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NASA NTRSDate not supplied

MISTRAL: Concept and Analysis of Persistent Airborne Localization of GHG Emissions

In this study, we propose a concept for localizing and mitigating greenhouse gases and other emissions from areas of various sizes, using a multiscale, airborne localization fleet, supported by information from satellites. Our objective is a rapidly implementable capability for the reduction of atmospheric methane by empowering the communities in proximity to emissions to actively contribute to methane mitigation via fleet operations by local governments, first responders, and individuals . Fully autonomous operations are economically desirable and technically feasible now. However, the development of regulations for autonomous operations is in early stages. Since rapid impact is critical to reducing emissions, the concept starts with remotely controlled vehicles. As regulations for autonomous operations reach maturity, the concept can be easily transitioned from the remotely piloted to autonomous mode. We report on a preliminary system analysis of the concept. Localization of methane emissions from the Permian Basin oil and gas production site serves as the design reference mission; however, the airborne concept is applicable to a broad range of use cases, and we describe several in this analysis. We consider leveraging satellite and ground-based resources within the airborne concept of operations. The study analyses uninhabited aerial system localization operations under the most conservative assumptions on resources and available information. Even under these assumptions, the concept is shown to be viable for local community operations.

airborne methane localization↗
NASA NTRS2015

Radioisotope Power Systems Reference Book for Mission Designers and Planners

The RPS Program's Program Planning and Assessment (PPA) Office commissioned the Mission Analysis team to develop the Radioisotope Power Systems (RPS) Reference Book for Mission Planners and Designers to define a baseline of RPS technology capabilities with specific emphasis on performance parameters and technology readiness. The main objective of this book is to provide RPS technology information that could be utilized by future mission concept studies and concurrent engineering practices. A progress summary from the major branches of RPS technology research provides mission analysis teams with a vital tool for assessing the RPS trade space, and provides concurrent engineering centers with a consistent set of guidelines for RPS performance characteristics. This book will be iterated when substantial new information becomes available to ensure continued relevance, serving as one of the cornerstone products of the RPS PPA Office. This book updates the original 2011 internal document, using data from the relevant publicly released RPS technology references and consultations with RPS technologists. Each performance parameter and RPS product subsection has been reviewed and cleared by at least one subject matter representative. A virtual workshop was held to reach consensus on the scope and contents of the book, and the definitions and assumptions that should be used. The subject matter experts then reviewed and updated the appropriate sections of the book. The RPS Mission Analysis Team then performed further updates and crosschecked the book for consistency. Finally, a second virtual workshop was held to ensure all subject matter experts and stakeholders concurred on the contents.

Lee, Young↗
NASA NTRS1990

Conceptual spacecraft and arcjet propulsion system design for the SP-100 interim reference mission

An arcjet propulsion system, delivering 7.5 N thrust, was defined for the SP-100 Space Reactor Power System Interim Reference Mission. Conceptual design trades and configuration studies of a spacecraft suitable for the mission were performed to the extent necessary to define propulsion system requirements and interfaces. The propulsion system design is based on the use of 30-kW, constricted arc, ammonia arcjet engines operating in parallel during orbit boost from low earth to geosynchronous orbit.

Zafran, S.↗
NASA NTRS2003

Mars Global Reference Atmospheric Model (Mars-GRAM) and Database for Mission Design

Mars Global Reference Atmospheric Model (Mars-GRAM 2001) is an engineering-level Mars atmosphere model widely used for many Mars mission applications. From 0-80 km, it is based on NASA Ames Mars General Circulation Model, while above 80 km it is based on Mars Thermospheric General Circulation Model. Mars-GRAM 2001 and MGCM use surface topography from Mars Global Surveyor Mars Orbiting Laser Altimeter. Validation studies are described comparing Mars-GRAM with Mars Global Surveyor Radio Science and Thermal Emission Spectrometer data. RS data from 2480 profiles were used, covering latitudes 75 deg S to 72 deg N, surface to approximately 40 km, for seasons ranging from areocentric longitude of Sun (Ls) = 70-160 deg and 265-310 deg. RS data spanned a range of local times, mostly 0-9 hours and 18-24 hours. For interests in aerocapture and precision landing, comparisons concentrated on atmospheric density. At a fixed height of 20 km, RS density varied by about a factor of 2.5 over ranges of latitudes and Ls values observed. Evaluated at matching positions and times, these figures show average RSMars-GRAM density ratios were generally 1+/-)0.05, except at heights above approximately 25 km and latitudes above approximately 50 deg N. Average standard deviation of RSMars-GRAM density ratio was 6%. TES data were used covering surface to approximately 40 km, over more than a full Mars year (February, 1999 - June, 2001, just before start of a Mars global dust storm). Depending on season, TES data covered latitudes 85 deg S to 85 deg N. Most TES data were concentrated near local times 2 hours and 14 hours. Observed average TES/Mars-GRAM density ratios were generally 1+/-0.05, except at high altitudes (15-30 km, depending on season) and high latitudes (greater than 45 deg N), or at most altitudes in the southern hemisphere at Ls approximately 90 and 180 deg. Compared to TES averages for a given latitude and season, TES data had average density standard deviation about the mean of approximately 2.5% for all data, or approximately 1-4%, depending on time of day and dust optical depth. Average standard deviation of TES/Mars-GRAM density ratio was 8.9% for local time 2 hours and 7.1% for local time 14 hours. Thus standard deviation of observed TES/Mars-GRAM density ratio, evaluated at matching positions and times, is about three times the standard deviation of TES data about the TES mean value at a given position and season.

