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

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NASA NTRS1991

Spacely's rockets: Personnel launch system/family of heavy lift launch vehicles

During 1990, numerous questions were raised regarding the ability of the current shuttle orbiter to provide reliable, on demand support of the planned space station. Besides being plagued by reliability problems, the shuttle lacks the ability to launch some of the heavy payloads required for future space exploration, and is too expensive to operate as a mere passenger ferry to orbit. Therefore, additional launch systems are required to complement the shuttle in a more robust and capable Space Transportation System. In December 1990, the Report of the Advisory Committee on the Future of the U.S. Space Program, advised NASA of the risks of becoming too dependent on the space shuttle as an all-purpose vehicle. Furthermore, the committee felt that reducing the number of shuttle missions would prolong the life of the existing fleet. In their suggestions, the board members strongly advocated the establishment of a fleet of unmanned, heavy lift launch vehicles (HLLV's) to support the space station and other payload-intensive enterprises. Another committee recommendation was that a space station crew rotation/rescue vehicle be developed as an alternative to the shuttle, or as a contingency if the shuttle is not available. The committee emphasized that this vehicle be designed for use as a personnel carrier, not a cargo carrier. This recommendation was made to avoid building another version of the existing shuttle, which is not ideally suited as a passenger vehicle only. The objective of this project was to design both a Personnel Launch System (PLS) and a family of HLLV's that provide low cost and efficient operation in missions not suited for the shuttle.

Source record↗
NASA NTRS2007

Contingency Operations during Failure of Inertial Attitude Acquisition Due to Star Tracker Blinding for Three-Axes-Stabilized Interplanetary Spacecraft

The three interplanetary ESA missions Mars-Express, Rosetta and Venus-Express (launched 2003, 2004 and 2005 resp.) are three-axes stabilized spacecraft (s/c) that estimate their inertial attitude (i.e. the attitude of the s/c w.r.t. the inertial frame) using measurements from a redundant set of star trackers (STR). Each s/c is equipped with four reaction wheels, a reaction control system based on thrusters and a redundant set of ring laser gyroscopes (gyros). The STR h/w layout of the three s/c is identical whereas there is a difference in the star pattern recognition algorithm of Rosetta which uses five neighbouring stars around a central star instead of star triads. The Rosetta algorithm has been implemented to cope with the presence of false stars which are expected to be seen during operations around the comet. The attitude acquisition capability from lost in space is different also in terms of AOCMS: The survival mode of Rosetta which is entered upon STR failure is presented. The AOCMS of Mars- and Venus-Express manages temporary STR outages during sky occultation by the planet not even by using redundancy. Though, a blinding of both STR during cruise lasting for the order of days confronts the ground operators with the limits of the AOCMS design. The operations and analyses that have been planned and partially been performed to compensate for the outage of the STR are demonstrated for Mars-Express. The caution measures taken before Venus orbit insertion of Venus-Express are detailed.

Keil, Joachim↗
NASA NTRS2024

Testing of Advanced Capabilities to Enable In-time Safety Management and Assurance for Future Flight Operations

In order to refine an initial Concept of Operations, explore Concepts of Use, and expose/validate requirements for future In-Time Aviation Safety Management Systems (IASMS), testing architectures were created, along with a set of capabilities and underlying information exchange protocols. These systems were conceived and developed based on hazards associated with two envisioned urban area flight domains: (1) highly autonomous small uncrewed aerial systems (sUAS) operating at low altitudes, and (2) highly autonomous air taxis. The initial scope of this development is described in [1]; this report provides an update, focusing on the subsequent developments and test activities. As stated in [1], it is important to note that there are many capabilities already in use by the industry (or soon to be in use) that will play critical roles in future IASMS designs. Those reported here were developed to address a gap in the current state-of-the-art regarding specific hazards/risks, and/or to allow for investigation of the interplay between and across hazard types — particularly regarding how overall safety risk can be reduced or managed effectively. Results of testing and development activities are organized by the operational phase wherein a particular capability would be employed (i.e., preflight, in-flight, and post-flight/off-line). Pre-flight: A set of capabilities were developed to help mitigate safety risk prior to flight (e.g., during flight and mission planning). Results of testing summarize (1) validation activities to raise the Technology Readiness Level (TRL) and (2) evaluation activities where the capabilities were applied to flight/mission planning procedures and used by operators/pilots. For the latter, flight plans were automatically assessed, and operators/pilots were notified of hazardous flight segments so as to enable adjustment of the flight plan and re-evaluation, and/or to better inform go/no-go decisions. Capabilities addressed hazards associated with power consumption, third-party risk, wind, navigation system performance, radiofrequency interference, and proximity to geo-spatial threats (e.g., buildings, trees, and no-fly zones). In-flight: Flight experiments tested capabilities that detect and respond to hazards encountered during flight. In the first series, safety hazards were monitored and assessed onboard, and system-generated mitigation maneuvers were recorded (but not acted upon by the vehicle). In the second series, mitigation maneuver commands directed the aircraft in response to safety hazards (i.e., auto-mitigation). The sUAS used for testing is described in full, as is the test architecture, which included commercial avionics, research avionics, and onboard software designed to detect, assess, and respond to hazards. The onboard system was designed as a run-time assurance framework, consistent with [2] and supportive of both supervisory and automated modes. The primary functions included: real-time risk assessment (RTRA), auto-pilot monitoring, constraint monitoring, and contingency select/triggering. RTRA performs integrated risk assessment considering data from several hazard-related monitors (e.g., battery, motors, navigation, communications, population density, and loss-of-control). Post-flight/off-line: Data monitored and recorded during flights can enable IASMS capabilities that execute after flights have completed (or “off-line”). These include: (1) the ability to identify anomalies and trends that may only be observable when comparing data spanning a number of similar flights; (2) the ability to update and validate pre-flight and in-flight capabilities and any underlying models to improve their performance; (3) the ability to report anomalies/off-nominals that may indicate design changes or maintenance actions are needed; and (4) the ability for humans involved in operations to report safety-relevant observations to help in understanding the flight data and/or the operational context of a flight. Progress on three such capabilities is summarized; the first investigates anomaly detection given a limited set of flight logs and applies an approach previously used for space operations. The second explores what could be identified using a larger set of flight logs, including from web-based forums where flight logs are posted by sUAS autopilot users. The third creates a new means of collecting information on UAS incidents and accidents via the Aviation Safety Reporting System (ASRS).

