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All sourcesNASA NTRSDOE OSTI
NASA NTRS2003

Experimental Verification of Electric Drive Technologies Based on Artificial Intelligence Tools

A laboratory implementation of a fuzzy logic-tracking controller using a low cost Motorola MC68HC11E9 microprocessor is described in this report. The objective is to design the most optimal yet practical controller that can be implemented and marketed, and which gives respectable performance, even when the system loads, inertia and parameters are varying. A distinguishing feature of this work is the by-product goal of developing a marketable, simple, functional and low cost controller. Additionally, real-time nonlinearities are not ignored, and a mathematical model is not required. A number of components have been designed, built and tested individually, and in various combinations of hardware and software segments. These components have been integrated with a brushless motor to constitute the drive system. A microprocessor-based FLC is incorporated to provide robust speed and position control. Design objectives that are difficult to express mathematically can be easily incorporated in a fuzzy logic-based controller by linguistic information (in the form of fuzzy IF-THEN rules). The theory and design are tested in the laboratory using a hardware setup. Several test cases have been conducted to confirm the effectiveness of the proposed controller. The results indicate excellent tracking performance for both speed and position trajectories. For the purpose of comparison, a bang-bang controller has been tested. The fuzzy logic controller performs significantly better than the traditional bang-bang controller. The bang-bang controller has been shown to be relatively inaccurate and lacking in robustness. Description of the implementation hardware system is also given.

Rubaai, Ahmed↗
NASA NTRS2011

A Super-Resolution Algorithm for Enhancement of FLASH LIDAR data

A novel method for enhancement of the spatial resolution of 3-dimensional Flash Lidar images is being proposed for generation of elevation maps of terrain from a moving platform. NASA recognizes the Flash LIDAR technology as an important tool for enabling safe and precision landing in future unmanned and crewed lunar and planetary missions. The ability of the Flash LIDAR to generate 3-dimensional maps of the landing site area during the final stages of the descent phase for detection of hazardous terrain features such as craters, rocks, and steep slopes is under study in the frame of the Autonomous Landing and Hazard Avoidance (ALHAT) project. Since single frames of existing FLASH LIDAR systems are not sufficient to build a map of entire landing site with acceptable spatial resolution and precision, a super-resolution approach utilizing multiple frames has been developed to overcome the instrument s limitations. Performance of the super-resolution algorithm has been analyzed through a series of simulation runs obtained from a high fidelity Flash LIDAR model and a high resolution synthetic lunar elevation map. For each simulation run, a sequence of FLASH LIDAR frames are recorded and processed as the spacecraft descends toward the landing site. Simulations runs having different trajectory profiles and varying LIDAR look angles of the terrain are also analyzed. The results show that adequate levels of accuracy and precision are achieved for detecting hazardous terrain features and identifying safe areas of the landing site.

Bulyshev, Alexander↗
NASA NTRS2019

SmallSat missions enabled by paired low-thrust hybrid rocket and low-power long-life Hall thruster

The capabilities of a SmallSat-class spacecraft targeting the outer solar system and using a combined chemical and electric propulsion system are explored. The development of compact hybrid rockets has enabled high-thrust engines to be packaged tightly enough to fit on cubesat and SmallSat spacecraft. These hybrid rockets provide 10’s-100 N of thrust depending on the propellant load & >300 s of specific impulse and have been demonstrated in both ambient and vacuum environments. Advancements in low-power long-life Hall thruster technologies have provided the potential for significantly greater propellant throughputs, enabling their use as a primary propulsion element on interplanetary spacecraft. In a recent characterization test campaign, the MaSMi-DM Hall thruster demonstrated power throttling from 150 – 1000 W with >1500 s of specific impulse available at >500 W and >40% total thrust efficiency available at >300 W; peak values of 1940 s and 53% were observed. A notional low-mass spacecraft employing a combined hybrid rocket and low-power electric propulsion system was designed and used for mission concept analysis targeting the outer solar system. Using an imposed wet mass limit of 400 kg, mission trajectories to Saturn and Uranus were generated. Orbit capture with >40% of the launch mass was shown to be possible at either target, with mission transfer times of 7.5 years and 13.5 years for Saturn and Uranus, respectively. Significant follow-on mission activities near Saturn (e.g. to Titan & Enceladus) were also possible by carrying extra propellant mass while remaining under the total wet mass limit.

