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Genesis is the fifth mission selected as part of NASA's Discovery Program. The objective of Genesis is to collect solar wind samples for a period of approximately two years while in a halo orbit about the Sun-Earth colinear libration point, L1, located between the Sun and Earth.
Shortly after the NASA Genesis Mission sample return capsule returned to earth on September 8, 2004, the science team discovered that all nine ultra-pure semiconductor materials were contaminated with a thin molecular organic film approximately 0 to 100 angstroms thick. The organic contaminate layer, possibly a silicone, situated on the surface of the materials is speculated to have formed by condensation of organic matter from spacecraft off-gassing at the Lagrange 1 halo orbit during times of solar exposure. While the valuable solar wind atoms are safely secured directly below this organic contamination and/or native oxide layer in approximately the first 1000 angstroms of the ultra-pure material substrate, some analytical techniques that precisely measure solar wind elemental abundances require the removal of this organic contaminate. In 2005, Genesis science team laboratories began to develop various methods for removing the organic thin film without removing the precious material substrate that contained the solar wind atoms. Stephen Sestak and colleagues at Open University first experimented with ultraviolet radiation ozone (UV/O3) cleaning of several non-flight and flown Genesis silicon wafer fragments under a pure flowing oxygen environment. The UV/O3 technique was able to successfully remove organic contamination without etching into the bulk material substrate. At NASA Johnson Space Center Genesis Curation Laboratory, we have installed an UV/O3 cleaning devise in an ambient air environment to further experimentally test the removal of the organic contamination on Genesis wafer materials. Preliminary results from XPS analysis show that the UV/O3 cleaning instrument is a good non-destructive method for removing carbon contamination from flown Genesis array samples. However, spectroscopic ellipsometry results show little change in the thickness of the surface film. All experiments to date have shown UV/O3 cleaning method to be the best non-destructive method for removing organic contamination from the surface of the Genesis materials. The UV/O3 cleaning process can also clean carbon contamination to levels below non-flight standards. This can be seen by comparing sample 60260's carbon 10667 cps with non-flight Si carbon 21675 cps. Therefore, surface carbon contamination should not hinder the analysis of solar wind.
The Advanced Concepts Office at Marshall Space Flight Center completed a brief spacecraft design study for the 8-meter monolithic Advanced Technology Large Aperture Space Telescope (ATLAST-8m). This spacecraft concept provides all power, communication, telemetry, avionics, guidance and control, and thermal control for the observatory, and inserts the observatory into a halo orbit about the second Sun-Earth Lagrange point. The multidisciplinary design team created a simple spacecraft design that enables component and science instrument servicing, employs articulating solar panels for help with momentum management, and provides precise pointing control while at the same time fast slewing for the observatory.
Spectrum Roentgen Gamma (SRG) is an X-ray astrophysical observatory, developed by Russia in collaboration with Germany. The mission will be launched in 2014 from Baikonur, by a Zenit rocket with a Fregat booster and placed in a 6-month-period halo orbit around L2. The scientific payload consists of two independent telescopes . a soft-x-ray survey instrument, eROSITA, being provided by Germany and a medium-x-ray-energy survey instrument ART-XC being developed by Russia. ART-XC will consist of seven independent, but co-aligned, telescope modules with seven corresponding cadmium-telluride focal plane detectors. Each will operate over the approximate energy range of 6- 30 keV, with an angular resolution of <1 ', a field of view of ~30 ' and an energy resolution about 10% at 14 keV. The NASA Marshall Space Flight Center (MSFC) will fabricate some of the mirror modules, to complement others fabricated by VNIIEF in Russia.