sUAS↗
NASA NTRS2013

Optimal Propellant Maneuver Flight Demonstrations on ISS

In this paper, first ever flight demonstrations of Optimal Propellant Maneuver (OPM), a method of propulsive rotational state transition for spacecraft controlled using thrusters, is presented for the International Space Station (ISS). On August 1, 2012, two ISS reorientations of about 180deg each were performed using OPMs. These maneuvers were in preparation for the same-day launch and rendezvous of a Progress vehicle, also a first for ISS visiting vehicles. The first maneuver used 9.7 kg of propellant, whereas the second used 10.2 kg. Identical maneuvers performed without using OPMs would have used approximately 151.1kg and 150.9kg respectively. The OPM method is to use a pre-planned attitude command trajectory to accomplish a rotational state transition. The trajectory is designed to take advantage of the complete nonlinear system dynamics. The trajectory choice directly influences the cost of the maneuver, in this case, propellant. For example, while an eigenaxis maneuver is kinematically the shortest path between two orientations, following that path requires overcoming the nonlinear system dynamics, thereby increasing the cost of the maneuver. The eigenaxis path is used for ISS maneuvers using thrusters. By considering a longer angular path, the path dependence of the system dynamics can be exploited to reduce the cost. The benefits of OPM for the ISS include not only reduced lifetime propellant use, but also reduced loads, erosion, and contamination from thrusters due to fewer firings. Another advantage of the OPM is that it does not require ISS flight software modifications since it is a set of commands tailored to the specific attitude control architecture. The OPM takes advantage of the existing ISS control system architecture for propulsive rotation called USTO control mode1. USTO was originally developed to provide ISS Orbiter stack attitude control capability for a contingency tile-repair scenario, where the Orbiter is maneuvered using its robotic manipulator relative to the ISS. Since 2005 USTO has been used for nominal ISS operations.

Bhatt, Sagar↗
NASA NTRS2007

Multi-Sensor Testing for Automated Rendezvous and Docking Sensor Testing at the Flight Robotics Laboratory

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as AR&D). The crewed missions may also perform rendezvous and docking operations and may require different levels of automation and/or autonomy, and must provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success of the Exploration Program. NASA has the responsibility to determine whether the Crew Exploration Vehicle (CEV) contractor proposed relative navigation sensor suite will meet the requirements. The relatively low technology readiness level of AR&D relative navigation sensors has been carried as one of the CEV Project's top risks. The AR&D Sensor Technology Project seeks to reduce the risk by the testing and analysis of selected relative navigation sensor technologies through hardware-in-the-loop testing and simulation. These activities will provide the CEV Project information to assess the relative navigation sensors maturity as well as demonstrate test methods and capabilities. The first year of this project focused on a series of"pathfinder" testing tasks to develop the test plans, test facility requirements, trajectories, math model architecture, simulation platform, and processes that will be used to evaluate the Contractor-proposed sensors. Four candidate sensors were used in the first phase of the testing. The second phase of testing used four sensors simultaneously: two Marshall Space Flight Center (MSFC) Advanced Video Guidance Sensors (AVGS), a laser-based video sensor that uses retroreflectors attached to the target vehicle, and two commercial laser range finders. The multi-sensor testing was conducted at MSFC's Flight Robotics Laboratory (FRL) using the FRL's 6-DOF gantry system, called the Dynamic Overhead Target System (DOTS). The target vehicle for "docking" in the laboratory was a mockup that was representative of the proposed CEV docking system, with added retroreflectors for the AVGS. The multi-sensor test configuration used 35 open-loop test trajectories covering three major objectives: (1) sensor characterization trajectories designed to test a wide range of performance parameters; (2) CEV-specific trajectories designed to test performance during CEV-like approach and departure profiles; and (3) sensor characterization tests designed for evaluating sensor performance under more extreme conditions as might be induced during a spacecraft failure or during contingency situations. This paper describes the test development, test facility, test preparations, test execution, and test results of the multi-sensor series of trajectories.