Karp, Ashley C.↗
NASA NTRS2019

Airspace Technology Demonstration 3 (ATD-3): Dynamic Routes for Arrivals in Weather (DRAW) Technology Transfer Document Summary Version 2.0

Airspace Technology Demonstration – 3 (ATD-3) is part of NASA’s Airspace Operations and Safety Program (AOSP) – specifically, its Airspace Technology Demonstrations (ATD) Project. ATD-3 is a multi-year research and development effort which proposes to develop and demonstrate automation technologies and operating concepts that enable air navigation service providers and airspace users to continuously assess weather, winds, traffic, and other information to identify, evaluate, and implement workable opportunities for flight plan route corrections that can result in significant flight time and fuel savings in en route airspace. In order to ensure that the products of this tech-transfer are relevant and useful, NASA has created strong partnerships with the FAA and key industry stakeholders. This summary document and accompanying technology artifacts satisfy the third Research Transition Product (RTP) defined in the Applied Traffic Flow Management (ATFM) Research Transition Team (RTT) Plan, which is Dynamic Routes for Arrivals in Weather (DRAW). This technology transfer consists of artifacts for DRAW Arrival Metering (AM) Operations delivered in June 2018, DRAW AM updates, and DRAW Extended Metering (XM) Operations. Blue highlighting indicates the new or modified deliverables. Some of the artifacts in this technology transfer have distribution restrictions that need to be followed. Distribution information is noted in each section. DRAW is a trajectory-based system that combines the legacy Dynamic Weather Routes (DWR) weather avoidance technology with an arrival-specific rerouting algorithm and arrival scheduler to improve traffic flows on weather-impacted arrival routes into major airports. First, DRAW identifies flights that could be rerouted to more efficient Standard Terminal Arrival Routes (STARs) that may have previously been impacted by weather. Second, when weather is impacting the arrival routing, DRAW proposes simple arrival route corrections that enable aircraft to stay on their flight plan while avoiding weather. The DRAW system proposes reroutes early enough to allow Time Based Flow Management (TBFM) to make the necessary schedule adjustments. As a result, metering operations can be sustained longer and more consistently in the presence of weather because the arrival schedule accounts for the dynamic routing intent of arrival flights to deviate around weather. The first DRAW tech transfer in June 2018 focused on arrival metering operations with the DRAW algorithm implemented in the NASA Center TRACON Automation System (CTAS) automation software. This tech transfer delivery includes updates for DRAW implemented in FAA’s TBFM 4.7 automation software and preliminary research into DRAW for XM operations.

Isaacson, Douglas R.↗
NASA NTRS2015

Space Launch System Co-Manifested Payload Options for Habitation

The Space Launch System (SLS) has a co-manifested payload capability that will grow over time as the rocket matures and planned upgrades are implemented. The final configuration is planned to be capable of inserting a payload greater than 10 metric tons (mt) into a trans-lunar injection trajectory along with the crew in the Orion capsule and the service module. The co-manifested payload is located below the Orion and its service module in a 10-meter high fairing similar to the way the Saturn launch vehicle carried the lunar lander below the Apollo command and service modules. A variety of approaches have been explored that utilizes this co-manifested payload capability to build up infrastructure in deep space in support of future asteroid, lunar, and Mars mission scenarios. This paper is a report on the findings from the Advanced Concepts Office study team at the NASA Marshall Space Flight Center, working with the Advanced Exploration Systems Program on the Exploration Augmentation Module Project. It includes some of the possible options for habitation in the co-manifested payload volume on SLS. Findings include module designs that can be developed in 10mt increments to support these missions, including overall conceptual layouts, mass properties, and approaches for integration into various scenarios for near-term support of deep space habitat research and technology development, support to asteroid exploration, and long range support for Mars transfer flights.

Smitherman, David↗
NASA NTRS2024

Theoretical Prediction of Thermal Expansion Anisotropy for Y 2 Si 2 O 7 Environmental Barrier Coatings Using a Deep Neural Network Potential and Comparison to Experiment

Environmental barrier coatings (EBCs) are an enabling technology for silicon carbide (SiC)-based ceramic matrix composites (CMCs) in extreme environments such as gas turbine engines. However, development of new coating systems is hindered by the large design space and difficulty in predicting properties for these materials. Density Functional Theory (DFT) has successfully been used to model and predict some thermodynamic and thermo-mechanical properties of high-temperature ceramics for EBCs, although these calculations are challenging due to their high computational costs. In this work, we use machine learning to train a deep neural network potential (DNP) for Y 2 Si 2 O 7 , which is then applied to calculate thermodynamic and thermo-mechanical properties at near-DFT accuracy much faster and using less computational resources than DFT. We use this DNP to predict phonon-based thermodynamic properties of Y 2 Si 2 O 7 with good agreement to DFT and experiments. We also utilize the DNP to calculate the anisotropic, lattice direction-dependent coefficients of thermal expansion (CTEs) for Y 2 Si 2 O 7 . Molecular dynamics trajectories using the DNP correctly demonstrate accurate prediction of the anisotropy of the CTE in good agreement with diffraction experiments. In the future, this DNP could be applied to accelerate additional property calculations for Y 2 Si 2 O 7 compared to DFT or experiments.