A novel concept is presented in this paper for a human mission to the lunar L2 (Lagrange) point that would be a proving ground for future exploration missions to deep space while also overseeing scientifically important investigations. In an L2 halo orbit above the lunar farside, the astronauts would travel 15% farther from Earth than did the Apollo astronauts and spend almost three times longer in deep space. Such missions would validate the Orion MPCV's life support systems, would demonstrate the high-speed re-entry capability needed for return from deep space, and would measure astronauts' radiation dose from cosmic rays and solar flares to verify that Orion would provide sufficient protection, as it is designed to do. On this proposed mission, the astronauts would teleoperate landers and rovers on the unexplored lunar farside, which would obtain samples from the geologically interesting farside and deploy a low radio frequency telescope. Sampling the South Pole-Aitkin basin (one of the oldest impact basins in the solar system) is a key science objective of the 2011 Planetary Science Decadal Survey. Observations of the Universe's first stars/galaxies at low radio frequencies are a priority of the 2010 Astronomy & Astrophysics Decadal Survey. Such telerobotic oversight would also demonstrate capability for human and robotic cooperation on future, more complex deep space missions.
Spectrum Roentgen Gamma (SRG) is an X-ray astrophysical observatory, developed by Russia in collaboration with Germany. The mission will be launched in March 2016 from Baikonur, by a Zenit rocket with a Fregat booster and placed in a 6-month-period halo orbit around L2. The scientific payload consists of two independent telescopes - a soft-x-ray survey instrument, eROSITA, being provided by Germany and a medium-x-ray-energy survey instrument ART-XC being developed by Russia. ART-XC will consist of seven independent, but co-aligned, telescope modules. The NASA Marshall Space Flight Center (MSFC) is fabricating the flight mirror modules for the ART-XC/SRG. Each mirror module will be aligned with a focal plane CdTe double-sided strip detectors which will operate over the energy range of 6-30 keV, with an angular resolution of less than 1′, a field of view of approximately 34′ and an expected energy resolution of about 10 percent at 14 keV.
This paper examines the benefits of navigating a crewed vehicle at the Moon using both ground tracking and satellite-to-satellite tracking, where the tracking satellite is stationed in a lunar halo orbit. Linked Autonomous Interplanetary Satellite Orbit Navigation (LiAISON) is a new technique that has been shown to dramatically improve the navigation of lunar satellites, libration orbiters, and Earth orbiting satellites using simple scalar satellite-to-satellite observations, such as range or Doppler. In this paper, LiAISON is applied to the problem of navigations a crewed vehicle in low lunar orbit. It has been found that adding LiAISON observations to a ground navigation solution improves the navigation enough to reduce the number of active ground tracking stations from six to three.
An increasing number of spacecraft are relying on the Global Positioning System (GPS) for navigation at altitudes near or above the GPS constellation itself - the region known as the Space Service Volume (SSV). While the formal definition of the SSV ends at geostationary altitude, the practical limit of high-altitude space usage is not known, and recent missions have demonstrated that signal availability is sufficient for operational navigation at altitudes halfway to the moon. This paper presents simulation results based on a high-fidelity model of the GPS constellation, calibrated and validated through comparisons of simulated GPS signal availability and strength with flight data from recent high-altitude missions including the Geostationary Operational Environmental Satellite 16 (GOES-16) and the Magnetospheric Multiscale (MMS) mission. This improved model is applied to the transfer to a lunar near-rectilinear halo orbit (NRHO) of the class being considered for the international Deep Space Gateway concept. The number of GPS signals visible and their received signal strengths are presented as a function of receiver altitude in order to explore the practical upper limit of high-altitude space usage of GPS.
An increasing number of spacecraft are relying on the Global Positioning System (GPS) for navigation at altitudes near or above the GPS constellation itself - the region known as the Space Service Volume (SSV). While the formal definition of the SSV ends at geostationary altitude, the practical limit of high-altitude space usage is not known, and recent missions have demonstrated that signal availability is sufficient for operational navigation at altitudes halfway to the moon. This paper presents simulation results based on a high-fidelity model of the GPS constellation, calibrated and validated through comparisons of simulated GPS signal availability and strength with flight data from recent high-altitude missions including the Geostationary Operational Environmental Satellite 16 (GOES-16) and the Magnetospheric Multiscale (MMS) mission. This improved model is applied to the transfer to a lunar near-rectilinear halo orbit (NRHO) of the class being con- sidered for the international Deep Space Gateway concept. The number of GPS signals visible and their received signal strengths are presented as a function of receiver altitude in order to explore the practical upper limit of high-altitude space usage of GPS.