Justus, C. G.↗
NASA NTRS1981

The Venus Orbiting Imaging Radar Mission

The scientific objectives and rationale for a Venus Orbiting Imaging Radar Mission are presented. A provisional science payload responsive to these objectives is described and a reference set of measurement requirements and their priorities is established. Those high priority measurements that are most demanding on spacecraft and mission design are used to develop a reference spacecraft design and a reference mission design. A discussion of mission performance issues is also included.

Dallas, S. S.↗
NASA NTRS1989

Mars Rover and Sample Return Mission design

The current reference Mars Rover and Sample Return mission is described. Technical issues are outlined, including high-resolution image acquisition and reconstruction, approach navigation, ground and flight systems operational complexity, rover autonomy, autonomous rendezvous and docking in Mars orbit, aerocapture and aeromaneuver, estimating the probability of mission success, and end-to-end information system design. Focus is placed on lander hazard identification and avoidance, pinpoint landing guidance and control, Mars ascent vehicle guidance and control, planetary protection and quarantine, sample acquisition and preservation, project management and control, systems requirements and interface control, and costing. In addition, program issues such as international participation, fiscal constraints, and launch-vehicle availability are considered.

Kwok, Johnny H.↗
NASA NTRS2013

A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions

The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.

Mercer, Carolyn R.↗
NASA NTRS2014

A Combined Solar Electric and Storable Chemical Propulsion Vehicle for Piloted Mars Missions

The Mars Design Reference Architecture (DRA) 5.0 explored a piloted Mars mission in the 2030 timeframe, focusing on architecture and technology choices. The DRA 5.0 focused on nuclear thermal and cryogenic chemical propulsion system options for the mission. Follow-on work explored both nuclear and solar electric options. One enticing option that was found in a NASA Collaborative Modeling for Parametric Assessment of Space Systems (COMPASS) design study used a combination of a 1-MW-class solar electric propulsion (SEP) system combined with storable chemical systems derived from the planned Orion crew vehicle. It was found that by using each propulsion system at the appropriate phase of the mission, the entire SEP stage and habitat could be placed into orbit with just two planned Space Launch System (SLS) heavy lift launch vehicles assuming the crew would meet up at the Earth-Moon (E-M) L2 point on a separate heavy-lift launch. These appropriate phases use high-thrust chemical propulsion only in gravity wells when the vehicle is piloted and solar electric propulsion for every other phase. Thus the SEP system performs the spiral of the unmanned vehicle from low Earth orbit (LEO) to E-M L2 where the vehicle meets up with the multi-purpose crew vehicle. From here SEP is used to place the vehicle on a trajectory to Mars. With SEP providing a large portion of the required capture and departure changes in velocity (delta V) at Mars, the delta V provided by the chemical propulsion is reduced by a factor of five from what would be needed with chemical propulsion alone at Mars. This trajectory also allows the SEP and habitat vehicle to arrive in the highly elliptic 1-sol parking orbit compatible with envisioned Mars landing concepts. This paper explores mission options using between SEP and chemical propulsion, the design of the SEP system including the solar array and electric propulsion systems, and packaging in the SLS shroud. Design trades of stay time, power level, specific impulse and propellant type are discussed.

Mercer, Carolyn R.↗
NASA NTRS2011

A Phobos-Deimos Mission as an Element of the NASA Mars Design Reference Architecture 5.0