rare earth silicates↗
NASA NTRS2017

Current Safety Nets Within the U.S. National Airspace System

There are over 70,000 flights managed per day in the National Airspace System, with approximately 7,000 aircraft in the air over the United States at any given time. Operators of each of these flights would prefer to fly a user-defined “4D” trajectory (4DT), which includes arrival and departure times; preferred gates and runways at the airport; efficient, wind-optimal routes for departure, cruise and arrival phase of flight; and fuel efficient altitude profiles. To demonstrate the magnitude of this achievement a single flight from Los Angeles to Baltimore, accesses over 35 shared or “constrained” resources that are managed by roughly 30 air traffic controllers (at towers, approach control and en route sectors); along with traffic managers at 12 facilities, using over 22 different, independent automation system (including TBFM, ERAM, STARS, ASDE-X, FSM, TSD, GPWS, TCAS, etc.). In addition, dispatchers, ramp controllers and others utilize even more systems to manage each flight’s access to operator-managed resources. Flying an ideal 4DT requires successful coordination of all flight constraints among all flights, facilities, operators, pilots and controllers. Additionally, when conditions in the NAS change, the trajectories of one or more aircraft may need to be revised to avoid loss of flight efficiency, predictability, separation or system throughput. The Aviation Safety Network has released the 2016 airliner accident statistics showing a very low total of 19 fatal airliner accidents, resulting in 325 fatalities1. Despite several high profile accidents, the year 2016 turned out to be a very safe year for commercial aviation, Aviation Safety Network data show. Over the year 2016 the Aviation Safety Network recorded a total of 19 fatal airliner accidents [1], resulting in 325 fatalities. This makes 2016 the second safest year ever, both by number of fatal accidents as well as in terms of fatalities. In 2015 ASN recorded 16 accidents while in 2013 a total of 265 lives were lost. How can we keep it that way and not upset the apple cart by premature insertion of innovative technologies, functions, and procedures? In aviation, safety nets function as the last system defense against incidents and accidents. Current ground-based and airborne safety nets are well established and development to make them more efficient and reliable continues. Additionally, future air traffic control safety nets may emerge from new operational concepts.

Smith, Brian E.↗
NASA NTRS2014

Asteroid Crewed Segment Mission Lean Development

The next generation of human spaceflight missions presents numerous challenges to designers that must be addressed to produce a feasible concept. The specific challenges of designing an exploration mission utilizing the Space Launch System and the Orion spacecraft to carry astronauts beyond earth orbit to explore an asteroid stored in a distant retrograde orbit around the moon will be addressed. Mission designers must carefully balance competing constraints including cost, schedule, risk, and numerous spacecraft performance metrics including launch mass, nominal landed mass, abort landed mass, mission duration, consumable limits and many others. The Asteroid Redirect Crewed Mission will be described along with results from the concurrent mission design trades that led to its formulation. While the trades presented are specific to this mission, the integrated process is applicable to any potential future mission. The following trades were critical in the mission formulation and will be described in detail: 1) crew size, 2) mission duration, 3) trajectory design, 4) docking vs grapple, 5) extravehicular activity tasks, 6) launch mass and integrated vehicle performance, 7) contingency performance, 8) crew consumables including food, clothing, oxygen, nitrogen and water, and 9) mission risk. The additional Orion functionality required to perform the Asteroid Redirect Crewed Mission and how it is incorporated while minimizing cost, schedule and mass impacts will be identified. Existing investments in the NASA technology portfolio were leveraged to provide the added functionality that will be beneficial to future exploration missions. Mission kits are utilized to augment Orion with the necessary functionality without introducing costly new requirements to the mature Orion spacecraft design effort. The Asteroid Redirect Crewed Mission provides an exciting early mission for the Orion and SLS while providing a stepping stone to even more ambitious missions in the future.