Brewster, L.↗
NASA NTRS2008

Multi-Sensor Testing for Automated Rendezvous and Docking Sensor Testing at the Flight Robotics Lab

The Exploration Systems Architecture defines missions that require rendezvous, proximity operations, and docking (RPOD) of two spacecraft both in Low Earth Orbit (LEO) and in Low Lunar Orbit (LLO). Uncrewed spacecraft must perform automated and/or autonomous rendezvous, proximity operations and docking operations (commonly known as AR&D). The crewed missions may also perform rendezvous and docking operations and may require different levels of automation and/or autonomy, and must provide the crew with relative navigation information for manual piloting. The capabilities of the RPOD sensors are critical to the success ofthe Exploration Program. NASA has the responsibility to determine whether the Crew Exploration Vehicle (CEV) contractor-proposed relative navigation sensor suite will meet the requirements. The relatively low technology readiness level of AR&D relative navigation sensors has been carried as one of the CEV Project's top risks. The AR&D Sensor Technology Project seeks to reduce the risk by the testing and analysis of selected relative navigation sensor technologies through hardware-in-the-Ioop testing and simulation. These activities will provide the CEV Project information to assess the relative navigation sensors maturity as well as demonstrate test methods and capabilities. The first year of this project focused on a series of "pathfinder" testing tasks to develop the test plans, test facility requirements, trajectories, math model architecture, simulation platform, and processes that will be used to evaluate the Contractor-proposed sensors. Four candidate sensors were used in the first phase of the testing. The second phase of testing used four sensors simultaneously: two Marshall Space Flight Center (MSFC) Advanced Video Guidance Sensors (AVGS), a laser-based video sensor that uses retroreflectors attached to the target vehicle, and two commercial laser range finders. The multi-sensor testing was conducted at MSFC's Flight Robotics Laboratory (FRL) using the FRL's 6-DOF gantry system, called the Dynamic Overhead Target System (DOTS). The target vehicle for "docking" in the laboratory was a mockup that was representative of the proposed CEV docking system, with added retroreflectors for the AVGS.' The multi-sensor test configuration used 35 open-loop test trajectories covering three major objectives: (l) sensor characterization trajectories designed to test a wide range of performance parameters; (2) CEV-specific trajectories designed to test performance during CEV-like approach and departure profiles; and (3) sensor characterization tests designed for evaluating sensor performance under more extreme conditions as might be induced during a spacecraft failure or during contingency situations. This paper describes the test development, test facility, test preparations, test execution, and test results of the multisensor series oftrajectories

Brewster, Linda L.↗
NASA NTRSDate not supplied

Preparation for an Earth Independent Medical Operations Demonstration using the Tempus ALS™ Medical Device

NASA’s exploration-class missions have severe resource constraints, long return trip durations, significant communication delays and limited resupply opportunities. Validation on the International Space Station (ISS) of medical devices that fit within Earth-Independent Medical Operations (EIMO) systems is a necessary preparation step. Key features of an EIMO medical system include: 1) technologies that support the prevention, diagnosis, and treatment of spaceflight medical events; 2) components that meet mass, volume, power and crew time/training constraints; 3) consideration of the medical skill level of the astronaut caregiver; 4) collection, storage and analysis of medical data within a central data architecture; and 5) incorporation of appropriate guidance and support tools that allow crew autonomy. The Human Research Program’s (HRP) Exploration Medical Capability (ExMC) Element and the Exploration Medical Integrated Product Team (XM-IPT) are planning an ISS technology demonstration to determine the feasibility of including a multifunctional medical device into an EIMO medical system. Demonstration Preparations: The Tempus ALS (Remote Diagnostic Technologies, Ltd., Philips Corp., Farnborough, UK) is a commercial-off-the-shelf medical device, with United States Food and Drug Administration clearance and is Conformité Européene marked in Europe. Several thousand units have been sold and are successfully operating in pre-hospital and in remote settings, such as by European Space Agency (ESA) flight surgeons and some of NASA’s commercial partners during post-flight medical operations. The Tempus ALS provides vital sign measurements such as blood pressure, electrocardiograms, heart rate, end tidal CO2, respiration rate, pulse oximetry and temperature. The device has ultrasound imaging and video laryngoscopy capabilities and has automatic or manual defibrillation modes for treating cardiac arrhythmias. It includes procedural guidance capabilities to assist in the collection of the vital sign measurements and it has various data transmission and report generation features. NASA’s HRP ExMC and XM-IPT are partnering with the ESA to demonstrate the Tempus ALS on ISS, with ESA manifesting the Tempus ALS and its accessories. ESA will compare performance of periodic health status exams and medical contingency drills performed nominally and with the Tempus ALS. NASA’s EIMO demonstration will include the use of Tempus ALS to diagnose a complaint of abdominal pain under increasingly independent circumstances. The caregiver must assess the present illness, collect vital sign measurements, and perform an abdominal ultrasound, under one of three communication situations, including real-time, with a several second delay and with a delay on the order of minutes. Expected Outcomes: Information will be gained about the feasibility, benefits, and challenges of using a multifunctional, all-in-one, medical device for medical diagnosis instead of separate devices with singular functionalities. Information will also be collected about performance differences as communication delays increase and available ground support decreases. Gaining this understanding will allow for further development of exploration medical system capabilities which takes the EIMO construct into consideration.