NASA has conducted a series of mission studies over the past 25 years examining the eventual exploration of the surface of Mars by humans. The latest version of this evolutionary series of design reference missions/architectures - Design Reference Architecture 5 or DRA-5 - was completed in 2007. This paper examines the implications of including a human mission to explore the moons of Mars and teleoperate robots in various locations, but not to land the human crews on Mars, as an element of this reference architecture. Such a mission has been proposed several times during this same 25 year evolution leading up to the completion of DRA-5 primarily as a mission of testing the in-space vehicles and operations while surface vehicles and landers are under development. But such a precursor or test mission has never been explicitly included as an element of this Architecture. This paper will first summarize the key features of the DRA-5 to provide context for the remainder of the assessment. This will include a description of the in-space vehicles that would be the subject of a shakedown test during the Mars orbital mission. A decision tree will be used to illustrate the factors that will be analyzed, and the sequence in which they will be addressed, for this assessment. The factors that will be analyzed include the type of interplanetary transfer orbit (opposition class versus conjunction class), the type of parking orbit (circular versus elliptical), and the type of propulsion technology (high thrust chemical versus nuclear thermal rocket). The manner in which each of these factors impacts an individual mission will be described. In addition to the direct impact of these factors, additional considerations impacting crew health and overall programmatic outcomes will be discussed. Numerical results for each of the factors in the decision tree will be grouped with derived qualitative impacts from crew health and programmatic consideration. These quantitative and qualitative results will be summarized in a pros/cons table as a summary for this analysis.

Hoffman, Stephen J.↗
NASA NTRS2019

Improvements to the Copernicus Trajectory Design and Optimization System for Complex Space Trajectories

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.

Murri, Daniel G.↗
NASA NTRS2011

Optical Design Trade Study for the Wide Field Infrared Survey Telescope [WFIRST]

The Wide Field Infrared Survey Telescope (WFIRST) mission concept was ranked first in new space astrophysics mission by the Astro2010 Decadal Survey incorporating the Joint Dark Energy Mission (JDEM)-Omega payload concept and multiple science white papers. This mission is based on a space telescope at L2 studying exoplanets [via gravitational microlensing], probing dark energy, and surveying the near infrared sky. Since the release of NWNH, the WFIRST project has been working with the WFIRST science definition team (SDT) to refine mission and payload concepts. We present the driving requirements. The current interim reference mission point design, based on the use of a 1.3m unobscured aperture three mirror anastigmat form, with focal imaging and slitless spectroscopy science channels, is consistent with the requirements, requires no technology development, and out performs the JDEM-Omega design.

Content, David A.↗
NASA NTRS2013

Wide Field Infra-Red Survey Telescope (WFIRST) 2.4-Meter Mission Study

The most recent study of the Wide Field Infrared Survey Telescope (WFIRST) mission is based on reuse of an existing 2.4m telescope. This study was commissioned by NASA to examine the potential science return and cost effectiveness of WFIRST by using this significantly larger aperture telescope. We review the science program envisioned by the WFIRST 2012-2013 Science Definition Team (SDT), an overview of the mission concept, and the telescope design and status. Comparisons against the previous 1.3m and reduced cost 1.1m WFIRST design concepts are discussed. A significant departure from past point designs is the option for serviceability and the geostationary orbit location which enables servicing and replacement instrument insertion later during mission life. Other papers at this conference provide more in depth discussion of the wide field instrument and the optional exoplanet imaging coronagraph instrument.

exoplanet imaging coronagraph↗
NASA NTRS1975

Jovian system science issues and implications for a Mariner Jupiter Orbiter mission

Science goals for missions to Jupiter in the early 1980's are reviewed and a case is made for the science community to play the key role in assigning relative priorities for these goals. A reference set of measurement requirements and their priorities is established and those high priority goals that are most demanding on spacecraft and mission design are used to develop a reference mission concept. An orbiter mission is required to satisfy a majority of the measurements, and a spacecraft data handling capability as least equivalent to the Mariner Jupiter/Saturn spacecraft is the major system design driver. This reference Mission Concept is called Mariner Jupiter Orbiter. The remaining measurement requirements are reviewed in light of the potential science return of this mission, and certain options are developed to augment this science return. Two attractive options fulfill high priority objectives not achieved by the reference Mariner Jupiter Orbiter mission alone: an atmospheric entry probe, released prior to orbit insertion; and a daughter satellite dedicated to particle and fields measurements, ejected into an independent orbit about Jupiter.

Beckman, J. C.↗
NASA NTRS2017

UAV Inspection of Electrical Transmission Infrastructure with Path Conformance Autonomy and Lidar-Based Geofences NASA Report on UTM Reference Mission Flights at Southern Company Flights November 2016