Gard, Joe↗
NASA NTRS2000

Navigation Strategy for the Mars 2001 Lander Mission

The Mars Surveyor Program (MSP) is an ongoing series of missions designed to robotically study, map and search for signs of life on the planet Mars. The MSP 2001 project will advance the effort by sending an orbiter, a lander and a rover to the red planet in the 2001 opportunity. Each vehicle will carry a science payload that will Investigate the Martian environment on both a global and on a local scale. Although this mission will not directly search for signs of life, or cache samples to be returned to Earth, it will demonstrate certain enabling technologies that will be utilized by the future Mars Sample Return missions. One technology that is needed for the Sample Return mission is the capability to place a vehicle on the surface within several kilometers of the targeted landing site. The MSP'01 Lander will take the first major step towards this type of precision landing at Mars. Significant reduction of the landed footprint will be achieved through two technology advances. The first, and most dramatic, is hypersonic aeromaneuvering; the second is improved approach navigation. As a result, the guided entry will produce in a footprint that is only tens of kilometers, which is an order of magnitude improvement over the Pathfinder and Mars Polar Lander ballistic entries. This reduction will significantly enhance scientific return by enabling the potential selection of otherwise unreachable landing sites with unique geologic interest and public appeal. A landed footprint reduction from hundreds to tens of kilometers is also a milestone on the path towards human exploration of Mars, where the desire is to place multiple vehicles within several hundred meters of the planned landing site. Hypersonic aeromaneuvering is an extension of the atmospheric flight goals of the previous landed missions, Pathfinder and Mars Polar Lander (MPL), that utilizes aerodynamic lift and an autonomous guidance algorithm while in the upper atmosphere. The onboard guidance algorithm will control the direction of the lift vector, via bank angle modulation, to keep the vehicle on the desired trajectory. While numerous autonomous guidance algorithms have been developed for use during hypersonic flight at Earth, this will be the first flight of an autonomously directed lifting entry vehicle at Mars. However, without sufficient control and knowledge of the atmospheric entry conditions, the guidance algorithm will not perform effectively. The goal of the interplanetary navigation strategy is to deliver the spacecraft to the desired entry condition with sufficient accuracy and knowledge to enable satisfactory guidance algorithm performance. Specifically, the entry flight path angle must not exceed 0.27 deg. to a 3 sigma confidence level. Entry errors will contribute directly to the size of the landed footprint and the most significant component is entry flight path angle. The size of the entry corridor is limited on the shallow side by integrated heating constraints, and on the steep side by deceleration (g-load) and terminal descent propellant. In order to meet this tight constraint it is necessary to place a targeting maneuver seven hours prior to the time of entry. At this time the trajectory knowledge will be quite accurate, and the effects of maneuver execution errors will be small. The drawback is that entry accuracy is dependent on the success of this final late maneuver. Because propulsive maneuvers are critical events, it is desirable to minimize their occurrence and provide the flight team with as much response time as possible in the event of a spacecraft fault. A mission critical maneuver at Entry - 7 hours does not provide much fault tolerance, and it is desirable to provide a strategy that minimizes reliance on this maneuver. This paper will focus on the Improvements in interplanetary navigation that will decrease entry errors and will reduce the landed footprint, even in the absence of aeromaneuvering. The easiest to take advantage of are Improvements In the knowledge of the Mars ephemeris and gravity field due to the MGS and MSP'98 missions. Improvements In data collection and reduction techniques such as "precislon ranging' and near-simultaneous tracking will also be utilized. In addition to precise trajectory control, a robust strategy for communications and flight operations must also be demonstrated. The result Is a navigation and communications strategy on approach that utilizes optimal maneuver placement to take advantage of trajectory knowledge, minimizes risk for the flight operations team, is responsive to spacecraft hardware limitations, and achieves the entry corridor. The MSP2001 mission Is managed at JPL under the auspices of the Mars Exploration Directorate. The spacecraft flight elements are built and managed by Lockheed-Martin Astronautics in Denver, Colorado.

Mase, Robert A.↗
NASA NTRS2005

Conceptual Design and Dynamics Testing and Modeling of a Mars Tumbleweed Rover

The NASA Langley Research Center has been developing a novel concept for a Mars planetary rover called the Mars Tumbleweed. This concept utilizes the wind to propel the rover along the Mars surface, bringing it the potential to cover vast distances not possible with current Mars rover technology. This vehicle, in its deployed configuration, must be large and lightweight to provide the ratio of drag force to rolling resistance necessary to initiate motion from rest on the Mars surface. One Tumbleweed design concept that satisfies these considerations is called the Eggbeater-Dandelion. This paper describes the basic design considerations and a proposed dynamics model of the concept for use in simulation studies. It includes a summary of rolling/bouncing dynamics tests that used videogrammetry to better understand, characterize, and validate the dynamics model assumptions, especially the effective rolling resistance in bouncing/rolling dynamic conditions. The dynamics test used cameras to capture the motion of 32 targets affixed to a test article s outer structure. Proper placement of the cameras and alignment of their respective fields of view provided adequate image resolution of multiple targets along the trajectory as the test article proceeded down the ramp. Image processing of the frames from multiple cameras was used to determine the target positions. Position data from a set of these test runs was compared with results of a three dimensional, flexible dynamics model. Model input parameters were adjusted to match the test data for runs conducted. This process presented herein provided the means to characterize the dynamics and validate the simulation of the Eggbeater-Dandelion concept. The simulation model was used to demonstrate full scale Tumbleweed motion from a stationary condition on a flat-sloped terrain using representative Mars environment parameters.

Calhoun Philip C.↗
NASA NTRS2017

NASA's Near Earth Asteroid Scout Mission

NASA is developing solar sail propulsion for a near-term Near Earth Asteroid (NEA) reconnaissance mission and laying the groundwork for their future use in deep space science and exploration missions. The NEA Scout mission, funded by NASA's Advanced Exploration Systems Program and managed by NASA MSFC, will use the sail as primary propulsion allowing it to survey and image one or more NEA's of interest for possible future human exploration. NEA Scout uses a 6U cubesat (to be provided by NASA's Jet Propulsion Laboratory), an 86 m2 solar sail and will weigh less than 14 kilograms. The solar sail for NEA Scout will be based on the technology developed and flown by the NASA NanoSail-D and The Planetary Society's Lightsail-A. Four ~7 m stainless steel booms wrapped on two spools (two overlapping booms per spool) will be motor deployed and pull the sail from its stowed volume. The sail material is an aluminized polyimide approximately 3 microns thick. NEA Scout will launch on the Space Launch System (SLS) first mission in 2018 and deploy from the SLS after the Orion spacecraft is separated from the SLS upper stage. The NEA Scout spacecraft will stabilize its orientation after ejection using an onboard cold-gas thruster system. The same system provides the vehicle Delta-V sufficient for a lunar flyby. After its first encounter with the moon, the 86 m2 sail will deploy, and the sail characterization phase will begin. A mechanical Active Mass Translation (AMT) system, combined with the remaining ACS propellant, will be used for sail momentum management. Once the system is checked out, the spacecraft will perform a series of lunar flybys until it achieves optimum departure trajectory to the target asteroid. The spacecraft will then begin its two year-long cruise. About one month before the asteroid flyby, NEA Scout will pause to search for the target and start its approach phase using a combination of radio tracking and optical navigation. The solar sail will provide continuous low thrust to enable a relatively slow flyby of the target asteroid under lighting conditions favorable to geological imaging. Once complete, NASA will have demonstrated the capability to fly low-cost, high Delta-V cubesats to perform interplanetary missions.