Beth Lewandowski↗
NASA NTRS2020

Lunar Transient Accelerations White Paper

Recently, the United States announced a plan to return astronauts to the moon by 2024 [8]. Lunar landing (and subsequent Mars landing) architecture was not considered when the current NASA standards and vehicle design requirements for crew injury risk were developed. Therefore a gap exists in protecting the crew in planetary landing scenarios. One of the interesting aspects of this design reference mission (DRM)is the consideration of having the crew stand during dynamic phases of flight. Although this approach was contemplated for the Apollo Lunar Module, current NASA standards do not address design solutions that allow the crew to stand. Currently, NASA uses several tools and associated limits to mitigate crew injury because of dynamic loads. Some of these tools are the Brinkley Dynamic Response Criterion (BDRC) model and Hybrid III Anthropomorphic Test Devices (ATDs) [10]; and these have several limitations for assessing spaceflight loading environments, as well as specific underlying assumptions that may not be applicable in planetary landing vehicles. The BDRC is a simple lumped mass parameter model developed by the U.S. military, and has been used primarily to evaluate injury risk associated with aircraft ejection systems. The model evaluates seat accelerations in each axis to determine injury risk. Because the model treats the human-seat-restraint system as a single system, it is contingent on a restraint system and seat with similar characteristics of the original test data underlying the model. In particular, the model requires a rigid seat with a minimum natural frequency of 15Hz, minimal seat pan padding, side supports, and multipoint harness. The BDRC model uses undamped natural frequency and damping coefficients based on these requirements and any deviations may render the model injury predictions void. In lunar landing, one expects that a minimal or even no seat with minimal restraints will be employed. In this case, the original model parameters are likely to not be applicable. For capsule-based spacecraft returning crew to Earth, the Hybrid III ATD in various sizes also is used to supplement the BDRC. This analytical tool was added to address limitations in the BDRC related to spacecraft landings while wearing a pressure garment and helmet. Although the Hybrid III ATD has additional measurement capability; head, neck,and lumbar spine responses were the only metrics included because of the limitations imposed by the model.Lunar landing acceleration limits must assume the crew is standing during landing, an orientation for which we have limited data.Although the Apollo missions did employ a standing orientation for the crew, much of the data is lost. Therefore data from othersources havebeen examined to inform lunar landing acceleration limits.

Jeffrey T Somers↗
NASA NTRSDate not supplied

VERNE: Revealing the Mysteries and Histories of Venus

Introduction: The three Venus missions that were recently selected for upcoming flight (VERITAS, DAVINCI+, and EnVision) will be incredibly valuable to our understanding of Venus’ history, geology, and atmosphere. However, even once completed, key gaps in our knowledge of Venus, and more generally the formation and active processes on rocky, Earth-like planets, will still persist. Remaining questions include 1) how global intrinsic magnetic fields might be maintained on rocky worlds, and how they could then go extinct, and 2) what role atmospheric sulfur chemistry plays in climates of Earth-like planets, which is an increasingly timely subject as Earth’s own atmospheric sulfur content is climbing due to human activity. These questions require in-situ observations from Venus’ cloud deck, at the altitudes at which the UV absorber exists. The Venus Environment Research and Novel Exploration (VERNE) mission will address these questions with an aerial platform that will drift around the equatorial region of the planet for 9 days. VERNE will collect data to determine the identity of the mysterious UV absorber, while also taking magnetic field measurements over the tesserae, the regions on Venus that are most likely to retain remanent crustal magnetization, in order to understand the potential role of a past intrinsically-generated global magnetic field on Venus. Mission Objectives: The two major science objectives that drive the VERNE mission are: 1) Determine the identity of the Venusian unknown ultraviolet absorber(s). First observed approximately a century ago [1], the composition of Venus’ ultraviolet (UV) absorber is one of the oldest mysteries in Venus atmospheric chemistry [2,3]. While several candidate UV absorbers (mostly sulfur species) have been proposed, no consensus has been reached on its composition and its specific interactions with the atmosphere. Determining the identity of the UV absorber will aid climate models by showing how and where incident solar energy is absorbed by the atmosphere and make progress towards understanding the chemical and energetic processes taking place above Venus’ upper cloud deck [4]. 2) Determine if Venus retains evidence of a past, internally-generated magnetic field. While Venus does not currently have an intrinsically-generated magnetic field, evidence for the existence of a past field on Venus and a timeline of its decay will fill in a more holistic picture of the evolution of Venus’ geological record and atmosphere. As the oldest geologic units on the surface, Venus’ tesserae may still have remanent crustal magnetization signatures within the rocky composition [5]. In any case, the signatures detected will provide insight into Venus’ past geological and core dynamo activity. Mission Summary: VERNE includes a 3-part flight system made up of 1) an entry system with a HEEET (Heatshield for Extreme Entry Environment Technology) aeroshell, 2) an orbiter for data relay, and 3) an in-situ balloon and gondola. After entry into Venus’ atmosphere, the balloon will be deployed within the upper cloud deck, at an altitude of 62 km, over a tesserae region. The in-situ data collection will last for the duration of 2 full circumnavigations of the planet, which will take ~9 days. Instrumentation Suite: The four instruments that comprise VERNE’s instrument payload will enable the identification of the unknown UV absorber and the detection of remanent crustal magnetization if it exists in the tesserae. The proposed instrument suite cycle during mission operations is shown in Fig. 1. The four instruments are described below: 1) Adams (ion neutral mass spectrometer) has a range of 18-257 AMU and will detect and distinguish the mixing ratios of O₂, H₂O, H₂SO₄, S, S₂, S₃, S₄, S₅, S₆, S₇, S₈, SO, SO₂, OSSO, SO₃, Cl₂, FeCl₃ and other trace sulfur and organic species. The spatial (longitudinal) and temporal (day/night) variations will be observed throughout 2 circumnavigations with a sample cadence of 12 minutes. 2) Shelley (nephelometer) will determine the size distribution of the aerosols (0.4 to 36 um) in the atmosphere. With a size resolution of <0.7 um, it can determine which mode of H2SO4 is present and characterize the large (>30 um) organic particles previously detected by the Venera and Galileo missions [6,7]. 3) Herbert (UV imager) will measure UV radiance at 283 nm (the wavelength of SO2 absorption) and 365 nm (the unknown part of the absorber). UV images will be taken concurrently with the INMS and nephelometer to correlate UV absorption with abundances of the UV absorber species. 4) Vonnegut (magnetometer) has a range of >600 nT and a precision and accuracy of 1 nT. If magnetized by a past field, the crust may have retained a magnetization of up to 3 A/m2 [5]. With a noise floor of 10 nT, the magnetometer will be able to detect RCM from an altitude of 62 km, even if the thickness of the magnetized crust is just 1 km (Fig. 2). Mission Concept Design: VERNE will be launched with a mass of 3300 kg in an intermediate-high performance class vehicle with a 4-m fairing. The 475-day mission includes 466 days for the cruise, coasting, and orbit initialization phases before entry, descent, and balloon deployment. During the 9-day science phase, the aerial platform will make 2 circumnavigations of the planet at an altitude of 62 km. The INMS and the nephelometer will acquire data for 2 hours during the daytime and nighttime during each circumnavigation, while the UV imager will be on for the duration of the daytime, and the magnetometer will be operational throughout the entirety of the science phase. Data will be stored and processed with the JPL-designed Sphinx command and data handling system. Data will be sent from the balloon to the orbiter using an S-band relay link, stored on the orbiter, and then forwarded to Earth where it will be received by the DSN. Conclusion: VERNE will fill key gaps in our understanding of the history and ongoing processes related to the geology and atmosphere of Venus and rocky worlds in general. Even with adequate flight system contingencies and expected costs below the $900M New Frontiers cost cap, VERNE is not without its risks and challenges. Further trade spaces to explore include 1) using solely battery power vs. including solar panels to increase the mission duration and 2) investigating the use of lightweight materials and 3D-printed structures to reduce the gondola mass, among others. Acknowledgments: We would like to thank the JPL Planetary Science Summer School, especially our mentors Troy Hudson, Karl Mitchell, and Leslie Lowes, as well as our Team-X study lead Al Nash and the members of Team-X. We’d additionally like to thank our review panel for asking insightful questions and providing valuable feedback. References: [1] Ross, F. E. (1928) Astrophysical J., 68, 57-92. [2] Rossow, W. B. et al. (1980) J. Geophysical Research, 85, 8107-8128. [3] Pinto, J. P. et al. (2021) Nature Communications, 12, 175. [4] Titov, D. V. et al. (2007) Cosmic Research, 21, 401. [5] O’Rourke, J. et al. (2019) Geophysical Research Lett., 46, 5768–5777. [6] Limaye, S. S. et al. (2018) Astrobiology, 18(9), 1181-1198. [7] Grinspoon, D. H. et al. (2013) Planetary and Space Sci., 41(7), 515-542. [8] Parker, R. L. (2003) J. Geophysical Research, 108, 5006. *The cost information contained in this document is of a budgetary and planning nature and is intended for informational purposes only. It does not constitute a commitment on the part of JPL and/or Caltech.