The purpose of this thesis is to design a machine learning algorithm platform that provides expanded knowledge of mission availability through a launch season by improving trajectory resolution and introducing launch mission forecasting. The specific scenario addressed in this paper is one in which data is provided for four deterministic translational maneuvers through a mission to a Near Rectilinear Halo Orbit (NRHO) with a 9:2 synodic frequency. Current launch availability knowledge under NASA's Orion Orbit Performance Team is established by altering optimization variables associated to given reference launch epochs. This current method can bean abstract task and relies on an orbit analyst to structure a mission based on an established mission design methodology associated to the performance of Orion and NASA's Space Launch System. Introducing a machine learning algorithm trained to construct mission scenarios within the feasible range of known trajectories reduces the required interaction of the orbit analyst by removing the needed step of optimizing the orbit to fit an expected translational response required of the spacecraft. In this study, k-Nearest Neighbor and Bayesian Linear Regression successfully predicted classical orbital elements for the launch windows observed. However both algorithms had limitations due to their approaches to model fitting. Training machine learning algorithms off of classical orbital elements introduced a repetitive approach to reconstructing mission segments for different arrival opportunities through the launch window and can prove to be a viable method of launch window scan generation for future missions.
The purpose of this assessment was to develop updates and new features for the NASA Copernicus Spacecraft Trajectory Design and Optimization analysis tool (version 5.0) for application to NASA programs and projects. These updates will significantly improve the ability to design and optimize complex trajectories over multiple trajectory phases; will allow the use of unique vehicle-specific guidance, control, and trajectory strategies and constraints; and the creation of an almost unlimited number of unique user-defined capabilities. The primary stakeholders for this assessment are the trajectory design and optimization analysts and engineers, and the chief engineers and project managers for existing programs, projects, and/or tasks that involve impulsive, finite burn, and/or continuous thrust trajectories (e.g., Sun, planet, comet, asteroid, halo orbit, Lagrange point, and distant retrograde orbit). The breadth of application spans the preliminary engineering and mission design concepts and optimization, to the development of candidate reference missions and integrated mission design for vehicle system design and operation, to the design and development of flight trajectories and associated propulsive maneuvers for real-time operations.
This paper describes and predicts the performance of a conceptual autonomous GPS-based navigation system for NASA's planned lunar Gateway. This system is based on the flight-proven Magnetospheric Multiscale (MMS) GPS navigation system, augmented with an earth-pointed high-gain antenna and with an option for an atomic clock. High-fidelity simulations, calibrated against MMS flight data and making use of GPS transmitter patterns from the GPS Antenna Characterization Experiment (ACE) project are developed for operation of the system in the Gateway Near-Rectilinear Halo Orbit (NRHO). The results indicate that GPS can provide an autonomous, realtime navigation capability with comparable, or superior, performance to traditional Deep Space Network approach with eight hours of tracking per day.
Motivated by the aggressive timeline of NASA’s Artemis Program, the feasibility of evolving a mid-sized, reusable and refuelable cargo lunar lander technology demonstrator into the descent element of the three- stage human lander was assessed with the goal of forming synergies between both acquisition programs. Requirements for such a concept are that it must be deployed on commercial launch vehicles that are expected to enter into service consistent with Artemis’ timeline. In order to assess this concept’s feasibility, a physics-based analysis of alternatives was conducted where mission and vehicle architectures are traded side-by-side. Mission trades considered include the impacts of traveling to near-rectilinear halo orbit quickly versus slowly; vehicle trades include fuel and oxidizer tank configurations, number of engines, and propellant combinations, as well as several technology options, e.g. reduced and zero boil-off strategies. Results and discussions are presented to facilitate the consideration of this concept. lunar lander, systems analysis, multidisciplinary design and analysis, space systems, analysis of alternatives,
In preparation of the NextSTEP-2 Appendix H awards, NASA’s Human Landing System (HLS) Program (HLSP) went through an abbreviated Design Analysis Cycle (DAC), or miniDAC, to characterize and quantify how a storable propellant design would affect and improve overall system performance of the Government reference design. As part of that miniDAC, HLSP ran multiple trade studies and assessments, including aspects of Mission Design, Flight Mechanics (FM), and Guidance, Navigation, and Control (GN\&C). The storable propellant designs were assessed through the baseline mission targeting the Near Rectilinear Halo Orbit (NRHO) and several alternate staging orbits. This paper focuses on the FM and GN\&C trades and assessments completed as part of the miniDAC as well as the GN\&C elements of the miniDAC Master Equipment List (MEL) developed for the storable propellant lander configurations.