Johnson, Les↗
NASA NTRS2020

Encealdus Orbilander: A Flagship Mission Concept for Astrobiology

"Whether life exists beyond Earth remains a fundamental question driving our exploration of the Solar System. At Saturn’s moon Enceladus, plumes of oceanic material vented into space allow the investigation of the astrobiological potential of an ocean world, hinted at by Cassini, without the necessity of drilling through kilometers of ice crust. The Enceladus Orbilander is a flagship ($2.56B in fiscal year 2025 dollars) mission concept created for the 2023–2033 Planetary Science Decadal Survey. Orbilander takes full advantage of the opportunity provided by Enceladus’ plumes to search for signs of life. A single spacecraft both orbits and lands, capturing samples from four distinct reservoirs offered by the plumes. These samples, both particulate and vapor, are then analyzed by the Life Detection Suite (LDS), a set of five instruments conducting complementary and orthogonal biosignature-seeking measurements. To provide the context that specifically enhances interpretation of LDS measurements, geochemical and geophysical investigations are conducted both in orbit and on the surface. These reveal the physio-chemical state of the ocean and core as well as the processes involved in ejection of plume material and how these affect the ocean material analyzed by the LDS. The Orbilander can be delivered to the Saturn system via several launch vehicle and trajectory options, including a direct trajectory (7-year cruise), a ∆V-EGA trajectory (9-year cruise) and several options using an inner cruise with Venus and Earth flybys (10-year cruise). Upon Saturn Orbit Insertion, a 4-year moon tour pumps down the Orbilander’s orbit to intercept Enceladus. The most optimal arrival times balance the Jupiter flyby opportunities of the late 2030s and solar illumination at the Enceladus high southern latitudes where plume material is most abundant. This mission concept therefore targets project start in 2030. Upon Enceladus Orbit Insertion, the Orbilander begins a 1.5-year-long campaign of landing site reconnaissance, remote sensing science, and collecting sufficient plume sample to run all but one of the LDS measurements. After successful landing, the Orbilander spends 2 years on the surface conducting multiple LDS measurements with all five instruments on actively and passively collected plume material, as well as seismic investigations. The schedule laid out here is well-defined, but the mission also has operational and resource flexibility should additional reconnaissance be needed. As part of the design study, mission and development risks were identified and mitigation strategies proposed. Technologies key to achieving the life detection science objectives include instrumentation matured under programs like COLDTech and ICEE-2, such as aspects of the sampling system and microfluidic devices, as well as well-known techniques like high-resolution and separation-capable mass spectrometers. Autonomous onboard navigation is planned to maintain a halo orbit around Enceladus to enable passive sampling from orbit as well as reconnaissance measurements for use in site selection and landing. Terrain relative navigation is included to ensure safe landing, given that targeted areas may contain landing hazards. Continued development of radioisotope thermoelectric generator (RTG) technology and long-life batteries is essential for this long duration mission. The Enceladus Orbilander represents an optimal point in the trade space of science value versus cost, taking advantage of the extensive knowledge of Enceladus provided by Cassini, how well Enceladus lends itself to a search for life in material from its ocean, and the flexibility afforded by the innovative design developed by the APL team. By taking full advantage of Enceladus’ plumes both in orbit and on the surface, Orbilander represents a robust search for life with complementary and orthogonal biosignatures as well as contextual geophysical and geochemical measurements, determining not only whether Enceladus is inhabited (at levels up to 500,000× scarcer than in Earth’s oceans) but also why. "

Exobiology↗
NASA NTRSDate not supplied

Enabling Mission Flexibility to Battery Driven Deep Space Endeavors With Generalized Battery-Health-Monitoring Using Physics-Based and Data-Driven Reduced-Order Models