H Alpert↗
NASA NTRS2015

Evaluation of the Oxygen Concentrator Prototypes: Pressure Swing Adsorption Prototype and Electrochemical Prototype

An oxygen concentrator is needed to provide enriched oxygen in support of medical contingency operations for future exploration human spaceflight programs. It would provide continuous oxygen to an ill or injured crew member in a closed cabin environment. Oxygen concentration technology is being pursued to concentrate oxygen from the ambient environment so oxygen as a consumable resource can be reduced. Because oxygen is a critical resource in manned spaceflight, using an oxygen concentrator to pull oxygen out of the ambient environment instead of using compressed oxygen can provide better optimization of resources. The overall goal of this project is to develop an oxygen concentrator module that minimizes the hardware mass, volume, and power footprint while still performing at the required clinical capabilities. Should a medical event occur that requires patient oxygenation, the release of 100 percent oxygen into a small closed cabin environment can rapidly raise oxygen levels to the vehicles fire limit. The use of an oxygen concentrator to enrich oxygen from the ambient air and concentrate it to the point where it can be used for medical purposes means no oxygen is needed from the ultra-high purity (99.5+% O2) oxygen reserve tanks. By not adding oxygen from compressed tanks to the cabin environment, oxygen levels can be kept below the vehicle fire limit thereby extending the duration of care provided to an oxygenated patient without environmental control system intervention to keep the cabin oxygen levels below the fire limits. The oxygen concentrator will be a Food and Drug Administration (FDA) clearable device. A demonstration unit for the International Space Station (ISS) is planned to verify the technology and provide oxygen capability. For the ISS, the demonstration unit should not exceed 10 kg (approximately 22 lb), which is the soft stowage mass limit for launch on resupply vehicles for the ISS. The unit's size should allow for transport within the spacecraft to an ill crewmember. The user interface needs to be designed for ease of use by the local care provider and with consideration to the limited amount of training available to the astronaut corps for medical equipment and procedures.