In preparation of the NextSTEP-2 Appendix H awards, NASA’s Human Landing System (HLS) Program (HLSP) went through an abbreviated Design Analysis Cycle (DAC), or miniDAC, to characterize and quantify how a storable propellant design would affect and improve overall system performance of the Government reference design. As part of that miniDAC, HLSP ran multiple trade studies and assessments, including aspects of Mission Design, Flight Mechanics (FM), and Guidance, Navigation, and Control (GN&C). The storable propellant designs were assessed through the baseline mission targeting the Near Rectilinear Halo Orbit (NRHO) and several alternate staging orbits. This paper focuses on the FM and GN&C trades and assessments completed as part of the miniDAC as well as the GN\&C elements of the miniDAC Master Equipment List (MEL) developed for the storable propellant lander configurations.
This paper discusses the Concept of Operations of the nine contingency strategies available for off-nominal Rendezvous, Proximity Operations, and Docking of Orion with the Gateway in a Near-Rectilinear Halo Orbit around the Moon. Explanations are provided for each contingency strategy, how they are initiated, when they can be initiated, and how they are designed to protect the crew. Then an overview is provided of the sequencing of the Guidance, Navigation, and Control flight software used to achieve each contingency strategy, in the form of Phases, Segments, Activities, and Modes. These off-nominal scenarios re-quire the definition of new Segments and Activities.
NASA’s Marshall Space Flight Center (MSFC) is developing the Solar Cruiser solar sail mission to continue to mature solar sail propulsion. Solar Cruiser is a 95kg spacecraft capable of deploying a 1654 square meter solar sail. Solar Cruiser will demonstrate science observation capabilities and navigation in sub Lagrange Point (L1) halo-orbits. Solar Cruiser sail technology can be scaled to larger solar sail missions, including attitude control system and sail membrane and boom technology. Larger solar sails could be used to image the solar poles, enabling Solar Polar Imager (SPI) missions currently prohibitively expensive for propellant-based propulsion methods. Similarly, solar sails can be used to create artificial equilibria and indefinite station-keeping at locations sunward of Lagrange Point one, L1, along the Sun-Earth line (SEL), which can provide space-weather monitor and prediction and reveal discoveries about our Sun and solar system. To achieve the high characteristic acceleration required for larger solar sail missions such as SPI, a solar sail area of approximately 7000 square-meters would be needed. In comparison, the largest solar sail flown to date is the JAXA’s IKAROS mission with a 196 square-meters sail. Attitude control an SPI-size sail poses challenges, including station keeping under solar sail induced torques and thrust vector pointing. In this study, the scalability of the Solar Cruiser sail attitude control architecture to larger sails is studied, such as pointing control performance, mass and power requirements. Solar Cruiser attitude control actuators include propellant-based, such as ion engines, and propellantless options which create differential solar pressures to generate spacecraft torques. Propellantless attitude control includes active translation of the center of mass, Reflectivity Control Devices (RCDs), and tip vanes. The study summarizes the scalability characteristics of the different attitude control architectures for larger solar sail missions.
The Genesis mission will launch in 2001, sending a spacecraft into an L1 halo orbit in the Sun-Earth system to collect solar wind samples.