The needs and requirements for an electrochemical energy storage for deep space exploration is well explored. It is often understood that different mission sites and environmental conditions require different battery chemistries or technologies. Additionally, various engineering solutions are deployed to overcome specific chemical challenges. One often overlooked need is the “health” monitoring of an electrochemical storage system. The term generalized health monitoring, as envisioned in this work, refers to the monitoring of various aspects such as electrode health, electrolyte health, reaction pathway health, cooling system health, sensor health, and BMS health [1]. Generalized health monitoring allows mission leads, engineers, and scientists to incorporate flexibility in mission designs, make on-the-fly mission changes, and extend the duration of science missions. Moreover, it enables automation and data-driven decision-making without compromising safety and performance. Recently, our group developed a hierarchy of thermal reduced-order models (TROM) by combining a physics-based modeling approach and data-driven model reduction techniques applied to flight data [2]. The resulting TROMs were found to be not only accurate but also identifiable from the flight data. Consequently, the coefficient of variance of the model parameters is small over the course of hundreds of flights, allowing for monitoring the parameter evolution trajectories as the battery ages and degrades. These parameters constitute the metrics of the generalized health of a battery. Monitoring their evolution allows such models to be used for anomaly detection and prognostics, improving early detection of abnormal behavior and thus enabling timely maintenance, longer battery life, and enhanced battery safety. For this presentation, the practicality of the thermal model will be validated on a pack of 14cells under various topology configurations such as 1S14P, 2P7S, 7S2P, and 1P14S. It is well known that manufacturing and non-uniform aging lead to variability in the performance of a cell, which is exacerbated by cell balancing during active load. Additionally, in extreme scenarios, the paramount objective is to complete the mission, regardless of the stresses on the battery. Topology-induced balancing issues further stress the battery. The goal of this study is to determine if the noise (identifiability) in the reduced-order thermal model parameters is sensitive to topology, cell spacing, cooling strategy, and manufacturing or age variability. The variability in cells is considered by assuming a multimodal distribution for microscopic parameters of a cell (such as porosity, tortuosity, reaction kinetics, volumetric thermal conductivity, and volumetric heat capacity). The compounded effect of manufacturing variability, topological selection, cooling strategies, and cell balancing ages each cell in a battery differently. The study aims to clarify whether the challenge in extracting maximum information depends on the minimum number of sensors or models used for data extraction.

Automation↗
NASA NTRS2021

ATD-2 Phase 3 Benefits Mechanism

NASA has been developing and demonstrating a suite of decision support capabilities for integrated arrival, departure, and surface (IADS) operations in a metroplex environment. The effort is being made in three phases, under NASA’s Airspace Technology Demonstration 2 (ATD-2) sub-project, through a close partnership with the Federal Aviation Administration (FAA), air carriers, airport, and general aviation community. ATD-2 Phase 1 & 2 have successfully demonstrated new technologies developed to manage the IADS capabilities at a single airport. The Phase 3 builds upon Phases 1 and 2 by extending the capabilities to a Metroplex environment where multiple airports are interacting and competing for resources at the terminal boundary. This document describes the metrics used to inform the flight operators, Air Traffic Control (ATC) tower, FAA, and even broader aviation community about opportunities to reroute aircraft and the metrics used to assess the performance of the ATD-2 Phase 3 system including benefits. This document provides definitions and detailed calculation methods of identified Phase 3 benefit metrics such as OFF Delay Savings, IN Delay Savings, and Aggregate System-Wide Savings. This document also describes the mechanism to translate delay savings metrics into fuel and emissions savings.

ATD-2 Phase 3↗
NASA NTRS2003

Ultrasail

Ultrasail is a complete sail system for the launch, deployment, stabilization and control of very large solar sails enabling reduced mission times for interplanetary and deep space spacecraft. Ultrasail is an innovative, non-traditional approach to propulsion technology achieved by combining propulsion and control systems developed for formation-flying microsatellites with an innovative solar sail architecture to achieve sq km-class controllable sail areas, sail subsystem area densities of 1 gm per sq m, and thrust levels equivalent to 400 kW ion thruster systems used for comparable deep space missions. Ultrasail can conceivably even achieve outer planetary rendezvous, a deep space capability now reserved for high-mass nuclear and chemical systems. Ultrasail is a Delta IV-launched multi-blade spin-stabilized system with blade lengths as long as 50 km, reminiscent of the MacNeal Heliogyro. The primary innovation is the near-elimination of sail supporting structures by attaching the sail tip to a rigid formation-flying microsatellite truss which deploys the sail blade, and which then articulates the blade to provide attitude control, including spin stabilization and precession of the spin axis. These tip microsatellites are controlled by a solar-powered 3-axis microthruster system (electric or cold gas) to maintain proper sail film tension during deployment and spin-up. The satellite mass also provides a stabilizing centrifugal force on the blade while in rotation. Understanding the dynamics of individual blades is key to the overall dynamics of Ultrasail. Forces and torques that must be modeled include those due to solar pressure, those generated by the microsatellite at the blade tip and by torques applied at the blade root. Centrifugal forces also play a significant role in the deployment and maintenance of the sail configuration. To capture the dynamics of the overall system, the equations of motion for the blades have been derived. Using these differential equations, a control law will be derived to maneuver Ultrasail. This law involves the pitching of the individual blades thereby moving the distribution of the radiation pressure on each individual blade and inducing a resultant torque on the system. The direction of the angular momentum vector and its rate of precession can be controlled through the pitch angle of the blades. The Ultrasail trajectory is also being studied. Optimal or near-optimal trajectories are being generated to showcase Ultrasail performance. Various missions, e.g. outer planet and solar polar missions for observation of the Sun, are currently being investigated to demonstrate the performance enhancements generated by Ultrasail technology. Calculus-of-variations-based optimization software is used to produce optimal Ultrasail trajectories. The performance of these trajectories is being compared to optimal results generated with other propulsion models, including chemical propulsion, ion propulsion, and competing solar sail concepts. Results of these studies will quantify the performance of Ultrasail compared to existing solar sail concepts for high energy missions.