Gilkey, Kelly M.↗
NASA NTRSDate not supplied

Lunar Surface Concept of Operations for the Global Exploration Roadmap Lunar Surface Exploration Scenario

The International Space Exploration Coordination Group (ISECG) is a voluntary, non-binding coordination forum of 26 space agencies. Building on the 2018 Global Exploration Roadmap (GER) and on growing global interest in space exploration, ISECG’s GER Supplement (Aug 2020) captures the latest developments in lunar exploration planning in an updated Lunar Surface Exploration Scenario. The updated Lunar Surface Exploration Scenario describes a phased approach to implementing infrastructure and exploration on the lunar surface to meet the goals and objectives defined by the ISECG. The updated scenario starts with Boots on the Moon, where space agencies focus on sending humans to the Moon beginning in 2024 and robotic exploration missions to support the 2024 goal and later phases. Next is Phase 2, Lunar Exploration—Expanding and Building, with an emphasis on completion of the proposed lunar surface objectives by exploring the lunar surface diversely and ultimately settling at the most beneficial site. Initial focus is on lunar surface exploration objectives pertaining to human landing and ascent, logistic cargo landers, and long-range traverses. The later focus is on lunar surface exploration objectives pertaining to long duration habitation, crew health and performance, and in-situ resource utilisation (ISRU). The third phase, Sustained Lunar Opportunities, envisages laying the foundation for a sustained and vibrant lunar presence in the coming decades through partnerships with international governments, academia and industry. During this phase, governments would shift their investment focus to further expand the frontier, including Mars exploration missions. This paper will describe the concept of operations by phase that supported the development of the updated scenario, along with refinements to those concepts of operations since the release of the updated scenario. By phase, the paper will describe the scenario, give an overview of the applicable architecture elements, identify functional requirements/needs of those elements, and include a discussion of nominal operations and contingency scenarios.

Markus Landgraf↗
NASA NTRS2006

Jet Propulsion Laboratory: Annual Report 2005

What an amazing host of new sights 2005 brought us. With impeccable choreography, one spacecraft sent an impactor slamming into a comet, reversing the traditional view of these space wayfarers by revealing it to be buried in deep drifts of a fine talcum-like powder. Another spacecraft delivered a probe from our European partners to the surface of Saturn's haze-shrouded moon Titan, disclosing a landscape eerily like Earth's, if we had methane rivers cascading down hillsides of ice. An orbiting observatory for the first time showed us the light from planets circling other stars, which astronomers previously knew to exist only from indirect clues. Throughout the year we also amassed continually expanding views of Earth as well as Mars, by far the most-explored planet after our own. In all, 18 spacecraft and five instruments were stationed across the solar system, studying our own world, other planets, comets and the deeper universe. These missions were enabled by the efforts of everyone at JPL. The Deep Space Network of communications complexes across three continents continued to experience a period of remarkable activity. Others were at work creating technologies both for NASA missions and other uses. JPL's contingent of scientific researchers was equally busy coordinating the science activities of our missions or pursuing independent investigations. None of this would be possible without the support of world-class business and administrative teams. All of our missions in one way or another support our nations Vision for Space Exploration, which envisages a gradually widening robotic and human presence across the solar system in the years ahead. The year was not without its challenges. NASA set forth to implement the Vision for Space Exploration, which resulted in some flight projects and technology efforts being terminated. To adjust to this new direction, it was necessary for us to reduce the JPL workforce by about five percent. Taking steps like this is painful, but we tried to make the process as orderly as possible. In the end, the adjustments made have left JPL on a healthy footing for the years ahead.

Jet Propulsion Laboratory (JPL)↗
NASA NTRSDate not supplied

Adaptive Optimization for System Performance and Combined Bernstein Polynomial, Optimal Reciprocal Collision Avoidance, Differential Dynamic Programming for Trajectory Replanning and Collision Avoidance for UAM Vehicles

The emerging urban air mobility (UAM) sector in aerospace is driving development of unconventional multi-modal vehicle configurations and autonomous flight. The combination of multi-modal vehicle dynamics, complex environment, requirements to deal with flight contingencies in an efficient and safe manner, as well as necessity for precise trajectory following and performance, are the driving influence behind adaptive optimization for system performance. We are interested in trajectory optimization algorithm that would system parameter estimation and identifying the optimal switching time between modes of hybrid dynamical systems. This presentation discusses a parameterized optimal control trajectory optimization algorithm that is an extended and generalized version of Differential Dynamic Programming (DDP), titled Parameterized Differential Dynamic Programming (PDDP). DDP is an efficient trajectory optimization algorithm relying on second order approximations of a system’s dynamics and cost function and has recently been applied to optimize systems with time invariant parameters. Experiments are presented applying PDDP to solve model predictive control (MPC) and moving horizon estimation (MHE) tasks simultaneously. In particular, PDDP is used to determine the optimal transition point between flight regimes of a complex urban air mobility (UAM) class vehicle exhibiting multiple phases of flight and to identify and compensate for actuation faults.

optimization↗
NASA NTRSDate not supplied

Connecting Community Scientific Hypotheses to Mars Sample Science

The proposed Mars Sample Return (MSR) program seeks to bring back to Earth a set of specially curated samples from the surface of Mars. The first segment of the MSR program is presently under-way; the Mars 2020 rover Perseverance has, to date, collected 24 samples of martian rock, regolith, and at-mosphere within Jezero crater. Its geologically diverse rock sample set can provide insight into the time evolu-tion of the Jezero environment, including the possible preservation of ancient biosignatures. Presently, a cache of 10 samples has been deposited in the Three Forks region on the floor of Jezero crater. These sam-ples, deemed scientifically return worthy by the Mars science community [1], are intended as a contingency sample set, as Perseverance has subsequently obtained more samples, and is therefore storing a more compre-hensive and scientifically valuable set of samples that will grow as the rover ascends to the rim of Jezero crater. The return of samples from Mars has been identi-fied as the highest priority planetary mission in the past three National Academies Decadal Surveys [2-4]. Sub-stantial work has gone into establishing science objec-tives for an MSR program, an overall framework for sample retrieval and analysis, and the expected impact on future Mars science and exploration (e.g., [5]). These objectives have been established at a relatively high level, e.g., “Interpret the primary geologic pro-cesses and history that formed the martian geologic rec-ord, with an emphasis on the role of water.” (Objective 1 in [5]) and are designed to encapsulate the broad sci-entific priorities of, and lay the groundwork for, inves-tigation strategies and measurements to be conducted by a future MSR program. It is from this overarching framework that the Connecting Community Scientific Hypotheses to Mars Sample Science workshop (‘The Workshop’) was conceived, with the purpose of distil-ling these high-level scientific objectives into discrete, testable hypotheses across a range of scientific disci-plines. The output of the workshop is meant to be a ‘hy-pothesis document’ serving as a starting point for cata-loging and condensing specific hypotheses for consid-eration by the scientific community. The template for this approach is based on the “driving hypotheses” de-veloped by the OSIRIS-REx team as part of their Sam-ple Analysis Plan [6].

