evaluation of human space habitats were proposed. An important finding from that study was that the necessary rotational diameter of the vessel has to be on the order of two kilometers to ensure comfort for humans under the artificial gravity conditions necessary to maintain physiological well being of passengers, beyond the level of mere survival.

Source record↗

NASA NTRSDate not supplied

The Gateway Program as Part of NASA’s Plans for Human Exploration Beyond Low Earth Orbit

This paper provides an overview and status of Gateway, humanity’s first space station in lunar orbit as a vital component of the NASA-led Artemis missions to return humans to the Moon as preparation for the first human missions to Mars. Gateway is an aggregation point in deep space for a variety of spacecraft, including the crewed Orion vehicle, the Human Landing System that will ferry astronauts to and from the lunar surface, logistics supply craft, and vehicles transiting further into deep space beyond the Earth-Moon system, such as to Mars. NASA is building on decades of partnership with space agencies on three continents and multiple commercial partners to design, build, and launch Gateway’s core elements to near-rectilinear halo orbit (NRHO) around the Moon, where it will operate for a minimum of 15 years. Gateway is humanity’s next in-space science utilization platform, and its first in deep space, with three science payloads already selected to study solar and cosmic radiation. This paper will provide an overview of the Gateway space station’s major components in various stages of development, including the Power and Propulsion Element (PPE), Habitation and Logistics Outpost (HALO), the International Habitation (I-Hab) module, ESPRIT Refueling Module (ERM), the planned airlock, advanced external robotics systems, Deep Space Logistics supply craft, and next-generation autonomous Vehicle System Manager software. It will also provide an overview of how Gateway will be utilized for science, and highlight the space station’s multilateral governance structure and international agreements.

Emma Lehnhardt↗