Burton, R.↗
NASA NTRSDate not supplied

Design and Technology Maturation of the Stratospheric Projectile Experiment of Entry Dynamics

The supersonic and transonic dynamic stability of blunt-body reentry vehicles currently poses large risks in all of NASA’s ongoing entry missions (MSR SRL, MSR EES, and Dragonfly). These projects have allocated millions of dollars to testing and modeling efforts to buy down risk by using the current state-of-the-art (SoA) facilities at NASA’s disposal. While these facilities have heritage in supplying dynamics data to reentry missions, their availability is severely limited – particularly with the high number of concur-rent projects requesting simultaneous testing– and are costly when considering the science density per dollar. None of the current SoA facility methodologies allow the test model to have the dynamics fully develop through a flight relevant free-stream profile and as such require extrapolations with resultant high uncertainties in order to relate the test dynamics to flight expectations. SPEED is a NASA Ames Center Innovation Fund (CIF) project that is developing a highly tailorable and cost-effective test methodology to better assess the dynamic stability of blunt-body reentry vehicles via a stratospheric balloon flight. This is accomplished by dropping a suite of instrumented capsules from a stratospheric balloon to gain a statistically relevant dataset of scaled reentry vehicles in mission relevant free-flight conditions. This presentation will walk through how the test methodology is being implemented specifically for the Mars Sample Return (MSR) Earth Entry System (EES) geometry in an awarded Flight Opportunities Program (FOP) test flight in early CY24. SPEED Application to MSR: SPEED consists of three main mechanical systems: the Drop Platform, the Projectile, and the test Capsule. SPEED is being developed as a set of guidelines and recommendations for how to test with the proposed Concept of Operations (Conops) since the specific design parameters will vary depending on the specific project’s reference trajectory and entry vehicle design. As such, this presentation will walk through the development time-line as shown in Fig. 2. This is meant to serve as a blueprint for further missions as desired. Mechanical and Avionics Design. The SPEED test platform designed for the MSR-EES capsule geometry with nominal entry parameters has the ability to carry 10 Capsules to altitude instrumented with: 1. 3-Axis Accelerometer 2. IMU 3. Gyroscope 4. Magnetometer 5. Pressure Transducer cruciform 6. Uplook and Horizon Cameras To package the avionics/instrumentation suite, the capsule is approximately 1’ in diameter with the Outer Mold Line (OML) centroid-scaled from the full EES design. The internal volume is gutted and custom-shaped to fit the desired instrumentation suite, as well as to allow for the positioning of ballast mass such that the Center of Gravity is analogous to the flight vehicle. All structural components in the Capsule and Projectile are 3D printed, which significantly reduces the cost of each flight unit to around $1500 including all instrumentation, avionics, and structural components. Flight Conops. The test Capsule is accelerated to the desired altitude and Mach number while stowed in the Projectile, a missile-like vehicle consisting of steel ballast in the nose, a low-drag OML, and an Ejection Mechanism to reliably release the Capsule into the free-flow supersonic conditions. For the MSR-EES design, the capsule employs ~3kg of ballast mass at the nose to accelerate the 1.25kg test Capsule to ~Mach 1.7 at 23km altitude. This requires an initial release altitude of 40km, the quoted limit of a 80kg payload by the FOP-contracted balloon provider. Once the Ejection Mechanism avionics detect the proper conditions, the spring-loaded Ejection Mechanism will release and – guided by the sabot – expose the test Capsule to the desired test conditions for ~5 seconds of free-flight in the supersonic/transonic regimes. Dynamics in the subsonic regime will also be captured with the instrumentation suite with post-flight recovery operations aimed at recovering the high-G-load capable SD cards after the planned hard impact landings. Testing and Development: In the few months the SPEED project has worked the development of MSR-EES flight test, the team has performed lab and drone based testing which this presentation will overview. After the first design phase, the team fabricated Engineering Demonstration Units (EDUs) of all subsystems to perform validation testing shown in Fig. 5. After validation was completed on the subsystem level, a drone-drop test was performed at the recreational flight ceiling of 400ft altitude to assess the SPEED systems in a flight environment. Parameters such as in-flight stability, hard impact landing performance, and avionics performance were quantified and qualified. The FY23 CIF will culminate in a helicopter drop test aboard an Air National Guard Blackhawk. This will prepare the team for the CY24 FOP stratospheric balloon flight that should provide the final verification to begin offering the test platform for mission support. Focus of Presentation: This presentation will outline the technology maturation path of the SPEED implementation to the MSR-EES capsule baseline as well as the details regarding the mechanical system, avionics and instrumentation, and flight operations. Note that a complementary presentation is being submitted for a methodology overview of the SPEED test platform, introducing the testing technique and benefits as well as the full application space of the technology.