Michael A Mischna↗
NASA NTRS2002

INTEX-NA: Intercontinental Chemical Transport Experiment - North America

INTEX-NA is an integrated atmospheric chemistry field experiment to be performed over North America using the NASA DC-8 and P-3B aircraft as its primary platforms. It seeks to understand the exchange of chemicals and aerosols between continents and the global troposphere. The constituents of interest are ozone and its precursors (hydrocarbons, NOX and HOX), aerosols, and the major greenhouse gases (CO2, CH4, N2O). INTEX-NA will provide the observational database needed to quantify inflow, outflow, and transformations of chemicals over North America. INTEX-NA is to be performed in two phases. Phase A will take place during the period of May-August 2004 and Phase B during March-June 2006. Phase A is in summer when photochemistry is most intense and climatic issues involving aerosols and carbon cycle are most pressing, and Phase B is in spring when Asian transport to North America is at its peak. INTEX-NA will coordinate its activities with concurrent measurement programs including satellites (e. g. Terra, Aura, Envisat), field activities undertaken by the North American Carbon Program (NACP), and other U.S. and international partners. However, it is being designed as a 'stand alone' mission such that its successful execution is not contingent on other programs. Synthesis of the ensemble of observation from surface, airborne, and space platforms, with the help of global/regional models is an important It is anticipated that approximately 175 flight hours for each of the aircraft (DC-8 and P-3B) will be required for each Phase. Principal operational sites are tentatively selected to be Bangor, ME; Wallops Island, VA; Seattle, WA; Rhinelander, WI; Lancaster, CA; and New Orleans, LA. These coastal and continental sites can support large missions and are suitable for INTEX-NA objectives. The experiment will be supported by forecasts from meteorological and chemical models, satellite observations, surface networks, and enhanced O3,-sonde releases. In addition to characterizing Atlantic-outflow and Pacific-inflow, INTEX-NA will characterize air masses transported between the U.S., Canada, and Mexico. INTEX-NA will be the first continental scale inflow, outflow, and transformation experiment to be performed over North America. It will provide the most comprehensive observational data set to date to understand the O3/NOX/HOX/aerosol photochemical system and the carbon cycle. One of the critical needs of the carbon cycle research is to obtain large-scale vertical and horizontal concentration gradients of CO2, throughout the troposphere over continental source/sink regions. INTEX-NA is ideally suited to perform this role. Coastal and continental operational sites will allow us to develop a curtain profile of greenhouse gases (e. g. CO2,) and other key pollutants across North America. Such information is central to our quantitative understanding of chemical budgets on the continental scale. We expect to provide a number of satellite under-flights over land and water to test and validate observations from the appropriate satellite platform (e. g. Aura). We plan to develop strong collaborations with other national and international observational programs. Results from INTEX-NA should directly benefit the development of environmental policy for air quality and climate change.

Singh, Hanwant B.↗
NASA NTRSDate not supplied

A Crew and Logistics Lander for the Common Habitat Architecture

The Common Habitat Architecture is a conceptual study that explores the use of a large habitat derived from the Space Launch System (SLS) liquid oxygen tank as a core habitation element that can be used for crew missions or training in 0g, 1/6g, 3/8g, and 1g. This is not part of current NASA Artemis mission planning, but instead represents an architecture that could potentially follow after the Artemis missions and initial human Mars landings. This architecture leverages Starship-derived vehicles for crew landing on and ascent from the Moon and Mars in support of long-duration surface missions with pressurized crew transfer baselined as a nominal capability. This is not the Human Landing System variant of Starship but represents a number of modifications to enable long-duration surface missions with an eight-person crew size. A rapid brainstorming study was conducted in February of 2021, using only public data, to identify options to deliver the 90-ton Common Habitat to the surface and emplace it at the intended habitation site. This study compared three lunar lander concepts and three Mars lander concepts before ultimately selecting the SpaceX Starship as the most viable lander. The Common Habitat Architecture assumes that this Starship variant can further be modified for crew and cargo delivery, assuming that the use of a common system will lead to cost benefits. It is assumed that the Starship will expend too much propellant in landing crew and logistics to be able to launch, given the assumed absence of surface propellant production. Consequently, a separable ascent stage is used for crew ascent. The pressurized elements of this modified Starship are discussed: Starship Ascent Module, Airlock, Transfer Tunnel, Pressurized Crew Transfer Module, and Logistics Modules. For each element, a description, dimensions, rough mass estimates, core capabilities, and design features are presented. Key mechanisms and internal structures of the starship are also discussed. This will include flame diverters for the ascend module, the orbital docking hatch, Pressurized Crew Transfer Module Garage door, Starship Ascent Module fairing, flame diverter blow-out panels, Pressurized Crew Transfer Module lift system, logistics module lift system, contingency crew ascent via the logistics module lift system, internal catwalks, and structural interfaces. Concept of Operations will be discussed for both Moon and Mars. This will include in-space crew rendezvous and transfer, crewed landing, shirtsleeve crew transfer, crew departure, contingency surface operations, and final element disposition. Key differences for Mars will be discussed, such as the Martian atmosphere, dust storms, the absence of crew handover, and Deep Space Exploration Vehicle rendezvous. Contingency microgravity maintenance access will also be discussed. This work will demonstrate viable pressurized crew transfer with a Starship-based lander architecture. Forward work includes Garage lighting and camera systems, Starship Ascent Module propulsion system and docking port trades, contingency habitation trades, and mass/power equipment estimation. Finally, forward work includes developing a heavy cargo return system derived from the Crew and Logistics Starship. The goal of this system is the return of at least 10,000 kg payloads from the surfaces of the Moon and Mars.