Flights at low altitudes in close proximity to electrical transmission infrastructure present serious navigational challenges: GPS and radio communication quality is variable and yet tight position control is needed to measure defects while avoiding collisions with ground structures. To advance unmanned aerial vehicle (UAV) navigation technology while accomplishing a task with economic and societal benefit, a high voltage electrical infrastructure inspection reference mission was designed. An integrated air-ground platform was developed for this mission and tested in two days of experimental flights to determine whether navigational augmentation was needed to successfully conduct a controlled inspection experiment. The airborne component of the platform was a multirotor UAV built from commercial off-the-shelf hardware and software, and the ground component was a commercial laptop running open source software. A compact ultraviolet sensor mounted on the UAV can locate 'hot spots' (potential failure points in the electric grid), so long as the UAV flight path adequately samples the airspace near the power grid structures. To improve navigation, the platform was supplemented with two navigation technologies: lidar-to-polyhedron preflight processing for obstacle demarcation and inspection distance planning, and trajectory management software to enforce inspection standoff distance. Both navigation technologies were essential to obtaining useful results from the hot spot sensor in this obstacle-rich, low-altitude airspace. Because the electrical grid extends into crowded airspaces, the UAV position was tracked with NASA unmanned aerial system traffic management (UTM) technology. The following results were obtained: (1) Inspection of high-voltage electrical transmission infrastructure to locate 'hot spots' of ultraviolet emission requires navigation methods that are not broadly available and are not needed at higher altitude flights above ground structures. (2) The sensing capability of a novel airborne UV detector was verified with a standard ground-based instrument. Flights with this sensor showed that UAV measurement operations and recording methods are viable. With improved sensor range, UAVs equipped with compact UV sensors could serve as the detection elements in a self-diagnosing power grid. (3) Simplification of rich lidar maps to polyhedral obstacle maps reduces data volume by orders of magnitude, so that computation with the resultant maps in real time is possible. This enables real-time obstacle avoidance autonomy. Stable navigation may be feasible in the GPS-deprived environment near transmission lines by a UAV that senses ground structures and compares them to these simplified maps. (4) A new, formally verified path conformance software system that runs onboard a UAV was demonstrated in flight for the first time. It successfully maneuvered the aircraft after a sudden lateral perturbation that models a gust of wind, and processed lidar-derived polyhedral obstacle maps in real time. (5) Tracking of the UAV in the national airspace using the NASA UTM technology was a key safety component of this reference mission, since the flights were conducted beneath the landing approach to a heavily used runway. Comparison to autopilot tracking showed that UTM tracking accurately records the UAV position throughout the flight path.

Moore, Andrew J.↗
NASA NTRS1973

The Jupiter orbiter satellite tour mission

Consideration of the problem of exploring the Jupiter-Galilean satellite system by a spacecraft in orbit about the planet. A simple patched conic algorithm is developed which identifies Jupiter orbiter trajectories that utilize gravity assist flybys of the Galilean satellites. A detailed reference mission is designed against specific science criteria for satellite exploration to demonstrate the possible science return of utilizing close satellite encounters. A navigation strategy is developed, including both planet and satellite relative orbit determination and a maneuver analysis; spacecraft delivery statistics are determined in the barycentric Jupiter-Galilean satellite coordinate system and mapped to the satellite target planes to determine feasibility. A simple maneuver sequence is employed which yields acceptable velocity requirements.

Beckman, J. C.↗
NASA NTRS2022

COMPASS Final Report: 2008 International Lunar Network (ILN)

The Compass team developed three designs for a soft lander for the International Lunar Network (ILN) using different power systems. These systems included solar arrays and batteries and two radioisotope power system (RPS) options scaled down from previous designs. These are a half Stirling radioisotope generator and a one-eighth Radioisotope Thermoelectric Generator (RTG). The COMPASS design consists of a soft-lander capable of network science (seismic, magnetospheric, laser reflector, and thermal transfer) that can land anywhere on the near side of the Moon and operate for at least six years. The baseline version, using only batteries and solar arrays, requires a very large battery (~80 kg) for the roughly 20 W night-time power requirements. The lander launch mass (with solid rocket breaking motor) is almost 900 kg. With such a large battery, the lander will require a launcher larger than the desired Minotaur V; a Taurus II (being developed for NASA Commercial Orbital Transportation Services (COTS)) or larger (Delta II or Atlas V) would be needed. The two RPS cases, half an Advanced Stirling Radioisotope Generators (ASRG) and one-eighth RTG have much smaller launch masses of 500 and 600 kg respectively. (The one-eighth RTG case lacks the power for the full science package, so a solar array battery is still needed.) Unfortunately, the Minotaur V’s launch capability is still not large enough (~430 kg to trans-lunar injection (TLI)) for the RPS systems. Consequently, a fourth case was developed to fly a limited science floor suite of instruments (seismometer only) which reduced the RPS Lander launch mass to fit on a Minotaur V.

Radioisotope Power Systems↗
NASA NTRS1981

Space platform reference mission studies overview

The design requirements for three major space platform systems are identified. The three were the Science and Applications Space Platform (SASP), the Geostationary Platform (GSP), and the Satellite Power System (SPS). Because the SASP and GSP were assumed to require no advanced technology for their development an advanced version of each was selected on which to base the design requirements. The SPS represented the opposite development state hence a nearer term test article was selected on which to base the requirements. The development period for these missions is estimated.

Harrison, J. K.↗