pitch damping coefficient↗
NASA NTRS2011

Base Flow Investigation of the Apollo AS-202 Command Module

In recent years, both Europe and the US are developing hypersonic research and operational vehicles. These include (re)entry capsules (both ballistic and lifting) and lifting bodies such as ExoMars, EXPERT, ARV, CEV and IXV. The research programs are meant to enable technology and engineering capabilities to support during the next decade the development of affordable (possibly reusable) space transportation systems as well as hypersonic weapons systems for time critical targets. These programs have a broad range of goals, ranging from the qualification of thermal protection systems, the assessment of RCS performances, the development of GNC algorithms, to the full demonstration of the performance and operability of the integrated vehicles. Since the aerothermodynamic characteristics influence nearly all elements of the vehicle design, the accurate prediction of the aerothermal environment is a prerequisite for the design of efficient hypersonic systems. Significant uncertainties in the prediction of the hypersonic aerodynamic and the aerothermal loads can lead to conservative margins in the design of the vehicle including its Outer Mould Line (OML), thermal protection system, structure, and required control system robustness. The current level of aerothermal prediction uncertainties results therefore in reduced vehicle performances (e.g., sub-optimal payload to mass ratio, increased operational constraints). On the other hand, present computational capabilities enable the simulation of three dimensional flow fields with complex thermo-chemical models over complete trajectories and ease the validation of these tools by, e.g., reconstruction of detailed wind tunnel tests performed under identified and controlled conditions (flow properties and vehicle attitude in particular). These controlled conditions are typically difficult to achieve when performing in flight measurements which in turn results in large associated measurement uncertainties. Similar problems arise when attempting to rebuild measurements performed in "hot" ground facilities, where the difficulty level is increased by the addition of the free-flow characterization itself. The implementation of ever more sophisticated thermochemical models is no obvious cure to the aforementioned problems since their effect is often overwhelmed by the large measurement uncertainties incurred in both flight and ground high enthalpy facilities. Concurrent to the previous considerations, a major contributor to the overall vehicle mass of re-entry vehicles is the afterbody thermal protection system. This is due to the large acreage (equal or bigger than that of the forebody) to be protected. The present predictive capabilities for base flows are comparatively lower than those for windward flowfields and offer therefore a substantial potential for improving the design of future re-entry vehicles. To that end, it is essential to address the accuracy of high fidelity CFD tools exercised in the US and EU, which motivates a thorough investigation of the present status of hypersonic flight afterbody heating. This paper addresses the predictive capabilities of after body flow fields of re-entry vehicles investigated in the frame of the NATO/RTO - RTG-043 Task Group and is structured as follows: First, the verification of base flow topologies on the basis of available wind-tunnel results performed under controlled supersonic conditions (i.e., cold flows devoid of reactive effects) is performed. Such tests address the detailed characterization of the base flow with particular emphasis on separation/reattachment and their relation to Mach number effects. The tests have been performed on an Apollo-like re-entry capsule configuration. Second, the tools validated in the frame of the previous effort are exercised and appraised against flight-test data collected during the Apollo AS-202 re-entry.

Walpot, Louis M. G.↗
NASA NTRS2018

Studies in Support of Venus Aerocapture Utilizing Drag Modulation

Aerocapture has been extensively studied and these studies have shown the benefit for planetary exploration missions. While the traditional approach to aerocapture with lifting configurations and lift-guided modulations have been assessed to be technologically feasible, aerocapture using purely drag modulation was proposed and studied by Prof. Braun and his students. These studies show that if one can assess the feasibility of aerocapture using drag modulation at Venus, and develop tall pole technologies needed at Venus, then this concept is much easier to execute at all other relevant destinations. Based on the above finding, partnered proposals were submitted by Adam Nelessen at JPL and Ethiraj Venkatapathy at Ames in collaboration with Prof. Braun at the University of Colorado, Boulder (UCB). Under this partnership, Ames Research Center (ARC) is working to address some of the key entry technology challenges associated with drag modulation aerocapture at Venus. Drag modulation aerocapture is a simple, scalable, and likely cost-effective way to enhance planetary science missions. The approach envisioned is to design a small spacecraft, that would most likely be a secondary payload, with a removable drag skirt. The vehicle would enter the atmosphere at Venus with a low ballistic coefficient, decelerate rapidly, drop the skirt resulting in a smaller vehicle with a higher ballistic coefficient which would skip out of the atmosphere and enter into a desired orbit. ARC's role in this collaboration is multifold. First of which is to perform design studies on various pre- and post-jettison geometries utilizing a 3-DOF trajectory code to determine the aerodynamics and aerothermodynamics of the vehicles and evaluate viable thermal protection material system designs. Once these design studies are complete, Ames will then perform higher fidelity CFD and TPS sizing to further design the vehicles. Second, the multi-body separation dynamics of the drag modulation event will be explored using both CFD simulations (CART3-D and US3D) as well as possible ballistic range testing. ARC's tools and expertise have been used to assess and advise on the selection of the separating configuration. In addition to the preliminary evaluation, ARC will provide tools and expertise to UCB team members to further assess aerodynamic interactions between the separating bodies and provide guidance as to the feasibility of stable transition.

Beck, Robin↗