Lunar Lander↗
NASA NTRS2008

Jet Propulsion Laboratory: Annual Report 2007

Many milestones are celebrated in the business of space exploration, but one of them that arrived this year has particular meaning for us. Half a century ago, on January 31, 1958, the Jet Propulsion Laboratory was responsible for creating America's first satellite, Explorer 1, and joined with the Army to launch it into orbit. That makes 2007 the 50th year we have been sending robotic craft from Earth to explore space. No other event before or since has had such a profound effect on JPL's basic identity, setting it on the path to become the world's leader in robotic solar system exploration. It is not lost on historians that Explorer 1, besides being America's first satellite, was also the first spacecraft from any country to deliver scientific results in its case, the discovery of the Van Allen Radiation Belts that surround Earth. Science, of course, has been the prime motivator for all the dozens of missions that we have lofted into space in the half-century since then. JPL has sent spacecraft to every planet in the solar system from Mercury to Neptune, some of them very sophisticated machines. But in one way or another, they all owe their heritage to the 31-pound bullet-shaped probe JPL shot into space in 1958. Although we have ranged far and wide across the solar system, we have a very strong contingent of satellites and instruments dedicated, like Explorer, to the environment of our home planet. JPL missions have been providing much of the data to establish the facts of global warming - most especially, the melting of ice sheets in Greenland and Antarctica. During the past year, JPL and our parent organization, the California Institute of Technology, have created a task force to focus the special capabilities of the Laboratory and campus on ways to better understand the physics of global change. While Earth is a chaotic and dynamic system capable of large natural variations, evidence is mounting that human activities are playing an increasingly important role. A central piece of this effort is a search for novel energy sources to replace fossil fuels, the combustion of which adds carbon dioxide to our atmosphere. All of this is supported by our worldwide Deep Space Network, which provides the communication link between spacecraft and the ground. In addition, missions are infused with technologies developed by researchers working on projects for non-NASA sponsors as well as on pure research.

Jet Propulsion Laboratory (JPL)↗
NASA NTRSDate not supplied

Using Open Standards and NASA Open Source Simulation Tools to Model Artemis Base Camp Mission Timelines

The United States’ National Aeronautics and Space Administration (NASA) has announced that the Artemis Program will return humans to the Moon, establishing a persistent presence with the Artemis Base Camp (ABC), and extend human exploration to Mars. The NASA Exploration Systems Simulations (NExSyS) team at NASA’s Johnson Space Center is using internationally developed simulation interoperability standards and NASA open source simulation tools to support Artemis concept, analysis, designs, development, training, and ultimately operations. The NExSyS team has been tasked to support early ABC architecture and mission analysis using mission time lines developed by the crew operations mission planning team. The NExSyS team is developing a distributed simulation framework with initial Artemis element implementations to model the ABC mission timelines using the international simulation interoperability standard High Level Architecture (HLA), the Simulation Interoperability Standards Organization’s Space Reference Federation Object Model (SpaceFOM), the NASA open source Trick Simulation Environment, and another NASA open source interface package called TrickHLA. The ABC architecture is composed of a number of key surface elements and resources. Some examples of modeled elements (also known as entities) are landers, habitats, rovers, logistics carriers, and astronauts. Some examples of modeled transferable and consumable resources are power, water, oxygen, nitrogen, scientific samples, and food. These entities and resources are modeled in a collection of individual simulations called Federates. A coordinated collection of interoperable federates is called a Federation and when these federates are tied together in a coordinated simulation run, it is referred to as a Federation Execution. The federates communicate through HLA using data exchange formats defined by a collection of machine readable files called Federation Object Models (FOMs). These FOM files are based on extensions to the SpaceFOM. This enables the instantiation and sharing of objects and interactions between federates in the federation. These provide for entity and resource tracking, object transfer, and data collection. Federate interactions are used to trigger events and notify federates of entity or resource transfers. For the initial implementation, the constituent federates are Trick-based simulations that use TrickHLA to provide the required HLA-base interoperability. These Trick-based simulations provide the required modeling for the individual Artemis elements along with the associated element resources. These federates provide a means to explore traverses between surface elements and exploration sites as scheduled in a mission timeline and explore the affects traverse times have on the overall mission timeline. The mission time lines are modeled using a Trick input file event handling capabilities. Each timeline operation is handled as individual simulation events, and triggered based on previous event status, time of operation, and simulated task completions. In addition, the ABC Federation can be used to perform Monte Carlo analysis. The Monte Carlo tool can vary the inputs, timings, and malfunctions to show how various contingencies in the mission can affect the mission timeline.

Keaton Craig Dodd↗